Helicopter Handling Qualities

NASA

CP

NASA Conference Publication 2219

c.1 1

Helicopter

Handling

Qualities

Proceedings of a Specialists TECH LIBRARY KAFB, NM

IIIyMIIuIIll~llllll

0093354

NASA Conference Publicaw,‘ LLI J

Helicopter

Handling

Qualities

Proceedings of a Specialists Meeting on Helicopter Handling Qualities sponsored by the NASA Ames Research Center and the American Helicopter Society and held at NASA Ames Research Center Moffett Field, California April 14-15,1982 National Aeronautics and Space Administration Scientific and Tdmial lnfommtion Branch

PREFACE

This conference publication contains the formal papers of a specialists' meeting

on helicopter handling qualities. The conference was co-sponsored by the American

Helicopter Society -

San Francisco Bay Area Chapter, and the NASA Ames Research

Center, and was held April 14 and 15, 1982 at the NASA Ames Research Center, Moffett

Field, California.

Helicopters are being called upon by the military and civilian communities to

perform more and more tasks, and extend operations into poor weather and at night.

Accompanying this increased use is a significant increase in pilot workload and a

need for better handling qualities. The ability to define handling qualities

required to perform such missions has not kept pace with the actual' uses. The

objective of this specialists'

meeting was therefore to develop an overview of the

status and problems in the development and specification of helicopter handling-

qualities criteria, and highlight topics for future research efforts by government

and industry.

The conference was divided into five sessions and a round-table discussion: Overview of Current Criteria - an overview of problems and needs for criteria, specification standards of helicopters and high-speed rotorcraft.

Agility and Maneuverability - the subject of agility-and maneuverability and how

to define maneuvers that can aid mission performance and survivability.

All Weather and Night - minimum requirements for stability, control, and dis-

plays in IFR landing approach and distinction that must be made for one- or two-pilot

operation.

Integrated Cockpit - techniques and criteria for improving the pilot-helicopter

interface and integration of control display devices with flight-control tasks

automation.

Handling-Qualities Technology - generic handling-qualities research techniques

and facilities.

Round-Table Discussion - an exchange of views on current handling-qualities

criteria and techniques, particularly, the problem areas and future needs.

Special appreciation is due to the Session Chairmen, Mr. Dale E. Hutchins, Mr. Tommie L. Wood, Mr. Bruce B. Blake, Mr. Dean E. Cooper, and Dr. James A. Franklin, for their efforts in developing the technical program. Also to the General Chairman,

Mr. C. Thomas Snyder for his guidance and support to the conference objectives, to

the Arrangements Chairman, Mr. George E. Tucker,

for handling all the local arrange-

ments for the conference activities, and to the Technical Information Division for

preparing and publishing the proceedings of the meeting.

David L. Key

Conference Technical Chairman

II IllI Ill II

CONTENTS

Page

PREFACE...................:. , . . . . . . . . . . . . . ..iii

1. VTOL AND VSTOL HANDLING QUALITIES SPECIFICATIONS

AN OVERVIEW OF THE CURRENT STATUS . . . . . . . . . . . . . .z . . . . . . . 1

Kevin W. Goldstein

2. CIVIL (FRE~4ctiju.s.) CERTIFICATION OF THE COAST Gum03 tit+65A 0AuPtiI~ . . . 9

J. C. Hart, J. M. Besse, and K. W. MeEZreath

3. BGEING 234 FLIGHT CONTROL DEVELOPMENT . . . . . . . . . . . . . . . . . . . 15

James J. Morris

4.

INFLUENCE OF MANEUVERABILITY ON HELICOPTER COMBATEFFECTIVENESS . . , . . . 23

.-

Michael FaZco and Dr. Roger Smith

5. FLIGHT TESTS FOR THE ASSESSMENTOF TASK PERFORMANCE

AND CONTROLACTIVITY . . . . . . . . , . . . . . . . . . . . , . . . . . . . 35

Heinz-J&gen Pauser and Dieter Hwnmes

6.

A HELICOPTER HANDLING-QUALITIES STUDY OF THE EFFECTS OF ENGINE RESPONSE

CHARACTERISTICS, HEIGHT-CONTROL DYNAMICS, AND EXCESS POWERON

NAP-OF-THE-EARTH OPERATIONS . . . . . . . . . . . . . . . . . . . . . . . , 47

LZoyd D. CorZiss

7. UNIFIED RESULTS OF SEVERAL ANALYTICAL AND EXPERIMENTAL STUDIES OF

HELICOPTER HANDLING QUALITIES IN VISUAL TERRAIN FLIGHT . . . . . . . . . . , 59

Robert T. N. Chen

8. AN ASSESSMENTOF VARIOUS SIDE-STICK CONTROLLER/STABILITY AND CONTROL

AUGMENTATIONSYSTEMS FOR NIGHT NAP-OF-EARTH FLIGHT

USING PILOTED SIMULATION . . . . . . . . . . . . . . . . . . . . . . , . . . 75

Kenneth H. Landis and Edwin W. Aiken

9. DEFINITION OF DISPLAY/CONTROL REQUIREMENTSFOR ASSAULT TRANSPORT

NIGHT/ADVERSE WEATHERCAPABILITY . . . . . . . . . . . . . . . . . . . . . . 97

R. Joseph MiZeZZi, Gary 'W. Mowery, and Carmen PonteZandoZfo

10. SOME PILOTING EXPERIENCES WITH MULTIFUNCTION ISOMETRIC SIDE-ARM

CONTROLLERSIN A HELICOPTER . . . . . . . . . . . . . . . . . . . . . . . . 109.

J. Murray Morgan

11. RESULTS OF NASA/FAA GROUND- AND FLIGHT-SIMULATION EXPERIMENTS

CONCERNINGHELICOPTER IFR AIRWORTHINESS CRITERIA . . . . . . . . . . . . , . 121

J. Victor Lebacqz, Robert T. N. Chen, RonaZd M. Gerdes,

Jeanine M. Weber, and Raymond D. Forrest

12. STATE-OF-THE-ART COCKPIT DESIGN FOR THE HH-65A HELICOPTER . . . . . . . . , 139

DanieZ E. Castleberry and Marsha Y. McEZreath

13. PERFORMANCE EVALUATION OF A KINESTHETIC-TACTUAL DISPLAY . . . . . . . . . , 145

Richard J. Jagacinski, John M. FZach, Richard D. Gilson, and Richard S. Dunn

V

14. SYNTHESIS OF.AN INTEGRATED COCKPIT MANAGEMENT SYSTEM . . . . . . , . . . . 151

Joseph A. Dasaro and Charles T. Elliott

15. THE ROLE OF VOICE TECHNOLOGYIN ADVANCED HELICOPTER COCKPITS . . . . . . . 163

Howard P. Harper

16. COCKPIT INTEGRATION FROM A PILOT’S POINT OF VIEW . . . . . , . . . . . . . 171

David L. Green

17. INTEGRATED COCKPIT FOR A-129 . . . . . . . . . . . . . . . . . . . . . . . 183

James A. Gracia, and Bryce W. Koth

Dott. Ing. FiZippo Reinu,

NEW DEVELOPMENTSIN FLYING QUALITIES CRITERIA WITH APPLICATION TO

18.

ROTARY WING AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . , . . , 193

Roger H. Hoh

19. HELICOPTER SIMULATION TECHNOLOGY: AN AMES RESEARCH CENTER

PERSPECTIVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199

Richard S. Bray

20. PAST APPLICATIONS AND FUTURE POTENTIAL OF VARIABLE STABILITY

RESEARCHHELICOPTERS . . . . . . . . , . . . . . . . . . . . . . . . . . . 209

WiZZiam S. Hindson

21 . A PILOT-IN-THE-LOOP ANALYSIS OF'SEVERAL KINDS OF HELICOPTER

ACCELERATION/DECELERATION MANEUVERS . . . . . . . . . . q . . . . . . . .

Robert K. HeffZey

22. APPLICATIONS OF PARAMETERESTIMATION METHODSTO THE

PREDICTION OF HELICOPTER STABILITY, CONTROL, AND

HANDLING CHARACTERISTICS . . . . . . . . . . . . . . . . . . . . . . . . .

G. D. PadfieZd and R. W. DuVai!

vi VIOL AND VSTOL HANDLING CUALITIES SPECIFICATIONS AN OVERVIEWOF THE CURRENTSTATUS Kevin W. Goldstein Aerospace Engineer Naval Air Development Center Warminster, PA The highlights of a canparative analysis Multi-Purpose System (LAtJPS) SHdOB, the Army Utility Tactical Transport Aircreft System between the current helicopter and VSTOL (UTTAS) UH-6OA, and the Advenced Attack hendlinp qualities specifications and four representative state of the art rotary wing Helicopter (AAH) all use advanced flight control systems for stability and control aircraft are presented.

Longitudin21, eugmentation. The need to adequately address lateral, and directional control power 2nd the flying qualities of these state of the dynamic stability cheracteristics were art vehicle/control systems has necessitated analyzed for hovering conditions. Forward the use of "type specific2tions1' or "prime flight static and dynamic stability were item developnnent specifications" uniquely anelyzed for the longitudinel and devised for e2ch new aircraft/control system.

lateral-directional axes. Results of the Pany papers have been written describing the analyses in terms of the numerous shortcanings of PIL-H-8501A in- applicability/utility of the VIL-H-85OlA realistically regulating handling aualities criteria are presented for each of the ebove of present and future helicopters2'6, areas. The review of the YIL-H-E5OlA indicating a very real need for 2n updeted criteria 2painst those in MIL-F-83300 and version of YIL-H-85OlA. A summary of the AGARD577 indic2ted many areas in which major problem ereas described by the ebove ML-H-8501A does not give 2dequate design papers is presented as a b2ckground and guidance.

overview of the current status.

To facilitate the development of revlised Notation criteria it is necessary first to canpile a deta bese of past and present helicopter Pitch Rate Damping (second-')

Mq

stability and control charlcteristics.

'Ihis Pitch Control Sensitivity (r2d/seaond2/inch) paper presents the beginning of such a M6B canpilation.

Six degree of freedom math models of the SP-60B and the CH53D single Yaw Rate Damping (second-') Nr rotor helicopters were an2lyzed ageinst the fundanental stability and control espects Yaw Control Sensitivity (red/second'/inch) addressed by PIL-H-8501A. Vertical control Worms1 Acceleration (feet/second') response and 2utorotation criteri2 were not included et this time.

Angle of Attack (radians) Flight test deta for the XH-59A Advancing Blede Concept (ABC), the Later21 Control Deflection (inch) XV-15 tilt-rotor, and the CH-46A tandem rotor were also included 2nd discussed.

Undamped Natural Period (r-ad/second) % 5 In the development of the present day Damping Ratio ,VSTCL handling qualities specifications, Roll Angle Attained within @l PIL-F-83300'2nd AGARD 577' extensive rotary One Second (degrees) CT@L Conventional Take-off and Landing wing pilot rating data were enalyzed to substantiete the finalized hover/low speed VSTOL Vertic2l/Short Take-off and Landing criteria. Although AGARD 577 is not intended to be a helicopter specification and ML-F-83300 has not been used by the Navy or Introduction Army for a helicopter development program, these specific&ions do supply alternative With the development of a new generation methods of addressing VTOL handling qualities Of rotary wing eircraft for military characteristics.

operations, The alternative criteria it has becane annarent that the from !JIL-F-E3300 2nd AGARD577 were directly present helicopter handling-qualities canpared with the criteria frun ML-H-8501A specification, ML-H-85OlA', cannot accuretely assess the characteristics of to highlight helicopter specification !deficiences and vehicle encmalies.

these aircraft. The fact that PIL-H-8501A was last updated 20 years ago only tends to amplify this point.

The Navy Light Airborne stabilization system. ‘Ihere is little YIL-H-8501A Deficiencies quantitative guidance available defining sufficient levels of control or stability.

As described ebove, the mejor military With the canplex augmentation systems being helicopter development programs since 1965 employed on the SH-60B and the CH-53E there have used type specifications designed is a need to set minimun quantitative levels exclusively for the flying qualities of degreded flying qualities for partial AFCS characteristics of a particular vehicle failures and single or dual engine failures.

mission 2nd rotor configuration. Although The three levels of flying qualities (see the type specifications were at first Table 2) used in the VSTOL and CTOL basically FIL-H-8501A with slight revisions, specifications could be incorporated in recent development of the SH-6OF and the AAH YIL-H-8501A to specify quantitative levels of was besed on type specificetions very degraded flying qualities for control different fran NL-H-8501A. ‘Ibis is due to response, static stability, and dynamic the need to address the increased mission stability in eny flight mode.

requirements of these helicopters. lhe launch and recovery of the SH-60B from a seaborne pl2tform .in up to Sea State 5 conditions is an example of these Table 1. Ex2mple of ML-H-85OlA criteria for stabilization system failures requirements. Recent work with the HXY type specificetion highlighted new problem areps, Helicopters employing eutometic including the need to address characteristics 3.5.9(d) that may be unique to a tilt-rotor stgbilization 2nd control or configuration. Through the past decade many stability augmentation equipment papers have been written describing specific or both shall possess a sufficient degree of stability and control areas in which PTL-H-85OlA is deficient.

Three of these areas are discussed in the with all the equipment disengaged to allow continuation of norm21 following paragrephs.

level flight and the meneuvering PIL-H-8501A presently addresses necessary to permit a safe 12nding helicopter flying qualities in terms of the under visual flight, condit.ions.

longitudinal, lateral, directional, and verticel 2xes. Thera is no systematic delineation between hover/low speed Teble 2. Flyjr2 qualities levels ch2racteristics and forward flight cheracteristics. In hover 2 helicopter pilot tends to use longjtudinal, lateral and Pilot FC! FO directional controls independently. For Rating Level Cescription example, in a station keeping task, transletion along the longitudinal 2nd 1.0-3.5 1 Flying oualities cleerly adequete leter21 exes is implemented by the respective for the mission Flight Phase cyclic input, while heading angle is controlled by pedel inputs. Forward flight 3.5-6.5 2 Flying Qualities adequete to cherecteristics of a helicopter tend to sccanplish the mission Flight resemble those of an airplene, thus the pilot Ph2sa but some increase in pilot needs to use lateral 2nd direction21 controls worklo2d or degr.zdation in mission in a coupled manner. Also meny single rotor effectiveness or both, exists.

he1 icopters show a coupled pitch-roll dynamic oscillation in hover, whereas in forward 6.5-9.0 3 Flying qualjties such th2t the flight 2 dutch-roll type response is often airplene c2n be controlled safely, A breakdown of the he1 icopter found.

but pilot workload is excessive or specification into hover/loh~ speed criteria mission effectiveness is inadequate, 2nd forward flight criteria (similar to or both.

VIL-F-83300) would be 2 me2ns to address the different axis couplings between hover and forward flight .

A third area that could benefit from 2 restructuring of tJILiH-8501A is in defining A sug;pest.ion bv Kev’ is that a criteria that are mission oriented. Tbe restructuring of FnIL-H-85OlA in line with relicopter specification currently uses a NL-F-83300 and VIL-F-8785C would allow for a weight paremeter for hover control power more thorough treatment of degredcd flying considerations that is the result of scaling qualities. YJL-H-8501A presently has laws and not meant to represent the qualitative criteria for feilures of power verietions in control response which may be boosted controls, automatic stabilization required for vehicle mission differences.

systems and engine failures. Table 1 Foth the VSTC!L and CTOL specificetions define presents one section of the criterion four classes of vehicles according to over211 eddressing failure of an autcm2tic mission requirements, although in ML-F-83300 the vehicle gross weight for all three the cless distinctions are only used for specifications are shown in Fig. 1. The control force limits and roll control lower bounderies of all three specifications effectiveness in forward flight.

Table 3 are substantiated by the level 2 rating given shows a general breakdown of mission as used to the XV-15 with augmentation off. There in tJIL-F-83300.

Shipboard recovery and ere two other major points to be raised from nap-of-the-earth (NCE) flight mission Fig. 1. First the CH-53D AFCS on response categories could be incorporated into these -has been described 2s quite adequate for the type of cl2ss divisions.

assault mission, yet the vehicle does not This then satisfy the VSTOL boundary.

substentiates the need for some type of weight dependency as used by ML-H-8501A. It Table 3. FIL-F-83300 clessification of aircreft is questioneble though whether or not pilots will eccept a lower response for extremely CLASFm DESCRIPTION lerge vehicles. For example, 2 vehicle in the heavy lift helicopter (HLH) gross weight I Sm211, light aircraft such as category (gw=130000 lb) would only need to - light utility attain a b2nk angle of 2.1 degrees within one - light observation second for a one inch leterel stick displacement to satisfy the tEIL-H-8501A II pedium weight, low-to-medium requirement.

maneuverability aircraft such as - utility The second point fran Fig. 1 is the - search and rescue large difference in roll response between the - enti-submarine similer weight SH-60B and CH-46A (ten degrees - 2ssault transport per inch versus four degrees per inch). lhe CH-46A has been described 2s having very III ,satisfactory response ch2racteristics for its Lerge, heavy, low-to-medium m2neuverebility aircraft such as assault and verticel replenishment missions.

- heavy transport The SH-6OB has been qualitatively described - heavy bomber as having just adequate response chzacteristics for a turbulent, high se2 IV High maneuverability aircraft state condition, indicative of the LACPS such as mission. Yet the SH-60B shows a response - fighter well ebove the visual flight rules WFR) or - atteck instrument flight rules (IFR) ML-H-850lA bounderies. The difference bethreen these two vehicles then raises the point of h2vin.g ettitude response criteria dependent on the The Navy h2s begun a program essessing vehicle mission 2nd weight. In particular the b2sic flying qualities criteria in the small landing pletforms and dynemic E"JL-H-85CJlA against the VSTOL specificetions atmospheric conditions Navy helicopters will (YIL-F-83300 and AGARD 577) 2nd be expected to 12unch and recover fran ere en representative present and future rotary wing exwple of a mission that may not be eircraft. The significent results fran the edequately designed for by the still wind, assessment of hover control power criteria out-of-ground effect control power criteria and dynemic response criteria are presented presently in ML-H-8501A.

in the following sections.

Hover Control Power Helicopter control power requirements are usually determined by the hover mission control requirements.

As described above, ML-H-P50lA uses a weight paremeter to specify attitude response within one second or 1QSS.

In an extensive review of YIL-H-P50lA, Walton 2nd Ashkenas2 suggest that the InIL-H-P5OlA weight dependency is too simplified to give adequate guidance for verious vehicle missions.

In canparison to ML-H-850lA the two VSTCL specifications define a constant limit of ettitude response.

Fig. 1. Hover roll response canparisons The boundaries for roll 2ttitude per inch of lateral control displecement 2s a function of To insure that the helicopter response Overall it was found th2t the is not initially too sensitive PIL-H-8501A YIL-H-8501A attitude response and enpuler also has minimum engular rate danping rate d2mping criteria gave minimal design criteria for the longitudinal, lateral, and guidence in canparison to the vehicles direction21 2xes. Using these d.zmping analyzed. Further enalysis and data are bounderies with the ebove attitude response needed to determine the effect of vehicle criteria, rate danping versus sensitivity mission and veried rotor configurations.

bounderies can be developed. Fig. 2 shows the ABC and tilt-rotor canpared to the VIL-H-8501A requirements for the yaw axis.

‘Ihe interesting point here is that neither Dynamic Stability aircraft satisfied the requirement yet. the ABC has been described in 2 recent Navy Following a disturb2nce (control or flight test progr2m ’ 2s having “crisp, 2tmospheric) to a helicopter in hover the predictable” yaw control 2nd that the “high rate danping criteria discussed above should yaw rates (in excess of 45 degrees per ensure 2n initiel satisfactory response.

second) that resulted fran one inch pedal After this initial response the aircraft may step inputs were well-d2mped 2nd easily still have an unacccpteble dynamic response.

errested , all owing 1erge, rapid heading In a precision hover task it is mandetory Changes.” The m-15 in canparison was that the pilot be 2ble to correct eesily for described as sluggish and not adequate. The unwanted oscilletory responses. Uncommanded point here is not that the ABC is good 2nd pitch or roll responses ten cause tracking or the XV-15 bad, but the differences in the two station keeping errors, plus any short period rotor configurations. The ABC develops yaw dynamic responses must be well-d2mmped so 2s control t,hrough differential collective of not to impede precise control of the the two rotor systems while the tilt-rotor helicopter.

develops yaw control via differential cyclic inputs. The results presented in Fig. 2 show Satisfactory boundaries for dynamic an epparent anomaly between tJIL-H-850lA and stability characteristics are defined by eech the different rotor configuretions of the ABC of the specifications reviewed through the and tilt-rotor. Fig. 3 shows the pitch use of second-order response parameters. The response charecteristics of the SH-60B, general trend is similar for 211 the CH-53D and the XV-15. Similer to the specificetions such that short period directional axis tJIL-H-85OlA 2dequately oscillations require a d,mped response while predicts the single rotor vehicle ratings for longer periods, neutral stability to (the SH-6OE and the CH-53D) but again the slight instebility is acceptable. Fig. 4 tilt-rotor shows a discrepancy. shows a plot of nondimensional damping ratio versus d2mped netural period with a canparison of the three specifications for pitch or roll hover dynamic responses. Note that only VIL-H-AOIA has a separete boundary for VFR conditions.

.6 .5 Ol 0 .25 .50 rdhc’ I6 p (TF-1 Fig. 2. yaw rate vs. sensitivity canparisons am Fig. 4. Hover longitudinal dynamic stability requirements Pig. 3. Pitch rete vs. sensitivity canparisons It should be noted that it is assuned within Each aircraft has also been given level 1 ML-F-83300 that “IFR capability is inherent ratings, in particular the SH-60B is in all military aircraft operational described as having excellent phugoid missions.” For the limited data aveilable damping. It should be noted that both VSTOL very few conclusions can be drawn about the specifications have additional requirements adequacy of the specifications boundaries. for short period oscillations such that the Of the three aircraft shown only the SH-609 damping ratio must be at least 0.3. AGARD shows a “conventional” phugoid mode. Within 577 defines a short period response such that reference 3 the point is presented that for the dwped period is less than 3 to 6 modern helicopters the MAIL-F-E3300 boundary seconds. MIL-F-83300 specifies short period shown in Fig. 4 is generally undemanding. requirements according to Fig. 6. Note that This is questionable considering the SH-6OB the frequency boundary is a function of the response that Navy pilots described 2s vehicle n/a ratio. The CH-53D was the only adequate for the LAMPS mission. Both the vehicle analyzed that showed 2 short period CH-53D and the XH-59A have also been type response, and it canpared favorably with qualitatively described 2s having level 1 the Fig. 6 boundaries (e.g. G > 0.3). For the characteristics. In particular the CH-53D vehicles canpared against YIL-H-850lA, the specification gives lenient but adequate has essentially dead-beat dynamic responses in hover. From the data analyzed it appears guidance for normal flight conditions.

that ML-H-85OlA gives adequate guidance for hover dynamic responses.

Just as in hovering conditions, it is necessary that a helicopter have satisfactory dynamic response characteristics in forward flight . For ex anple , in contour flying or mine sweeping missions, a slowly divergent phugoid response with a gradual altitude loss would be objectionable. YIL-H-8501A specifies VFR and IFR dynamic response criteria for the longitudinal axis (the same as the above hover requirements), while only stipulating IFR criteria for the lateral-directional axes.

Looking first at the longitudinal criteria, Fig. 5 shows a canparison between the VSTOL and helicopter specification Fig. 6. VSTOL specificet,ion short period boundaries. lhe he1 icopter specification is reouirement’s by far the most lenient in specifying stability requirements,. in particular for The lateral-directional dynamic long period responses (>20 seconds) under VFR conditions. stability requirements as specified by thQ In contrast., the VSTOL specifications do not 2110~ divergent long VSTPL and helicopter specifications are shown period dynamic responses. With augmentation on Fig. 7. The same general trend is followed by each criterion. Note that on, the three vehicles shown on Fig. 5 easily YIL-H-8501A has no requirement for VFR satisfied all the specifications.

lateral-directional dynamic stability.

-Fig. 5. Forward flight longitudinal dynamic stability requirements guidance to address the differences in The cluster of open symbols shows a camnon handling qualities charecteristics drmped dutch roll response for the single between hovering and forward flight rotor helicopters (SHdOB, CH-53D, SH-3A) conditions.

analyzed. This type of yaw-roll coupled dynamic response has been given ML-H-8501A has very limited guid2nce unsatisfactory ratings for single rotor for degraded flying qualities, he1 icopters. Thus there should at least be a especially towards defining minimun base1 ine criteria limiting all owabl e characteristics for AFCS failures.

divergent responses for VFR conditions. For 2ugment2tion on the responses are all ‘Ihe hover control power criteria well-damped over 2 wide range of frequencies.

(2ttitude response and rete dwping An interesting canparison between v2ried criteria) inabequat.ely address v2ried rotor configurations is shown on Fig. 7 as mission characteristics or rotor the ABC has a dutch roll response that falls configuration differences.

right on the ML-F-83300 level 1 bound2ry.

Pilots described the ADC as having very Dynamic response criteria are in general satisf2ctory later21-directional forward 2dequate but very lenient, in particular flight characteristics that were very similar for VFR mission reouirements where no to a fixed wing aircraft. A Sikorsky report guidance is given for (reference 10) on the ABC compared this leteral-directional responses.

response to lr”IL-F-8785, the fixed wing flying qualities snecification. The ABC aaain Analyses in the 2reas of height control appears as an anomaly in canparison-to the response, aerodynamic and gyroscopic helicopter sprcificetion boundary. For the cross-coupling characteristics, and vehicles analyzed PIL-H-8501A gives adequate autorotation criteria are underway.

guidance for IFR 12teral-direction21 dynamic responses but has no guidence for VFR conditions.

r -.5 20 25 5 0 3 10 15 Damped satur.1 period (5ecoml.)

Pig. 7. Forward eight lateral-direction21 dynemic stability requirements References Conclusions 1. Anonymous, “He1 icopter Flying 2nd Ground Although the need to update UL-H-850lA Handling Qualities; Generel Requirements hes been known for many yeers, very little for,” PIL-H-85O?A, 7 Sep 7961.

systematic work has been directed towards developing modern criteria. A step towards 2. Walton, R. P. and AshkenPs, I. L., this goal is the future Army-Navy progrzn “Analytical Review of. k?litary Helicopter designed to develop an updated rotary-wing handling qualities specification. This paper Flying Qualities, ” Systems Technology, Inc., Technical Report No 143-1, Aug 1967.

bes presented the major deficiences in YIL-H-85Q1A es cited by many previous papers 3. Green, D. L. and Richards, R. F. Jr., as well 2s the signific2nt results of a “Review of KL-F-83300 ‘Flying Qualities of preliminary Navy assessment of YIL-H-P5OlA.

Piloted VSTOL Aircraft’ to Assess the In particular: Applicability to Helicopters,” Pacer Systems, Inc., Technical Report No PWR-054-72, Yar ML-H-8501A does not give edequate 1972.

p. Anonymous, 'VSTOL Handling Qualities 4. Key, D. L., "A Criticue of Handling Criteria, Part I - Criteria and Discussion Qualities Specifications for US Military Dee 1970; Part II - DocldTlentation," NATO Helicopters," AIAA Paper 80-1592, Aug 19PO.

AGARD Report 577, Jun 1973.

5. Dooley, L. W., "Handling Qualities 9. MacDonald, LCDR T. L., USN, and Kolwey, Considerations for NOE Flight," American "Advanced Helicopter Rotor Systems Second S., Helicopter Society Journal, 22, (41, Ott NAVY Evaluation of the XH-59A Advancing Plede 1977.

Concept (ABC) Demonstrator Aircraft," NATC Report RW-39R-80, 24 Dee 19PO.

6. Pitt,D. M. and Heacock, F. E., "Advanced Scout Helicopter Flying Qualities 10. Ruddell, A. J., et al, "Advancing Flade Requirements, How Realistic Are They?" AHS Concept (ABC) Technology Demonstrator," Report 79-Z, May 1979.

.USAAVRADCCY-TR-Pl-D-5, Apr 1981.

"Flying Qualities of Piloted 7. Anonymous, VSTOL Aircraft," MIL-F-83300, Dee 1970.

llllllll1llll I I Ill I

CIVIL (FRENCH/U.S.) CERTIFICATION OF THE COAST GUARD'S HA-65A DAUPHIN J. C. Hart Manager of Flight Test and Certification, SRR Aerospatiale Helicopter Corporation Grand Prairie, Texas J. M. Besse Director of Flight Test Department Helicopter Division, Societe National Industrielle, Aerospatiale Marignane, France .: ,; K. W. McElreath Manager Systems Design and Verification Flight Control Systems Department Rockwell Collins Government Avionics Division Cedar Rapids, Iowa ABSTRBS;T One of the requirements this program a dynamic In imposed by the Coast Guard for the simulator was designed and new Short Range Recovery helicopter constructed by Rockwell Collins is it be FAA certified.

that The Government Avionics Division to Aerospatiale HH-65A Dauphin is in support and verify the dynamic the certification process, both in avionics aspects of the system, France and in the United States.

pa;ticularly the Automatic Flight The basic aircraft/enqine Control System (AFCS). The role of combination is being certified- in Dynamic Simulator in this the France for VFR daytime operation program will be discussed.

with FAA compliance under FAA The Brussells. night Category II IFR certification is being UCTIQN conducted at Grand Prairie, Texas under cognizance of FAA Southwest Coast In June 1979 the U.S.

Region. Guard signed a contract with Aerospatiale Helicopter Corporation (AHC) of Grand Prairie, Texas, for This paper will describe both 90 HH-65A helicopters. These certification programs with aircraft are intended to be used as particular emphasis on handling Short Range Recovery helicopters, qualities requirements for each. replacing the current HH-52's.

Completion of VFR Derived from the Aerospatiale SA the Type Certification is scheduled to be 365N civil helicopter, the HH-65A be- FAA certified completed late this year and the is required to IFR certification in the United before the first delivery to the States in August 1982. The authors Coast Guard in late 1982. This includes both the will attempt to identify certification d'ifferences, if any, in aircraft and its avionic systems the certification resuirements of the and will encompass VFR, dual-pilot IFR, and Category II ILS two countries. -This program is unique in that the Automatic Fliqht requirements. Furthermore, the rescue mission Control System is a four-axis nature of the including stabilization demands that the aircraft and its system through the collective control. systems be designed to perform low-speed and hover Thus, stabilized flight in the low extended speed regime will be an integral operations, thus causing the effort to address part of the development flight test certification program. capabilities heretofore not available.

lllllllllllIlI I I I I II

&licowter Cou~l& FJL Director The FAA certification criteria Transport Category Rotorcraft, for -0 FAR part 29, and FAA Southwest Region's Airworthiness Criteria for are The FDS complements the pilot Helicopter Instrument Flight providing automatic path the primary governing documents for by following or maneuvering through the HH-65A.

the certification of the AFCS. The pilot selects the In addition, all of the aircraft desired FD mode to perform the and avionic systems will be desired maneuver automatically. He certified to perform their intended or not then may modify that mode by using functions, whether regulatory criteria exist. beep/sync switches on the control stick or by making manual control stick inputs. The five modes designed especially for low speed helicopter operations are the Approach mode (A-R) I the Transition-to-Hover mode (T-HOV), the Airspeed and Vertical RescueMission Speed/Altitude hold mode (IAS-VS), the Hover Augmentation mode (HOV AUG) and the Takeoff/Go-Around mode The avionics system allows the (GA). As a reversionary feature, and helicopter to confidently crew the pilot may fly the pitch, roll perform operations which would be and collective FDS steering difficult or impossible otherwise.

commands on the Attitude Director Although only the Automatic Flight or AFCS, directly Indicator to continue a task in Control Svstem, case a partial failure of the AFCS aircraft handling impacts the - occurs.

qualities and stability, the total impact of sensors, displays and guidance computations on the pilot workload and performance is equally . . . .

the lSSJ&.l m ComDuter great. In particular, all of these integration of to automatically perform elements The Mission Navigation given tasks, such as an approach to Computer acts as a full-time hover at a rescue site, contribute to crew navigator on board the' helicopter, significantly automatically fixing the aircraft effectiveness, safety and mission position, managing the navigation success. Singularly important system features are the following: computing radios and sensors, courses and even fliqht plan generating precise search patterns to be automatically followed by the Four-Ax+ &KS (PitcL Roll. m FDS through the AFCS. In addition, Collective1 the computer provides synthetic three-dimensional approach to hover guidance at any point where the The four-axis AFCS provides The pilot pilot desires to hover.

full-time stability and command final hover indicates his desired augmentation over the entire flight position by pressing a button when envelope for maneuvers in all overflying the point. The computer pitch, roll and yaw. It also then generates a lateral course and provides automatic trim, hands-off 5 degree descending approach path attitude and heading hold, and just downwind of the to a point coupled following of the flight similar to ILS guidance.

target, director commands. It . flight director's APPR Using the fail-passive and allows the oil:: and T-HOV modes, the pilot may to make manual control inputs at approach accomplish an automatic time. The design goal of the transition to a stabilized any and AFCS was to enhance the natural hover.

handling qualities of the aircraft, making them more consistent, but not substantially altering them.

Because of the innovative nature To complement the guidance and and advanced capability of the avionics system for IFR fliaht the pilot's task of monitoring the flight and aircraft situation, the and low-speed, remote a;ea operation, along with the attendant HSVD system displays various modes associated with a given mission impact upon crew workload and phase or task. Besides performance, AHC and Rockwell-Collins conventional HSI, MAP and RADAR or planned to reduce the development and certification FLIR video presentations, the HSVD also has a hover display mode for schedule risk initially by operations. Significant evaluating the avionics system on a low-speed data displayed in this mode are fixed-base "dynamic simulator.'

This computed wind speed and direction, engineerinq development the current omnidirectional simulator -combined actual avionics airspeed vector and the flight flight hardware, disolavs and computers director commanded longitudinal and with a-simulated-cockpit lateral speed reference for and aircraft response model. The automatic hoverins. Such aircraft model was programmed to information apprises- the pilot cover the entire fliqht envelope.

continuallv of the aircraft flisht from 20 knots rearward to 140 knots condition -and allows him to make forward fliqht and UP to +1500 feet hover minute- verticai speed.

decisions based on known The per data. Both sideward and cockpit incorporated the aircraft can be and forward/rearward flight avionics controls and displays carefully controlled and used to to perform a total mission profile the best advantage during low speed with realistic scenarios. In addition, or hover operations. This display the actual aircraft mode then complements the automatic control system with properly hover capability of the FDS and emplaced AFCS servos and feel/trim AFCS confident units duplicated the proper feel for safe, maneuvering. and pilot-AFCS interaction. Thus, the - avionic equipment interfaced and performed exactly as it later Although all of the above would in flight. Two objectives were addressed and avionic capabilities normally met using the coordinated fashion, dynamic simulator: (1) The system operate in a they independently provide operation was verified and refined to reversionary capability in case of reduce the flight test schedule failure. Thus, the and schedule risk. That this goal single any pilot can still safely continue the was successfully achieved was flight or task if single manifested when the AFCS any element fails. successfully stabilized and controlled the aircraft the first time it was engaged in flight.

(2) . . . . . .

erification Certification The test pilots and U.S. Coast ELI Guard personnel could evaluate the suitability of the avionics system and the integrated system operation The VFR, daytime, Type for the intended missions, Certificate for the HH-65A is being especially for search and rescue.

issued by the French civil aviation The early use of the simulator authority, or DGAC, to Aerospatiale enabled many aspects of the system Division Helicopters of France, to be refined and modified while with FAA compliance via FAA the program schedule impact was Brussels. Then AHC of Grand still minimal.

Prairie is requesting Supplemental Certification (STC) of the izKFliahtTestina Type night, IFR, and Category II The HH-65A flisht testinq of operations, including all avionics and mission equipment, through the the avionics and mission equipment The AFCS, FAA's Southwest Region in Fort commenced in July 1981.

HSVD Worth, Texas. flight director system, system, multifunction display com/nav system and omnidirectional lllllllllllllll IIIll1 II I airspeed system have completed the engineering development testing.

The testing of the mission computer functions and the combined mission The currently projected date suitability of the various systems for delivery of the first aircraft to the Coast Guard is is in progress. The lower end of September 1982. The FAA this IFR speed envelope is limited certification primarily by the static stability process began in February with characteristics of the aircraft. submission of system functional, The HH-65A, as is typicai of most interface and fault analysis data helicopters, exhibits deteriorating to the Southwest Region, along with static stability characteristics on meetings and presentations.

Three the backside of the power curve. critical safety items were of During the handling qualities particular interest to the FAA: 1) survey, a reversal in the cyclic the fail-passive design of the stick position versus airspeed AFCS, which is intended for use in hands-off automatic hovering; 2) curve was noted below 40 knots.

This, of course, is contrary to FAA the survivability of those system standards. functions which are redundant; and 31 the qualification of disital software for the multiple data-bus, the HSVD Low-speed IFR flight potential system and the mission computer functions.

further limited by the lack of Zsplayed ems information. This data is processed by the flight director computer for the The major remaining milestones prior generation of steering commands to FAA certification are the production conformity inspection, during T-HOV and GA maneuvers.

Demonstrably safe IMC approaches to the Inspection Authorization Type and departures from hoverinq flight approval, and the FAA certification are possible either in coupled-or flight testing. From the manual flisht. This capability standpoint of handlins qualities and - crew should lead; at least in theory, to workload,- <he IFR evaluation will examine all landing minimums significantly normal those currently and degraded modes of operation for lower than available. This, unfortunately is IFR suitability.

not the case at the present time.

Expansion of the low-speed IFR CONCLUSION flight envelope will require, first all, a means of either of satisfying FAA static stability The requirement for FAA standards or modifying those certification has meshed well with requirements. Secondly, inclusion the originally stated Coast Guard mission and system performance of an airspeed ind.icating system, requirements.

Several special such the recently developed configuration changes have occurred CollinsasASI-800 is necessary.

due to the FAA involvement; This system aiLplays both lateral namely, the routing of the wiring OADS information and longitudinal cables and the independence of speeds from rearward flight to and certain displayed information.

the forward flight limit of the However, the overall process and aircraft. This indicator utilizes outcome reflects how similar the oiux3 data at speeds less than 40 Coast Guard's mission and aircraft knots, a blend of OADS and pitot requirements are to the typical pitot from 40 to 60 knots, and sophisticated offshore or corporate information only above 60 knots.

helicopter operators.

Hopefully, further technical adbances-. and experience gained, as a result of this and future As a result of the HH-65A will make certification programs, program, the groundwork has been IFR certification to zero speed laid for the application of many possible in the near future.

advances in helicopter avionics and integrated systems technology to the civil helicopters of the 1980's. Notable achievements are the four-axis AFCS, the low-speed coupled flight director system, multifunction CRT displays, omnidirectional airspeed system, computerized automatic navigation and mission aids other and a multiplex data bus interconnect system. With the groundwork of FAA certification once accomplished, the rapid introduction of these and other similarly advanced concepts is greatly facilitated.

I llllllllllIIIIlIIlIlII III1 II II I I

BOEING 234 FLIGHT CONTROL DEVELOPMENT James J. Morris Technology Manager, Commercial Chinook Boeing Vertol Company Philadelphia, Pennsylvania Abstract helicopter from Sumburgh to the Dunlin platform. The 234 handling qualities and flight control systems have been designed The Boeing 234 is the commercially certified derivative of the to accommodate this type of mission under IFR, VFR, day, and CH-47 Chinook. The automatic flight control system and flight night conditions. This paper willdescribe the 234 flight con- director with coupler have been designed to reduce pilot work- trol system, the criteria which led to this system, and the load for missions of approximately six hour duration during results of testing the aircraft.

VFR, IFR, day and night conditions. The AFCS system for the 234 is essentially the same system as developed for the CH-47D.

which has airspeed hold, attitude hold, and maneuver enhance- ment in all three axes. The system also has the capability to couple to the Sperry Helcis flight director system which pro- vides for enroute navigation and landing approaches. Certifica- tion testing has been completed, by both the FAA and CAA, to FAR Part 29 for Transport Category Rotorcraft and BCAR Section G: Rotorcraft. The aircraft was certified for civil operation in June 1981.

Introduction The Boeing 234 is the commercial derivative of the CH-47 Chinook tandem rotor helicopter which has accumulated over one and half million hours of flight. The aircraft (as shown in Figure 1) is designed to carry 44 passengers for a distance of 574 nautical miles with IFR fuel reserves at its maximum in- ternal load gross weight of 48,500 pounds. FAA and CAA certification was received in June of 1981. Revenue service with the initial customer, British Airways Helicopters (BAH), began on July 1, 1981.

Figure 2. Typical Offshore Oil Mission Flight Control System Development Criteria Development of the 234 flight control system encompassed three considerations which had to be satisfied in order to achieve a satisfactory system for civil operation.

The first consideration was that the applicable civil regulations had to be satisfied since it was a program requirement to be certified by FAA and CAA. The relevant FAA documents were FAR Part 29 and the interim I FR standards dated December 15, 1978. The applicable CAA documents were BCAR Section G: Rotorcraft, and for IFR working papers, number 612 (Instru- ment Flight) and number 615 (Automatic Flight Control and Stability Augmentation Systems). These defined a minimum level of flight characteristics during normal and degraded mode operations, the criteria used to judge the system during certi- Figure 1. Boeing 234 Heicop ter fication flight testing, and a minimum level of reliability for system operation. Table I summarizes the pertinent criteria (in Figure 2 shows a typical offshore oil mission for the 234 in a general sense) to which the 234 was certified and shows that, which advantage is taken of its range capability to go directly for the basic stability and controllability criteria, the FAA and from Aberdeen, Scotland to the Dunlin oil platform. This CAA requirements are very similar. For AFCS failures, the replaces the previous practice of flying fixed wing from Aber- CAA requirement is significantly more stringent with a maxi- deen to Sumburgh in Shetland Islands, and then flying by mum time delay of five seconds compared to the FAA maxi- mum of three seconds. With regard to what failures are to be evaluated, the FAA is more stringent with their 10mg probability of occurrence compared to the CAA requirement of 10m7.

llllllllllllllllIIlII I II II II I

stick gradient with airspeed. The CH-47B maintained the same TABLE I. HANDLING QUALITIES CERTIFICATION CRITERIA FOR VFR AND IFR FLIGHT basic configuration of the SAS but modified the signal shaping in the pitch axis and the yaw axis. On the CH-47C a Pitch CAA ITEM FAA Stability Augmentation System (PSAS) was incorporated which The rotorcraft shall be safely added the features of airspeed and pitch attitude feedback for controllable and maneuverable improved pitch stability. The CH-47D and 234 are the current Must be able to trim out steady step in the automatic flight control system (AFCS) development.

forces No undesirable dirontinuitier As shown on Table II, continuous airspeed and pitch attitude in control force gradients hold and stability has been added; bank angle hold logic has Static Longitudinal Must demonstrate static longi- Must demonstrate static longi.

been changed from a wings level hold to a capability to hold Stabilitv tudina, stability from hover tudinal stability from hover through VNE through VNE any bank angle; heading and altitude hold and maneuver en- static Lateral Must demonstrate static lateral- Must demonstrate lateral direr- hancement in the pitch, roll, and yaw axes has been added.

tiona, stability throughout IFR Ddrectiona, Stability directional stability throughout Cross coupled feedbacks have been added to the pitch, roll, and IF R envelope zelope yaw extensible links for improved failure characteristics. With Dynamic Stability Murt demonstrate dynamic Must demonrtrate dynamic stability characteristics stability characteristics this generation of the Chinook automatic flight control system Stabihty level varies with Stability level varies with a hands off capability has been obtained.

frequency of oscillation frequency of oscillation AFCS Failures Must demonrtrate failures with Must demonstrate failures wiIh TABLE II. CHINOOK AUTOMATED FLIGHT time delays varying from normaI time delays varying from 1.5 pilot reaction to 3 seconds, seconds to 5 seconds. dependin!

CONTROL SYSTEM CAPABILITY

_.. _ -

depending on flight condition on flight condition’ HANDLING Extremely remote IlO-‘) CH-47A CH.47B

F

QUALITIES (SASI ISAS) FUNCTION Al, Axes All Axes Rate Damping

I--.-

Pitch Attitude The second consideration was that two specific operational Hold and Stability Hold requirements had to be satisfied. The first requirement was Roll Atlitude Hold NO Any Sank Any Bank Angle Angle to provide for an approximate six hour flight over water NO NO NO Heading Hold Yes Yes during IFR and VFR, day, and night conditions. The second Airspeed Hold NO NO About Trim Continuous Continuou was to be able to operate in adverse atmospheric conditions - NO NO Altitude Hold NO Yes Yes specifically cross winds of up to 50 knots. This criteria is Maneuver Enhancement Yaw iTurn Yaw (Turn Yaw (Turn Pitch, Roll. Pitch, Roll especially necessary for operation in the North Sea, where in Entry Only) Entry Only Entry Only1 and Yaw and Yaw Cross Coupled Feedbacl NO NO NO Pitch, Roll. Pitch. Roll winter months it is not unusual to have winds UP to 50 knots.

for Improved Failure and Yaw and Yaw Characteristics The final consideration was the systems had to be acceptable

Flight Director NO NO No NO

I I ye=

by pilot qualitative evaluation. Included in these pilot evalua- tions were handling qualities evaluation during VFR, IFR, and Automatic Flight Control System (AFCS) Configuration flight in turbulence for normal AFCS operation and degraded mode operation. During these evaluations trimability, stability, The control system configuration is shown schematically in control cross coupling, and dynamic stability were evaluated.

Figures 3 through 7. The AFCS is in general dualized; i.e., The criteria of acceptability by pilot qualitative evaluation input signals, signal conditioning, and differential actuation.

determined the signal paths and the gain levels and shaping in The collective parallel actuator has not been dualized.

each of the sianal oaths for the hnsic AFCS Philosophy of the system mechanization has been to dualize Evaluation of these requirements resulted in the definition of an AFCS system which provides full time attitude and airspeed differential actuation paths which influence basic handling qualities and maintain a single system for parallel paths which hold, maneuver enhancement, and vernier trim capability, and includes the incorporation of a flight director and coupler do not affect basic aircraft handling qualities. Mechanization of the system is such that the handling characteristics of the which has navigation capture and tracking, approach guidance capture and tracking, altitude hold, vertical speed hold, heading aircraft are essentially unchanged on single or dual AFCS.

hold and course select.

The pitch axis of the AFCS (Figures 3 and 4) is comprised of two parts - dualized high rate, low authority, hydraulically Chinook Flioht Control Historv powered, extensible links for pitch damping; and dualized low rate, high authority, electromechanical Differential Airspeed The Boeing 234 configuration is a result of over 20 years of de Hold (DASH) actuators for pitch attitude and airspeed hold.

velopment of the flight control system as well as the evaluation Dual system authority for the extensible links is f 25 percent of requirements. The history of the development of the of cockpit control and for the DASH is 50 percent of cockpit Chinook flight control system is interesting in that it parallels control. Also included in the control laws for the DASH sys- the development of the state of the art of flight control systems.

tem is a longitudinal stick pick off for maneuver enhancement.

Data signals used by the AFCS for the pitch axis are pitch at- Table I I summarizes the features of the automatic flight con- titude (from which pitch rate is derived), airspeed, and longi- trol system for the CH-47A. B, C, and D. On the CH-47A the tudinal stick position.

SAS was essentially a rate damping system which improved the stability characteristics and provided a short term hands-off The roll axis of the AFCS is shown on Figure 5.

It consists of capability through a pseudo pitch attitude hold (lagged pitch signal paths for roll rate damping, roll attitude hold with syn- rate). A scheduled airspeed input into a differential collective chronization logic so that any commanded bank angle can be pitch actuator was provided to obtain a positive longitudinal held, lateral stick position feedback for maneuver enhancement, AFCS

9oLL

COMPUTER -- jj=-- IUAL LWRLE iZ~ ------J -+ IMMCNEU&& LATERAL[ROLL]AXIS PROYIOES: . STABILITYAUGNENTATION IMOTIONOAMPINCI . LATERALCONTROL RESPONSE OUICKENING . ROLLATTITUOEHOLO l ROLLAXIS TRIM LONOITUOINAL (PITCHI AXIS PROVIOES: l AUTOTIJRN~TO~NEAOING l STABILITY AUGMENTATION (MOTION OAMPINOI Figure 3. Pitch A FCS Mechanization Figure 5. Roll AFCS Mechanization AFT “IX AFCS CONPUTER -9-l UPPER I OIRECTIONAL [YAWIAXISPA . STAMLITY AUGMENTAT . PEOALCONTROL AESPO AIASPEEG/PITCHATT. HOLO . “EAOINE “OLD POSITIVESTICK GRAOIENT CONTROLAUGMENTATION Figure 6. Yaw Axis Mechanization Figure 4. Pitch AFCS (Dash) Mechanization roll attitude beep for vernier attitude control, and flight director ing, sideslip angle from sideslip transducers on the nose of the steering commands. Actuation for the roll axis AFCS is pro- aircraft, yaw rate from rate gyros, roll rate derived from roll vided by dualized hydraulically powered extensible links. attitude, and pedal position. Logic for the operation of the Dual system authority for the extensible links is f 26 percent yaw damping and holding hold signal paths varies in the yaw of cockpit control authority. Data signals used by the roll axis depending upon flight condition. For yaw damping at low AFCS are roll attitude (from which roll rate is derived), lateral speed there is full time rate damping. At high speed (above stick position, flight director steering command;and beep trim 54 knots) the rate damping is washed out with a four second command. time constant so that the yaw AFCS is not saturated during turns. For heading hold at low speed, the heading hold func- tion is synchronized with pedals out of detent so that the air- The yaw axis of the AFCS is shown on Figure 6. It consists of craft may be turned with pedals and relatches when the pedals signal paths for yaw damping, heading hold, sideslip stability, are returned to detent and the yaw rate is less than one and a pedal position for maneuver enhancement, and roll rate into half degrees per second. At higher speed (above 54 knots) the yaw for turn entry coordination. Actuation for the yaw axis heading hold circuit synchronizes for lateral stick or pedals out is provided by dualized hydraulically powered extensible links.

of detent so that the aircraft can be turned with the stick or Dual system authority for the extensible links is f 30 percent sideslipped with pedals. Heading hold relatches when the stick of cockpit control. Data signals used by the yaw axis are head- lateral steering commands through the aircraft roll control and vertical steering command through the collective parallel actua- tor. The system can be used in the coupled or uncoupled mode.

The flight director system consists of a mode selector panel (Figure 9) and a computer with sensor inputs provided from radio navigation, attitude, acceleration, and air data devices.

The flight director mode selector panel provides the controls for engaging/disengaging and displaying the status of any avail- able flight director mode. The modes provided are as shown in Table III.

TABLE Ill. FLIGHTDIRECTOR SELECTOR PANEL MODES Lateral Collectiv Axis Axis Heading Select Mode (HDG) X Navigation Mode (NAV) X Instrument Landing System Mode (I LS) X X . F/D VERTICAL STEERING NO, 1 AFCS COMPUTER Back Course Mode (BC) X Figure 7. Collective Axis Mechanization Go-Around (GA) X X and/or pedals are back in detent, roll rate is less than one and a VOR Approach Mode (VOR APR) X half degrees per second, yaw rate is less than one and a half Altitude Hold Mode (ALT) X degrees per second, and roll attitude is less than one and a half Vertical Speed Hold Mode (VS) X degrees.

- -

standby Mode ISBY) !

The collective axis of the AFCS is shown on Figure 7. It con- sists of a parallel actuator with signal paths for flight director commands and for altitude hold. The collective axis is not Heading Select Mode (HDG) redundant and operates through the Number 1 AFCS unit.

The Heading Select Mode is selected by pressing the HDG The flight director with its interfaces is shown on Figure 8. button on the mode selector. In the HDG mode, the flight di- The flight director is the Sperry Helcis system which provides rector computer provides inputs to the roll steering pointer to for enroute navigation and landing approach with the capability command a turn to the heading indicated by the heading bug for coupling into the automatic flight control system for on the HSI. When HDG is selected, it overrides the NAV. BC, GYRO r ROLL EXT LINK NO. 1 ROLL EXT - AFCG LINK NO. 2 UNIT NO. 2 t

pi!EGip

I 1 Figure 8. Flight Direc tar/A FCS Interface

ia

Instrument Landing System Mode (ILS) The I LS Mode is used to make an I LS approach. Pressing the I LS button with a LOC frequency tuned, arms both the loca- lizer and glideslope modes. In the I LS mode, both the NAV and I LS modes are armed to capture the localizer and glide- slope, respectively. The initial localizer capture angle is set using the heading bug similar to the VOR mode.

With I LS mode armed, the collective axis can be in any one of the other collective modes, except Go-Around. When reaching the vertical beam sensor trip point, the system automatically switches to the glideslope mode. The collective mode and I LS ARM annunciators extinguish and I LS GS annunciator illumi- nates. At capture, a command is generated to intercept the glideslope beam. Capture can be made from above or below the beam.

Glideslope mode is interlocked so the localizer must be cap tured prior to glideslope capture. If the glideslope receiver is not valid prior to capture, the vertical beam sensor will not // \\ trip and the system will remain in the existing collective mode.

Figure 9. Flight Director Mode Select Panel After capture, if the glideslope receiver or vertical gyro become invalid, the collective steering pointer will bias out of view.

and I LS modes. In the event of a loss of valid signal from either Back Course Mode (BC) the vertical or directional gyros, the roll steering pointer is biased out of view.

The Back Course Mode is selected by pressing the BC button on the Mode Selector. Back Course operates the same as the Navigation Mode (NAV) LOC mode with the deviation and course signals locked out when in the BC mode. When BC is selected outside the lateral The NAV Mode provides steering commands for both VOR and beam sensor trip point, BC ARM and HDG will be annunciated.

localizer navigation.

At the capture point, BC CAP will be annunciated with BC ARM and HDG extinguished.

VOR Mode Go-Around (GA) Pressing the NAV button on the F/D MSP with the navigation receiver tuned to a VOR frequency engages the VOR mode.

The Go-Around Mode may be engaged by pressing either the When outside the lateral bracket sensor trip point, the roll GA button on the mode selector or the remote GA button steering pointer receives a heading select cominand as describ- located on the pilot’s collective pitch lever. When selected, ed above, and both the NAV ARM and HDG mode annuncia- all other modes are reset and the GA annunciator is illuminated.

tors are illuminated. Upon reaching the lateral bracket sensor The roll steering cue receives a roll zero command while the trip point, the system automatically switches to the VOR collective cue commands a positive rate of climb of 500 fpm.

mode - HDG and NAV ARM annunciators extinguish and the NAV capture NAV CAP annunciator illuminates. At capture, a command is generated to capture and track the VOR Approach Mode (VOR APR) selected VOR course. When passing over the station, an over- station sensor detects station passage, removing the VOR Pressing the VOR APR button on the mode selector with the deviation signal from the command until it is no longer er- navigation receiver tuned to a VOR frequency engages the ratic. While over the station, course changes are made by VOR Approach Mode. The mode operates identically to the selecting a new course on the HSI.

VOR mode with the gains optimized for a VOR approach.

If the NAV receiver is not valid prior to the capture point, the lateral beam sensor will not trip and the system will re- Altitude Hold Mode (ALT) main in the HDG mode. After capture, if the NAV receiver, The Altitude Hold Mode is selected by pressing the ALT button compass data, or vertical gyro go invalid, the roll steering pointer will be biased out of view. on the mode selector. When ALT is selected, it overrides the I LS GS, GA, or VS modes and the altitude at time of selection Localizer Mode will be maintained. In the ALT mode, the collective command is proportional to altitude error relative to the engage reference.

The Localizer Mode is selected by depressing the NAV button Once engaged in the altitude hold mode, the collective steering on the MSP and being tuned to a lot frequency. Mode selec- pointer will bias out of view if either the VG or altitude sensor goes invalid, and the collective AFCS will revert to manual tion and annunciation in the LOC mode is the same as the VOR mode. control.

.9 Vertical Soeed Hold Mode (VS) HEAVY GROSS WEIGHT AFT CENTER OF GRAVITY 3,000 FOOT DENSITY ALTITUDE The vertical speed hold mode is engaged by pressing the VS button on the mode selector. When VS is selected, it over- CRUISE AT 0.9 VNE rides the I LS, GA, and ALT modes. A vertical speed reference FWD ?I!

is set by the bug on the pilot’s vertical speed indicator. Once engaged, if either the vertical gyro or altitude sensor go invalid, LONG.UN.) ;:!j the collective steering pointer will be biased out of view.

0.0 FIT 2.0 1.0 Standby Mode (SBY) 0.0 LAT IIN.) -l.o Pressing the SBY button on the Mode Selector resets all the -2.0 other flight director modes and biases both flight director FIT 2.0 command bars from view. While depressed, SBY acts as a 1.0 OIRIIN.) 0.0 lamp test, causing all mode annunciator lights to be lit and the -1,o flight director warning flag on the ADI to come in view. When -2.0 the button is released, all the other mode annunciator lights 40 60 80 100 120 140 160 TRUE AIRSPEED - KNOTS extinguish and the flight director warning flag retracts from SLOW CRUISE AT 1.1 VMIAIFR view.

Handling Qualities Characteristics I I I I 0.0 The handling qualities of an aircraft are quantitatively evalu- LONG.(IN.)-0.5 ated by its static and dynamic stability characteristics. Be- -1.0 I I I I I cause of the tandem rotor design and the configuration of the AFCS the handling characteristics of the 234 are essentially independent of variations in gross weight, center of gravity, and density altitude. Representative static longitudinal sta- bility characteristics areshown in Figure 10. There are several

RT ::o” I

characteristics which should be noted. First the longitudinal I I I 1 1 _I stick gradient is essentially independent of flight condition OIRUN.) 0.0 -,.. w ;, ; f+ 1 40 60 80 100 120 140 160 is a longitudinal axis task only - there is no significant cross- TRUE AIRSPEED - KNGTS coupling with the lateral or directional axis. Third, the true stability level of the aircraft to external disturbances is masked APPROACH due to the longitudinal stick pickoff which cancels part of the FWO 1.0 0.5 airspeed feedback. The stability level to external disturbances L0NG.UN.j 0.0 is approximately four times that shown. Representative lateral -0.5 directional static stability characteristics are shown in Figure -1.0 11. Note that the stability level is essentially independent of FIT 2.0 I I I I I 1.0 flight condition and that there is no cross coupling with the al,;1 LAT (IN.) 0.0 B ; I I longitudinal axis. Dynamic stability characteristics in cruise I I I I I -1.0 I I I I I I are shown in Figure 12. The response to pitch, roll, and yaw I I -2.0 I control pulses are well damped. Note that in the pitch axis RT 20 1.0 the system is designed for a return to trim capability, but the DIR (IN.) 0.0 logic in the roll and yaw axis is such that when the stick is re- -1.0 turned to detent a new attitude is held.

-2.0 40 60 80 100 120 140 160 TRUE AIRSPEEO - KNOT6 AFCS failures from dual system operation are mild due to MCP CLIMB the cross coupled feedback scheme on the extensible links FWO 1.0 (Figure 13). The effect of this mechanization is to obtain 0.5- 1 immediate relief from a number 1 or number 2 system actua 0.0 L ! !

LONG.IIN.) -0.5 tor hardover. Figures 14 and 15 show typical failure charac- I I I I I I -1.0 I teristics in cruise in the pitch, roll, and yaw axes. Failures RT 2.0 I I I I I I 1 from single system operation are more abrupt and are I I I I I 1.0 &I; ^ :^ characterized by delay times of one to two seconds and I I 0.0 ! 1 1 I LAT (IN.) -1.0 maximum pitch, roll, or yaw rates in the axis of failure of -2.0 ---- 10 to 15 degrees per second.

AT 2.0 1.0 Crosswind trim characteristics are shown in Figure 16. The 0.0 DIR (IN.1 -1,o we- 3 tandem rotor configuration is especially suited for this type -2.0 . ‘. . . . . . . ‘. . 1.‘. ” ‘. ‘. .B of operation because of its insensitivity to wind direction.

40 60 80 100 120 140 160 Note from the figure that low speed control is a one axis TRUE Al RSPEEDi- KNOTS control task - lateral stick. There is essentially no cross coupling with longitudinal or directional control. Figure 10. Static Longitudinal Stability HEAVY GROSS WEIGHT AFT CENTER OF GRAVITY HEAVY GROSS WEIGHT AFT CENTER OF GRAVITY 3,000 FEET DENSITY ALTITUDE PITCH FWO LONGITUOINAL 0 CONTROL (IN.1 AFT -5 NU PITCH RATE IoEGIsEcI NO -25 PITCH ATTITUOE (oEG) ON.1 NO -25 ROLL MT RT ON.1 LATERAL NT CONTROL OIR (IN.1 ON., -10 0 10 20 -20 SlOESLIP ANGLE - GEG AT ROLL RATE (OEGKEC) ROLL ATTITUOE (OEG) -20 -IO 0 10 20 SlOESLIP ANGLE - GEG YAW RT DIRECTIONAL CONTROL UN.)

!$ R2T5 YAW RATE IOEGISECI -10 0 10 20 -20 -‘2: SIGESLWANGLE--0EG Figure 11. Static Lateral Directional Stability RT YAW ATTITUOE 0 LOEGI -‘2!

Figure 12. Cruise Dynamic Stability HEAVY GROSS WEIGHT AFT CENTER OF GRAVITY CRUISE AT 135 KNOTS ROIL YAW I OIRECTIONAL ” LATERAL IT PEOAL D STICK POSITION POSITIbid Lo (IN.1 1IN.l I ‘: 0.25 LIO YAW SAS NO. 1 ROLLSAS NO. 1 ACTUATOR I ACTUATOR 0 TllTAL AFCS EXTENSION EXTENSION MECHANICAL ON. ACT.1 -0.~ IIN. ACT.1 -au MOTION a54 YAW SAS NO. 2 ROLLSASN0.2 ACTUATOR 0 ACTUATOR 0 EXTENSION EXTENSION IIN. ACT.1 -0.10 UN. ACT.1 -~.a LECENO zs ROLL YAW - ELECTRICAL 0 0 RATE SIGNALS RATE IOEG/SECJ IOEG/SECI - MECHANICAL -*I -ax couPLlMG 54 II ROLL SIOESLIP ATTITUOE D ANGLE a Figure 13. Extensible Link - Cross Coupled Feedback IOEGI IOEGI -25 YAW ATTITUDE 0 HEAVY GROSS WEIGHT AFT CENTER OF GRAVITY IDEG) -n CRUISE AT 135 KNOTS DASH Figure 15. AFCS Failure Characteristics - Roll and Yaw Axis LONGITUOINAL HLAVY GROSS WEIGHT MID-CENTER OF GRAVITY FORWARD o LONGITUDINAL STICK -1. I” I I p POSITION IINCHESI

I r i”I

-2 RIGHT 2 LATERAL STICK POSITION (INCHES) OASH NO. 2 PTI I I I -2 RIGHT , DIRECTIONAL PEDAL 1 b I b 1-1 POSITION (INCHES) -1 ACTUATOR up 5 EXTENSION COLLECTIVE LEVEL 4 POSITION (INCHES) SAS NO. 2 “,, z ACTUATOR 50 40 30 20 10 0 10 20 30 40 50 RIGHT LEFT STANDARD VELOCITY (KNOTS) Figure 16. Slow Speed Controllability Conclusions Development of a flight control system for today’s helicop- ters must consider the certification criteria for the countries ATTITUOE in which it is to be operated; diverse operational criteria which include I FR, VFR, long duration missions, and high wind conditions; and addition of pilot aids such as a flight Figure 14. A FCS Failure Characteristics - Pitch Axis director and coupler to minimize the overall niission work- load. For the Boeing 234 these criteria and needs have re- sulted in an AFCS with airspeed and attitude hold and maneuver enhancement and a flight director with a coupling capability.

INFLUENCE OF MANEDVERABILITY ON HELICOPTER COMBAT EFFECTIVENESS Michael Falco Senior Staff Scientist Grumman Aerospace Corporation Bethpage, New York Dr. Roger Smith Aerospace Engineer U.S. Army Aviation Research and Development Command St. Louis, Missouri Abstract concert with the other system attributes in an An investigation was conducted to quantify equally detailed way. It is necessary for the the impact of maneuver capability on the combat methodology to develop an optimal probability of effectiveness of current and advanced design kill or survival solution for all relative geometries for which combat can be initiated.

helicopters in one-on-one engagements against specific threats. Solution optimality is important for consistent A newly developed casputa- tional procedure employing a stochastic learning effectiveness comparisons between aircraft/ weapon concepts and serves to minimize the method in conjunction with dynamic simulation of effect of maneuver strategy prejudgments and helicopter flight and weapon system operation other preliminary bias factors introduced by the was used to derive helicopter maneuvering stra- tegies. The derived strategies maximize either analyst.

survival or kill probability and are in the form Application of modern optimal control and of a feedback control based upon threat visual differential game theory methods seems well or warning system cues. Maneuverability para- suited to these problems at first sight. How- meters implicit in the strategy development ever, the pioneering effort of Isaacs (Ref. l), included maximum longitudinal acceleration and followed by those of Breakwell and Merz deceleration, maximum sustained and transient (Ref. 2), indicate that there does not appear to load factor turn rate at forward speed, and be a general systematic method for solution of maximum pedal turn rate and lateral acceleration even some simply structured pursuit-evasion at hover.

Results are presented in terms of games. This difficulty has led applications- probability of kill for all combat initial oriented investigators (Ref. 3, 4, 5) toward conditions for two threat categories. In the consideration of discrete game approximations first category the use of maneuverability is which circumvent the analytical problems of the examined in a defensive role against an anti- continuous theory, and still offer some form of tank guided missile (ATGM) launched by a threat suboptimal solution in more realistic combat helicopter, The second category is concerned models.

with the impact of maneuverability in both defensive and offensive roles against a gun This paper presents a partial summary of armed helicopter threat. recent computational experience gained in military helicopter design applications using Introduction variations of a stochastic learning method first reported in Ref. 4. Representative ccmputa- tional results are presented for two important In the early stages of military helicopter conceptual design, there is a need for method- categories of one-on-one helicopter air ccmbat: ology to better quantify combat effectiveness in the first, a study of maneuver capability in terms of the major aircraft/weapon system attri- defending against an anti-tank guided missile butes such as design maneuver capability and (ATGM) launched by a threat helicopter; the maximum speed, weapon capability, passive/active second maneuverability employed defensively and survivability equipments performance, detect- offensively against a gun-armed threat heli- ability, and threat warning. To analyze the copter. An explanation of the maneuver strategy maneuver capability contribution to combat development and effectiveness assessment metho- effectiveness against various threats, the dology is given in both case studies. The associated models are required to be of high representative results reported here limit fidelity in terms of the dynamical simulation of Ihelicopter maneuvering to constant altitude solutions using variable altitude helicopter flight and yet permit the maneuver flight paths; contribution to be assessed either singly or in maneuvering with terrain constraints in the air-to-air gun study were not available in time reasonable measure of survivability against mis- for inclusion in this publication. Geographical sile threats. The aircraft in these investi- terrain features have not been considered in gations are assumed to employ an active radar these studies; the ground is modeled as a plane.

warning system supplying relative range and azimuth information regarding the incoming The same approach has been extended to threat. The baseline configuration for this problems of land warfare, particularly armored warning receiver model employs 12 azimuth gates vehicle maneuver effectiveness and survivability and 7 range gates from 0.25 km out to a maximum against anti-tank missile threats. Corrobora- detection range of 5'km, as shown in Fig. 2.

tion of the computer derived solutions for This configuration is indicative of the warning specific threat cases has been obtained in inde- receiver performance levels that are projected pendent field trials with the actual systems.

for operational systems in the near future.

Additional effort must be dedicated to flight At each threat warning contingency (represented trial verification of the model approximations by one of the 7 x 12 = 84 range/azimuth cells), and computed solutions. Continued research is the aircraft is allowed a choice from a finite warranted in the application of optimal control, number of elemental maneuvers. Five elemental differential game, and the stochastic processes maneuver choices are shown in Fig. 2.

The branches of applied mathematics to provide choices may be comprised of maximum performance effective numerical procedures for helicopter turns, longitudinal acceleration, deceleration, combat analyses.

and a straight ahead constant speed policy. In vertical plane maneuvering studies climb and pushover maneuver choices would be added. An Maneuverability in Air-to-Air Missile Avoidance aircraft evasive maneuvering strategy is the selection of an elemental maneuver for each Missile Threat threat warning cell. An optimal strategy is a strategy which maximizes aircraft survivability The threat is an optically tracked, wire for all launch initial conditions.

guided missile employing a semi-automatic com- mand to line of sight guidance system. This threat was primarily designed as an anti-tank guided missile (ATGM), but has air-to-air It is assumed to have a 245 application as well.

m/s sustainer velocity, maximum range of 4 km, In addition, and maximum flight time of 16.3 s.

it is assumed to have a 4 g maximum lateral man- euver capability, and that the launch aircraft is at co-altitude with the target. The low com- bat flight altitude of the target (dictated by detection and masking considerations) allows the survivability results to be safely extrapolated to ground launched cases as well. This threat is normally equipped with a shaped-charge con- Figure 1. Warhead Lethality tact fuse warhead for armor penetration. How- ever, proximity fuze warheads employing expand- ing rod or fragment kill mechanisms are also indicated to be adaptable to this missile air- frame, and two of these types were considered in this investigation. The contact fuse warhead lethality model utilizes a probability of kill, PR = 1.0 for missile contact anywhere on the helicopter fuselage envelope. Two proximity fuse warhead models are described in Fig. 1.

Warhead A denotes an expanding rod warhead as used in short range air-to-air missiles. War- head B is the largest blast/fragment warhead that can be accommodated by the missile airframe and propulsion configuration. The kill effec- tiveness, PR, of these two warheads is given 0413002P as a function of detonation distance RpET Figure 2. Aircraft Warning System & Maneuver Strategy (from the target eg). The data shown represent an average of all warhead/target detonation aspects; however, functional dependence upon aspect is considered in the studies. Stochastic Learning Method Threat Warning and Maneuver Strategy The stochastic learning method is comprised of two phases: a reinforcement learning phase, Earlier investigations have postulated the in which the optimized evasive strategy is need for evading aircraft to be equipped with a ultimately derived, and a statistics phase. The threat warning system in order to achieve a learning phase involves the development of a decision table that consists of a probability are then selected and trajectories dynamically distribution used in the selection of an ele- simulated. In a manner typical of Monte Carlo mental maneuver for each warning contingency. approaches, the averaged probability of kill and That table is shown in its initial form at the missile warhead detonation distance statistics upper right of Fig. 3. The column indices 1, are computed for each warning (or launch) cell.

. . . . . 5 under the control caption are the five elemental maneuver choices. The row indices, Elemental Maneuvers labled R, ranging fran 1, . . . . 84 represent the threat warning contingencies. Initially, the In this paper, the helicopter maneuver choice of maneuver for each contingency is choices are restricted to those which maintain a governed by sampling from the equally likely low constant altitude. The maneuver vectorgram, discrete distribution, as shown. labeled control set I in Fig. 4, is aimed at quantifying the impact of longitudinal and turn maneuver capability in constructing an effective LEARNING PHASE: OBTAIN OPTIMIZE0 STRATEGY evasive maneuvering strategy throughout the whole speed range from hover to maximum level of flight speed. At forward speed, the helicopter can command maximum transient (or sustained) load factor turns, labeled (1) and (5); maximum longitudinal acceleration, (2); or maximum longitudinal deceleration, (4); as well as maintaining the current speed and heading, (3).

At very low forward speeds including hover, the load factor turns are replaced with maximum rate pedal turns.

2. STORE: SEOUENCE IFLCI: 165.41 153.21. ; OUTCOME = MISS R ?

3. MODIFY DlClSlDNTASLE 0 6[ 0413-003P CONTROL SET I Figure 3. Stochastic Learning Method CCELERATION A random initial condition for the ccmbat is selected and both aircraft and threat trajecto- ries dynamically simulated. The aircraft employs a selected maneuver within the initial contin- gency cell until a second cell is entered and another maneuver choice is made. Threats may be launched outside the range of the warning space.

In this case, the aircraft maintains its current speed and heading until the threat first enters PE.DAL TURN the warning space at which time the control selection process begins. This simulation pro- cess is continued until warhead detonation or flyby, and a kill or survival event is calcu- lated using the probability of kill distribution derived frw the warhead lethality function. In the process of simulating the trajectories, the sequential contingency/control pairs employed by the aircraft are temporarily stored. Based upon the kill/survival event, the probability asso- ciated with those control choices made for each contingency are modified by a reinforcement rule. For the survival event, the probabilty of employing the same elemental maneuver for each stored contingency is increased, and is decreased for the kill event. The trajectory Figure 4. Elemental Maneuvers simulation and table modification process is repeated over all possible threat launch range and azimuth initial conditions using a random selection method. Approximately 100 launches per warning cell or 8400 total trajectories are numerically simulated to produce a converged decision table.

The 8400 trajectories require The maneuver vectorgram at the right in approximately 20 minutes CPU time on IBM 370/168 Fig. 4, captioned control set II, is aimed at systems.

quantifying the impact of lateral acceleration (sideward flight) and pedal turn capability in In the statistics phase the converged deci- constructing a maneuver strategy at or near sion table is fixed.

Random starting conditions hover speeds only. Choices (1) and (5) represent llllllI I IllI IIIII maximum performance pedal turns; choices (2) and missile against an optimally maneuvered evader.): (41, maximum performance lateral accelerations; Threat launches were initiated fras 72 of the 84 and choice (3) maintains current lateral speed range/azimuth cells within the 5 km maximum at the current aircraft heading. Similarly, range in both learning and statistics phases.

vertical or composite vertical/horizontal man- No launches were simulated from the 12 cells euver models can be constructed and investigated making up the inner range ring (range less than without change in the basic methodology. 0.25 km) due to severe missile guidance tran- sients at very short target ranges. It should be noted that in all results presented, the attack- Helicopter Maximum Maneuver Capability ing aircraft is assumed to maintain a speed Figure 5 graphically summarizes the sea equal to the initial speed of the target, and fly a pure pursuit navigation course toward the level maximum maneuver capability data associated with the elemental maneuver models of target during missile flyout.

Fig. 4, for a conceptual enhanced performance Figure 6 shows the kill effectiveness of the version of a current helicopter design. The maximum commanded turn capabilities shown at ATGM equipped with the expanding rod type war- head. Because of left-right symmetry considera- upper left are employed for choices (1) and (5) tions, only half of the warning space need be in control sets I and II. For the case of max- shown. Four levels of kill effectiveness (PE) imum transient turn, the associated longitudinal are given to simplify the presentation. The transient deceleration is shown at the upper legend at lower center is employed throughout right. The maximum longitudinal acceleration and deceleration capabilities utilized for this section. The origin of each semicircular plot corresponds to the helicopter position at choices (2) and (4) in control set I, are given missile launch, and the aircraft initial heading in the two lower diagrams. The lateral acceler- (0") is shown by the helicopter symbol. Head-on ation required for choices (2) and (4) of con- trol set II is given in the diagram at lower launches correspond to 0" to 30" aximuth sec- tors, and tail aspects launches 150" to 180", left. These studies employ first order models for the aircraft transient response to the respectively. The kill results are presented maximum acceleration and rate canmands. for four helicopter initial speed condition groups, beginning with hover at upper left, and progressing clockwise to maximum speed at the CURRENT DESIGN ENHANCED PERFORMANCE) lower left. Within each of the four speed TURN RATE TRANSlENT DECELERATION the left semicircle, labeled nonman- groups, euver, represents missile kill effectiveness when the aircraft maintains its current speed :‘;m jpy-j and heading. This case is important for quanti- fying target speed effects without maneuver, and is useful for establishing baseline survivabil- ity measures without use of threat warning and 0 100 200 0 100 optimal maneuver.

200 Clearly, a scan of the non- “ELOC’W - KN VELOCITY - KN maneuver cases for the four initial speeds indicates improving survivability in longer LONGlTUDlNAL & LATERAL ACCELERATlON LONGlTUDlNAL DECELERATlON range rear aspect launches with increasing speed, but at the expense of reduced surviva- - LONGlTUDlNALl ‘I bility in the corresponding forward launch cases. In addiition, a small window of improv- ing survivability for short range beam launch cases can be seen developing with increased speed; this is due to guidance transients associated with high line of sight rate targets.

0 100 200 0 100 200 The nonmaneuver cases show that speed alone VELOCITY - KN VELOCITY - KN (equivalent to no threat warning) does not 0413-005P provide sufficient survivability against the Figure 5. Maximum Maneuver Capability AIGM with Warhead A. The semicircles labeled OPT I in each of the four speed groups quanti- Effectiveness of Maneuverability fies the survivability improvements that can be achieved with the 84 cell warning system, The aircraft survivability or equivalently together with an optimal maneuvering strategy the missile kill effectiveness results (PE) derived from control set I. In the four results for the ATGM threat for all launch conditions labeled OPT I, the helicopter employed its are calculated and presented in the helicopter maximum transient load factor turn performance warning space coordinate system for convenience. for choices (1) and (5). One can see that survivability is still poor with combat initi- In this case the maximum effective launch range of the threat (4 km) was less than the maximum ated at hover, although small improvements exist detectable range of the warning system (5 km). for tail launches at the 4 km range. This is (The results could also be presented in a space due to helicopter acceleration away from the relative to the launch aircraft and would rep- oncoming missile and the missile maximum range resent the effective launch envelope for that limitation. However, at higher initial speeds, optimal maneuvering, employing transient load optimal strategy development.

The nonmaneuver factor performance can provide high surviva- and optimal survivability results for Warhead A bility. The lack of effectiveness of control are repeated at lower left. Corresponding sur- set II (lateral acceleration and pedal turns) in vivability results for the contact fuzed warhead constructing an optimal maneuver strategy from are shown upper center; those for Warhead B are hover is shown by the shaded semi-circle labeled shown at the lower right. The norrmaneuver OFT II. This result, together with that for OPT results are statistically equivalent in all I to the immediate left, indicate the low sur- cases and typify the small miss distances vivability afforded by maneuver against the ATGM achievable by the missile guidance system with Warhead A at hover flight speeds. against constant velocity targets. The heli- copter can be made completely survivable against The sensitivity of survivability of the en- the contact fuzed ATGM using optimal maneuvering hanced performance helicopter to variations in at this initial aircraft speed. However, the ATGM warhead type and lethality is shown in corresponding result for Warhead B indicates The three warhead types: contact, that optimal maneuver would be completely Fig. 7.

proximity Warhead A, and proximity Warhead B, ineffective. These results indicate the strong have been examined at the helicopter minimum interplay between missile warhead lethality and power required initial speed. The helicopter guidance, and the need for carefully timed deployment of the aircraft's maximum maneuver -employs control set I with maximum transient capability to generate adequate miss distances turns for elemental maneuvers (1) and (5) in the against this threat.

AIR-TO-AIR THREAT/WARHEAD A NONMANEUVER OPT I OPTII NONMANEUVER OPT I HOVER MIN POWER REO’OSPEED NONMANEUVER OPT I NONMANEUVER OPT I TRANSIENTTURNS m l.0>FK>0.9~ m 0.9 > PK > p.0 m 0.5 > P< 5 0.3 0 0.3 > PK 5 0.0 II MAX SPEED MAX RANGE SPEED Figure 6. Helicopter Survivability AIR TO AIR THREAT, WARHEAD COMPARlSONS NONMANEUVER OPT I . MIN POWER REQUIRED SPEED . TRANSIENT TURNS WARNING 4 SPACE 3 CONTACT NONMANEUVER OPT I NONMANEUVER OPT I WARHEADS WARHEAD A 0413-007P Figure 7. Hsl&optsr Survivability Three optimal evasive trajectories from Visual Model hover using maneuver set I against the contact fused warhead are shown in Fig. 8. The surviv- The visual model employed in the gun combat ability results for nonmaneuver and optimal / studies is displayed in Fig. 9. Each combatant maneuver are presented at the upper left of the is assumed to have a visual contact volume figure. For each case illustrated, only the extending to a maximum range of visual detecta- bility. Within this volume each combatant is terminal portion of the missile path and the entire helicopter path are shown because of permitted to select a maximum performance tac- scale effects. The head-on case at upper right tical maneuvering strategy for flight path and beam aspect case at lower right illustrate control of the aircraft. For these studies the pedal turns immediately following launch, maximum range has been arbitrarily set at 3 km for both combatants.

followed by straight accelerated flight and This is consistent with a maximum performance load factor turn line of sight visual capabilities at low alti- finally, near termination. The tail aspect launch at tudes in typical rolling terrains. Aircraft lower left employs only the acceleration segment size, paint/camouflage, and background contrast followed by the load factor turn at termination. factors have been neglected. A helmet mounted the aircraft maneuvers to sight operational tracking volume associated In all cases shown, with a turreted gun fire control system is also achieve a tail aspect to present its minimal fuselage envelope dimension at missile flyby. considered as illustrated. Gun firing opportu- Launches within 2 km cannot be made highly nities exist only when the target is within the survivable because the missile flight time tracking volume limits.

termination is too short to permit adequate HELMET forward acceleration and load factor turn SIGHT maneuvers to avoid fuselage hits.

TRACKING VOLUME CONTACT WARHEAD 3-km I I I MAX - RANGE *o---.

.’ #’ /’ e<-- .,- ---(‘\,, ,ij :,,;,, - -jTjg .

MAXIMUM MANEUVER TACTICAL VOLUME 0413009P Figure 9. Visual Model The maximum maneuver volume of each combat- ant is decomposed into a finite set of tactical contigencies by an assignment of thresholds involving the relative positions, velocities, and other observables during the combat. For the constant altitude maneuvering model, each Figure 6. Evasive Maneuvers (From Hover) combatant is assumed to measure relative range, angle off and relative heading as depicted in Fig. 10. Relative range has been divided into 5 Maneuverability in Air-to-Air Gun Combat cells fran zero to 3 km; angle-off into eight 45" sectors from 0" around the compass to 360"; This section concerns quantifying the impact and relative heading divided into the four quad- of aircraft maneuverability, gun capability and rants as shown in Fig. 10. These thresholds ballistic hardening in air-to-air visual range divide the maximum maneuver volume into 160 con- gun combat effectiveness. Three blue (friendly) tingencies for the constant altitude combat helicopter design concepts are separately case.

evaluated against the same red (threat) heli- The first blue aircraft, called the copter.

baseline, is representative of a current opera- Gun Model tional attack helicopter design, and the second, Both blue and red aircraft are assumed to be an advanced light helicopter concept (LHX) hav- equipped with a turreted gun with target track- ing greater maximum maneuver capability and The third concept aircraft ing accomplished by a helmet mounted sight.

level flight speed.

Fire control lead prediction employing target is a variant of the second; employing equivalent range, range rate, angular rate in flight data maneuverability but with improved ballistic together with specific projectile ballistics is hardening.

considered in the armament simulation. Depend- Table I for that threat projectile. The areas have been normalized by the numerical value of ing upon the gun and projectile, a firing oppor- tunity requires satisfaction of the following: the vulnerable area in the side aspect for the target entry into the tracking volume; a baseline aircraft. The N-shot burst probability.

"pipper" settle time delay associated with of kill for each combatant is employed at each and target range within a prespecified entry; step in the trajectory numerical integration maximum firing range. process to determine the termination event; kill by red, kill by blue, mutual kill, and no kill by either.

RELATIVE HEADING BLUE Table 1. Aircraft Relative Vulnerability RED RANGE km 1.0 0.50 0.25 BOTTOM TO BLUE .64 .59 .71 .72 .26 .26 .47 .77 I I I I I I OSSERVABLES # THRESHOLDS . RELATIVE RANGE (5) CONSTANT ALTITUDE CASE . ANGLEOFF (81 160 CONTINGENCIES . RELATIVE HEADING (41 Maneuver Strategy Development . RELATIVE SPEED (3) * . RELATWE ALTITUDE (3) The constant altitude maneuver strategy for both combatants employs the elemental maneuver 0413.01OP set labled "control set I" in Fig. 4. The as- Figure 10. Maximum Maneuver Volume Thresholds sociated maximum maneuver capabilities of the The probability of kill associated with an blue and red aircraft are summarized in Fig. 11.

N-shot gun burst is developed from single shot The transient response of all combatants to max- imum canmanded rates or accelerations is repre- considerations as follows: sented by a family of first order models as shown at the lower right of Fig. 11. The time , SINGLE SHOT constant associated with longitudinal commands is given by TAUPIT; load factor turn cam-sands by TAUROL; and pedal turn commands by TAUYAW.

dy2 AV ‘KSS = TERMY I 2r(TERMX’TERMY)“2 ‘exp Each combatant's maneuvering strategy is repre- sented by a choice of an elemental maneuver for each contingency cell of the maximum maneuver 2 AV 0 TERMX=o; volume shown in Fig. 10. The stochastic learn- *‘TX +z ing methodology is easily extended to the two 2 AV canbatant case as depicted in Fig. 12. In con- 0 TERMY =oD +OGy+ 1;; trast to the single decision table learning phase described in the missile avoidance appli- 0 AV GIVEN FOR SPECIFIC VIEWS cation of Fig. 3, a blue and red decision table are now sequentially modified to produce optimal N-SHOT BURST maneuver strategies for both combatants.

= 1 - (&PK )N Helicopter/Armament System Combat Effectiveness

pKN

SS The oneon-one gun canbat problem requires In the above CD is the dispersion error of that one determine the domains of combat initial projectile; (TTX, (JTy the casposite target conditions (positions, velocities) for which tracking errors in x, y coordinates; and Av each of the combatants has a unilateral capabil- the ballistic vulnerability of the aircraft to ity in deciding the outcane of the canbat. The the threat projectile (measured in terms of comparative size of these domains furnishes a vulnerable area). Other N-shot vulnerability 'quantitative measure of superiority of one air- models (such as the salvo fire model) can easily craft/armament system over the other. To deter- have been employed in these studies without mine these domains, the computational method is alteration of the basic methodology but are not' first employed with each side maximizing his reported here. In the caaputational results to kill probability, and secondly, with one canbat- follow both blue and red aircraft were assumed ant maximizing kill probability with the other equipped with a 25 mm gun. The respective maximizing survivability. These separate vulnerabilities of the aircraft are given in solutions determine domains where each vehicle TRANSIENT TURN RATE TRANSIENT DECEL SUSTAINED TURN & PEDAL RATE TRANSIENT RESPONSE (ALL AIRCRAFT) LONGITUDINAL ACCEL LONGITUDINAL DECEL Figure 11. Sea Level Maximum Maneuver Performance is best operated offensively, and where each centered on the blue combatant as shown in should operate defensively with survivability as Fig. 13. The probabilities of kill for each corn- the main goal. batant and other important terminal statistics are computed for each discretized initial con- dition region as shown.

DlscRETlZED 1 lN,T,Al. CONDlTlON ’ SPACE RESULTS . EVENT PROBABlLlTlES . AVERAGE SHOTS BY BLUE a RED . AVERATE TERMlNATlON TlME Figure 13. Format for the Computational Resdts 0413.012P Two representative canputational solutions Figure 12. Maneuvering Strategy Development employing the initial condition polar format of Fig. 13 are given in Figs. 14 and 15.

The opti- Each of the offensive/defensive canputa- mal solution in Fig. 14 considers the case of tional results emerging from the stochastic the blue LHX aircraft in an unarmed defensive learning solution methodology is presented in role against an offensive red adversary equipped with a 25 mm, 1500 spm turreted gun.

terms of a discretized initial condition space This solu- tion considers combat initial speeds of 87 kn for both combatants with both helicopters employ- ing sustained turn for their load factor turn elemental maneuver choices. The four half-polar charts (due to initial condition symmetry) give the probability of kill for red in terms of relative range, angle-off, and relative heading.

The result at upper right corresponds to the coincident heading case, as schematically repre- sented by the B and R vectors in small auxiliary The remaining three heading cases are diagram.

interpreted with the aid of the rotated R vector in the auxiliary diagrams. The cells of high kill probability for Red (PKR) are shaded according to the accompanying legend. The solu- tion in Fig. 15 considers the LRK in the offen- sive role against an offensive red adversary for the same initial speed case of 87 kn. TheLRK is equipped with identical turreted gun armament and fire control as the red helicopter. The initial condition cells of high kill effective- ness are shown for each combatant using the legend as indicated.

pKB' pKR Figure 15. LHX Offensive Solution, 87 kn e LHX OFFENSIVE SOLUTION R BLUE AVG SHOTS/KILL 0413-014P / Figure 14. LHX Defensive Solution, 87 kn % d BLUE A more ccmpact bar summary format enabling convenient combat effectiveness canparisons between helicopter/armament systems is shown in Fig. 16. The total of high kill and mutual kill PERCENT INITIAL area for both combatants as a percent of total AREA area within a fixed radius of initial conditions DOMINATED for Fig. 15 is now plotted on the vertical scale at the right. The fixed radius is taken as 1.5 km representative of ranges associated with change encounter initiation of helicopter c engagements. The data shown in the circles at % top and bottom of the bar graph indicate the RED average shots/kill achieved by each combatant in the high kill and mutual kill areas. The per- RED-AVG cent of total initial area dominated by each SHOTS/KILL 0413.016P cwbatant is a quantitative measure of his combat effectiveness or air superiority.

Figure 16. Bar Summary Format

llllllllllllllllllllllllllll II I lllll III

RED 1 1.5 km RADIUS 1 + NO PERFORMANCE PENALTY FOR “R *DEFENSIVE ROLE Figure 17. Helicopter/Armament Combat Effectiveness Comparisons smaller asnsunition expenditure obtained by the A comparison of combat effectiveness of the baseline and LHX aircraft including variations assumed 6 mil error fire control tracking error in ballistic hardening, gun mount, and shot rate performance of a limited sweep HUD system.

characteristics is shown in Fig. 17. All solu- tions shown are for combat initiated at 87 kn The fifth column corresponds to the base- line helicopter equipped with the same turreted for both combatants with maximum sustained turn gun as the red adversary. The % area ratio for capability employed as the load factor turn elemental maneuver choice. The first three bar measuring dominance is nearly 1:l against red.

The sixth coluum shows the LHX capability with graphs (from the left) correspond to defensive solutions for various blue helicopters against the same turreted gun against the red threat.

the offensive red adversary. The red threat The gain in combat effectiveness of the higher employs a 25 mm, 1500 spm, turreted gun with maneuverability LHX compared to the baseline is appreciable, *90 degree azimuth capability, and +21" and but is somewhat offset by the higher ballistic vulnerability of the LHX.

-50 elevation capability. The reduction in red Coluam seven quantifies the gains achievable by kill effectiveness achieved by the more maneu- the LHX if superelevation of the turreted gun to verable LHX and LHX with ballistic hardening can be canpared with the baseline aircraft. +50" were permissible (rather than +21" because of rotor clearance). Bar graph eight illus- Bar graphs four through nine consider var- trates the impact of ballistic hardening ious blue aircraft/armamsnt configurations in an improvements to the turreted gun LHX.

Compari- offensive role against the offensive red threat son with the standard LHX results in column six In the fourth case, previously described. indicates an applicable reduction in the kill labeled (LHX/FLEX) the LHX aircraft was equipped effectiveness area of red while improving the % with a limited sweep (*6" elevation and azi- area of highest kill probability (9% improved to muth) gun mount. The caaposite tracking error The last column on the right illustrates 27%).

in this case was assumed to be (7TX = 6 mil and the high canbat effectiveness achieved with a The gun the projectile dispersion UB = 5 mil. 3000 spm turret gun equipped LKX design incor- caliber and shot rate were assumed equivalent to porating ballistic hardening.

These results that employed by the threat. (Note: for all 25 illustrate the significant interdependence of mm turreted gun applications, both blue and red, maneuverability, armament, and ballistic hard- the canposite tracking error was assumed to be ening factors for friendly and threat heli- u TX = 20 mil. and the dispersion UB = 5 copters that enter the canbat effectiveness mil). The low shots/kill by blue reflects the evaluation.

listic hardening, and armament capability can be LHX Maneuver Effectiveness considered in design studies and that cwbat effectiveness assessments can be accomplished In the design concept phase, it is often with reasonable cusputer time budgets.

important to quantify the sensitivity of canbat effectiveness to maneuver parameter variations Although the results show that canbat on a one at a time basis while holding other effectiveness is strongly dependent upon the aircraft and armament parameters fixed. As an integrated use of the above factors, a maneuver example of this, the original sea level maximum capability advantage can provide sizable gains longitudinal acceleration parameter of the LHX in survivability in the defensive role and kill (labeled LHA A in Fig. 18) was enhanced to that The effectiveness in the offensive role.

given by the function labeled LHX B.

All other results presented here have canbat maneuvers maneuver, ballistic hardening, and armament limited to constant altitude, however, similar parameters were held fixed. The corresponding ccmputational models which include vertical improvement in canbat effectiveness for the blue plane maneuvering are currently under investiga- offensive/red offensive case for the 87 kn ini- tion with results available in the near future.

tial speed is shown in Fig. 19.

The bar graph on the left is the result originally obtained Acknowledgments for the LHX turret case first illustrated in Fig. 17.

The authors would like to thank the fol- .a ./-.\ lowing personnel and their organizations for their technical contributions in the development / ‘\ of the helicopter/armament system computer mod- .6 /* 1 ‘\ els utilized in these studies: Mr. J. Means and Dr. L. Feaster; U.S. Army AVRADCOM, St. Louis, MO

~s,:~~,.~

Mr. E. V. Merritt and Mr. J. Anderson; U.S.

Army AVRADCOM, Ft. Eustis, VA

& 100 250 Mr. P. Townsend and Mr. T. Hung; U.S. Army

150 200 ARRADCOM, Dover, N.J.

VELOCITY, KN 041J-ol*P Mr. R. Bruce and J. Wagner; General Elec- Figure 18. Sea Level Maximum Longitudinal Acceleration Variation tric Co., Burlington, VT Mr. F. Sobierajski and Mr. H. Hinz; Grumman Aerospace Corp., Bethpage, N.Y.

REFERENCES Differential Games, John Wiley, 1. Issacs, R., New York, 1965 2. Breakwell, J.V. and Merz, A.W., "Toward a Complete Solution of the Homicidal Chauffeur Game", Proceedings of the First Interna- tional Conference on Theory and Applica- tions of Differential Games, Amherst, Mass., October 1969 3. Baron, S., Kleinman, D., Serben, S., "A Study of the Markov Game Approach to Tacti- cal Maneuvering Problems", NASA Report CR-1979, February 1972 0419-019p Il.5 km SUMMARY 1 Figure 19. Sensitivity of Combat Effectiveness 4. Falco, M., Cohen, V., "Strategy Synthesis in Aerial Dogfight Game Models", AIAA 11th Aer- Conclusions ospace Sciences Meeting, Washington, D.C., AIAA Paper 73-233, January 1973 This paper has sketched the development and npplication of a digital simulation technique 5. Falco, M., and Carpenter, G., "Survivability incorporating optimization and game theory con- Analysis of Air and Land Vehicles to Missile cepts for assessment of combat helicopter man- Threats", 4th Symposium on Vulnerability and euver effectiveness in the one-on-one setting. Survivability, American Defense Preparedness The numerical experience to date suggests that a Association, Tyndall AFB, Florida, March respectable amount of detail regarding the 1979, (and Grumman Research Department integrated use of maneuver, threat warning, bal- Melaorandum).

FLIGHT TESTS FOR THE ASSESSMENT OF TASK PERFORMANCE AND CONTROL ACTIVITY Heinz-Jiirgen Pausder and Dieter Hummes Deutsche Forschungs- und Versuchsanstalt fiir Luft- und Raumfahrt e.V. (DFVLR) Institut fiir Flugmechanik, Braunschweig, FRG of simulation and flight test results per- Abstract formed in the Fifties. The development in Helicopter flight tests were conduc- the field of helicopter design which hap- ted to look at the influences of pilot and pened during the last two decades has not helicopter system on the performance in been taken into consideration. Implementing NOE-flying. A 'visual dolphin course was rotors with non-hinged blades and divergent built up. The tests were performed with the flying characteristics of the helicopter helicopters BO 105 and UH-1D. Closely con- pointed out that the application of the netted with tactical demands the six test specifications can vield wrong conclusions..

pilots' task was to minimize the time and Although the requirements could not be met, the altitude over the obstacles. The data acceptable flying qualities have been in- reduction yields statistical evaluation ferred from pilot evaluations.* parameters describing the control activity of the pilots and the achieved task perfor- As a consequence of these discrepancies mance. The results are shown in form of several attempts have been made to revise evaluation diagrams. Additionally dolphin the specifications. The specification MIL- tests with varied control strategy were F-833003 was published for VISTOL-systems.

performed to get more insight into the in- Closely following the structure and format fluence of control techniques. From these of the specification for conventional air- craft,the different demands of maneuvers are test results recommendations can be derived to emphasize the direct force control and considered by coordinating requirements to to reduce the collective to pitch cross- three categories of flight phases. MIL-F- coupling for the dolphin.

83300 attempted to include the helicopter systems,but the essential criticism of its Introduction application for helicopters implies that the specific problems of helicopter flying The military or civil user of a heli- qualities and helicopter missions are not copter primarily emphasizes the demand for taken into account sufficiently.

a successful completion of his special mis- sion. This requirement is an unsufficiently The application of helicopters are greatly extended. This includes the expan- defined basis for the design efforts of a helicopter system. TherefoFe the need ex- sion of mission types, and the specific de- ists to transform tactical demands in mands of mission phases. In Fig. 1 typical standards of technical terminology. With mission parts are skeletonized as they are such diagnostic tool the contractor is thus presently being discussed in the F.R.G.: enabled to check the achieved adaptation of the overall requirement of adequate mis- 1) Anti tank.

sion performance during the design phase. 2) Combat rescue.

Once the helicopter has been built, the 3) Air to air.

question about the qualities of the flying characteristics is also asked. If it doesn't The mission phases can be characteri- meet the mission demands, the question is: zed by the required low altitude above Why not? Consequently flying quality spe- ground and by the flight speed. Demanding cifications should contain technical scales of the pilot/helicopter system to use the terrain as cover against exposure and to for a quantitative evaluation and they' should set the standards for 1) checking obtain superiority in direct contact with the tactical missions demands, 2) transfor- the enemy high agility is required of heli- ming these demands in measurable data, 3) copter system. Current specifications for proving the efforts during the design phase, flying qualities regarding the demands of and 'I) a quantitative evaluation during the such specific missions phases don't exist.

certification Dhase. Alternativelv an assessment of adequate flying characteristics can be performed due Nevertheless, the specification MIL- to pilot evaluations given in flight or si- H-8501 A1 is valid still nowadays. The ba- mulation tests. But this approach doesn't sis of these requirements was constituted get rid of the urgent problem of formulat- _ ing requirements for fiying qualities in parameters. Those can be registered by mea- surements and can be applied to a certifi- cation as evaluation scales.

I llllllllllllllllllll1llll I I I lllllllll FLIGHT TEST ft COMBAT RESCUE ALTITUDE - A/C STATE "ARlABLES I I - PILOT CONTROLHOYEHENTS - A,C POSITlO,, "ARIABUS (SUBJECTIVE DATA) AL1 DATA ANALYSIS . COLLECTlON OF PARAMETERS RELLYANT TO EVALUATlON I I 100 kt 200 Fig. 2 Evaluation approach SPEED evaluation. The statistical analysis of the Fig. 1 Helicopter missions test data calls for a definition of the flight test task that allows a reproducable test performance. This includes clear in- Primarily, the objective of flight structions of task conditions and test ob- mechanical investigations is to constitute jectives for the test pilot. On the other a data base for deriving recommendations for flying quality criteria. The different hand? the test conditions should as much as tactical mission demands of today require possible correspond to the realistic condi- a mission- and task-oriented adaptation of tions of the selected mission element. In the evaluation. In several institutions order to obtain a broad data base, the test configurations are varied. This can be a- activities exist with this objentive.4,5 chieved by variation of: At the Institute for Flight Mechanics of the DFVLR a technical approach was also de- veloped with the overall objectives as fol- 1) Helicopter characteristics.

lows: 1) to investigate task-oriented fly- 2) Environmental conditions.

ing qualities; 2) to constitute a flying qualities data base for the assessment of In order to register and eliminate the in- quantitative requirements for helicopter dividuality of the pilots, several pilots should be engaged in the tests.

systems.

DFVLR Evaluation Method The following data are measured and recorded: 1) state variables of the heli- In this paper a brief review of the copter, 2) control inputs of the pilot, and DFVLR evaluation method shall be given. 3) position variables of the helicopter.

More details are presented in Ref. 6. The One test is composed of a number of runs ulterior motive is the correlation of pi- with equal conditions. Additionally the pi- lot ratings from flight tests with para- lot ratings and comments are collected of meters obtained by a statistical analysis each test. The applied rating scale corre- of measured data. The statistical parame- sponds to the Cooper-Harper scale', but it ters, which include a good correlation with is slightly modified by dividing them into three groups of questions referring to the the ratings, are collected in the data ba- se. workload of the pilot? the task performance and the handling qualities of the helicop- Fig. 2 shows the general approach. ter. This yields a redundancy of the rat- Starting from a given mission, elements ings. By adding the pilot comments the in- are selected which are representative for terpretation of the ratings is facilitated more.

this mission and which include the demands which are critical for a flight mechanical MISSION 1 I PILOT ! COCKPIT ! ztI;;;TROL 1 ENVIRONMENT / sWSX’;~~~C~~.

In a detailed statistical analysis the and technically displayed information.

measured data are then reduced to parame- To define the conditions of flight ters characterizing the dynamic behaviour of the closed-loop system. By correlating testing and to interpret the test results, the pilot ratings and comments with the those many influence factors have to be statistical parameters the relevance to the taken into consideration. This involves evaluation is being checked. Parameters 1) a well defined test task, 2) clear pi- lot instructions, 3) a qualified selection with high correlation to the ratings are collected in the flying qualities data ba- of test pilots, 4) well defined environ- se. mental conditions, and 5) a definition of helicopter state and pilot information.

Influence Factors Description of Experiments ~----- The task performance and the control activity depends on various influences, Evaluation Task described in the pilot-in-the-loop diagram of Fig. 3. Proceeding from the defined The starting point for the DFVLR stu- flight task, dies was the German anti tank mission. This the elements of the loop with partly time dependent characteristics are mission includes phases with high portions of precise hovering, quick stop maneuvers passed through to get the task performance as the result. and flying near the ground making use of The main influence factors result form the pilot and the helicopter the terrain as a cover. With close refer- ence to the tactical demands, the DFVLR system.

evaluation tasks were defined (Fig. 4).

The pilot transfers the task instruc- First studies were conducted with a hover- tions in a conception of an adequate task ing tracking task.8,g performance as reference for the system For the dolphin task a course was performance achieved in the test. The human pilot adapts to the task, the characteris- built with two obstacles, as shown in Fig.

tics of the helicopter and the subsystems, 5. The distance between the obstacles was 350 m. The run started 200 m before the and the environmental state by means of a control strategy appearing optimal to him. first obstacle and ended 200 m behind the The obstacles had an al- In doing so, he profits his high capability second obstacle.

titude of 15 m. They were built up as to of adaptation. The handling qualities of the helicopter system represent the limi- put the pilot into a - as much as possi- ble - realistic situation which he has to tations for the pilots adaptation. The feedback to the pilot with different types For safety aspects the last 3 m deal with.

of information includes influences inten- are consist of bushes. The centerline of sifying with extreme environmental states the course is marked on the ground to fa- Flight Tests The flight tests were conducted with the helicopters BO 105 of the DFVLR and UH-1D of the German Forces Flight Test Cen- ter (Fig. 6). Different test configurations were achieved by varying the test parame- ters s@eed and gross weight. Table 1 shows the test matrix for the dolphin test with the BO 105.

Table 1.

Test configurations dolphin Gross weight Speed G G min max 40 kt X X X 60 kt X 80 kt X X 100 kt X X Fig. 4 Definition of evaluation tasks In order to obtain a broad spectrumof pilot's behaviour, six test pilots werein- volved (3 of the Flight Test Center, 1 of the DFVLR and 2 tactical test pilots of the Army). The control strategy tests were per- formed with both helicopters and threetest pilots. Table 2 reviews the configurations.

Table 2. Test configurations control strategy BO 105/UH-1D (Gmin) Fig. 5 Dolphin course Comb. Cyclic Collect.

Speed control control control cilitate the heading for the pilots. In the dolphin tests, the pilots are instructed 40 kt X X X to traverse the course while minimizing thetime and the altitude over the obstac- 60 kt X X X les. The altitude of entering and finish- ing the course is 15 ft over ground and 80 kt X X X the pilots have to align the helicopter on 100 kt X X X the 15 ft altitude between the obstacles if possible. The speed was defined in the test conditions and had to be flown in the beginning, between the obstacles and at The testing procedure was always the the end of the course. Deviations of head- same to guarantee reproducable test results.

lateral position, and bank attitude After explaining the objectives the task iw , was to be avoided. instructions were given to the pilot. Al- though all pilots were experienced in NOE- To study the influence of control stra- sufficient time was given to them flying, tegythe task is slightly varied. The pi- to train the course. Subsequently, each lotshave to perform the course at firstwith test was conducted by flying seven isolated a control combination of the pilot's con- runs of the same kind. A quicklook was in- secondly with using only stick in- stalled in the ground station to control =ept , puts and thirdly using primarily collec- the training and the test.

tive inputs and the longitudinal cyclic for minimizing the pitch attitudes due to Data Acquisition coupling.

After each test, the pilot had to an- swer a questionnaire relating only to the test performed. The questions concerned the Fig. 6 Test helicopters pilot workload, the quality of task per- formance and the handling qualities. In ad- ditionto that the pilots commented on the reasons for their ratings. The data acqui- sition was provided by an analog magnetic tape recording in the ground station. Re- corded variables included control inputs, attitudes, rates, accelerations, air speed, altitude, torque and rotor speed. The heli- copter position data relativ to the obsta- cles was measured by a laser position tracking system and was recorded time syn- chronized with the helicopter state and control data. To register these data in the helicopter and to transmit them to the ground a programmable multipurpose instru- mentation system was used.1° The concept made it possible to reach a quick adapta- tion to the test technique (helicopter direction of flight). The data were type, digitized online in the ground station and were available for data analyzing, sampled with a frequency of 20 Hz.

Discussion of Dolphin Results Fig. 7 Pi .ot ratings of dolphin tests correlation between the ratings for the The pilot ratings of the task perfor- workload and the task performance. The pi- mance and handling qualities are compared lot will give high ratings, if he doesn't with the ratings of workload as shown in come up to the wanted system performance Fig. 7. There are clear differences of up in spite of high effort.

to three points between the ratings eva- luating the handling qualities and the The information content of the signal workload with the tendency to give the data are summarized for each run by the handling qualities a better rating. Apply- statistical analysis. The run parameters ing the original Cooper-Harper scale a are checked with a confidential test and close relation is suggested between thede- averaged for each flight test. By correlat- mand on the pilot and the aircraft charac- ing the statistical parameters with the teristics.Indeed, the precondition for this ratings of the pilot's workload, the eva- assumption is observing exactly a tenta- luation parameters are determined. They are tively defined task performance. In more as follows for the dolphin tests: complicated flight tasks as the dolphin the pilots yield a variation of perform&g 1) Sum of the standard deviations of the task. In this way, the ratings forthe longitudinal and collective control inputs handling qualities are influenced too.

(control activity).

This behaviour also accounts for the good

Ill Ill Ill Ill I I I I

Adapting the dolphin instructions, the 2) Time integral of altitude over the pilots choose a control strategy characte- obstacles (evaluation area).

rized by a combination of the main controls 3) Peak-to-peak value of pitch atti- longitudinal cyclic and collective. Devia- tude.

4) Peak-to-peak value of acceleration tions in state and position due to coupling are compensated with the other controls.

in z-direction.

Fig. 8 shows the control strategy depending on the pilot. Pilot 1 tends to a relative Table 3. Correlation of parameters high activity in the collective, while pi- with pilot ratings for workload lot 2 flies the dolphin with high longitu- dinal control inputs for example. Fig. 9 points out the tendencies of standard de- viations of controls with speed. The pi- Pilot Evaluation lot's behaviour in the control strategy parameter 1 2 3 4 5 6 illustrates the broad spectrum of pilot adaptation.

Evaluation .72 .ll .60 .56 --"-79 .63 The influence of the pilots also af- area fects the resulting system performance.

Peak to peak .62 .89 .22 .20 .87 .20 The levels of accelerations obtained inthe tests have a linear dependency on thespeed.

pitch attit.

The peak-to-peak of pitch attitude and the evaluation area parameter includes the di- Peak to peak .57 .78 .74 .86 .96 .84 vergent weighting of the pilots for flying z-acceler.

the dolphin. While some pilots keep theeva- luation area constant this parameter in- Control .51 .48 .IO .67 .85 .03 creases with speed for the others. The pi- activity lots influence upon the pitch attitude va- lues seems to be still higher. Depending on the pilot the attitude level increases, Table 3 shows the correlation coefficients. decreases or keeps constant.

Divergences exist between the pilots. The The test results of all pilots can be reasons for low correlation values are the summarized in the evaluation diagrams. The pilot hasn't altered the parameter with relation between the control activity and speed and gross weight and/or the relation- ship between parameter and rating isn't the evaluation area as the main task per- linear. formance parameter is shown in Fig. 10. The pilot ratings describe clear tendencies for the defined evaluation scales. Accordingly evaluation boundaries are inserted in the diagrams. The ratings deteriorates with in- creasing evaluation area and the pilots have experienced a higher workload. With area values over the additionally drawn boundary the pilots have substantiated their workload with minimizing the time and alti- tude over the obstacles. Relative to the control activity there exists an evaluation optimum. A relation between evaluation and the separated activity in the controls can't The causes for the disor- be constituted.

ientation of some test results in Fig. 10 are due to the values of other parameters (see Fig. 11 and 12). With higher levels of pitch attitude the workload of the pi- lots increases, because high attitudesren- der more difficult the orientation of the pilots in the course. The reasons for the relative good ratings of pitch levels I I higher than 40 deg are high accelerations.

0.05 0.1 tllm,lllm STANDARD DE”lATlON LONG. SilCK CONTROL Moreover the pilots mentioned acceleration levels over 1.6g as a reason of theirwork- load in the comments. But also lower accel- Fig. 8 Control strategies of pilots erations influence the pilot ratings work- load.

Summarizing the evaiuation diagrams characteristics of the helicopter systems Fig. 9 Tendencies of evaluation parameters Fig. 10 Pilot rating trends with control activity and evaluation area ’ A ,,,,fl,,,,,,,,, L4 m ,,1,111~1~.,, fi ’ #’ .,J--- 6’ I O- 0.2 0.1 n-m,"" CONTROL ACTIVITY PILOT RATINGS WORKLOAD PILOTS COMMENT 0' OL 40 0 20 1 2 0 dw B PEAK TO PEAK OF Z-ACCELERATION PEAK TO PEAK OF PITCH ATTITUDE Fig. 12 Pilot rating trends with Fig. 11 Pilot rating trends with pitch attitude and evaluation area z-acceleration and evaluation area can be recommended related to the flown tegies: Independent of the helicopter state, task.

the fulfillment of the dolphin task should 1) Combination of longitudinal stick be possible for a satisfactory evaluation: and pitch control.

1) with low altitudes and short times above 2) Stick control.

the obstacles (evaluation area lower than 3) Primarily pitch control.

3 mmsec), 2) with a maximum level of z- As an example Fig. 13 shows test data of acceleration of + 0.45 g, and 3) with a low level of pitch attitude. For the nec- the BO 105. The curves give an impression essary control activity an optimum exists, of the signal contents in amplitude and but a separated influence of longitudinal dynamics. The cross correlation of control cyclic and collective controls can't be and a/c state signals is evident.

achieved by the test results.

The realization of the defined strate- Discussion of Control Strategy Results gies by the pilots is skeletonized in Fig.

.- 14 and 15. All test configurations were The objective of the additionally con- feasible for the pilots and were accepted ducted control strategy tests is to assess after sufficient time of exercise with the the influence of different control combi- exception of the 40 kt stick configuration nations on the task performance in the of the UH-1D. When comparing the helicopters dolphin. More than that the NOE-flying it was noted that in the UH-1D tests the must give an answer to the questionwhether pilots used higher collective inputs, es- a moment control or a force control of pecially for the stick configuration. The helicopter offers the better technique to UH-1D requires a lower stick activity on fly closely over obstacles. As mentioned account of the lower collective to pitch above the task includes three controlstra- cross-coupling. For the BO 105 the coupling Fig. 15 Control strategy verification (UH-1~) 0.2 is nearly doubled with the speed from 40to the mission. This technique has been ap- 100 kt. Two pilots tried to compensate this plied to a dolphin task that is derived behaviour with appropiate longitudinal cy- from the German Anti Tank Helicopter mis- sion. From the test results the following clic inputs.

general tendencies and conclusions are To evaluate the achieved task perfor- noted: mance the resulting parameters are drawn versus the ratio of the standard deviations 1) The described method leads to an for collective and longitudinal cyclic. acceptable assessment of task performance Fig. 16 and 17 summarize the test results. and control activity.

All parameters show quite the same tenden- 2) The combinations of the parameters cies with increasing of control ratio ex- yield flying qualities recommendations for cept the deteriorating values of pitch helicopters related to the dolphin task.

attitude and evaluation area for extreme 3) The parameters for a quantitative high pitch to stick control ratio for the evaluation are the evaluation area, the level of pitch attitude, the level of ver- BO 105 due to coupling.

tical acceleration, and the activity in To derive recommendations from these longitudinal and collective control.

results for the helicopter system the dol- 4) With emphasis on collective control phin task can be performed in a better way (direct force control) the dolphin task with emphasis on collective control that can be performed in a better way. There- means direct force control consequently. fore a low collective to pitch cross-coup- ling is necessary.

Naturally a decrease of the collective to pitch cross-coupling is assumed. This can be mainly achieved by the design compro- References mise of the rotor system or by additional feedback systems. Carrying out the dolphin 1. "Helicopter Flying and Ground Handling task with high control moment capacity is Qualities; General Requirements for not adequate, but produces relativehigher Military Specification", MIL-H-8501 values of attitude, acceleration, and eva- A, Sept. 1961.

luation area. For the application and a- daptation of a direct force control addi- 2. Reichert, G., and Delker. P., "Hand- tional studies including engine dynamics ling Qualities with the Bolkow Ri- have to be performed. gid Rotor System", Paper No. 218, American Helicopter Society 24th Annual Forum, Washington, D.C., Conclusions May 1968.

A test and analysis technique has been developed at DFVLR and has been proven as "Flying Qualities of Piloted V/STOL 3.

a valuable tool for the evaluation of clo- Aircraft; Military Specification", sed loop flying qualities with regard to MIL-F-83300, Dec. 1970.

Ii--

I I I o- ’ 0.01 0.1 1 10 STANDARD DEV. PITCH CONTROL STANDARD DEV. LONG. STICK CONTROL Fig. 16 Effect of control ratio on evaluation parameters (Bo 1135) 7. Cooper, G.E:, and Harper, R.P., "The Use of Pllot Rating in the Evalua- tion of Aircraft Handling Qualiti- es", AGARD Report No. 567, April 1969.

8. Sanders, K., Pausder, H.-J., and Hummes, D., "Flight Tests and Sta- stistical Data Analysis for Flying Qualities Investigations", Paper No. 56, 6th European Rotorcraft and Powered Lift Aircraft Forum, I I n Bristol, England, Sept. 1980.

9. Gmelin, B.L., and Pausder, H.-J., "The Impact of Helicopter Flight Mechanics on Mission Performance", AGARD-CP-313, Paris, France, April 1981, pp. 15-1-15-14.

10. Karmann, R., "Programmable Multipur- pose Flight Test Instrumentation System", AGARD-CP-299, Geilo, Norway, Oct. 1980, pp. 22-1-2213.

Fig. 17 Effect of control ratio on evaluation parameters (UH-1~) Corliss, L.D., and Carico, G.D., "A 4.

Preliminary Flight Investigation of Cross-Coupling and Lateral Damping for Nap-of-the-Earth Heli- copter Operations", Paper ~0.81-28, American Helicopter Society 37th Annual Forum, New Orleans, La, May 1981.

Tomlinson, B.N., and Padfield, G.D., 5.

"Piloted Simulation Studies on Helicopter Agility", Paper No. 30, 5th European Rotorcraft and Po- wered Lift Aircraft Forum, Amster- dam, The Netherlands, Sept. 1979.

6. Pausder, H.-J., and Gmelin, B.L., "Flight Test Results for Task Oriented Flying Qualities Evalua- tion", Paper No. 80-29, American Helicopter Society 36th Annual Fo- rum, Washington, D.C., May 1980.

A HELICOPTER HANDLING-QUALITIES STUDY OF THE EFFECTS OF ENGINE RESPONSE CHARACTERISTICS, HEIGHT-CONTROL DYNAMICS, AND EXCESS POWER ON NAP-OF-THE-EARTH OPERATIONS Lloyd D. Corliss U.S. Army Aeromechanics Laboratory U.S. Army Research and Technology Laboratories (AVRADCOM) NASA Ames Research Center Moffett Field, California Abstract fuselage-Ftical damp- ZWfus ing, set A ground-based simulation study was conducted inflow vr:tical damp- on a large-scale motion simulator to study the gwinflow effects in the vertical axis of engine response ing, set characteristics on handling qualities for a nap-of- stability augmenta- the-earth (NOE) operating environment. This study =,S concentrated specifically on the helicopter con- tion vertical damping, figuration with an rem-governed gas-turbine engine set-l and expands previous work by focusing on aspects peculiar to rotary-wing and NOE operations. A collective sensitivity z& wide range of engine response time, vehicle damp- g/in ing and sensitivity, and excess power levels was studied. The data are compared with the existing AT pure time delay, set handling-qualities specifications, MIL-F-83300 and AGARD 577, and in general show a need for higher damping ratio minimums when performing such NOR maneuvers as a dolphin and bob-up task. equivalent first-order Teng engine time-constant, Nomenclature (enE:ne response time), set Kn.Ku.Ke.Kl,K2 engine parameters (see thrust time constant, %T,%,\,Yy,Kq,~e,~p 1 Fig. 3) set-l engine gas generator Nl rotor speed, rad/sec speed, rpm fuel flow, lb/hr engine power turbine Nil speed, rpm second-order engine fre- radlsec quency , NOE nap-of-the-earth PR Cooper-Harper pilot rating Introduction maximum torque, ft-lb 'lim The potential for improving helicopter flying qualities through the use of electronic fuel- power turbine torque, QPT control devices on helicopter gas turbine engines ft-lb has led to a renewed interest in the study of coupling effects due to engine dynamics on the torque required, ft-lb Q ra vehicle height and yaw responses. An understanding and quantification of these engine coupling effects thrust, main rotor, lb Tmain is essential for the successful exploitation of the use of such controls. It is equally important to thrust-to-weight ratio T/W determine excess power requirements for specific -1 tasks such as those pertaining to nap-of-the-earth vertical damping, set =W (NOE) operations.

z equivalent vertical damp- W Earl studies in the area of VTOL flying -1 ing, set qualitiesys3 have provided a foundation for under- standing fundamental effects such as the thrust z aerodynamic vertical wa response time-constant and excess power require- damping, set-l ments. These studies involved ground-based simu- lation experiments that considered the near-hover tasks of station-keeping and rapid ascent and descent. Later studies4r5 expanded this work to consider the coupling effects of thrust response time with vertical velocity damping. These 2.5-3.0 for the tasks considered in this experi- studies - together with the results of Ref. 6, ment. The stability derivative matrices for the which considers vertical damping only - form the baseline augmented configuration at 40 knots and basis for the vertical-axis handling-qualities hover are shown in Table 1.

specifications found in MIL-F-83300 and AGARD 577.

Engine Model Reference 7 provides a good summary of the above studies and criteria; however, it is impor- The basis for the gas turbine engine model comes from a model developed for real-time simula- tant to preface that those engine coupling studies considered a fixed wing VTOL (aircraft) tion by Bell HelicopterlO and represents an XT-53 for which the engine response time (Teng) and engine with the inertias for a UH-1C rotor and transmission system. A block diagram of the adap- thrust response time (TT) were the same as depicted in Fig. 1. Such is not, in general, the tation of that model for this study is shown in Provisions are included for a pure time situation for a rotary-wing aircraft with an rpm- Fig. 3.

governed rotor response, as shown in Fig. 2. delay At and torque limiting Qlb at the power turbine stage. By ignoring the nonlinearities in This thrust response is influenced by a combina- tion of the energy stored in-the rotor, engine a transfer function with a second- AT and Qlim, and the h damping resulting order denominator can be generated (Fig. 3). As governor response, from rotor inflow. Thus, while the engine indicated in Ref. 10, most of the terms of that response of Fig. 2 may be simplified to a first- expression vary as a function of the gas generator a rangeiof 60-95% on N1 order time constant, the thrust response, as a speed N1; for example, cannot be. Reference 8 does, in a limited for the XT-53 engine results in a range of frequen- rule, address this problem; however, what is 4-8 rad/sec and a range of damping sense, cies of w, = needed is a review of the existing criteria and of 5 = 0.6-1.1. In this experiment, the engine terms the appropriateness of these requirements for were held constant for a given configuration and rotary-wing vehicles. It is also essential that the configurations studied varied over a range of specific mission tasks be addressed (e.g., NOE = 2-10 rad/sec and 5 = 0.3-1.0.

In addition to % operation) so that the criteria may be more frequency and damping, Qlim was varied to provide directly applied to the design of modern military a steady state (T/W) max. in hover ranging from helicopters. 1.025 to 1.25. Bear in mind that actual transient thrust can exceed these limits via the stored This paper describes a ground-simulation energy in the rotor system.

experiment that considered a wide range of engine response times and a wide range of vehicle verti- Task and Simulation Set-Up cal damping and collective control sensitivities fora helicopter model powered by an rpm-governed The determination of an appropriate task gas turbine engine. Several levels of available required the selection of one that would be mini- engine torque were also evaluated. The tasks per- mally affected by such simulation limitations as formed were the NOE tasks of dolphin, quick-stop, limited field of view and limited motion cues and and bob-up, and the study was performed on the yet one that would place large demands on the Ames five-degree-of-freedom Vertical Motion Simu- engine and vertical axis. Both requirements were lator (VMS) that used a model terrai.n-board visual sufficiently satisfied by flying the course out- An aural cueing system was used which The task consists of a constant- system. lined in Fig. 4.

provided the pilot with the sound of rotor over- speed (40 knots) berm-hopping maneuver (called a speed and underspeed, blade slap, and transmission "dolphin") followed by a deceleration to hover and noise, and was based on an approach used in Ref. 9. then a bob-up maneuver. The pilot was requested to The real-time simulation mathematical model con- change altitude during the dolphin maneuver, pri- sisted of a nine-degree-of-freedom helicopter marily through collective control inputs. He was model coupled to a simplified engine model, which instructed to maximize his masking by crossing the included the first-order dynamics of the governor, four berms with minimal clearance and staying low gas generator, power turbine, and rotorftransmis- between the berms. Because of a protective probe sion inertias. The data obtained for this wperi- on the terrain board camera, a minimum scaled ment are compared with previous studies and, where clearance of 17 ft was necessary. The pilot was possible, with the existing criteria. provided with a software-generated radar altimeter reading to assist him in determining his altitude.

After the fourth berm, the pilot performed a Description of Experiment deceleration of his choosing in preparation for The hover bob-up required the bob-up maneuver.

A requirement for this study was the develop- the sighting of three objects through 45" direc- ment of a real-time engine model and the estab- tional turns while maintaining maximum masking by lishment of a meaningful task. The test matrix the trees. The course was completed after the bob- consisted of variations in the vehicle s-axis down and reestablishment of a steady hover.

the engine response dynamics dynamics (Zw, Z6c), and available torque or excess power The pilot provided two Cooper-Harper pilot Wn) 9 /k;w max); the remaining vehicle characteristics ratings for each run, one for the dolphin portion were unchanged. The vehicle simulated was an of the course and one for the bob-up portion.

8000-lb, two-bladed teetering-rotor helicopter, Evaluation of the deceleration segment was combined sufficiently augmented to yield a pilot rating of with the bob-up maneuver during the experiment when changes were being made to the engine dynamics increasing the h rise time. The thrust responses only and was evaluated separately during the time all exhibit an immediate maximum thrust because of when changes were being made to the vehicle dynam- stored energy in the rotor followed by the tran- its. The latter was necessary since 2, varies sient behavior of the engine response and the h damping owing to the rotor inflow and augmentation.

as a function of speed and could only be specified Note that since the maximum thrust is achieved at 40 knots and hover.

almost immediately, all the thrust responses The cockpit instrument panel is shown in satisfy the 0.3-set level 1 Vertical Flight Char- Fig. 5. The primary instruments the pilot included acteristics (par. 3.2.5.2) criteria of MIL-F-83300 in his scan were radio altimeter, torque, rpm, and and the 0.5-set rise time criteria of AGARD 577.

airspeed; an rpm warning light was added. The However, the resulting pilot ratings for these pilot also had an rpm "beep" trim switch, for his governors in the bob-up varied from 3.0 for the used on the collective grip.

ideal governor to 6.5 for engine configuration E27 (i.e., wn = 2.0 radlsec, 5 = 0.7). Based on pilot Five pilots - two NASA test pilots, two Army commentary, these ratings reflected not only the test pilots, and an Army tactical pilot - partici- changes to the vehicle response resulting from the pated in this experiment. Most configurations engine dynamics but also reflected the attention were evaluated by at least three of the pilots and required for undesirable governor droop and over- were often repeated; there was a total of about speed.

200 data runs.

It should be noted that the data for the MIL-F-83300 thrust-response criteria were extracted Discussion of Results from experiments based on configurations similar to the one shown in Fig. 1; therefore, they do not The discussion that follows is based primarily account for effects of rotor-speed control, stored on averaged pilot ratings and is presented in three or inflow damping. Hence, since helicopter energy, subsections. Variations in the engine dynamics as described by Fig. 2, are quite thrust responses, with the vehicle characteristics held at it is reasonable to expect only, different in nature, those described in Table 1, are discussed first. that additional criteria to cover these responses Data for variations in vehicle height damping Zw Perhaps a criterion based on are necessary.

and collective control sensitivity Z&=, with the vertical acceleration or on a frequency-domain engine dynamics held constant, are discussed second, approach would be more appropriate.

and trade-offs between engine response time and height damping for the bob-up maneuver are dis- Several engine configurations (i.e., w, = 2, cussed last. Excess power requirements for spe- and 10 rad/sec) were studied at various 4, 6, cific tasks are also discussed in each subsection. levels of excess power, ranging from a steady state T/W = 1.025 to 1.25. Figure 9 shows how pilot rat- KfLects of Engine Dynamics ings varied with changes in engine dynamics and T/W for the hover bob-up. The vertical damping As was shown in Fig. 3, the engine model in was held at a fixed augmented level of this study can be represented by an expression with Z, = -0.65 see-l and Z6c = 0.38 g/in. for hover.

a second-order denominator. It was through this The engine response is depicted in terms of both representation that the engine response time and frequency wn (at 5 = 0.7) and the equivalent damping (i.e., w, and 5) were controlled. Altera- first-order time-constant ~~~~~ Also shown in tions in wn and 5 in this model can be thought Fig. 9 is a line below which it was found that the engine power or torque limiting would likely occur of as changes in the power train inertias, gas and speed governor or power sometime during the run. These data indicate that generator dynamics, turbine gains. No attempt was made to isolate a satisfactory flying-qualities boundary is formed and an engine response of these terms specifically; instead the engine by T/W > 1.1 reng < 0.2 set (i.e., w, 2 7.0 radlsec). Based parameters were varied to provide an overall gov- erned response in terms of the desired w, and 5. on time histories and pilot commentary, the lower bound on T/W was influenced by excessive power Figures 6 and 7 present the average pilot limiting, and the bound on engine response time ratings for the engine configurations as a func- ~~~~ was dictated by excessive engine overspeed tion of wn and 5 for unlimited T/W. Figure 6 and underspeed, as well as sluggish response.

presents data for the constant-speed dolphin maneu- ver, and Fig. 7 shows the results for the decelera- Effect of Vehicle Characteristics tion and hover bob-up maneuver. Also shown on these figures are the pilot ratings for the ideal In this segment of the experiment, variations governor (i.e., 0 held constant); this case re- and collective control in the vertical damping Zw sulted in pilot ratings of 2.5 for the dolphin sensitivity Zg, were studied. During this phase and 3.0 for the bob-up.

a highly responsive engine governor (w,, = 10 rad/ thus keeping the effects of the set) was used, Time histories of the thrust and torque re- engine response minimal and yet realistic. Verti- sponses to a 0.5-in. collective step for several was varied through stability aug- cal damping Zw engine models are shown in Fig. 8. The rpm and mentation of the basic speed-dependent aerodynamic rate of climb responses are also shown; as can damping which was -0.25 set-l in hover. A range be seen, the slower governors have an effect of 0 to -4 set-1 in hover was'studied.

of z,= Figures 10 and 11 show how pilot ratings questionable. However, what can be said of the varied with Z, and Qc. The results for the 40- data shown is that the required level of T/W does knot dolphin task are shown in Fig. 10 along with depend on Zw and is minimized at a total damp- an approximate,pilot rating (PR) = 3.5 fit to the ing of Zw = -0.8 to -1.0 see-1 in hover.

data. Also shown are the characteristics of the basic simulation model and several current genera- Trade-Offs between Vertical Damping 2, and tion helicopters.ll As can be seen, all of these Engine Response 'eng basic configurations lie outside of the PR = 3.5 region determined by this experiment. These data A final segment of this experiment studied the indicate a need with this task for a higher damp- trade-off between engine response ~~~~~~ and damp- ing and sensitivity than currently provided. ing (2,) on the overall height response of the vehicle. First consider the representation given The results for the hover bob-up maneuver in Fig. 1. This configuration consists of two are shown on Fig. 11. The PR = 3.5 contours for cascaded first-order systems, which can be approxi- these data along with those of several previous mated by a single first-order time-constant and is The current Several lines of con- near-hover studies are also given. shown by rw, in Table 2.

results describe a subset of the previous results, stant Z;; resulting from that table are plotted in general, higher sensitivities. The Also shown in Fig. 14 are the results favoring, in Fig. 14.

current results and the low-speed handling- of Ref. 5, which show a satisfactory boundary (i.e., qualities criteria given in MIL-F-83300 and AGARD PR c 3.5) for a trade-off between ~~~~ and 2,.

577 are compared inFig. 12. Also shown are the Although that study did not address the idea of an characteristics of several helicopters including equivalent Z;;, it can be seen that the boundary the unaugmented model used in this experiment. The lies along a constant Z, of -1.0 see-l. This results from this study do fall within the MIL-F- treatment implies that by maintaining an equivalent 83300 Level 1 boundaries; however, for the hover damping of greater than -1.0 see-I, satisfactory bob-up task, they indicate a need for a higher flying qualities can be obtained. Such a trade-off minimum for both damping and sensitivity. of z, represents a considerable depar- for reng ture from the MIL-F-83300 Level 1 criteria shown The effects of vertical damping (Z,) on in that figure.

excess power requirements (T/W) has been addressed in Refs. 1, 3, and 4, and form the basis for the Now consider the representation given in criteria given in MIL-F-83300. Figure 13 shows Fig. 2. Exploring the possible trade-off between the data from this experiment for the hover bob-up engine response and vertical damping for a heli- maneuver. The solid lines on that figure are the copter is not as straightforward because of the criteria as given by MIL-F-83300. These criteria complex nature of the thrust response, which, in are for a vehicle whose vertical damping is com- general, cannot be characterized by a first-order posed of an aerodynamic contribution only (i.e., A closer look at the time time-constant TT.

Zw The damping of the vehicle in this histories in Fig. 8, however, shows that the engine = &?,I.

experiment is represented by both an aerodynamic governor does have an effect on the h response and stability augmentation contribution (i.e., and hence on the effective damping Z,. Specifi- z In the case of a heli- cally, the engine configuration E67 (reng = 0.23 - Zwa + ZAC OQc/w)).

cEp;er model, however, the aerodynamic damping can set) causes an increase in 1; rise time (i.e., be further broken down into at least the inflow time to 63%) of from 1.5 set, for the ideal case, and fuselage contributions (i.e., Zwa = Zwfus + This results in a decrease in effec- to 1.8 sec.

tive damping of from -0.65 see-1 to -0.56 set-1.

Zwinflow) where in hover the inflow damping is predominant (i.e., Zwa = Zwinflow). For the model A further decrease in effective damping can be used in this experiment the aerodynamic damping in noted in the distorted h response for the engine hover is -0.25 set-1 and hence Zwa = Zwinflow configuration E27 (reng = 0.7 set-1). Thus a trend = -0.25 see-l. From the time histories shown in in equivalent or effective damping exists for Fig. 8 and from the diagram shown in Fig. 2, it Fig. 2 which is similar in nature to that shown can be seen that the inflow damping and stability for Fig. 1. However, the results for this case, augmentation damping cause the thrust response to which are shown in Fig. 15, indicate a far more decay, and since the steady-state value of thrust restrictive trade-off between and Z, than 'leng returns to its original level, it can be concluded is shown in Fig. 14. As was indicated earlier, an Since that Zfus = 0 (i.e., Zw = Zws + Zwinflow). upper limit on exists which is determined reng the criteria of MIL-F-83300 is intended for com- more by tolerable levels of engine overspeed and parison with the portion of damping which does not underspeed than by resulting ii response.

cause thrust decay (e.g., Zwfus). one is led to conclude that the data from this experiment should be compared with boundaries based on an inherent Conclusions These MIL-F-83300 bound- damping equal to zero.

aries are shown in Fig. 8. However, a further The effects of vertical axis response on the look at the time histories in Fig. 8 indicates that handling qualities of an rpm-governed helicopter while the thrust response returns to the original operating in an NOE environment were studied. The level, the torque response (i.e., engine output) results from this motion-based simulation show This peculiarity, along with the does not. several areas where present handling-qualities stored energy in the rotor, makes a comparison of criteria need extension or modification. The fol- helicopter data with the MIL-F-83300 boundaries lowing trends or conclusions are summarized: 1) An engine governor response of 0.2 set or 3. Gerdes, R. M., "A Piloted Motion Simulator faster is required for satisfactory flying quali- Investigation of V/STOL Height-Control ties and rpm control for the tasks performed in Requirements," NASA TN D-2451, 1964.

this experiment.

4. Kelly, J. R., and Garren, J. F., Jr., "Flight 2) In addition to engine response time an Investigation of V/STOL Height-Control excess power level of T/W > 1.1 is required Requirements for Hovering and Low-Speed during the bob-up. This excess power level is a Flight under Visual Conditions," NASA TN function of Z, and is minimized at a Z, of D-3977, 1967.

between -0.8 and -1.0 set-1.

5. Vinje, E. W., and Miller, D. P., "Analytical 3) For satisfactory flying qualities there and Flight Simulation Studies to Develop is a restricted trade-off between engine response Design Criteria for VTOL Aircraft Control time and vehicle damping; however, increases in AFFDL-TR-68-165, Apr. 1969.

Systems," engine time-constant are limited by poor rpm over- speed and underspeed control. 6. A'Harrah, R. C., and Kwiatkowski, S. F., "A New Look at V/STOL Flying Qualities," 4) The results from this experiment indicate Aerospace Engineering, 20 (7), July 1961, that higher minimums for both Zw and Z6c are 22-23, 86-92.

PP.

required for these NOE tasks than are specified by MIL-F-83300 and AGARD 577.

7. Hoh, R. H., and Ashkenas, I. L., "Development of VTOL Flying Qualities Criteria for 5) The thrust response for an rpm-governed Low Speed and Hover," NADC-77052-30, helicopter cannot be compared directly with the Dec. 1979.

thrust response time-constant criteria of MIL-F- 83300. The helicopter thrust response is com- 8. Vinje, E. W., and Miller, D. P., "Flight Simu- posed of a combination of stored energy, governed lator Experiments and Analyses in Support of response, and inflow damping, and hence cannot be Further Development of MIL-F-83300 - V/STOL characterized as a first order; thus a new Flying Qualities Specification," AFFDL-TR- criterion is needed. 73-34, June 1973.

Parrish, R. V., et al., "Empirical Comparison 9.

References of a Fixed-Base and a Moving-Base Simula- tion of a Helicopter Engaged in Visually 1. Gerdes, R. M., and Weick, R. F., "A Prelimi- Conducted Slalom Runs," NASA TN D-8424, 1977.

nary Piloted Simulator and Flight Study of Height Control Requirements for VTOL Aic- 10. Sonneborn, W., and Torres, I., "UH-lC Data for craft," NASA TN D-1201, 1962. Hybrid Computer Simulation," Report 204-099- 892, Bell Helicopter Company, undated.

2. Garren, J. F., Jr., and Assadourian, A., "A VTOL Height-Control Requirement in Hover 11. Heffley, R. K., et al., "A Compilation and as Determined from Motion Simulator Study," Analysis of Helicopter Handling Qualities Data," Vol. I, NASA CR-3144, 1979.

NASA TN D-1488, 1962.

Table 1. Baseline Augmented Configuration F matrix is: 40 Knots V P PHI R U W THETA Q -.670943-01 .633033-02 .125323 02 -.19605E 02 -.935503-03 -.942083 00 .54309E-01 .187243 01 -.289473 01 -.160873 00 -.105553 01 .92916E 02 .208983 02 -.262103-01 -.970033 00 .554383 00 .115423-01 .280873-02 -.290613 01 -.221983 01 .149083-03 .14730E 00 -.44039E 02 .370283-02 .OOOOOE00 .OOOOOE00 .OOOOOE00 .lOOOOE 01 .OOOOOE00 .OOOOOE00 .OOOOOE00 .OOOOOE00 -.1455OE-01 -.37659E-02 -.239833 00 .61808E 00 -.159473 00 -.43990E 01 .22246E 02 -.630773 02 -.56975E-02 -.10780E-02 -.25133E 00 .501353 00 -.92668E-02 -.495743 01 -.62792E 01 .693723 00 .OOOOOE 00 .OOOOOE 00 .OOOOOE00 .OOOOOE00 .OOOOOE00 .lOOOOE 01 .OOOOOE00 .48996E-01 .524193-02 .166253-02 -.25227E 00 .31119E 00 .232763-01 -.875973 00 -.864693 00 -.40001E 01 G matrix is: DELTA C DELTA P DELTA E DELTA A .67605E 00 -.39219E-02 .223163-01 -.199033 01 -.123633 02 -.11006E-01 -.38370E 01 -.330183-02 -.581163-02 -.69883D-03 .35294E 00 -.489413-03 .OOOOOE 00 .OOOOOE00 .OOOOOE 00 .OOOOOE00 ii = FX + GU -.135033 00 -.77366E-01 .18137E 01 -.99206E 00 .907813-02 -.572863-01 .108683 01 -.251493 00 .OOOOOE00 .OOOOOE00 .OOOOOE00 .OOOOOE00 VI .18877E 00 -.34315E-02 .24901E-01 .726043 00 N Hover F matrix is: U W THETA V P PHI R Q -.73200E-01 -.172363-01 .16619E 02 -.183973 02 .76353E-03 -.992323 00 .178483-01 -.676493-01 .376403-01 -.65021E 00 .143573 00 -.138803 01 -.395583-01 .674013-01 -.327543 00 -.250433-01 .11312E-01 .648093-02 -.27200E 01 -.221713 01 .19824E-03 .15302E-00 -.827933-03 -.389763-01 .OOOOOE00 .OOOOOE 00 .lOOOOE 01 .OOOOOE 00 .OOOOOE00 .OOOOOE00 .OOOOOE00 .241973-02 -.651853-02 -.560823-02 -.96860E 00 .526453-02 -.135653 00 -.690563 01 .223183 02 .375983 01 -.55066E-03 .288173-02 -.81417E 00 .11131E-02 -.145283-01 -.455523 01 -.626383 01 .624473 00 .OOOOOE 00 .OOOOOE 00 -.10730E-03 .OOOOOE00 .OOOOOE00 .lOOOOE 01 .OOOOOE00 .443423 01 .715763-02 .145233-01 -.554453 00 .778323-04 .138363-01 -.10268E 01 -.92055E 00 -.367673 01 G matrix is: DELTA E DELTA C DELTA A DELTA P -.220403 01 .535813 00 .OOOOOE 00 .884403-04 -.95908E-02 -.12090E 02 -.31445E 03 .OOOOOE00 .357203 00 -.31973E-02 -.12692E-02 .786613-02 .OOOOOE 00 .OOOOOE00 .OOOOOE 00 .OOOOOE00 -.344033-01 -.168053 00 .180643 01 -.10006E 01 .934183-01 -.97817E-01 .10870E 01 -.253643 00 .OOOOOE 00 .OOOOOE00 .OOOOOE 00 .OOOOOE00 .245613 00 .37786E-01 .35604E-01 .73235E 00 Table 2. Equivalent damping for configuration in Fig. 1.

- Damping - Zwa (set-1) w, at c = 0.7 ?eng C-c) (rad/sec) 0 -0.5 -1.0 -1.5 -2.0 -3.0 -4.0 1.4 1 0 -0.3 -0.42 -0.52 -0.6 -0.71 -0.71 .; 0.71 .; 0.71 2 0 -0.42 -0.59 -0.73 -0.84 -1.03 -1.4 H H I z z 0.5 0.5 3 0 -0.5 -0.71 -0.87 -1.0 -1.22 -1.73\ g g 4 0.35 -0.5 -1.04 -1.2 -1.46 4 0.35 4 0 -0.84 -2.07 % m m g g Iti 0.23 -1.47 Iti 0.23 6 0 -0.5 -1.0 -1.3 -1.81 -2.55/ .rl .rl 2 2 0.18 0.18 8 0 -0.5 -1.0 -1.45 -1.67 -2.04 -2.89 w w 0.14 10 0 -0.5 -1.0 -1.48 -1.9 -2.31 -3.27 Case III Case I Case II 1 1 - > 5 Z", - = Z" Z > s- a "a - Teng =eng Teng - Z wa 1

z, = zwa

4 Teng =eng

d-

ENGINE ('ENG)

rl

A/C il T I * (ZWa) I I I I i I -B-s- ----- Fig. 1. VTOL (jet lift) vertical control. Fig. 2. Helicopter vertical control.

ROTOR/TR:NSMISSION %N,, s + I/T, .zFff KPT Kn Kcc K1 Ke + KP K” Kc, K2 + Kq S* + (l/Kp + 1he + KpT K, Kc, K2 + K,) S + + + 7e re re e P t \ J J v v *.rw, n . T-53 ENGINE 5 * 0.6- 1.1 o, * 4 - 8 radhec WITH UH-1C Fig. 3. Engine model.

DASH QUICK STOP 40 knot DOLPHIN BOBUP AND DOWN Fig. 4. Simulation task.

6.5 4.3 .

.

I’

1.0

/II’

.8 111’ 3.p II 3;6 .

Jl I i= d .6 ,,iI’ SATIS- 4.0 .

,,,I FACTORY p E .4.

3.4 3.75 ,,1”‘3:cl 3.0 . . .

,,I’ p .2. 11” IDEAL ENGINE P.R. = 2.5 2, = -1.05 set-’ I , , 8 0 2 4 6 8 10 FREQUENCY wn, radhec Fig. 6. Effects of engine frequency Fig. 5. Cockpit instrumentation.

and damping - dolphin.

I - IDEAL GOVERNOR E67 - w, = 6 radhc, 1 = 0.7 E27 - w, = 2 radhac, 5 = 0.7 # A 100 f

r

2 4 6 8 10 W”, rxvsec .E 5 t Fig. 7. Effects of engine frequency rg and damping - quick stop/bob-up.

2, = -0.65 see-’ 2500 1 1 .oo TORQUE LIMITING , P I I ADEQUATE . 5.5 I 0 1.1 1.3 T/W Fig. 9. Engine response time Fig. 8. Engine response time histories; versus T/W - bob-up maneuver. collective steps (Zw = -.65 set-1).

I

-2.5 2;6 / 6;3 4;9 3.9 3;9 2.6 , - -2.0 3.5 - .

/ / SATISFACTORY 4.7 3.5 3.6 3.6 4.25 -1.5 -1.5 II’ l E -1.0 N -1.0 7 PRESENT STUDY PR = 3.5 / I z 3.51 T * / N 5 -.

-. 5 REF. 6 4.8 .

3.6 3.7 9 : cl I I I I I I 1 1 I I 0 .2 .4 .8 .8 1.0 .4 .4 .8 .8 1.0 1.0 0 .2 .2 Zg,. din.

Fig. 11. Vertical damping and collective Fig. 10. Vertical damping and collective sensitivity - bob-up maneuver.

sensitivity -'dolphin maneuver.

AGARD SATISFACTORY -2.5 - BOUNDARY PRESENT STUDY -2.0 - * 2.6 -2.0 l- -1.5 - 7 ADEQUATE SATISFACTORY u / I 8.5 8 6.2 4.7 . . * 3.5 / z fADJUSiED N Zs -OPTIMIZED -1.0 - L2 83300 C Ll . 3.0 * 2.6 / -. 5- 4.9 . 2.75 * 3.2 I 1.0 0 .2 .4 .6 .8 Zgc, din.

Fig. 13. Vertical damping and T/W - Fig. 12. Comparison of bob-up data with existing criteria. bob-up maneuver.

7.25 .

Zg -OPTIMIZED c ADEQUATE 5.25 5.25 . .

4.0 .

216 3i5 SATlSFACTORY I I I 0 -2 -4 -6 -2 -6 -8 0 -4 Z, -1lsec z,, SC-’ Fig. 15. Trade-off between i& and =eng Fig. 14. Trade-off between I;, and 'c~ for the hover bob-up.

for configurations of Fig. 1.

UNIFIED RESULTS OF SEVERAL ANALYTICAL AND EXPERIMENTAL STUDIES OF HELICOPTER HANDLING QUALITIES IN VISUAL TERRAIN FLIGHT Robert T. N. Chen NASA Ames Research Center, Moffett Field, California Abstract tasks in visual meteorological conditions were performed.*-" The influence of A series of helicopter handling- engine dynamics and excess power on these qualities studies--analyses, piloted tasks was also examined.'l In addition, groundrbased simulations, and a flight research is in progress to investigate the experiment--is described. The studies, effect of flight directors, vision aids, conducted at Ames Research Center, were and side-stick controllers on performance undertaken to investigate the effects of of these terrain-flying missions in rotor design parameters, interaxis cou- instrument meteorological or night condi- and various levels of stability and tions.12r'3 pli9, control augmentation on the flying quali- ties of helicopters performing low-level, The first visual terrain-flight terrain-flying tasks in visual meteorologi- experiment' was conducted on a fixed-based cal conditions. Some unified results are simulator to explore the effects on the presented, and the validity and limita- handling characteristics of basic single- tions of the flying-qualities data obtained rotor helicopters of large variations in are interpreted. Selected results, related rotor design parameters, such as flapping- to various design parameters, provide hinge offset, flapping-hinge restraint, guidelines for the preliminary design of blade inertia (or Lock number), and pitch- rotor systems and aircraft augmentation flap coupling. In the second ground-based systems.

simulation experiment, representative con- figurations from the first experiment were Introduction evaluated on a moving-base simulator [the Flight Simulator for Advanced Aircraft In recent years, the Army helicopter (FSAA)] to examine the effect of motion mission has placed considerable emphasis cues3 and the effects of various levels of on terrain-flying tactics for purposes of stability and control augmentation.' A survival and effectiveness in modern combat more sophisticated stability and control environments.'

The terrain-flying tasks in augmentation system (SCAS) was also synthe- these missions place strong demands on the sized, using linear optimal control theory agility and precision control capabilities to meet a set of comprehensive performance of the helicopter and have raised questions criteria.' This system, designed expressly concerning the flying qualities needed for for a hingeless-rotor helicopter, was sub- such tasks and the means of achieving them. sequently evaluated in the third piloted The existing flying-qualities specification ex eriment on the FSAA.6 ground-simulator P for military helicopters, MIL-H-8501A, is A flight experiment was conducted on the a 1961 update of a 1951 document: it does variable stability UH-lH/VSTOLAND heli- not address specifically such present-day copter14 to verify some selected configura- requirements of terrain flying. tions from the first two ground experiments, to explore additional configuration varia- To answer these flying-qualities ques- tions, and to investigate the effect of tions, a joint NASA/Army research program field of view on helicopter flying quali- was established at Ames Research Center. ties for nap-of-the-Earth (NOE) operations.

A series of analyses, piloted ground-based To relate directly some of the results of simulations, and flight experiments involv- these flying-qualities experiments to the ing terrain-flying tasks and low-altitude design parameters of the helicopter, an tactical missions has been and is still analytical study'r" was conducted to being conducted. Studies and exper.iments develop a design rule for the selection of designed to examine the effect of aircraft some primary rotor parameters to decouple design parameters, interaxis coupling, and the longitudinal and lateral motions of the levels of stability and control augmenta- helicopter.

tion on the flying qualities and man- machine performance of the low-level flying The purposes of this paper are to consider this set of flying-qualities data for visual terrain-flying tasks in a uni- fied framework, to interpret the validity and limitations of these data, and to relate the results directly, where possible, to design parameters, thus making them lllllllllllllll I I II II available as guidelines for use in the owing to external wind/turbulence distur- preliminary design of basic helicopters bances or to uncommanded control inputs and their stability and control augmenta- from the pilot.

tion systems.

As a result of these .requirements, In what follows, we discuss the there are many factors that influence heli- flying-qualities factors considered in copter agility: the basic performance ca- designing the experiments, describe the pabilities of the aircraft and the engine/ conduct of the experiments, and discuss governor dynamic characteristics, as well the main results and their design implica- as the flying qualities discussed above.

tions. The sequence of experiments described in the next section was designed to examine Factors Influencing Pilot-Vehicle only the flying qualities while holding the Performance and Pilot Workload performance factors and propulsion system in Visual Terrain Flight characteristics constant. However, the effects of the latter two factors on the In terrain flight, especially in NOE pilot-vehicle performance and pilot work- flight, the pilot is often called upon to load have also been examined recently at fly complicated and rapidly changing Ames.l' flight-path trajectories. These trajec- tories are generated, for example, from Design and Conduct of Experiments the need to avoid obstacles vertically or horizontally and to unmask and rapidly The simulation models and experimental remask by accelerating and decelerating variables, the flight simulation facili- the aircraft vertically, longitudinally, ties, the evaluation tasks, and the acqui- or laterally. The quickness, ease, and sition of the experimental data for this precision with which the pilot is able to series of experiments (outlined in Table 1) fly these trajectories are essential if are described in this section.

mission performance is to be enhanced with a concomitant increase in endurance. Helicopter Mathematical Model Training, particularly in navigation skills, is of critical importance in NOE The generic real-time helicopter simu- the characteristics or flight; however, lation model (ARMCOP) developed at Ames for this series of piloted ground-simula- qualities of the helicopter that permit the pilot to fly these complicated trajec- tion experiments2-6 consists of five tories easily, precisely, and quickly are modules describing aerodynamic force and the key to safe and successful operation. moment contribu.tions of the main rotor, These qualities or characteristics may be tail rotor, fuselage, vertical tail, and horizontal stabilizer. The main-rotor and defined as "agility."

tail-rotor modules are discussed in Ref.

To fly these NOE trajectories quickly, 15. The rotor model was derived from a the helicopter must be able to change linearly twisted rigid blade with an offset rapidly the magnitude and direction of its flapping hinge, a spring restraint about and pitch-flap coupling.

velocity vector in space. It must, there- the flapping hinge, fore, be able to rotate quickly the thrust For the first two experiments,2-4 a common vector of the main rotor and to change its fuselate, tail rotor, and empennage with characteristics similar to those of an magnitude to overcome drag and gravita- tional forces. Adequate control powers in AH-1G helicopter were used; the main-rotor pitch, roll, and yaw are therefore characteristics were varied. For the third the generic mathematical model required to make possible the rapid rota- experiment,' tion of thrust vector necessary to achieve was configured to simulate a hingeless rotor helicopter with characteristics the desired direction of the aircraft velocity vector; adequate thrust capa- similar to those of a BO-105.

bility, installed power, and responsive- ness of the engine/governor system are The ARMCOP model also includes a needed to meet the demand for rapid change general form of SCAS (Fig. 1). The aug- in thrust magnitude. mentation system employs a complete state feedback and a control mixinq structure To fly these complex NOE trajectories that facilitates implementation of control cross-feed4r5 and control-quickening from easily and precisely, the helicopter must possess satisfactory flying qualities. each of the four cockpit control inputs.

adequate damping in consonance with Also, the augmentation system gains may be Thus, appropriate control sensitivity is needed programmed as functions of flight param- A limited attempt in pitch, roll, yaw, and heave; interaxis eters such as airspeed.

was made to validate the generic model, as cross-coupling must be minimized so that unnatural or complicated control coordina- discussed in Refs. 2 and 3.

tion is not required; and adequate sta- bility must be provided to damp out upsets Experiment Variables investigate the effects of variations in roll damping, roll sensitivity, and pitch- The general objective of experiment I roll cross-coupling on the helicopter fly- (Ref. 2) was to explore the effects on ing qualities for NOE operations and to terrain-flight flying qualities of large correlate the results with the ground- variations in four primary rotor design based experiments, I and II.

parameters: flapping-hinge offset, Flight Simulation Facility flapping-hinge restraint, blade Lock num- ber, and pitch-flap coupling. Forty-four combinations of the four parameters, which A fixed-base simulator, in conjunc- cover the teetering, articulated, and tion with a Redifon closed-circuit tele- hinged rotor system families, were con- vision system, was used in experiment I.

The simulator consisted of a Bell UH-1A figured in the generic mathematical model ARMCOP, using a common fuselage, tail cabin section facing a shrouded screen and rotor, and empennage. To investigate sys- TV projector. The UH-1A control system was used with working hydraulics, bungee tematically both the major and interactive these configurations were de- cords, and magnetic brake. A 1:400 scale effects, signed and related to three sets of flying terrain model was used in this simulation.

The Ames Flight Simulator for Advanced qualities parameters: damping and con- trol sensitivity in pitch and roll axes; Aircraft (FSAA), a six-degree-of-freedom pitch-roll cross-coupling owing to air- moving-base simulator (Fig. Z), was used craft angular rate; and longitudinal in experiments II and III. The pilot was static stability. again provided with conventional pedals, and collective controls, and cyclic stick, In experiment II (Ref. 4), the objec- a basic set of flight instruments, as tive was to investigate the use of various shown in Fig. 3. The visual scene was generated from the same terrain model used levels of SCAS to improve the flying qualities in terrain flight. Five basic the scene was presented in experiment I; single-rotor helicopters - one teetering, through the cab window on a color TV moni- > two articulated, and two hingeless - which tor with a collimating lens.

were found to have major deficiencies in the flight experiment,7 experiment I were selected as baseline Experiment IV, The major handling- was conducted on the NASA/Army variable- configurations.

qualities deficiencies included inadequate stability UH-1H helicopter, which incor- The damping and sensitivity in pitch and roll; porates a V/STOLAND avionics system.

excess pitch-roll coupling; and excess V/STOLAND system, equipped with two digital was designed for flight pitch and yaw coupling resulting from col- flight computers, The SCAS that were de- control, display, navigation, and guidance lective input.

signed and evlauated included simple con- research. The flight control portion of trol augmentation systems (CAS) to de- the V/STOLAND system was used in this experiment. Each control channel uses a couple pitch and yaw responses caused by collective input and to quicken the pitch combination of a limited-authority (20% to and roll control responses; rate-command- 30%) series servo and a full-authority designed to optimize the sensi- parallel servo. In the research mode, the type SCAS, tivity and damping and to decouple the left cyclic stick, controlled by the is mechanically discon- pitch-roll caused by aircraft angular rate; evaluation pilot, and attitude-command-type SCAS. The gen- nected from the right stick and operated eral form of the augmentation system in in a fly-by-wire status. The safety pilot the ARMCOP was used to configure the above on the right retains'control of the air- types of SCAS. craft through the standard UH-1H cyclic and cockpit instruments. The fixed-based simulator facility used for experiment I The objective of experiment III 6) was simply to conduct a compara- can be tied directly to the V/STOLAND (Ref.

tive evaluation to determine the extent hardware and was used in software devel- opment and checkout for this flight exper- to which the handling qualities of a basic hingeless-rotor helicopter can be improved iment.

by incorporating a sophisticated SCAS Evaluation Tasks designed on the basis of linear optimal control theory.5 Again, the basic air- Experiment I comprised three tasks: craft and the SCAS system were implemented on the ARMCOP model. The mechanization the longitudinal dolphin task - flying was done in such a way that two levels of over a sequence of barriers (hurdles) augmentation could be evaluated: sta- placed at irreaular intervals: a lateral task- flying a slalom course of trees bility augmentation only, and complete spaced similar to the barriers in a stability and control augmentation.

and a combined longitudina straight line; Experiment IV, the in-flight simula- and lateral-directional task - flying a tion experiment, was conducted to course of barriers combined with trees

I lllllllllllllllllllllllllllllllllllllllllllllllll lllll IllIII

placed down the centerline of the bar- qualities of the helicopter in visual ter- riers. Only the combination course (Fig. rain flight. For this paper, only the 4) was used in experiments II and III. Cooper-Harper Pilot Rating (CHPR) data A slightly different scaling was used in will be used to quantify the flying- experiment I; it resulted in somewhat qualities results: other experimental data larger trees (75 ft instead of 50 ft), pertaining to the pilot comments and the larger barriers (50 ft instead of 33 ft), task performance will not be discussed.

The latter have been discussed else- and a correspondingly longer spacing between barriers (700 to 1400 ft). The where'-" in the results of each indi- pilots were given instructions to fly as vidual experiment.

low as possible and as fast as possible through the courses, banking alternately Sensitivity and Damping in Pitch and Roll left and right around the trees and drop- ping down between the barriers. The tasks The combined effects of control started at a trimmed, level-flight initial sensitivity and damping were expected to conditions of 40 knots at about 100 ft AGL have a significant influence on NOE flying for experiment I (60 knots for Exp. II, qualities, since they determine the short- and 100 knots at 500 ft AGL for Exp. III). term characteristics of the pitch and roll Minimum vertical obstacle clearance was responses to cockpit cyclic controls.

taking together all the pilot limited to about 17 ft bv a device de- However, ratings for this series of experiments, signed to protect the television camera optics from inadvertent impact with the the results indicate that the relationship model terrain. Generally, each pilot was of the sensitivity and damping in pitch allowed a limited number of runs with a and roll alone is not a predominant factor standard configuration at the beginning of for the tasks evaluated. Other factors, his simulation test period in order to such as yaw damping, pitch-roll coupling allow him to become reaccustomed to the caused by aircraft angular rate, and col- simulator and task. Wind and turbulence lective input couplings to pitch and yaw were not introduced in these tasks. also were found to be important.

L For the flight experiment (Exp. IV), Figure 6 shows the results of the the task was to fly through a prescribed pilot rating data for configurations with slalom course over a runway at the NASA low yaw damping (N, = -1.2 set-l) and a Flight System Research Facility at Crows low level of pitch-roll coupling caused by Landing, California (Fig. 5). The pilots aircraft angular rate (I Lq/Lpl < 0.3).

were asked to fly through the course while Most of the configurations covering a wide maintaining speed and altitude constant at range of sensitivity and damping combina- 60 knots and 100 ft AGL, respectively. tions in roll received ratings of accept- Most of the evaluations were conducted in able (CHPR < 6.5) for the lateral task.

calm-air conditions or with winds below In terms of the change in roll attitude at 10 knots at directions of no more than 40° the end of 1 set in response to an inch- to the centerline of the course runway. step input in the lateral stick, A$,, these configurations extend from about 4O Data Acquisition to 30". It is noted, however, that the extreme low sensitivity and low damping Data collected from these experiments combinations were found to be unacceptable.

1) Cooper-Harper Pilot were of two types: These configurations were brought into the and verbal comments recorded at "clearly acceptable" ratings in RatingsI region of the conclusion of each evaluation; and 2) experiment II by increasing the damping time histories of helicopter trajectories, and sensitivity to a level of motion variables, and control usage for L - -5 set-l and L&a = 1.4 rad/sec'/in, real-time monitoring and for postflight rgspectively (and with slight augmenta- analysis. Two pilots participated in tion in yaw damping from Nr = -1.2 to experiment I and completed a total of 172 -1.6 set-I).

evaluations. A total of 127 evaluations Increasing the yaw damping to a high were achieved in experiment II by three participating pilots. In experiment III, value (Nr = -3.5 set-I) while reducing the two pilots completed a total of 21 NOE pitch-roll coupling owing to angular rate evaluations in addition to evaluations for to near zero improved the pilot rating tasks other than terrain flight. A total considerably, as shown in Fig. 7. Never- the improvement for the low of 150 evaluations were achieved by four theless, sensitivity and low damping combinations participating pilots in experiment IV.

was insufficient to achieve a rating Results and Discussions better than marginally acceptable. Lim- itations of in-flight simulation capa- The results of this series of experi- bilities hindered the exploration of a wider range of sensitivity and damping ments are combined and grouped in terms combinations in experiment IV. Based on of major factors influencing the flying as well as on the pilot this set of data, from unacceptable or marginally acceptable commentary, it appears that there is a to at least acceptable was achieved when level of sensitivity and damping combina- the coupling was reduced or the damping was increased or both.

tion below which a precise roll control may not be achieved without a tendency to overcontrol or to develop pilot-induced The results from experiment IV (Ref.

oscillations. The data also suggest that 7) pertaining to the effect on pilot rating a minimum roll damping of about -3 set-1 of the pitch-roll cross-coupling are shown with Ael from 4' to 300/in in 1 set in Fig. 9 for three levels of roll damping results in clearly acceptable flying with sensitivity held constant. With pitch qualities. and roll sensitivities fixed, the pilot commented that the aircraft was a little The flight experiment (Exp. IV) did oscillatory with low damping and sluggish not examine the effect of sensitivity- with high damping. Increasing the cross- However, damping combinations in pitch. coupling ratio degraded significantly the based on the result of experiments. I and pilot rating for the highest damping, but a minimum pitch damping (Mq) of about only slightly for the low- and medium- II, damping cases. In particular, when the -1.5 see-1 with AB, (which is the change in pitch attitude, at the end of 1 set, in most favorable combination of sensitivity response to an inch-step input in longi- and damping (Lpi = -4, Lga = -0.55, Abl = 6) tudinal stick) in the range of 4O-25' may the degradation of flying qualities with be appropriate for acceptable flying quali- cross-coupling was not as severe as ties for the longitudinal task. observed in the simulation experiments.

Pitch-Roll Cross-Coupling Resulting from Collective Input Coupling - Aircraft Angular Rate The effects of collective input cou- Unlike fixed-wing aircraft, for which pling to pitch and yaw were expressly Data pertaining pitch-roll coupling is rare except in high- examined in experiment II.

angle-of-attack operations, helicopters to these effects can also be extracted from generally exhibit undesirable pitch-roll the results of experiment III. The benefit coupling because of aircraft angular of reducing the collective input to yaw motion. For example, in response to a roll coupling was found to be dependent on the rate to the right, the tip-path plane (TPP) level of yaw damping. For a moderate yaw tilts to the left with respect to the rotor damping (Nr = -1.6 set-'), an improvement hub to provide desirable roll damping; of about one rating point was achieved in the TPP response can also include however, experiment II (see Fig. 10) by decoupling tilt in the fore-aft direction which pro- yaw to collective response. When the yaw duces an undesirable pitching moment. damping was high (Nr = -3.5 set-I) such as This coupling characteristic, for a general in some configurations examined in experi- configuration, is a result of combined ments III and IV, the results suggest that effects of gyroscopic and aerodynamic only a slight improvement is realized by moments acting on the rotor system. this decoupling.

The ratio of the roll moment result- In the speed range flown for the eval- ing from pitch rate to the roll moment uation tasks (40 to 80 knots), the coupling resuiting from roll rate, Lq/Lp, for to pitch from the collective input became example, plays an important role in deter- substantial for hingeless rotor or stif- Experi- mining the roll-rate-to-pitch-rate ratio fened hinged-rotor configurations.

in the short-term aircraft response to a ments I, II, and III indicate that this step input in the longitudinal stick; sort of coupling has a significant effect Figure 11 shows similarly, the ratio determines on the flying qualities.

Mp'Mq the ratio of pitch rate to roll rate in the effect on pilot rating of doubling and the short-term response to a step input in eliminating the collective input coupling the lateral stick. Figure 8 shows the to pitch (M6c), and a combined effect of variation of the pilot rating with eliminating both pitch and yaw coupling for ForLg'LP from experiments I, II, and III.

a hingeless-rotor helicopter examined in comparison purposes, the boundaries dis- experiment II.

cussed in Ref. 17 are also shown in the figure. The boundaries indicate that if Type of Flight Control System the value of the coupling parameter exceeds 0.3, ratings better than acceptable cannot As shown in Table 1, two types of be achieved. (Values greater than 0.5 flight control systems in the pitch and imply unacceptable flying qualities.) In roll axes were examined in this sequence of 1) a rate type (including the experiment I, adverse comments on this experiments-: kind of coupling were made by the pilots basic aircraft, considered in experiments I, exceeded 0.25. In experi- and IV, and 2) an attitude type, exam- when ILq/Lpl II, Taking ment II, improvement in the pilot rating ined in experiments II and III.

all the experiments together, the results and lessons learned are discussed in the do not indicate a clear preference by the following paragraphs.

pilots for either of the two types of This control system for the tasks flown. Elimination of Interaxis Coupling was reported previously in the results of experiment II and was further substanti- Pitch-Roll Decoupling Figure 12 ated in experiments III and IV.

shows the results for a pilot (pilot A) A design rule '#lo has been developed who participated in all four experiments. for the selection of the design parameters of the rotor systems to reduce the undesir- It should be emphasized that the able pitch-roll coupling caused by aircraft result is valid only for the tasks evalu- angular rate in pitch and roll. The basic ated. The tasks were flown at an airspeed idea of the design rule is to cancel per- In this in the range of 40 to 80 knots. fectly in hover the inertia and aero- flight regime, the pilot can perform the dynamic factors that contribute to the precision flight-path control task equally steady-state coupling in rotor tip-path- well and with ease with either a properly plane (TPP) response to the aircraft designed rate-type or attitude-command- angular rate in pitch and roll. In type control system in pitch and roll. essence, the method is to "tune" the flap- This result should not be extrapolated, ping frequency ratio, P however, to include other NOE tasks such as precision hover over the ground in turbulence. For these other precision position control tasks near hover an atti-

(1)

tude system or another type of control

‘I

such as a velocity-command type, system, pzeferred to the angular rate-type may be to the decoupling flapping frequency ratio sys tern.

given by PD Effect of Longitudinal Static Stability Limited consideration was given in g (+ - f)(i - fE + $) * (2) experiment I to investigating the effect PD= 1+ of variations in longitudinal static sta- eM bility with respect to angle of attack 21++ [ The effect of (M,) using a &3 hinge.

( ) variations in longitudinal static sta- bility with speed (MU) was not investi- through use of a pitch-flap coupling gated in this series of experiments, = tans,) or a flapping restraint 63 (Kl because the tasks evaluated in the ground or both for a given hinge offset e.

K0 simulations did not call for precise speed (1) and (2) above, y is the Lock In Eqs.

The result obtained from experi- control. number of the rotor blade; s is the ratio ment I suggests that, for the demanding to rotor radius: R is the angular of e tasks evaluated, some longitudinal static velocity of the rotor system; and Mg and instability with angle of attack, such as 1~ are,respectively, the blade mass moment.- is the case for some hingeless-rotor heli- and moment of inertia of the blade about copters in forward flight, appears accept- the flapping hinge.

this result must be able. However, qualified somewhat because the tasks were The values of pitch-flap coupling In turbulence, de- flown in calm air. required to achieve pitch-roll decoupling graded flying qualities caused by static are generally moderate, as shown in Fig.

instability may be expected. 13, even for extreme combinations of E They are effective in reducing the and KS.

Design Guidelines the coupling ratio "$;,y;p;%-1, ;; hover and in forward The experimental results clearly 14) and they result in well-behaved Fig.

indicate that the interaxis coupling, such TPP transient response. Figure 15 shows as pitch-roll cross-coupling and collec- an example of the TPP transient response tive input coupling to pitch and yaw, and to a unit change in roll rate (and pitch levels of sensitivity and damping are rate) at hover and at an advance ratio of major factors influencing the flying E = 0.05, y = 12, 0.3 for a rotor with qualities of the helicopter in terrain with and without the use of decoupling 6,.

flight. Analytical studies were performed to relate some of the experimental results Decoupling pitch and roll caused by to the design parameters of the rotor sys- aircraft angular rate may also be achieved tem and aircraft augmentation systems: using feedback control, as was done in this was done to develop means of improv- experiment II by feeding the pitch rate to ing the flying qualities. Some results lateral cyclic and roll rate to longitudi- nal cyclic control.

Decouplihg Collective to Yaw and Pitch In experiment II (Ref. 41, the design of the rate-type SCAS used Ael = 7.5', The yawing moment resulting from col- A#1 = lo', and AQ1 = 7.5O, approximately, lective input, Ndc, which exists in all and in experiment III (Ref. 5) the sensi- conventional single-rotor helicopters, tivity criterion used for the SCAS design should be eliminated, particularly when was 3 < A8, 5 ZOO, 4 5 Ael 5 ZO", and the yaw damping of the aircraft is low. for pitch, roll, and yaw, 6 < AJI1 < 23“ The yaw coupling can be eliminated simply respectively. The designs resulted in by cross-feeding collective to the pedals. pilot ratings of satisfactory for the The gain is a nonlinear function of air- tasks flown.

speed, the shape of which is similar to the familiar power required curve." Care The minimum acceptable damping must be exercised, however, in deriving required for the tasks considered in the the cross-feed gain, especially when experiments appears to be about small-perturbation derivatives are used.

M -1.5 to -2 set-l, Lp = -3 to -4, and Nq = Control derivatives such as can be -1.6 to -2, respectively for pitch, r= N6c a strong function of the magnitude as well roll and yaw. The pitch and roll damping as direction of perturbations, as shown in may be obtained by appropriately choosing Fib. 16. Modifications to the initial the design parameters of the rotor system design were required in experiments II and such as flapping-hinge offset, flapping III to accommodate this kind of non- A cursory restraint, and Lock number.'

linearity. survey indicates, however, that yaw damp- ing may be inadequate for many production Increased control power obtained helicopters for terrain flight; an aug- through hinge offset or a stiffened flap- mentation in yaw damping is thus desirable.

ping hinge produces a coupling in pitching moment caused by collective input, which Attitude SCAS Design increases with airspeed. This pitching moment can be eliminated simply by cross- A few combinations of the two major feeding the collective to the longitudinal design parameters associated with the cyclic and scheduling the gain with air- attitude command system in pitch and roll, speed. Again, care must be exercised in namely the sensitivity in aircraft atti- mechanizing the system so as not to intro- tude, change per unit stick deflection, duce the undesirable effect of reducing and the bandwidth, were examined in the longitudinal static stability with experiment II. As expected, these param- speed." eters had significant effect on the flying qualities for the tasks evaluated. The Selection of Sensitivity and Damping "optimized" sets of these two parameters in Pitch, Roll, and Yaw for the pitch and roll axes, as shown in Table 2, provide a guide for future design The wide range of acceptable sensi- of such SCAS systems.

tivity in pitch and roll axes, as exempli- fied in Figs. 6 and 7, makes it somewhat Finally, it is of interest to note difficult to select this parameter in the that for a hingeless-rotor helicopter, it preliminary design stage. However, a has been found beneficia15r6 to feed back proper selection may be accomplished by pitch-rate and pitch-attitude signals to judiciously relating the sensitivity collective pitch in addition to the longi- requirement to the task demands: Because the avail- lower tudinal cyclic pitch.

sensitivity for demands with smaller atti- able pitching moment resulting from tude excursions, higher for tasks demand- collective pitch increases with speed, the ing larger attitude excursions. For gains to collective pitch must be sched- example, to clear the obstacles in a uled with airspeed accordingly; however, slalom course, the radius for banked the gains to the cyclic pitch may be held turns must be smaller than one half of constant, because of essentially constant the spacing between two obstacles. The control effectiveness with the cyclic turn radius is a function of the speed of pitch for the hingeless-rotor helicopter.

flight and the bank angle, as shown in Fig. 17. For a spacing of 1000 ft, as Conclusions used in experiment IV, bank angles of about 30' or more are required if a speed A series of analytical and experi- of 60 knots is maintained.

Had the task mental studies investigating the effect of been flown at 80 knots or with the spac- rotor design parameters, interaxis cou- ina reduced to 500 ft. the bank angle and levels of stability and control pling, required would have been about 50' or augmentation on the flying qualities of more; the lower roll sensitivity of the helicopter in visual terrain flight A$J~ = 4.5", which received good pilot has been conducted. The evaluation tasks ratings (see Fig. 71, might have been used in the experimental studies consisted down-rated for the more demanding task. of a longitudinal dolphin task, a lateral and a combined longitudinal 2 Chen, R.T.N., and Talbot, P.D., "An slalom task, and later-directional task: all tasks were Exploratory Investigation of the Effects of Large Variations in flown in the airspeed range of 40 to 80 The following conclusions were Rotor System Dynamics Design knots.

reached: Parameters on Helicopter Handling Characteri.stics in Nap-of-the- Minimum levels of damping and Earth Flight," Preprint No.

1) sensitivity in pitch and roll are required 77.34-41, 33rd Annual National to achieve clearly acceptable or better Forum of the American Helicopter flying qualities (CHPR < 5). For damping, Society Washington, D.C., May a minimum of about -3 set-l for roll and 1977.

-1.5-l for pitch are appropriate: for in terms of the change in 3. Talbot, P. D., Dugan, D. C., Chen, sensitivity - R.T.N., and Gerdes, R. M., attitude at the end of 1 set following an inch-step input in cyclic stick - a mini- "Effects of Rotor Parameter Vari- mum of about 4O for both pitch and roll is ations on Handling Qualities of suggested for the tasks at the flight con- Unaugmented Helicopters in Simu- ditions noted. lated Terrain Flight," NASA TM- 81890, 1980.

To achieve satisfactory flying 2) cualities, the absolute value of the ratio 4. Chen, R.T.N., Talbot, P. D., Gerdes, of roll moment caused by pitch rate to R. M., and Dugan, D. C., "A roll dampinq must be less than 0.35. This Piloted Simulator Investiga- coupling-ratio can be reduced to nearly tion of Augmentation Systems to zero using a design rule developed in this Improve Helicopter Nap-of-the- Earth Handling Qualities," Pre- series of studies.

print No. 78-29, 34th Annual In forward flight, the large National Forum of the American 3) pitching moment resulting from collective Helicopter Society, Washington, input associated with rotors having a D.C., May 1978.

large flapping-hinge offset and a stiff flapping hinge can be detrimental to fly- 5. Miyajima, K., "Analytical Design of a Signifi- High Performance Stability and ing qualities in terrain flight.

cant improvement in pilot ratings has been Control Augmentation System for a achieved by cross-feeding longitudinal Hingeless Rotor Helicopter," cyclic from collective input. Preprint No. 78-27, 34th Annual National Forum of the American 4) The coupling to yaw caused by Helicopter Society, Washington, collective input can be objectionable, May 1978.

D.C., especially when damping in yaw is low.

Augmenting the yaw damping or cross- 6. Miyajima, K. and Chen, R.T.N., feeding collective input to the pedals to "Analytical and Experimental decouple the yawing moment substantially Study of an ADvanced Stability improves the pilot rating. and Control Augmentation System for a Hingeless Rotor Helicopter," NASA TM, in preparation.

Properly designed, both rate- 5) command and attitude-command SCAS made 7. Corliss, L. D. and Carico, G. D., "A substantial improvements in terrain-flight flying qualities in otherwise unacceptable Preliminary Flight Investigation helicopter configurations; no evidence was of Cross-Coupling and Lateral found for a clear-cut preference for Damping for Nap-of-the-Earth Heli- either type of augmentation for the tasks copter Operations," Preprint No.

81-28, 37th Annual National Forum flown.

of the American Helicopter The design of attitude-type SCAS Society, New Orleans, LA, May 6) for hingeless-rotor or stiff-hinged-rotor 1981.

helicopters should include the feedback of pitch rate and pitch attitude to collec- 8. Gerdes, R. M., "A Pilot's Assessment tive pitch, as well as their feedback to of Helicopter Handling-Quality the longitudinal cyclic pitch. Factors Common to Both Agility and Instrument Flying Tasks," References NASA TM-81217, 1980.

1. U.S. Army Field Manual l-l, Oct. 1, 9. Chen, R.T.N., "Effects of Primary 1975.

Rotor Parameters on Flapping NASA TP-1431, 1980.

Dynamics," I II II III 11111 10. Chen, R.T.N., "Selection of Some 14. Baker, F. A., Jaynes, D. N., Corliss, D ., Rotor Parameters to Reduce Pitch L. Liden, S., Merrick, R. B., Roll Couplinq of Helicopter and Dugan, D. C., "V/STOLAND Flight Dynamics," Preprint No. Avionics System Flight - Test Data on a UH-1H Helicopter," NASA TM- I-6, National Specialists' Confer- ence on Rotor S;stems Design of 78591, 1980.

the American Helicopter Society, "A Simplified Rotor Philadelphia, PA, Oct. 1980. 15. Chen, R.T.N., System Mathematical Model for 11. Corliss, L. D., "The Effect of Heli- Piloted Flight Dynamics Simula- copter Engine Response Dynamics tion," NASA TM-78575, 1979.

and Excess Power on NOE Handling Qualities," American Helicopter 16. Cooper, G. E., and Harper, Jr., R. P., Society H.Q. Specialists' Confer- "The Use of Pilot Rating in the ence, Ames Research Center, NASA, Evaluation of Aircraft Handlinq Moffett Field, Calif., Apr. 1982. Qualities," NASA TND-5153, 1969.

12. Aiken, E. W. and Merrill, R. K., 17. Huston, R. J., and Ward, J. F., "Results of a Simulator Investiga- "Handling Qualities and Structural tion of Control System and Display Characteristics of the Hingeless- Variations for an Attack Heli- Rotor Helicopter," Proceedings of copter Mission," Paper No. 80-28, the V/STOL Aircraft Conference, 36th Annual National Forum of the Apr. 1966.

American Helicopter Society, Washington, D.C., May 1980.

18. Forrest, R. D., Chen, R.T.N., Gerdes, R. M., Alderete, T. S., and Gee, 13. Landis, K. H., and Aiken, E. W., D. "Piloted Simulator Investi- R., "An Assessment of Various Side- gation of Helicopter Control Stick Controller/Stability and Systems Effects on Handlinq Quali- Control Augmentation Systems for ties during Instrument Flight," NOE Flight Using Piloted Simula- Paper No. 79-26, 35th Annual tion,” American Helicopter Society National Forum of the American H.Q. Specialists' Conference, Helicopter Society, Washington, Ames Research Center, NASA, D.C., May 1979.

Moffett Field, Calif., Apr. 1982.

Table 1. Summary of terrain flight experiments Experiments Tasks Simulator Rotor type Control system type Objective Basic helicopter I To determine effect Longitudinal vertical Fixed base Teetering of large variations in task (Ames S-19) Articulated (rate-type in pitch, Lateral slalom task Hingeless roll, and yaw) rotor design parameters Combined task SCAS Input Decoupling II To assess effect of Combined task Moving base Teetering Rate command various levels of SCAS (Ames FSAA) Articulated Hingeless Attitude command in pitch and roll Combined task Moving base Hingeless SCAS III To evaluate a sophisti- (Ames FSAA) Attitude and rate cated SCAS for hinge- Stability augmen- less rotor helicopter tation Control augmenta- tion To investigate roll Prescribed lateral In-flight Teetering Rate-type in pitch, IV roll sensi- slalom course over (UH-lo/ roll, and yaw damping, VSTOLAND) tivity, and pitch-roll a runway cross-coupling and correlate results with Experiments I and II.

Table 2. Partially optimized characteristics of attitude SCAS in pitch.and roll.

Pitch Roll Frequency and damping ratio wn ,rad/sec 1.9 to 2.0 1.8 to 2.0 5 0.9 to 1.0 1 to 1.2 Attitude sensitivities A0/6, , deg/in 5 to 10 20 to 22 A$/Sa I dedin DIRECTIONAL CONTROL

I

>

Al RCRAFT STATE, X CROSS-FEED AND FEEDBACK GAINS X = (u. w. q. Af’. v. P. A$. dT FEED FORWARD GAINS Fig. 1. General stability and control augmentation system structure of the AFQKOP model.

Fig. 2.

The flight simulator for Fig. 3. Instrument configuration in advanced aircraft.

simulator cab.

I I IA 1 I I 0 '8000 10000 12000 14000 16000 h DISTANCE FROM REF.

Fig. 4.

Layout of nap-of-the-Earth Fig. 5. Slalom-course task for the terrain-avoidance obstacle flight experiment (Crows course.

Landing, Calif.).

o EXP. I (N, =I-1.2sec-‘) A EXP. IV (Lq/Lp = 0; Nr = -3.5 WC-‘) 0 EXP. II (Lq/Lp = 0; Nr = -1.6 set-‘) q EXP. II (N, = -1.6sec-‘1 A EXP. IV (Lq/Lp = 0; N, = -1.2 se-‘) A EXP. IV (Lq/Lp = 0.25; N, = -1.2%-‘1 A @, ; 2.5’ ,I 4.25 loo I 20 / ,/:75 < ,//3.5 2o” / 0 d I ,’ / /’ 3.75 / ’ ;J4 / ,/ 425 ,,/ / 0 jj l5 / 45 /’ 3o” / I ,‘o ,/ ./ 40$= /’ A’ 4 10 /C j.Q / 505 /’ /HO I 96.2 ;‘$255.76/ 3.y I -g-- O//3.75 ,R’ /4.6 ,‘O I I I ~.5*@.2 $4.5’ ,o” .5 1 1.5 2 /’ ;ee”;f$-N;- Lg,. radlsec2/in.

0 1 2 3 4 5 6 Fig. 7. Effect of roll damping and sen- sitivity on average pilot rating, L6a, rad/sec2/in.

= 0; N, = -3.5 set-‘.

LCJLP Fig. 6. Effect of roll damping and sen- sitivity on average pilot rating, EXP. IV Lq/Lp < 0.3; Nr = -1.2 set-'.

60 knot SLALOM TASK = -3.5 set-’ Nr L6, = 0.55 rad/sec2/in.

EXP.1 PILOTAm,PILOTB 0 II AVERAGE PILOT RATING 0 7- III AVERAGE PILOT RATING 0 INADEQUATE _._-- ---- COMBINATION TASK 6- ADEQUATE

R E

h ,--------,r- ---- -- BOUNDARY SATISFACTORY (REF. 2, 17) I I I .25 .50 .75 I.80 -.60 -.40 -.20 0 .20 .40 .60 .80 RATIO OF COUPLING IMplMql, ILq/Lpl LqfLp Fig 8. Pilot rating vs. Lq/Lp. Fig. 9. Trends of pilot rating with ratio of coupling (from ref. 7).

I I llllllllllllllllllllllllll lllllllllllllll II II II I lllll

EXP. II 0 ARTICULATED ROTOR HELICOPTER . TEETERING ROTOR HELICOPTER

r

UNACCEPTABLE PI LOT OA 7- AB 0 c g 6- ACCEPTABLE 8r (BUT UNSATISFACTORY) E =.- 4 0

-1 -0

3- SATISFACTORY 2- l-

PITCH-ROLL - - -an)

BASIC SATISFACTORY A/C DECOUPLING PITCH-ROLL - I I I I I J AUG. DECOUPLING 2M sc X 2 BASIC Mg, = 6 Mgc = 6 COLLECTIVE N 6c = 0 TO YAW Fig. 11. Effect of pitch and yaw due to Fig. 10. Effect of pitch-roll coupling and yaw resulting from collec- collective input on pilot rating, tive input on pilot rating. hingeless rotor, all pilots.

6l EXP. II HINGELESS ROTOR Cl ARTICULATED ROTOR TEETERING ROTOR n V EXP. III HINGELESS ROTOR A EXP. IV TEETERING ROTOR (NO COLLECTIVE INPUT DECOUPLING;) 10 - 9- UNACCEPTABLE 8- A 7- QT gfj-: - ACCEPTABLE 5 5- (BUT UNSATISFACTORY) : 4- A &y;-A++ 3- SATISFACTORY 2- ( I I I I I I I I I -50 ATTITUDE COMMAND ’ BASIC RATE COMMAND 0 2 4 6 8 10 12 14 16 18 20 SCAS INPUT SCAS INPUT A/C LOCK NUMBER, 7 DECOUPLING DECOUPLING Fig. 13. Pitch-flap coupling required to Fig. 12. Effect of SCAS mode on pilot decouple tip-path plane tilt rating, pilot A.

for extreme values of flapping restraint and hinge offset.

cc WITH PITCH-FLAP COUPLING -0 BASED ON NO PITCH-FLAP --0.3 COUPLING DECOUPLING RULE f -l.Ol 8 8 v 1 ’ ’ I I , 0 4 8 12 16 20 0 4 8 12 16 20 LOCK NUMBER, ‘7 Fig. 14. Effect of decoupling rule on Lq/L,.

E = 0.05 WITH 63 ACCORDING TO DECOUPLING RULE +y = 12 -.- WITHOUT s3 SL = 30 radlsec ADVANCE RATIO HOVER p = 0.3 LATERAL TPP TILT, rad s-------e------ .08 .---*-----*- LONGITUDINAL TPP TILT, rad -.08 1 I 0 .iO .40 .60 0 .20 .40 TIME, set Fig. 15. Effect of decoupling rule on TPP transient response to 1 rad/sec step change in roll rate.

HINGELESS ROTOR HELICOPTER -%,, % G.W. 4630 lb, MID C.G.. 60 knots it-lb/in. ft-lb/in.

SEA LEVEL STD -%,. %c.

lb/in. ft-lb/in.

TF VI ‘200 1200 + UPWARD DOWNWARD C I T I -- -2.0 -1.0 1.0 (T:lM, A tic, in.

Fig. 16. Nonlinear effect of collective control derivatives.

t

L R .

k

I I I 4 I I I 0 10 20 30 40 50 BANK ANGLE, deg Fig. 17. Turn radius vs. bank angle in a slalom course.

An Assessment of Various Side-Stick Controllerjstability and Control Augmentation Systems for Night Nap-of-Earth Flight Using Piloted Simulation Kenneth H. Landis Engineering Specialist Boei n9 Vertol Company Philadelphia, Pennsylvania Edwin W. Aiken Aerospace Engineer Aeromechanics Laboratory U. S. Army Research and Technolooy Laboratories (AVRADCOM) Ames Research-Center Moffett Field, California ABSTRACT PFCS Primary Flight Control System PNVS Pilot Night Vision System A series of piloted simulator experiments was conducted to assess the interactive effects of Stability and Control Augmenta- SCAS side-stick controller characteristics and level of tion System stability and control augmentation on attack Several night helicopter handling qualities. Side-Stick Controller ssc nap-of-the-earth mission tasks were evaluated Visual Meteorological Conditions VMC using a helmet-mounted display which provided a limited field-of-view image with superimposed flight control symbology. A wide range of sta- INTRODUCTION bility and control augmentation designs was in- vestigated. Variations in controller force-de- The Army’s Advanced Digital/Optical Control flection characteristics and the number of axes System (ADOCS) Program is aimed at developing controlled through an integrated side-stick con- a battlefield-compatible advanced flight control troller were studied. In general, a small dis- system which can substantially increase aircraft placement controller was preferred over a stiff- mission effectiveness in part through decreased stick controller particularly for maneuvering pilot workload and improved handling qualities.

flight. Higher levels of stability augmentation The objectives of the program are: (I) the were required for IMC tasks to provide hand- development of the technology required for a ling qualities comparable to those achieved for digital optical flight control system, (2) the in- the same tasks conducted under simulated visual tegration of the new technology with advanced flight conditions. flight control concepts into a demonstrator air- craft, and (3) the demonstration of the advan- NOTATION tages of the system in the areas of: mission effectiveness, handling qualities, flight safety, ‘AAH Advanced Attack Helicopter cost, weight/volume, survivability/vulnerability, and reliability/maintainability. The program is ACC/AFCS Advanced Cockpit Controls/ divided into two phases: the first involves the Advanced Flight Control System development of component technology for a digital optical flight control system while the ADOCS Advanced Digital/Optical Control second is devoted to the development of the System ADOCS demonstrator system. The first flight of the demonstrator aircraft, a UH-GOA Black BUCS Back Up Control System Hawk, is scheduled for the fall of 1984.

DOCS Digital Optical Control System This paper presents the results of a conceptual FLIR Forward-Looking Infrared design and piloted simulation study of the cock- Heavy Lift Helicopter HLH pit controller configuration, flight control laws, and display logic required to achieve satisfac- Integrated Helmet and Display I HADSS tory handling qualities for the mission defined Sighting System for the ADOCS demonstrator aircraft: an attack helicopter mission conducted under both day and Instrument Meteorological Condi- IMC night/adverse weather conditions. The simula- tions tion, as part of the Advanced Cockpit Controls/ Advanced Flight Control System (ACWAFCS) NOE Nap-of-the-Earth element of the ADOCS program was conducted using the Boeing Vertol Flight Simulation Facil- ity. Although both day VMC and night IMC missions were simulated, this paper emphasizes the low-speed night NOE segments of the ADOCS mission. ,and, assesses the interactive effects on handling qualities of the integrated side-stick interactive effect on handling qualities of each controller characteristics, flight control laws, variation in an element of the system is keot in and helmet-mounted display symbol dynamics. perspective. A discussion of important iisues to be considered within each primary system EXPERIMENT DESIGN element follows, including specific details about the controller/SCAS/display characteristics eval- Pilot workload and the level of performance uated.

achieved during a specific attack helicopter mis- sion task are influenced by combined elements SCAS DESIGN- of the helicopter control/display system design.

COMMAND/STABILIZATION The primary elements considered during this CHARACTERISTICS simulation program were: LV/LV Side-stick Controller (SSC) Configuration -

(1)

Stiff or displacement type, and level of in- ATILV tegration ranging from a fully-integrated four-axis side-stick controller to a 2+1+1 AT/AT arrangement; i.e., a two-axis side-stick for pitch and roll control with small-dis- placement directional pedals and collective lever.

Stability and Control Augmentation System (2) (SCAS) Characteristics - Several generic types of feedback stabilization and feed- forward command shaping in each of the four control axes (pitch, roll, yaw, and vertical).

Visual Display - Either day VMC with the CAXIS CONTROLLER (3) simulator four-window, wide angle field-of- ATTITUDE CMD/Al-fITlJDESTAB view visual system, or night IMC using a simulated FLIR image and superimposed YAH-64 Pilot Night Vision System (PNVS)l symbology presented on a helmet-mounted display.

Figure 1 Three-Dimensional Flight Control GENERAL APPROACH System Description The approach to the systematic investigation of these elements is illustrated in Fiqure 1. The INTEGRATED SIDE-STICK CONTROLLER overall investigation was directed toward defin- ing those combinations of SSC, SCAS, and dis- Fly-by-wire or fly-by-optics flight control sys- tems allow flexibility not only in the synthesis play that produce Level 1, 2, and 3 handling of the control laws but also in the design of the qualities ratings2.

pilot’s controllers. The potential benefits of In applying this general approach to the spe- employing an integrated, multi-axis, side-stick the blocks defined in Figure 1 controller include: cific problem, improved visibility, enhanced crashworthiness, were broken down further into more detailed easier ingress and egress, a For example, each side- reduction in cockpit space requirements, and an ,configuration matrices.

increased potential for single-pilot operations.

stick controller configuration block contains var- iations in force/displacement relationships as Generic con- well as ergonomic characteristics. Related Research and Development Programs trol laws can be mechanized in several different Dis- Handling qualities research examining the effects ways with significantly different results.

of the characteristics of a two-axis side-stick play symbology involves a myriad of variations in parameters, format, scaling, and logic. controller was conducted in support of th.e development of the F-16 aircraft.

In a flight investigation Degraded modes can also be visualized in Fig- of the effects of variations in force-deflection ure 1. Since the selected controller configura- characteristics for certain fighter aircraft tasks3, ,tion will be part of the primary flight control it was concluded that a all allowable degraded modes will lie in small amount of side-stick motion provided im- system, For example, the control-law/display-law plane. proved flying qualities over those achieved with a fixed controller.

certain failures such as FLIR loss will affect the The results of this and while loss of a ground veloc- other similar flight experiments were incorpo- display axis only, rated in a design guide for two-axis side-stick ity signal may seriously affect the system con- controllers used in fighter aircraft4; included in trol law and display symbology.

the guide are recommendations for stick neutral ‘By considering the overall system design as a position, breakout forces, and force-deflection characteristics series of matrix levels of increasing detail, the in both the longitudinal and lateral axes.

and vertical) on the side-stick con- troller, Research involving the use of side-stick con- trollers in Army helicopters began in 1968 with Three-axis side-stick for (2) 3+1 (Collective): the Tactical Aircraft Guidance System (TAGS) pitch, roll and yaw control, and a programs. The system implemented in a CH-47B separate collective lever for vertical aircraft initially included an integrated four-axis control, large-displacement controller. Because of cou- pling problems between the longitudinal and side-stick (3) 3+1 (Pedal): Three-axis for vertical axes, a three-axis controller was even- roll and vertical control, and pitch, tually implemented with vertical control effected pedals for directional control, and through a standard collective lever. On the Heavy Lift Helicopter (HLH)“, a four-axis dis- (4) 2+1+1: Two-axis side-stick for pitch and placement controller was implemented at the roll control, with separate collective load-controlling crewman’s station in conjunction lever for vertical control, and pedals with a ground velocity command and stabiliza- for directional control.

tion system.

Side-stick control of single-rotor helicopters has been implemented in a production aircraft - side- stick cyclic control at the copilot’s station of the AH-l series of aircraft - arid investigated using both ground- and in-flight simulation. In a three-degree-of-freedom moving-base simula- tion of the unaugmented Lynx helicopter at RAE Bedford, a two-axis displacement side-stick was compared to the conventional cyclic controller for eleven different flight tasks7. When a suit- able control sensitivity was selected, the side- stick compared favorably with the conventional controller and, in fact, was preferred for cer- tain of the tasks.

A feasibility study of a four-axis isometric side- stick controller was recently conducted in the Canadian National Aeronautical Establishment a variable stability Bell Airborne Simulator, Model 205A-1, for a wide range of flight tasks’.

Two primary side-stick configurations, a four- axis controller and a three-axis controller with normal pedal control, were evaluated together with three SCAS variations: rate command/ attitude hold in roll and pitch with augmented Figure 2 Controller Configurations yaw rate damping; augmented rolt, pitch and rate damping; and the basic 205 with yaw stabilizer bar removed and horizontal stabilizer Force/Deflection Characteristics fixed. With appropriate gains, shaping, and prefiltering applied to the pilot’s force input in A definition of acceptable/unacceptable ranges each controlled axis, pilot ratings comparable to of force/deflection gradient for each controller controllers those obtained with conventional configuration option (4+0, 3+1, or 2+1+1) was were achieved by both primary side-stick con- necessary. The determination of force-deflec- figurations.

tion characteristics was performed using three 4-axis side-stick controllers: These investigations indicate that a comprehen- sive evaluation of multi-axis side-stick control (I) A stiff-stick force controller, for an attack helicopter mission must include variations in: 1) the number of axes controlled (2) A small-deflection controller with two force/ through the side-stick device, 2) the force-de- deflection gradient configurations, and flection characteristics of the controller, and 3) the attendant SCAS characteristics.

(3) A large-deflection controller with an as- sortment of springs which provided inde- Level of Integration (Number of Axes) pendent adjustment of force/deflection gradients and breakout forces in each Four variations in controller configuration rep- axis. This controller is a modified load- resenting different levels of controller integra- controlling crewman’s controller used tion were investigated. Figure 2 shows the during the HLH program.

controller configurations including: All controllers are a base-pivot type for pitch

(1) 4+6: All control axes (pitch, roll, yawl

and roll motion. Fore-aft force produces longi- tudi’nal control input and right-left force a lat- LONGITIJD~NAL eral control input. Yaw control is obtained by 13- twisting about the grip centerline, and vertical, control through application of pure up and down 12- forces. Figure 3 shows the three controllers.

SMALL OEFLN (I) F-16 14.4 LWJEG (AFT) MSI STIFF STICK 3 Four-Axis Side-Stick Controllers Figure APPLIED FORCE (LB) The selection of pitch and roll force/deflection gradients was guided by a review of previously Figure 4 Longitudinal Axis Force-Deflection described published data. References 4 and 9 Characteristics defined preferred regions of longitudinal and lateral force/deflection gradient developed from Air Force flight test evaluation of a two-axis variable force-deflection side-stick controller.

Figure 4 shows the recommended force/deflec- in addition to five specific tion gradient range, longitudinal controller force/deflection configur- ations evaluated during this study. The grad- ients were chosen to cover a range from a “stiff” force gradient with very small deflection to a “soft” force gradient with large deflection (?I2 degrees). The F-16 side-stick controller design is also shown for comparison.

Complete force/deflection characteristics for the five 4-axis controller configurations utilized Table 1 Four-Axis Controller Force- during this simulation are presented in Table 1.

Deflection Characte?lstics Operating force range, maximum deflection, and force/deflection gradient are given for the four Yaw and vertical controller com- STABILITY AND CONTROL AUGMENTATION control axes.

pliance for both small-deflection configurations SYSTEM (SCAS) CHARACTERISTICS were relatively “stiff” compared to the pitch and roll axes. In contrast, the medium- and large- The segments of the attack helicopter mission deflection configurations were evaluated with considered to be critical from a handling quali- ties point-of-view are those spent in nap-of- lighter yaw and vertical force/deflection gradi- the-earth (NOE) flight; those inherently high ents for harmony with pitch and roll.

workload tasks include low-speed point-to-point Evaluation of the (3+1) collective, (3+1) pedal, maneuvering using dash, quick stop, and side- and 2+1+1 controller configurations was per- ward flight techniques, masked hover in ground formed using a conventional collective lever and effect, and unmasked hover out of ground effect The simulator vari- directional pedal controls. including target search, acquisition, and weapon delivery. This simulation was designed to pro- able force-feel collective lever was implemented vide a preliminary definition of flight control as a “stiff” force controller with small deflection.

A pedal force control system was configured laws and SCAS mode switching logic require- using a mechanical spring capsule attached di- ments for the various mission phases. In addi- tion, the effects on both handling qualities and rectly to the pedals. The directional pedal con- flight safety of degraded SCAS modes were to figuration selected had a force/deflection grad- be determined. The effect of the side-stick con- ient of 40 Ibs/inch with a force breakout of 6.0 troller configuration under degraded SCAS mode Ibs.

conditions is important, since high levels of o Use of a control response model provides vehicle stability may mask undesirable charac- forward loop commands to tailor the short teristics of some controller options. SCAS re- and long term responses to pilot control dundancy requirements also need to be weighed inputs as required to achieve satisfactory in final selection of a controller configuration.

pilot ratings and performance. Any desired For example, a 3+1 axis controller configuration control response can be obtained by appro- requiring only rate stabilization may be more priate feed-forward shaping regardless of cost effective than a 4-axis side-stick controller the level of stabilization.

requiring attitude stabilization to achieve Level 2 handling qualities.

o Pilot display symbology is driven by the same sensor set used for flight control.

Results presented herein are for the first of For some failure modes, redundant signals two scheduled simulation phases, which concen- may be available in the AFCS as backup trated on the low speed portion of the NOE inputs to the symbology display.

mission, that is, airspeeds below approximately 50 knots. High speed control laws and transi- tion requirements will be evaluated during the second simulation phase.

Figure 5 presents a block diagram of the flight control system design concept developed for the ADOCS Demonstrator Program. The use of this system formulation allows for development of handling qualities requirements while still con- sidering aspects of hardware design and redun- dancy management. Major advantages of this system design concept are: L--------t--~----- I o Satisfactory unaugmented flight is attained by providing feed-forward command aug- mentation and shaping as an integral part of the primary flight control system (PFCS) Control mixing and prefiltering are in- cluded in the PFCS to reduce pilot work- load to an acceptable level for unaugmented Figure 5 ADOCS Demonstrator-Flight flight.

Control System Concept o Stabilization feedback loops are optimized solely for maximum gust and upset rejec- Various control system concepts were formulated tion. This allows use of high full-time to accomplish the attack helicopter low speed/ stabilization gains required for good atti- hover maneuvers. The generic SCAS configura- tude or velocity hold during NOE maneu- tions chosen for evaluation are identified in Fig- ure 6 in the form of a comman’d response/sta- vering or tight position hold for precision hover tasks . Also, aircraft attitude ex- bilization matrix. A simple identification code (Figure 6) was established. For example, a cursions are minimized for improved target system with angular rate command and attitude acquisition and weapon delivery. No com- promise for control response is necessary. stabilization in pitch and roll was identified with IDENTIFICATION COOE PITCH/ ROLL YAW VERTICAL VERTICAL ANGULAR ACCELERATION AC y - ANGULAR RATE RA 5, - ANGULAR ATTITUDE AT h : h LINEAR ACCELERATION LA - NA h’ LINEAR VELOCITY LV - LINEAR POSITION LP - hH EXAMPLE: AA/AT ANGULAR RATE COMhlANO/ATTlTUOE STABILIZATION 3NH YAW RATE COMMANOlHEAOlNG HOLD Figure 6 Generic SCAS Configurations-Command Response/Stabilization Matrix the letter code ‘RA/AT. It should be noted that A six degree-of-freedom small-perturbation model the longitudinal and lateral control axes were of the helicopter was used to develop the com- always evaluated with the same command re- mand response model for each axis. The analy- sponse and stabilization system. tical study established control response model .gains for cancellation of undesirable roots of The method of SCAS implementation used for the vehicle characteristic equation. Control re- the simulation is illustrated in Figure 7 for the sponse model feedforward parameters were de- lateral axis.

All control axes were implemented fined for each of the response types previously in a similar manner. described.

The stabilization gains For example, Figure 9 shows the shown on the diagram were selected prior to the lateral response to step force input for a rate, piloted evaluation phase using the helicopter/ attitude, and velocity response type. For the stability augmentation system model shown in angular rate command model, identical response Figure 8. characteristics were provided for both the atti- Elements of the model include trans- fer functions to represent the dynamics of the tude and velocity stabilized systems.

Similarly, basic helicopter, rotor and actuators as well as the attitude response model characteristics were a computational time delay. Nichols chart tech- the same regardless of the level of stabilization.

niques were used to select feedback gains.

Multiple feedback paths, each increasing overall During this preliminary control response design stability, were closed around the model based process, information from available literature, on a damping ratio design criteria (< = 0.7). as well as related experience, was used to de- The stabilization loop gains derived by this velop design criteria and quantitative guidelines.

method were similar to gains of previously’ Design guidance for SCAS intended for low developed aircraft systems (i.e., TAGS, HLH). speed and hovering flight is contained in Refer- ence 10 which develops tentative VTOL aircraft flying qualities criteria from the existing experi- mental data base. Requirements for generic SCAS such as angular rate command, attitude command, and translational rate command are proposed together with suggested vertical aug- mentation system characteristics. In addition, I I IDI11.1bOl214,6ll TIME ,SEC, TIMElSECl TIMEWC, Figure i' Lateral Axis Stability and Figure 9 Lateral Response to Step Force Input Control Augmentation System UPPER TIME AQOCS BOOST ACTUATOR ROTOR PLANT DELAY ACTUATOR (0.067s + II IS2 + 29.4s + 26741 IS’ + LX + 2381 L” I LP I Figure 8 Helicopter/Stability Augmentation System Model-Longitudinal Axis deicribed and’ the symbols which appear for the the use of velocity command system for the pre- three low-speed mission modes used during this cision hover task was flight demonstrated on investigation are identified.

the HLH Program (References 11 and 12), and the desirability of this control concept was con- In a simulator investigation of a night-time firmed based on study results published in attack helicopter mission which included a head- References 13 and 14.

up display of the PNVS symbology13, it was The preliminary analytical study established found that the dynamics of the symbology used baseline response characteristics to begin piloted to aid the pilot in achieving a precision hover Final characteristics evaluations. response at night had a significant effect on the handling developed during the initial phase of simulation qualities of the vehicle. As a result, because are presented in the Experiment Results section.

of the wide variation in candidate SCAS concepts to be investigated, it is necessary also to en- iw DISPLAY (IHADSS) sure compatibility of the symbol dynamics with the varying dynamic characteristics of the aug- Since the ADOCS mission is to be flown in mented helicopter.

night/adverse weather conditions as well as in VMC, it is necessary to consider not only the Variations to the baseline AH-64 symbology were effects of the controller and SCAS characteris- made based on Reference 13 as well as a review tics but also the impact on handling qualities of of reported display system characteristics imple- the pilot’s night vision aids. For this program mented on the PNVS surrogate trainer flown at it is assumed that the pilot is provided with the the U.S. Army Test Proving Ground, Yuma, AH-64 Pilot Night Vision System (PNVS) and Changes were incorporated in the Arizona.

include a helmet- associated avionics1 which programmed symbology primarily to improve low mounted display of flight control and fire con- speed maneuvering and hover hold task perfor- limited trol symbology superimposed upon a as well as to reduce pilot workload.

mance, field-of-view monochromatic image of the outside These changes, evaluated during the prelimi- world slaved to the pilot’s head motions.

nary IHADSS check-out testing, were as follows: The display system selected for simulation of Velocity vector sensitivity was decreased

(I)

the IMC mission is the Honeywell Integrated by a factor of two for all modes - from 6 Helmet Mounted Display/Sight System (I HADSS) knots to 12 knots full scale in the hover developed for the Army’s YAH-64 Advanced and from 60 knots to and bob-up modes, Attack Helicopter (AAH). The IHADSS permits 120 knots full scale in the transition and NOE, low level, and contour flight under IMC.

cruise modes.

Since the Helmet Mounted Display (HMD) is Hover position sensitivity was decreased (2) coupled to the pilot’s head, he is able to scan a for the bob-up mode from a full scale de- wide field-of-view without being constrained to flection of 44 feet to 88 feet.

a head-down or look-forward position. The pilot’s line of sight is tracked with a Helmet A horizon line was included in the symbol- (3) Mounted Sight (HMS) that provides closed-loop ogy format for all modes. The AH-64 has command signals to point the sensors.

the horizon line in the transition and cruise modes only.

The importance of superimposed flight control Lateral acceleration was used to drive the (4) symbology to the enhancement of handling quali- “ball” display instead of sideslip angle to ties with a limited field of view FLIR image of augment the simulation turn coordination the outside world has been reported in Refer- cues at low speed.

ence 15. Baseline display laws and information format used for this investigation were defined The cyclic director, or longitudinal and (5) based on the AH-64 Pilot Night Vision System lateral acceleration cue, approximated by (PNVS)l. The selectable display modes, which washed-out pitch and roll attitudes, re- are used to meet the operational requirements quired different sensitivity and time con- for various AAH mission tasks, are: stant values as a function of the command response system type, i.e., rate, attitude, Cruise: high-speed level flight enroute. to (1) or velocity. Values were established in the forward edge of the battle area; the same manner discussed in Reference 13.

Transition: low-speed NOE maneuvers sucl- (2) as dash, quick stop, and sideward flight; EXPERIMENT ACTIVITIES Hover: stable hover with minimum drift; (3) and To reduce the large number of possible SSC/ SCAS combinations to a manageable set of con- Bob-up: unmask, target acquisition, and (4) figurations for evaluation, the experiment was remask maneuvers over a selected ground designed with two major. phases of simulation position.

activity as shown in Figure 11. Phase I accom- plished IHADSS familiarization and controllet: Figure 10 presents the display mode symbology Phase 2 concentrated on evalua- development.

divided into three categories - central, periph- tion of controller/SCAS configuration combina- eral, and weapon delivery/fire control symbol- tions.

The characteristics of each symbol are WY.

MODES INFORMATION CRUlSElTRANS HOVER BOB-UP CENTRAL SYMBOL I Fixed reference for horizon line velocity 1. Aircraft reference vector. hover position, cyclic director, X X X

I I and fire control svmbols

Pitch and roll attitude with respect to 2. Horizon line aircraft reference (indicating nose up X X X pitch and left roll)

I I

Horizontal Doppler velocity components 3. Velocity vector (indicating forward and right drift X X X

I velocities). Senritivitv varies with mode

HORIZON LINE 4. Hover position Designated hover position with respect to aircraft reference symbol IindicatIng aircrafr X X

I forward and to right of desired hover position)

~~~---- Cycltc stnck command with respect to hover 5. Cyclic director posntion symbol (indicating left and aft cyclic (Acceleration Cue1 stick required to return to designared hover X X position). Approximated by washed out pitch/roll attntude MODES INFORMATION CRUISE/TRANS 1 HOVER 1 BOB.“P PERIPHERAL SYMBOL W 30 33 N 3 6 E 6. Aircraft heading Movnng tape nndlcatoon of heading (nnd~car

I I I I I I I

,nq North1 6 7 50 m 12 1 200 Headong at f,me bob up mode selected Head’ng’error 1 (indncatins 030, I I Ix1 Radar altitude 1 Hedght above ground level I” both analog and

I 6. dnantal form (indicatnna 50 ft)

40 11 9 100 Rate of Cl,mb Mowng ponnter wath full scale deflectnon of 9.

9 1.000 ft/mln (nndicatnng 0 fl mlnl

x IxIX

bl 10. Lateral acceleratlo” Inclinometer indlcatlon of snde force X x I x I 10 3 11. AIrspeed Digital readout I” knots T0rCjlle Engine torque I” percent X X X 12.

’ x I x I-+I MODES FIRE CONTROL SYMBOL INFORMATION CRUISE/TRANSI HOVER 1 BOB-UP W 30 33 N 3 6 E 13. Cued line of sight Overlays designated target positon on back- X lllrlllllllll ground video when target is in display field of view 15.

Designated target positron with respect to 14. coarse target display fneld of view (inner rectanglel and X locarion se”5.0, limits tourer rectanglel

I 17 I

IX Designated target bearnng (indicating 330” Target bearing or 30’ to left of current headlnu)

I= I I 1x1

;ci Target location dots lllumnnation of two adjacent dots nndxates 16.

-i- display quadrant in which designated target is located 5= 17. Missile launch Limits wth respect to arcraft reference for con*frainfs successful weapon lock.on to desionated taraet Figure 10 Display Mode Symbology teristics for a conventional collective lever Specific steps followed for controller develop- ment were: and pedals configured as force controllers.

A best 4-axis controller design was selected (1) Evaluation of the 4-axis stiff-stick control- This design was ler to determine best individual axis re- based on the above results.

characteristics and desired used to evaluate the 4+0, (3+1) pedals, (3+1) sponse/force collective, and 2+1+1 configurations for the pri- non-linear response shaping requirements.

mary and secondary controller/SCAS configura- (2) Evaluation of the three 4-axis deflection tion matrices as follows: controllers to define effect of force/deflec- (1) Primary configuration matrix - Variations tion gradient on pilot task performance.

to the pitch and roll SCAS with a fixed (3) Comparison of the stiff-stick and deflection directional and vertical command/stabiliza- controllers for various pitch and roll SCAS tion system (yaw rate command/heading configurations. hold and vertical rate command/altitude hold).

(4) Definition of desired response/force charac- (2) Secondary configuration matrix - Variations o Conventional helicopter flight and perfor- to the vertical and directional SCAS for a mance instruments, and a SCAS mode limited portion of the primary configuration select panel.

matrix with emphasis given to the less highly augmented pitch and roll systems, o Conventional helicopter collective and direc- particularly the rate/attitude (RA/AT) and tional pedals implemented as small-displace- attitude/attitude (AT/AT) systems.

ment force controllers, and three 4-axis side-stick controllers. An adjustable PHASE 1 mounting bracket attached to the armrest allowed orientation of each 4-axis side-stick SlMULATlONlMATH MODEL CHECKOUT controller for comfort and to minimize TASK DEFINITION inter-axis control inputs. A forward tilt IHADSS FAMlLlAAlZATlON of six degrees and a counter-clockwise rotation of five degrees relative to the armrest was selected.

I CONTROLLER DEVELOPMENT lVMCllMCI o Xerox Sigma 9 digital computer to drive I I ALTERNATE the entire simulation. The Sigma 9 was STIFF DEFLECTION CONFIGURATIONS programmed with a UH-60 full-flight envel- ope math model and easily variable SCAS I 1 I 1 1 1 configurations for this study.

(4 + ass (4 + also, f4+olyo (4 + OIL0 13t llc 13+ lip MSI HI HI HI HI MSI o Four-camera wide-angle television/terrain model visual display system for the simu- lation of terrain flight under either: 1 1 I 1 , OETERMINE BEST RESPONSE/FORCE VMC - Four-window cockpit visual dis- CHARACTERISTICS FOR ALL AXES I OEFINE EFFECT OF FORCEIOEFLECTION covering a field-of-view play - - GRAOIENT 125ox75o, or I SELECT4-AXIS DEFLECTION CONTROLLER FOR PHASE 2 IMC - FLIR image with superimposed symbology presented by a Honeywell helmet mounted display and sight system (IHADSS) in- PHASE 2 cluding head tracker.

YAW AXIS ITCH AN0 ROLl The FLIR sensor signal was simulated using the CMO/STAS OMMANO &E center window video channel to provide a 40°x300 TAB SYSTEMS ‘h . outside world field-of-view display. A Gaertner PRIMARY SCAS ihH CONFIGURATION Symbology Generator was utilized to overlay com- 14 + Olss LV/PH SECONOARY

1 S/6

puter generated symbols (Figure 10) on the LVILV SCAS I4 + Olso2 video picture. The ability to compare directly ATILV CONFIGURATION ATIAT (3 + II c VMC and IMC handling qualities with a specific RA/LV controller/SCAS combination was a unique fea- Kl+ Ijp RA,AT ture of this simulation.

RAIRA ,2+,+11 AC/GA Figure 11 Simulation Experiment Flow Diagram CONDUCT OF THE EXPERIMENT EQUIPMENT This experiment was conducted at the Boeing Vertol Flight Simulation Facility. Major elements of the facility shown in Figure 12 include: Figure 12 Boeing Vertol Flight Simulation o Single-seat cockpit cab mounted on a six- Facility degree-of-freedom limited-motion base.

EVALUATION TASK DESCRIPTION Nap-of-the-Earth (NOE) - A multi-axis control task requiring the pilot to fly through three legs of a narrow canyon (125 feet wide and 50 Evaluation of total system (pilot, controllers, SCAS, displays) performance was accomplished feet high), having two sharp turns (70” left using four specific low speed tasks - the slalom, and 80° right) and two obstacles (50 feet high), acceleration/deceleration, nap-of-the-earth, and to reach a termination hover area. During the No secondary duties (e.g., first leg of the course, an acceleration to 50 bob-up task.

communication, or navigation system knots is performed before crossing a road, armament, management) were required during the perfor- followed by a deceleration to 25 knots while mance of each task. For this experiment a maintaining a lateral ground track and an alti- 2OO:l scale model board (l-1/8 mile long by 3/5 tude of 30 feet. After executing a coordinated left turn to enter the second leg, the pilot must mile wide) with an existing 3000 ft airport run- control altitude to fly over an obstacle and re- way was modified as shown on Figure 13 to in- clude terrain features and obstacles necessary mask to 30 feet in as short a time as possible to perform the planned maneuvers. while attempting to maintain an airspeed of 25 knots. Following a sharp right turn, the pilot Low-speed lateral avoidance maneuver flies over a second obstacle, controls altitude Slalom - requiring the pilot to fly around 50 ft. high back to 30 feet, and decelerates to a hover obstacles placed 400 feet apart on the runway point in the termination area.

centerline. From a hover at 30 feet AGL, the pilot accelerates the helicopter to an airspeed of Bob-Up - A multi-axis task consisting of a ver- The pilot appropriately controls bank tical unmask maneuver from 25 feet ‘to 100 feet, 30 knots.

angle and heading to coordinate turns around a heading turn to acquire a target, and a ver- the obstacles while maintaining a constant air- tical remask to the original hover height. The speed of 30 knots and an altitude of 30 feet pilot attempts to hold a fixed horizontal ground throughout the maneuver. position throughout the vertical unmask/remask and heading turn maneuvers.

Acceleration/Deceleration - Forward translation of the helicopter while holding a lateral ground TEST PILOT BACKGROUND AND PARTICIPA- From an initial track parallel to the runway. TION hover position offset from the runway, the pilot Five experimental test pilots with extensive acclerates the helicopter to a forward speed of 50 knots, followed by a deceleration maneuver flight experience participated in this simulation to arrive at a desired hover position near the study - one each from Boeing Vertol, NASA, The pilot attempts to and the National last runway obstacle. Aeronautical Establishment (NAE) of Canada, and two pilots from the U.S.

hold lateral ground track and altitude, as well as complete the task in minimum time. Army assigned to NASA. Table 2 presents an Terrain Model for Evaluation Tasks Figure 13 a4 experience. summary for each evaluation pilot participated the first week, helped to define the including total flight time broken down by heli- specific tasks used for the remainder of the copter and fixed wing time.

After the initial experiment.

phase of simulation development, two pilots (A and B) were given 3 hours of IHADSS flight DATA COLLECTION AND ANALYSIS training on the PNVS Surrogate Trainer at the U.S. Army Yuma Proving Ground.

Experimental data collected for this investigation consist of both qualitative pilot evaluation data and quantitative system performance data. Pilot Cooper-Harper ratings and commentary were re- corded for each controller/SCAS/display/task combination evaluated. At the end of each eval- uation run, the pilot assigned a numerical Cooper-Harper rating to the task according to a structured decision making process defined by x Reference 16. The pilot’s comments were used to aid data analysis by identifying areas or parameters that most strongly influenced each

.L

rating.

Qualitative pilot rating data is emphasized in Table 2 Summary of Pilot Experience this paper. Quantitative measures of system performance and/or pilot workload are being cal- culated using statistical analysis programs. For A total of 204 simulation flight hours was accu- instance, the mean and standard deviation of mulated during this simulation experiment.

helicopter flight parameters relative to a refer- Sixty-three percent of the total time was utilized ence position or desired flight path are being for VMC evaluation, and thirty-seven percent computed as a measure of system performance.

for IMC evaluation. A breakdown of the total As an indication of pilot workload, the mean and hours by controller configuration and pilot is standard deviations of control command move- given in Table 3. Pilots A, B, and C were the ments are being analyzed.

Certain time indices primary evaluators with Pilot A having the are also being evaluated as an indication of largest flight time (54%) since he participated helicopter/pilot performance. Where applicable, during all eight weeks of the experiment.

Pilot ’ time to perform the entire task or portion of a B participated for three weeks of the study, task, i.e., unmask time, is used as a perfor- and pilot C participated for four weeks. Pilot mance index.

who had significant IHADSS experience on D, the AH-64, participated for one week and OTHER EXPERIMENTAL CONSIDERATIONS assessed the realism of the simulated IMC system compared to real life hardware. Pilot E, who Certain factors which might affect the outcome of the evaluations were identified. An effort was made, where possible, to account for these CONTROLLER CONFIGURATION VMC factors.

Specific examples are given below: . 4 AXIS STIFF STICK 33 . 4 AXIS SMALL DEFLN (1) 19 (I) The exact stabilization level selected for . 4 AXIS SMALL OEFLN (2) 11 each evaluation run was not revealed to . 4 AXIS MEDIUM DEFLN 5 the pilot.

. 4 AXIS LARGE OEFLN . 3 AXIS + PEDALS (2) The command response type (e.g., angular . 3 AXIS + COLLECTIVE 13 pitch/roll rate versus attitude) was revealed to eliminate surprises and to reduce effects . (2+1+1) PEDALS + COLLECTIVE 14 on pilot rating and performance caused by . CONVENTIONAL 7 re-learning a certain response character- TOTAL HOURS 129 istic.

(3) Established habit response patterns occa- PILOT FLIGHT HOURS 1 VMC 1 IMC 1 TOTAL 1

I

sionally had a noticeable effect when 68 42 110 changing to a different controller config- 21 13 34 uration. For instance, after many years 26 12 40 of flying conventional pedals, an adjust- 6 5 13 ment period to adapt to control of yaw 4 3 7 from the side-stick was common for all pilots. Likewise, after flying side-stick I TOTAL HOURS I 129 I 75 1 204 I twist to control yaw for several flight hours, converting back to the pedals was not always done with ease. A similar effect was noticed when switching vertical con- Table 3 Summary of Simulation troller configurations, that is, changing Flight Hours from side-stick to conventional lever or TASK: SLALOM vice-versa. If any. configuration change SCAS CONFlWIATIOI: AT/LV resulted in poor performance, the pilot re- peated the run and the best one was used DATA FROY ALL PILOTS for valid data.

Learning the IHADSS concept and symbol- (4) NOTATION: ogy took a significant period of time.

The 0 (4 +o)ss 4 (4+0) MO rate of improvement of pilot ratings with n (4+O)Sol l (4+O)LD IMC simulation flight time was much slower than for VMC flight time. IMC data pre- sented in this paper were obtained during the second simulation phase when the pilots demonstrated a more consistent level of proficiency with I HADSS.

EXPERIMENT RESULTS 4- Experimental results are based on an analysis of pilot ratings and comments, and discussion of 3- these results is organized according to the major activity phases - controller development and OVER CONTROLIN ROLL primary-secondary matrix evaluation. Results ROLL AXIS TOO SENSITIVE are summarized using average pilot ratings to I I I indicate general trends; the statistical validity 1 2 3 4 5 6 7 6 9 10 11 12 of this simplified approach is not implied and it ROLL ATTITUDE SENSITIVITY (OEWLB) is understood that care must be used in the interpretation of results, particularly when a large range of ratings is averaged.

Figure 14 Control Response Selection Process CONTROL RESPONSE CHARACTERISTICS Before different controller configurations were sitivity range. Best ratings were achieved with the small-deflection and medium-deflection con- evaluated, a set of control response character- trollers in the range from 5.5 to 7.5 degrees/lb.

istics for the four control axes and the generic system types (Figure 6) were defined through a series of mini-experiments. Response time con- The same procedure was followed to select pitch/ stants and sensitivities were varied within the roll rate and longitudinal/lateral velocity re- sponse characteristics. Table 4 summarizes the command model and effects on controllability final selected control response characteristics.

evaluated. A set of best response values was selected, Except for the acceleration command response, initially for the stiff controller, and the same set of values was then evaluated using characteristics are approximated by an equiva- the three alternate 4-axis deflection controllers. lent 1st order system response. The pitch and Additional variations were made about the roll acceleration response system was designed to provide a short-term rate response, with a nominal response values to define the effects on long-term acceleration response to automatically pilot ratings and task performance.

eliminate steady control forces required for This control response selection process is de- helicopter trim. This trim function was accom- picted by Figure 14. Roll attitude sensitivities plished with a low-gain integral feed-forward were evaluated for the slalom maneuver with the path. Higher integral feed-forward gains were Pilot comments in- used in the yaw and vertical axes to obtain various 4-axis controllers.

dicated a range where the roll control sensitivity purer acceleration command responses as indi- was too high producing a tendency to overcon- cated in Table 4 by the ratio of steady-state to trol. In contrast, low roll attitude sensitivities initial response.

less than 4.0 degrees/lb. resulted in heavy con- To provide acceptable response characteristics trol forces and sluggish response characteristics.

The best pilot ratings were obtained when all for small precision control tasks and large controllers had a roll attitude sensitivity of maneuvers, as well as to minimize the effect of approximately 6.0 degrees/lb. Figure 14 also inadvertent inter-axis control inputs, non-linear shows that pilot ratings of the large-deflection control response shaping (Figure 15) was used.

controller were generally degraded compared Each force command signal was passed through with the other configurations, and demonstrated a shaping function that allowed variation of a rapid degradation as control response sensi- deadzone, initial sensitivity gradient, break- tivities were reduced and/or control forces point, and high sensitivity gradient. Pitch, became heavy. The same tendency to degrade roll, and yaw control response shaping was quickly was evident with the stiff-stick. The .

symmetrical, whereas the vertical control shap- small-deflection controller was much more ing was asymmetric with a smaller breakout and tolerant to changes in sensitivity as indicated higher response sensitivity in the down direc- by the relatively shallow slope in the high sen- tion.

.

SENSITIVITY/POUND (SENS) AND TIME CONSTANT (TC) AN( LAR RESPONSE ACCEL (AC) RATE 1RA) STEAOY INITIAL STATE SENS SENS AXIS CONTROLLER (OEG/SEC*) (OEG/SEC21 LONGITUDINAL SIDE-STICK 4.0 0.2 LATERAL SIDE-STICK 10.0 1.0 3.5 1 0.25 0.4 SIOE-STICK 1.6 4.1 2.2 OIAECTIONAL PEDALS 0.6 0.44 -.,-- :

+5.0

SIUESTICK

-1.1

VERTICAL + UP NA - DOWN COLLECTIVE f1.6 LEVER

I

Selected Control Response Characteristics Table 4 Figure 15 Force Control Response Shaping Figure 16 E.ffect of Side-Stick Controller CONTROLLER DEFLECTION/FORCE GRADIENT Deflection/Force Gradient on Pilot Ratings Various side-stick deflection/force gradients were evaluated using the 4-axis stiff controller Stiff Controller: and three 4-axis deflection controllers described Task performance with each earlier in Table 1.

0 “Defining best control sensitivities was more controller was rated for both rate and attitude Figure 16 difficult and more critical with a stiff con- command systems in pitch and roll.

troller than deflection controller.” shows the best pilot ratings obtained as a func- tion of controller average deflection/force gra- dient. The small-deflection and medium-deflec- 0 “Inter-axis force harmony/sensitivities ap- tion controllers achieved the best pilot ratings. peared to ‘be more critical, especially dur- ing larger amplitude maneuvering.” Commentary from three pilots who compared the stiff-stick and small-deflection controllers was 0 “Tendency to over-control, particularly very consistent. All agreed that task perfor- during high frequency manipulative control mance improved substantially with the introduc- tasks. ‘I Typical comments were as tion of deflection.

follows : “Tendency to release forces abruptly and create inadvertent sharp acceleration re- sponse.” Small-deflection Controller: Two pilots evaluated the large-deflection con- troller and gave degraded ratings compared to 0 “This controller has a softer feel of actua- the small-deflection controller. Comments indi- tion than the stiff controller, and control cated a more sluggish pitch control response inputs seem to be smoother in application.” and less precise control of attitude for high- frequency inputs.

0 “Very noticeable improvement over stiff- Ability Based on these results, a second 4-axis small- stick using the same sensitivities.

to shape control commands during large deflection controller design, (4+O)SDZ, having a 50% amplitude maneuvers and control reversals higher deflection/force gradient, was se- lected for evaluation of the primary and secon- was a major improvement. ” dary controller/SCAS configuration matrices.

“This controller gave an immediate and very obvious improvement in handling qual- PRIMARY/SECONDARY CONFIGURATION MATRIX ities. Subjectively, I felt much more ‘in the loop’. While tendencies to cross couple (compared to stiff controller), A basic primary matrix - consisting of five con- remained they were far depressed below the primary troller and five pitch/roll SCAS configurations - control task and were insignficant. Con- was evaluated for all four tasks under both IMC and VMC.

trol inputs seemed much more natural and, The matrix for the bob-up task also included two although the response seemed to be more velocity command systems - one this effect was quite tolerable.” with velocity stabilization and the other with sensitive, position hold. For both IMC and VMC, a total Acceptance of the medium-deflection controller of 220 possible task/controller/SCAS combina- tions was evaluated.

was mixed. One pilot gave the controller de- graded ratings because height control was diffi- Figure 17 presents a matrix of data gathered cult due to a high force breakout in the verti- A second pilot gave the same con- for the NOE task and performed under IMC with cal axis.

the IHADSS.

troller improved ratings compared to the small- Each matrix element contains an average rating for each oilot who evaluated the deflection controller because he felt more in particular configuration control during large maneuvers. combination, as well as TASK: NOE COURSE MC PRIMARY SECONDARY YAW/VERTICAL SCAS CONFIGURATION: I#&,, h/h (4 + O),,,. PITCH/ROLL SCAS: RAIAl LEGEND: LEVEL 1 SATISFACTORY LEVEL 2 ;. 1::: : :‘;. .. .:: :‘,.., ..: .y..

;, .;,., :.Y III ACCEPTABLE 6; tl, 4;/v, G/-J, LEVEL 3 YAW SCAS !a (3 + 1lP. PITCH/ROLL SCAS: AT/AT UNACCEPTABLE ACIRA RAIAT AAILV AT/AT ATILV PITCH/ROLL SCAS CONFIGURATION I I I I 6;/G G;Iv-J” iJ&J” YAW SCAS Figure 17 Controller/SCAS Configuration Matrices was required to compensate for an inadvertent the number of test data points included in the control input.

average rating. A mean of the individual aver- age ratings in each block is also calculated.

The IMC bob-up task (Figure 18) was essen- Various levels of the handling qualities rat,ing scale are shaded on the matrix to emphasize tially an instrument reference task with neces- sary information such as velocity vector, X-Y where the major change from acceptable to un- position, acceleration cue, and altitude provided acceptable occurs. It can be seen that Level 1 flying qualities was not achieved for the NOE by the display symbology. Marginal Level 2 ratings were obtained with an AT/LV system.

task under IMC for any controller configuration.

Level 1 ratings were achievable with a velocity The interaction of SCAS/controller configurations command system having either velocity or posi- can be determined from the matrix.

An attitude tion stabilization.

command system achieved Level 2 ratings regard- less of the stabilization type for all controller In contrast to the ratings assigned for the other configurations with the exception of the 4-axis .stiff-stick. tasks under VMC, ratings for the bob-up task A RA/AT system exhibited marginal were more degraded.

Level 2 flying qualities for the 2+1+1 and (3+1) This degradation in VMC ratings was caused by lack of good visual space collective configurations.

references at altitudes above 75 feet in the bob-up location. In fact, VMC performance Secondary SCAS matrices are also shown on measured by X-Y position hold during the bob- Figure 17. An improvement from Level 3 to Level 2 ratings occurred when a yaw accelera- up was significantly degraded over the IMC task.

tion command was implemented for directional Because of inadvertent cross-coupled inputs, control in place of yaw rate command for the -the 4-axis side-stick controller received poorer (4+O)SD and RA/AT combination. In contrast, pilot ratings for the bob-up task. Separation the (3+1) pedal and AT/AT combination de: of the controllers, particularly vertical, im- graded to a Level 3 rating when vertical accel’ proved pilot ratings significantly. The best eration command was used in place of vertical ratings were achieved using a (3+1) collective rate command.

configuration combined with a velocity stabilized system.

Average pilot rating data contained in the pri- mary SCAS/controller matrices are presented for The IMC acceleration/deceleration task (Figure the four tasks in Figures 18 and 19. I nterac- 19). orimarilv a sinale-axis lonaitudinal maneu- tive effects of task, controller, and SCAS con- figurations are more easily seen by this method ver. with altit’ude hold and headi;g hold selected, was the easiest of the four IMC tasks. Level 2 of presentation, and are described in the follow- ratings of approximately 4.0 were obtained with ing discussion.

all controllers except for the (3+1) pedal config- uration. Workload and task performance were CONTROLLER/DISPLAY EFFECTS influenced primarily by the following factors: (I) tendency to couple pitch control into side- The NOE task (Figure 18) was the most difficult stick vertical control, vertical control of the low-speed maneuvering tasks. (2) Primary coupling into lateral-directional requiring pilot factors causing higher workload and degraded compensation, (3) pilot disorientation during a flight path performance for the NOE task under IMC were: nose-up maneuver, and (4) poor resolution of (1) inability to precisely control longitudinal/lateral positioning during decelera- height, (2) tendency to couple side-stick vert- tion to hover. Precise control of aircraft ical control imputs into pitch and/or roll, (3) position during the deceleration to hover was difficult coordination of lateral-directional con- difficult due to poor resolution of longitudinal, trol in turns, and (4) tendency to over control thought to be speeds and rate of closure, roll in high workload situations.

caused by the small field-of-view and limited peripheral cues. Small lateral speeds were dif- The most serious deficiency reported was poor ficult to discern from small yaw rates especially height and vertical speed resolution due to the at slow forward speeds.

small field-of-view, lack of peripheral cues, and/or lack of surface texture/picture detail.

Performance of the slalom task (Figure 19) under Weak motion cues as well as a lack of rotor/ IMC with altitude and headina hold selected was drive system noise may have contributed to a primarily a two-axis IateralIdirectional control tendency for overcontrol of the vertical axis.

task. Pilot ratings were degraded by approxi- The pilot had to rely almost totally on display mately one point compared to the acceleration/ information for vertical speed with no accelera- deceleration task. Task performance was tion lead cues.

judged principally on the ability to execute ‘coordinated turns and achieve a desired curvi- The 4+0 axis controller received poorer ratings linear path around obstacles at constant speed.

for the NOE course where collective control in- Primary factors which increased workload and puts were required to clear the obstacles.

degraded pilot ratings were: (I) tendency to Inadvertent inputs to pitch and roll increased couple side-stick yaw control inputs into roll the workload required to maintain airspeed and and/or pitch, (2) difficult turn coordination due flight path control.

Overcontrol in roll was to lack of peripheral cues with the IMC visual occasionally experienced when corrective action display, and (3) tendency to become disoriented NOE COURSE IMC VMC LEVEL 3 \ I I- I I I I I I I I ATILV ACIRA RA/AT RA/LV AT/AT AT/LV RA/AT RA/LV AT/AT ACIRA PITCH/ROLL SCAS CONFIGURATION BOB-UP LEVEL 3 9 L-

c \ LEVEL 2 -\ '11

LEVEL 1 -3 f =2 S*T,SF*CTOrf~ t I 1 I I 1 I I I 1 I RA/AT AT/AT AT/LV LV/LV RA/LV RAfAT RA/LV AT/AT AT/LV LV/LV LV/PH PITCH/ROLL SCAS CONFIGURATION Effect of Primary SCAS/Controller Variations Figure 18 on Pilot Ratings - NOE and Bob-Up Tasks with IHADSS when head movements were made control on the side-stick provided improved to locate desired flight path projection. It was lateral-directional control for IMC.

difficult to distinguish head response from air- craft response.

PRIMARY SCAS EFFECTS - LONGiTUDINAL/ acceleration/deceleration slalom For the and LATERAL tasks, the (3+1) pedal configuration received more degraded pilot ratings than all other con- For the most difficult IMC tasks (NOE and the acceleration command/rate stabili- figurations. If large errors were allowed to Slalom), build up, precise corrective control inputs with zation system (AC/RA) exhibited Level 3 hand- the pedals were difficult to achieve, and over- With the addition of attitude ling qualities.

control of yaw often resulted. Precision yaw stabilization, the RA/AT system received mar-

ACCEL-DECEL

IMC

VMC

g g

LEVEL 3

z 8 UNACCEPT*BIE

9 7

-------------

= f 6 LEVEL 2 9 5 ACCEPTABLE = 4

:%ziz?-Q ‘I--

k

-+.....:.~

LEVEL 1 d Q 2 SATl5FhClORY

K

r I I ACAIA RA/AT AT/AT AT/LV ACIRA RA/AT AT/AT ATfLY 0 (4+O)SS PITCH/ROLL SCAS CONFIGURATION l (a+o)so

SLALOM A (3+1)P

v (3+1)C

VMC

4 2+1+1

MAC

-- LEVEL 3 UNACCEPTABLE LEVEL 2 AcCEPT*BLE - LEVEL 1 S*T,SFACTOPY c I I I 1 I I I AC/RA RA/AT AT/AT AT/LV AC/RA AT/AT RA/AT AT/LV PITCH/ROLL SCAS CONFIGURATION Figure 19 Effect of Primary SCAS/Controller Variations on Pilot Ratings - Acceleration/Deceleration and Slalom Tasks significant degradation in pilot ratings (Figure ginal Level 2 ratings for IMC with high workload particularly noticed in turn maneuvers.

required to achieve adequate performance. It 181, compensation to achieve Pilot workload and was extremely difficult to maintain precise flight lateral-directional coordination were noticeably control parameters (airspeed, lateral ground higher, .possibly indicating an inherent concep- track, sideslip, etc.). Continuous pulse-type tual design problein with this combination (i.e., control inputs were required for best perfor- having the stabilization type more than one mance. When velocity stabilization was com- integration away from the command type).

bined with a rate command system (RA/LV), for all low-speed maneuvering tasks there was a A large improvement in IMC ratings for all tasks l MC was obtained with an attitude command system.

l ALL DATA AVERAGE0 With the same level of attitude stabilization, an l PILOT A ONLV attitude command system (AT/AT) improved pilot ratings an average of one rating point when MEAN Rmrus compared to the rate command system (RA/AT), A similar improvement occurred in the VMC ratings.

Pilot comments indicated that the attitude command system exhibited a noticeably stronger feel of “apparent” stability.

The pilots- felt more continuous in the control loop g X,Y,= INITIAL POSITION N = NUMBEROF DATA POINTS with a strong force/attitude (force/linear accel- z i eration) relation. By having- more precise con- trol of attitude, maintenance of airspeed and ground track and execution of coordinated F turns were performed with lower workload.

I There was also less tendency to overcontrol with an attitude command system particularly for iz large maneuvers and/or control reversals.

E AVERAGE COOPER-MOPER When combined with an attitude command RATING system, velocity stabilization improved pilot _ ratings for maneuvering tasks by ‘about half a f rating point for both IMC and VMC.

It was J most noticed by the ease of maintaining airspeed I and effecting turn coordination during slalom RA/AT AT/AT AT/LV LV/LV and NOE tasks, and by the ease of varying airspeed and maintaining lateral ground track PITCH/ROLL SCAS CORFI6URATION during the acceleration/deceleration and NOE tasks.

Figure 20 Bob-Up Task Performance The influence of SCAS configuration on pilot ratings for ,the bob-up task is shown in Figure SECONDARY SCAS/CONTROLLER EFFECTS - 18. The attitude command system yielded pilot DIRECTIONAL/VERTICAL ratings in the low Level 2 region (CHPR ?I 4.5).

Use of a velocity command/velocity stabilzation Directional and vertical SCAS configurations system reduced pilot workload, improved task were varied for the RA/AT, AT/AT, and AT/LV performance, and achieved Level 1 pilot ratings systems of the primary SCAS matrix. All con- for the bob-up task with all controllers except troller configurations were evaluated.

In gen- the 4-axis small-deflection configuration. eral, the yaw rate command/heading hold system provided the best pilot ratings with the pitch/ Velocity command response characteristics were roll attitude command systems for all controller reported to be more jerky than the attitude re- configurations and tasks. Turn coordination sponse system, however, small position changes and lateral ground track could be controlled The addition of position easily, particularly for VMC.

could be made easily. A yaw accelera- evaluated only with 4-axis con- tion command system made it more difficult to stabilization, trollers, made the bob-up task a series of execute precise heading changes or to establish single-axis control maneuvers. Level 1 ratings a zero yaw rate at a desired heading. LOW were achieved and excellent position hold was speed turn coordination and lateral ground track generally achieved. were also degraded due to this inability to modulate or vary yaw rate precisely, partic- Figure 20 presents an example of bob-up task ularly with the pedals.

performance achieved as a function of SCAS con- figuration. Deviations in longitudinal and An important interactive directional SCAS/con- lateral position from the initial/desired hover troller effect is shown in Figure 21 where yaw location are used to calculate a mean radius, control on the 4-axis side-stick is compared to a circle containing one-half the total num- i.e. yaw control with the (3+1) pedal configuration ber of data points. Data are presented for Pilot for the slalom task. Yaw acceleration command A and five controller configurations as a func- from the (3+1) pedal or 2+1+1 configuration de- Compared tion of pitch/roll SCAS configuration. graded pilot ratings with all pitch/roll SCAS to the rate command system, a large improve- configurations when compared to the yaw rate command system. When yaw acceleration com- ment in performance and pilot rating can be Best per- mand was implemented on the side-stick, either seen for an attitude command system.

formance was achieved with a velocity command a (3+1) collective or 4+0 configuration, pilot system (mean radius 5 12 feet) for all controller ratings were degraded with the pitch/roll atti- Data for the 4-axis controllers tude command system, configurations. but improved with the rate command system.

show degraded performance and pilot ratings, For the low speed coor- dinated turn maneuver, yaw acceleration com- particularly for the attitude command system.

the vertical control axis degraded control mand improved control capability by eliminating accuracy and necessitated pulse control inputs the requirement for steady forces to control to achieve the best flight path performance.

yaw rate.

It is difficult for the pilot to mod- ulate forces in one or two axes (pitch/roll rate Figure 22 compares vertical control on the side- control) while holding a steady force in another stick and conventional collective lever. Vertical axis (yaw rate command for turn coordination).

acceleration command on the collective lever de- The yaw acceleration command system provided graded the IMC handling qualities to Level 3.

improved control harmony for lateral-directional As with yaw control on the side-stick, vertical maneuvering when implemented with the pitch acceleration command on the side-stick offers and roll rate command systems.

the benefit of eliminating the need to hold steady vertical control forces to achieve a steady vertical rate. However, based upon the results, the benefit of altitude hold and vertical rate command apparently offset the requirement LEVEL 1 to hold vertical control forces. These particular results may be biased by the lack of strong rep-; < ‘--’ vertical motion and rotor/drive system noise cues in the simulator.

---------------.

SUMMARY OF PILOT RATINGS In order to summarize task and SCAS config- uration effects on pilot workload and perfor- mance, all data were reorganized into a task/ LEVI”0 Pilot rating data for all controller SCAS matrix.

A SIti” . $I& . &r4, configurations were averaged for each task/SCAS combination. Figure 23 presents the results of Figure 21 Effect of Yaw SCAS In addition to the effect of SCAS this analysis.

Variations on Pilot Ratings configuration, there was a significant effect of task on pilot ratings for IMC. The IMC display effects showed an additive degradation of pilot workload/performance as task difficulty in- Also shown on Figure 21, the yaw acceleration creased. In comparison, VMC pilot ratings were command/yaw rate stabilization system generally predominantly affected by SCAS configuration received better pilot ratings than the yaw accel- and, except for the bob-up task where visual eration command/heading hold system. As pre- cues become weak, task had little effect. When viously noted for the primary SCAS RA/LV comparing IMC results to VMC, the mean in- system, a degradation of task performance was crease in pilot rating points for each task was: observed if the stabilization level was more than NOE course 2.3, slalom 2.0, acceleration/decel- one integration away from the command type.

eration 1.2, and bob-up 1.3.

The vertical rate command/altitude hold system achieved the best pilot ratings for all pitch/roll 10 - SCAS systems and controller configurations.

Vertical rate command provided good control of 5 - vertical speed and precise control of altitude, s : s- particularly for VMC. Acceleration command in 7 - c z f 5- TASK: NOECOURSE MC VERTICAL COHTAOL VERTICAL COHTAOL 2 5 ,or an COLLECTlYE LEVER(St?, c on S1DESTICI,I +o, so L z 4- ” - f s- - 2- l- I 1 I I AC/CA Al/AT LVlP" OAIAT ATlLV LVAV PITCHIADLL SC15 CD”FR”“ATIO” Summary of Task Effect on Pilot Figure 23 Ratings The effect of primary SCAS configuration on pilot ratings for the slalom, acceleration/ decel- eration, and NOE tasks is summarized in Figure Figure 22 Effect of Vertical SCAS Pilot ratings from the three tasks were 24.

Variations on Pilot Ratings combined into a single primary SCAS/ controller matrix, thereby tending to average out the ef- A comparison of VMC with IMC is tasks under IMC (Figure 24). However, for feet of task.

also shown. The average degradation of IMC the VMC tasks, the (3+1) pedal ratings ranked ratings compared to VMC ratings for all SCAS in the middle and received improved ratings configurations is 1.8 on the Cooper-Harper rat- when compared to the 4-axis configuration.

ing scale. For each SCAS configuration, the The 2+1+1 controller configuration in general range of pilot ratings from the best to worse achieved good pilot ratings for all three IMC controller configuration was an average of one low-speed maneuvering tasks. For the IMC and one-half rating points for both IMC and bob-up task (Figure 18); the 2+1+1 configura- VMC.

tion ranked better than the 4+0 but worse than IMC WC the 3 + 1 configurations. The 2+1+1 configura- tion achieved the best ratings for all the VMC maneuvering tasks.

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1L 1 , , 1 , , I AClCA CAlAT ATlAT ATlLV AC,“A “AlAT AT/AT ATlLV PITCH/ROLLSCAS CO”FlO”RA,lO” ALL TASKS (EXCEPT SOS-UP) lWCL”OE0 DATA AVERAGED : - AVEIAPE~ = 1.0 FOR ALL OATA (WC A”0 IYCl Table 5 Controller Configura !ti on Ranking Figure 24 Summary of Primary SCAS/Controller Effects CONCLUSIONS Figure 24 shows that an acceleration command/ rate stabilization system (AC/RA) exhibited Piloted simulation investigations of the effects Level 3 ratings for the IMC tasks, and the on handling qualities of variations in side-stick addition of attitude stabilization with a rate com- controller configuration and stability and control mand response system (RA/AT) received mar- augmentation system characteristics for both day ginal Level 2 ratings. With the same level of VMC and night IMC terrain flight were conducted attitude stabilization, an attitude command using the Boeing Vertol Flight Simulation Facility.

system (AT/AT) improved both IMC and VMC pilot ratings by over one rating point. When velocity stabilization was combined with an Conclusions from these investigations are organ- attitude command system, pilot ratings for the ized according to the major elements of the sim- maneuvering tasks improved an average of half ulation study: side-stick controller design, con- a rating point for IMC and VMC. troller configuration, SCAS design, and IMC display effects.

Based on average pilot ratings, from Figures 18 and 19, a ranking of controller configurations SIDE-STICK CONTROLLER DESIGN was determined for each task as shown in Table 5. Each task was weighted equally to obtain an A small-deflection side-stick controller is pre- overall IMC and VMC ranking for each controller ferred for low speed NOE maneuvering and pre- configuration. cision hover tasks when compared to a stiff-stick controller for the following reasons: The (3+1) collective controller configuration pro- vided the best overall pilot ratings for all IMC (1) It is easier to modulate force control tasks. A tendency to cross-couple directional inputs, particularly during large man- control into roll was observed during coor- euvers and control reversals. In high dinated lateral-directional turn maneuvers, par- workload situations, there is less ten- However, ticularly during initial evaluations. dency to over-contol and/or cross- this cross-coupling tendency diminished quickly couple control inputs.

and pilot adjustment to yaw control on the side- stick was easily made. (2) Pilot ratings with a deflection control- ler are less sensitive to variations in Pilot ratings for the (3+1) pedal configuration control response/force gradient. As were more degraded than other controller con- a result, it would be easier to design acceptable control response character- figurations for lateral-directional maneuvering side-stick yaw acceleration and verti- istics for a wider range of pilot cal acceleration command systems to preferences if a small-deflection de- eliminate the requirement to hold vice were implemented.

steady forces during multi-axis ma- neuvers.

CONTROLLER CONFIGURATION For rigid or small-deflection force con- The (3+1) collective configuration achieved the (6) trollers, elimination of steady forces best overall pilot ratings for all IMC tasks, fol- for steady-state helicopter trim must lowed in rank order by the 2+1+1 and 4+0 or be automatic through design of the (3+1) pedal configurations. This particular con- primary control system and/or AFCS troller configuration provides the following sig- control response laws.

nificant IMC terrain flight: The build-up advantages for of long-term steady forces is unac- ceptable.

(I) A separate collective controller elim- collective to inates unintentional IMC DISPLAY EFFECTS pitch/roll coupling common to the 4-axis and (3+1) pedal configurations.

Pilot ratings for the most difficult IMC maneu- vering task were degraded by approximately two (2) Directional control on the side-stick points when compared to the same task under provides more precise heading control VMC; degradation in both longitudinal and than the pedals. There is a tendency lateral handling qualities was caused by the to inadvertently couple yaw control to limited field-of-view available from the helmet- roll; however, all pilots adjusted easily mounted display.

to eliminate or minimize this character- istic. The (3+1) pedal configuration siqnificantlv degrades pilot ratings because of’ yaw controllability for the RECOMMENDATIONS IMC tasks. The limited field-of-view helmet-mounted display had a strong Continued simulation studies and design effort effect on lateral-directional control.

should be directed toward: In contrast, the 2+1+1 and (3+1) pedal configur- Improvement of vertical axis control (1) ations achieved the best pilot ratings for VMC.

using a (3+1) collective or 4-axis con- With good peripheral visual cues, directional figuration. Emphasis should be given control becomes a less demanding task.

to human factor aspects such as grip design, side-arm support, and con- SCAS DESIGN troller orientation.

A trend of Development of a (3+1) collective con- handling qualities improvements (2) attainable by various generic SCAS configura- figuration using a left-hand side-stick tions vertical controller instead of a conven- was defined. Conclusions based upon these results are as follows: tional collective lever. Consideration should be given to having both con- Level 1 handling qualities were not trollers available for vertical control.

(1)

achieved for any of the controller/ The 4-axis controller would be used SCAS combinations investigated for for low workload situations, i.e., level the maneuvering tasks conducted in flight and contour flying to free the IMC. left hand for cockpit adjustments and secondary functions. The separate A longitudinal and lateral velocity left-hand controller would be available (2) command system provided Level 1 for high workload flight maneuvers, handling qualities for the bob-up task. I MC/VMC nap-of-the-earth e.g., maneuvers, autorotational landings, A pitch and (3) roll attitude command emergency situations.

system with longitudinal and lateral velocity stabilization generally pro- Refinement of control laws to achieve (3) vided the best pilot ratings for the Level 1 pilot ratings for IMC.

Pos- low-speed maneuvering tasks con- sible SCAS modifications include: ducted in IMC. Automatic low-speed turn coordina- tion, inter-axis control paths to de- Altitude and heading stabilization couple responses, and alternate con- (4) were beneficial for all tasks and con- trol response shaping characteristics.

troller configurations.

Assessment of the effect of large (4) Yaw rate and vertical rate command (5) motion cues on vertical SCAS/control- systems are generally preferred for ler design and overall pilot ratings all tasks and controllers.

However, using the Vertical Motion Simulator at with a pitch and roll rate command NASA Ames.

system, there exists a preference for Investigation of the effects of turbu- 6. Hutto, A. J.: “Flight Test Report on the (5) lence on system performance and pilot Heavy-Lift Helicopter Flight-Control Sys- workload.

tern”. AHS Preprint No. 961, 31st Annual National Forum of the American Helicopter Comparison of alternate configurations Society, Washington, D.C., May 1975.

(‘3) for SCAS off or degraded mode condi- tions.

7. Padfield, G. D.; Tomiinson, B. N.; and Wells, P. M.: “Simulation Studies of Heli- ACKNOWLEDGEMENTS copter Agility and Other Topics”. RAE Technical Memorandum, 1978.

The experiment described in this paper was conducted by the Boeing Vertol Company for 8. Sinclair, M. and Morgan, M.: “An I nves- the Aeromechanics Laboratory, U. S. Army tigation of Multi-Axis Isometric Side-Arm Research and Technology Laboratories (AVRA- Controllers in a Variable Stability Helicop- DCOM) under NASA Ames Research Center ter” . NRC, NAE LR-606, 1981.

Contract NASE-10880 as part of the Army’s Advanced Digital/ Optical Control System Pro- 9. Smith, Stephen B. and Miller, Charles M., gram managed by the Applied Technology Lab- “An Evaluation of Sidestick Captain, USAF, oratory, Fort Eustis, Va.

Force-Deflection Characteristics on Aircraft Handling Qualities”, AFFDL-TR-78-171, pp. 395-413, December 1978.

The authors gratefully acknowledge the contri- butions of P. Dunford and S. Glusman of Boeing IO.

Vertol for their able assistance in gathering and Hoh, R. H. and Ashkenas, I. L.: “Devel- preparing the data presented herein and the opment of VTOL Flying Qualities Criteria services of participating test pilots: for Low Speed and Hover”. NADC-77052-30, 1979.

Boeing Vertol L. Freisner 11. “Heavy Lift Helicopter Flight Control Sys- U. S. Army R. Merrill tem , Volume I I l-Automatic Flight Control National Aeronautical M. Morgan System Development and Feasibility Demon- Establishment, Canada stration”, USAAMRDL-TR-77-40C, Septem- ber 1977.

U. S. Army-Ames P. Morris 12. Davis, J. M.; and Landis, K. H.; Leet, NASA Ames G. Tucker J. R., “Development of Heavy Lift Hand- ling Qualities for Precision Cargo Opera- tions.” AHS Preprint No. 940; 31st Annual REFERENCES National Forum of the American Helicopter Society, Washington, D.C , May 1975.

and Kelley M. B.: 1. Troubanos, C. M.

13. Aiken, E. W. and Merrill, R. K.: “Pilot Night Vision System (PNVS) for Ad- “Re- AHS sults of a Simulator Investigation of Con- vanced Attack Helicopter (AAH)“.

78-16, 34th Annual National trol System and Display Variations for an Preprint No.

Forum of the American Helicopter Society, Attack Helicopter Misson”. AHS Preprint Washington, D.C., May 1978. No. 80-28, 36th Annual Forum of the American Helicopter Society, Washington, D.C., May 1980.

2. “Military Specification - Flying Qualities of MIL-F-83300.

Piloted V/STOL Aircraft.” 14. Hoh, R. H. and Ashkenas. I. L., “Handling “Flight and Smith, R. E.: Quality and Display Requirements for Low 3. Hall, G. W.

Investigation of Fighter Side-Stick Force-De- Speed and Hover in Reduced Flight Visi- bility.” flection Characteristics”. AFFDL-TR-75-39, AHS Preprint No. 79-29; 35th 1975. Annual National Forum of the American Helicopter Society, Washington, D. C., May and Moorhouse, D. J.: 1979.

4. Black, G. T.

“Flying Qualities Design Requirements for Sidestick Controllers”. AFFDL-TR-79-3126, 15. Keane, W. P., Shupe, N. K., Sun, P. B., 1979. Robbins, T., and Camoaana, R. W., “A Versatile Display for ‘NCE Operations. ‘I and Freisner, A. L.; AHS Preprint No. 77.33-24; 33rd Annual 5. Deardorff, J. C.; Albion, N. : “Flight Test Development of National Forum of the American Helicopter the Tactical Aircraft Guidance System”. Society, Washington, D.C., May 1977.

AHS Preprint No. 761, 29th Annual National 16. Cooper, G. W. and R. P. Harper, Jr., Forum of the American Helicopter Society, “The Use of Pilot Rating in the Evaluation Washington, D.C., May 1973.

of Aircraft Handling Qualities.” NASA TND-5153, 1969.

FOR ASSAULT TRANSPORT NIGHT/ADVERSti WEATHER CAPABILITY DEFINITION OF DISPLAY/CONTROL REQUIREMENTS R. Joseph Milelli and Gary W. Mowery Program Engineers Martin Marietta Corporation Orlando, Florida Carmen Pontelandolfo HNVS Project Engineer Naval Air Development Center Warminster, Pennsylvania terrain avoidance. Several types of systems Abstract were examined, such as night vision goggles, pyroelectric vidicon, active gated TV, low light The U.S. Marine Corps is currently develop- ing a Helicopter Night Vision System (HNVS) to level TV, and forward looking infrared (FLIR) devices. Related Army and Navy studies have improve low-altitude night and/or adverse weath- concluded that FLIR devices perform better than er assault transport capabilities. Martin other electro-optical systems on a significantly Marietta Aerospace, under contract to the Naval greater number of occasions. Consequently, a Air Development Center, has performed a number of man-in-the-loop simulation experiments in its FLIR system capable of being configured and fntegrated into the assault transport helicopter Simulation and Test Laboratory (STL) to define was selected as the night vision system with the the minimum display and control requirements for the assault transport mission. These simulation best potential for satisfying HNVS mission requirements.

studies have investigated forward looking infra- red (FUR) sensor requirements, along with alternative displays such as panel mounted dis- The HNVS concept, shown in Figure 1, is plays (PMD), helmet mounted displays (HMD), and Also explored based on a FLIR system that is mounted on the integrated control display units.

were navigation requirements, pilot/copilot forward section of the assault helicopter. FLIR imagery is provided on panel mounted displays interaction, and overall cockpit arrangement.

(PMDs) or helmet mounted displays (HMDs) for the Based on pilot performance and opinion data, pilot and copilot.

pilot use of an HMD and copilot use of a PMD The FLIR permits the pilot to operate under conditions of total darkness, appear as both the preferred and most effective night navigation combination. and flight symbology superimposed on the FLIR imagery minimizes the pilot's and copilot's scan patterns. In addition, support avionics (such Introduction as a self-contained navigation system, radar altimeter, aircraft transducer, central com- State-of-the-art forward looking infrared puter, and control panels) are also required.

(FLIR) systems make it possible for transport The entire system will be designed to enable the helicopters to conduct missions under conditions mission to be performed safely with a minimal that would normally preclude operations. The workload for both pilot and copilot.

transport mission requires the transport heli- copter to fly at extremely low altitudes at the highest speed possible. Pilots must also approach and land in unimproved landing zones. Prior simulation experiments conducted in Martin Marietta's man-in-the-loop facility using Personnel and equipment must be quickly offload- ed because the aircraft must depart to permit a six-degree-of-freedom motion base concentrated primarily on basic system design parameters and landing of the remaining formation. This mis- aircrew interaction using panel mounted displays sion must be accomplished day and night and in during the enroute portion of the transport adverse weather conditions.

The Navy continued the simulation mission.

The United States Marine Corps is presently studies to 1) further expand, verify, and refine developing and evaluating design requirements the data base during the approach and landing portion of the transport mission; 2) to evaluate for a Helicopter Night Vision System (HNVS) that would improve transport helicopter low-level alternative displays; 3) to further refine the overall cockpit configuration, and 4) to evalu- night and reduced visibility capabilities.

ate incorporation of a control display unit to In support of this effort, tradeoff analy- support the navigation requirements of the mis- ses and system alternative studies were conduc- sion. Results of the these studies were com- ted to determine which type of night vision sys- pared with data obtained in actual flight tests tem could provide pilots with precise visual of FLIR and helmet display technologies at Yuma cues required as an aid to navigation and for Proving Grounds.

PILOTS DISPLAYS FLIR VIDEO AND PILOT’S SYMSOLDGY GENERATOR COPILOTS DISPLAYS ELEVATION rr- F igure 1. EINVS Block Diagram Man/Machine Simulation Objective CH-53D aerodynamic characteristics.

The Auto- matic Flight Control System (AFCS), the Stabili- The objective of the man/machine simulation ty Augmentation System (SAS), and the outerloop experiment was to obtain human factors data for attitude and heading hold modes were modeled on low-level assault transport operations using analog computers. McFadden Systems three-axis night vision sensors and ancillary hardware dur- force control loaders were used to duplicate the ing the approach and landing portion of the mis- control svstem's mechanical characteristics.

sion. These data were collected and analyzed, then recommendations were developed that have been reviewed for incorporation into flight test evaluations and HNVS system specifications.

Approach The HNVS simulation experiment used classi- cal modeling, validation, and experimentation techniques. A highly realistic CH-53D cockpit developed for prior enroute simulations was used for these studies. Thirty-five operational fleet Marine pilots participated as subjects in the experiment and represented both CH-53 and CH-46 squadrons. These pilots had between 270 and 4000 helicopter flight hours, with an aver- age of 695 hours. Experiments were conducted to investigate aircrew performance during approach Figure 2. Cockpit on Motion Base and landing using different display combinations of panel mounted and helmet mounted displays, including the copilot's use of a Virtual Head Up Additionally, pilot performance Display (HUD).

was investigated using a control display unit to assist in the visual navigation requirements. A variable landing zone size was used to increase workload as a a measure of system performance.

CH-53 Cockpit The CH-53D cockpit is shown on a six- degree-of-freedom motion base in Figure 2. The interior of the cockpit (Figure 3) was precisely modeled to CH-53D dimensions using consoles and control panels from a stricken aircraft. A special-purpose rotorcraft simulator modeled the Cockpit Interior Figure 3.

9a Helmet Mounted Display (RMD) HNVS Controls The Integrated Helmet and Display Sight The HNVS controls, readily accessible to System (IHADSS), shown in Figure 4, was install- both pilots, were arranged to provide rapid and ed in the cockpit for both pilot and copilot.

accurate control selection and actuation. Sen- The sight determined the pointing directions of sor control was provided on both collectives and the pilot's line of sight (LOS), and the HMD Symbology select and on the center console.

provided both pilot and copilot with collimated field-of-view (FOV) select were provided on both video displays. The IHADSS was used to slave cyclic controls and the center console.

the HNVS sensor to the pilot's LOS and display HNVS imagery to both pilot and copilot HMDs.

Panel Mounted Displays (PMD) Since copilots might find it objectionable or become disoriented with the HMD continually The instrument panel (Figure 3) was modi- presenting sensor imagery while they scanned fied with the installation of two nine-inch instruments in the cockpit, a virtual HUD pre- diagonal CRT displays for presentation of simu- sentation was included for the copilot.

As the lated terrain imagery to the pilot and copilot.

copilot turned his helmet away from a 30- by The displays were located so that the pilot and 40-inch window located straight ahead, the ter- copilot design eyepoints were at the outside rain image moved off the HMD as if he were look- edge of the display from a viewing distance of ing at a stationary HUD. The 50-degree FOV pro- 34 inches. Brightness and contrast controls vided on the HMD yielded a 1:l real-world view were located directly below each display. A red to the pilots.

filter was installed over the display for simu- A 50-degree FOV provi- lated night operations.

ded on the nine-inch monitor, at the design eye distance OE 34 inches, yielded a 0.30:1 mini- fication of the real-world view.

INTEGRATED HELMET AND IHADSS - DISPLAY SIGHT SYSTEM INTEGRATED HELMET AND ‘HADSS - DISPLAY SIGHT SYSTEM SEU - SIGHT ELECTRONICS UNIT SSU - SENSOR SURVEYING UNIT IHU - INTEGRATED HELMET UNIT DEU- DISPLAY ELECTRONICS UNIT DAP -- DISPLAY ADJUST PANEL BRU -BORESIGHT RETICLE UNIT IHADSS System Diagram Figure 4.

Control Display Unit (CDU) The CDU (Figure 5) was the primary man/ machine interface for navigation initialization and mode control. It consists of a CRT display, master function switches, line key, and an alpha- numeric key set that enables the copilot to view either the mission flight plan, or the naviga- tion plot showing fly-to-point data, reference points, and aircraft position along a projected course.

SYMBOL NAME

J

1 AIRCRAFT SYMBOL 2 HORIZON/PITCH BARS 3 RADAR ALTITUDE IANALOGI RADAR ALTITUDE ,OlGlTAL, VELOCITY VECTOR 6 IR SENSOR 7 TORcl”E 8 GROVNOSPEEDlAlRSPEED 9 AlRCRAFT HEADING 10 NAVIGATION STEERING 11 DISTANCE TO GO 12 ALTlTVDE REFERENCE BAR 13 VERTICAL SPEED 15 TIME TO GO 16 AIRSPEED INDICATION 17 POINT OF INTEREST ,S FA,L”RE WARNING INDICATOI 19 CORRIDOR BAR Figure 6. Flight Symbology Format Figure 5. Control Display Unit SYMBOL NAME Symbology AIRCRAFT SYMBOL : HORIZON BARS Two symbology formats were provided. The 3 RADAR ALTITUDE IANALOG) Flight Symbology format (Figure 6) was developed 4 RADAR ALTITUDE IDIGITALI 6 IR SENSOR as a piloting aid during enroute flight and 7 TORQUE commencement of approach to hover. The Hover/ GROUNDSPEED/AIRSPEED 9 A,RCRAFT HEADING Transition Symbology format (Figure 7) was de- 10 NAVIGATION STEERING signed as an aid to transition the aircraft. from 11 DISTANCE TO GO 12 POSITION BOX forward flight to hover and as a precise hover 13 VERTICAL SPEED Numerous symbology formats were evaluated aid.

16 TIME TO GO 16 AIRSPEED INDICATION during the simulation. The Flight and Hover/ 17 POINT OF INTEREST Transition Symbology formats provided the best IS HOVER VELOCITY IS HOVER ACCELERATION pilot performance for airCraft control during the entire enroute and hover portions of the Figure 7. Hover/Transition Symbology Format mission.

Hybrid Computing System performance data tends to support pilot opinion data, but it is not as pronounced, a result typ- ical in simulation programs.

The simulation was controlled by a hybrid computing system consisting of two Sigma 5 digi- tal computers, three EAT 231-RV analog com- Display Combination Evaluation puters, appropriate instrumentation, and inter- face and peripheral equipment. The computer To determine the effects of display combi- nations on crew performance during approach and arrangement controlled the aerodynamics, pro- cessed position commands to the terrain model landing, three treatment conditions were tested: (Figure 8) and TV, handled operational mode 1) pilot and copilot using PMDs, 2) pilot using HMD and copilot using PMD, and 3) pilot and logic and switching functions, generated com- mands to position symbology on the cockpit dis- copilot using HMDs. Each combination was evalu- and stored performance data. ated in landing zones with two difficulty lev- plays, els. The large zone was 3.5 rotor diameters (difficulty level 2) or more, and the small landing zone was 3.4 rotor diameters or less (difficulty level 1).

PMD-PM0 HMO-PM0 HMDHMD Figure 8. Terrain Model Experimental Procedures Figure 9. Experimental Hatrix for Approach and Landing PUD/lMD/CDU Evaluation Pilots were given an orientation to Martin Marietta's Simulation and Test Laboratory (STL), a system briefing, and an experiment briefing.

The data matrix for this evaluation is Ground school was conducted on HNVS cockpit con- shown in Figure 9. A Greco Latin Square design trols and displays. The pilot groups then pro- allowed order effects to be evenly distributed gressed through fixed and motion-base familiar- across all subjects and treatments.

ization flights, and finally progressed to training configurations that mirrored the data Touchdown and Approach Data for Display acquisition procedures. When all pilots Combination Evaluation approached their learning asymptote, as evi- denced by their performance, data collection The touchdown performance data were analy- commenced. Before each session of data runs, a zed on five dependent variables (landing time, briefing structured to resemble an air intelli- radial landing error, X drift, Y drift, and 2 gence briefing was held. Pilots were given a drift) and three approach variables (percent map of the area, a flight card, and a simulated under 100 feet, average altitude, and average 8 by 10 inch black and white reconnaissance groundspeed). The independent variables were photos of checkpoints and the landing zone.

three display configurations: PMD-PMD, PMD-HMD, Route legs and checkpoints were presented on the and HMD-HMD. Table 1 shows the levels of sig- map. The pilots participated in informal nificance resulting from this analysis. The debriefing sessions at the conclusion of data significant difference in landing error was run sets, and they completed extensive debrief- expected as a function of zone size. Although ing questionnaires when they completed all data no display combinations resulted in significant sessions. The informal debriefing sessions and performance differences, trends in favor of the questionnaires were designed to obtain subjec- HMD combinations do appear. The HMD-PMD combi- tive information from the participants on rele- nation had the greatest time under 100 feet, the vant HNVS issues.

lowest mean radar altitude, the only mean alti- Pilot Performance Data tude under 100 feet, and the least amount of Z drift. Tables 2 through 5 show the touchdown A large number of pilot performance mea-

and approach results, along with relative rank-

sures was gathered during the data runs, and ings of these results between display configur- several measurements of pilot performance were ations. The HMD-HMD combination had the short- taken as part of each evaluation. Pilot est landing time and the best overall ranking on These results indicate touchdown performance.

that the pilot's display affects performance most significantly, and performance is better with the HMD.

Pilot Performance in EHD-PM0 Table 1.

Evaluation: Touchdown and Approach INDEPENDENT “*RIABLES DlFFlCVLTY STANDARD DISPLAY DEVIATION CONFIGURATION LEVEL INTERACTION DEPENDENT VARIABLES OVERALL MEAN TOUCHDOWN: 234.528 100.70 NS” NS NS LANDING TIME 34.41 NS p = 0.046 NS RADlAL LANDING ERROR 31.57 FT NS -X DRIFT -1.61 FT,S 1.88 NS NS 2.41 FT,S 3.18 NS NS NS +x DRIFT NS NS -Y DRlFT -1.88 FT,S 1.89 NS I.24 FT,S 1.24 NS NS NS +V DRIFT NS NS 2 DRIFT 416 FT,S 2.88 NS APPROACH: PERCENT “NDER 100 FEET 38.68 % 28.81 NS LANDING ZONE DOES NOT AVERAGE GROUNDSPEED 59.57 KN 12.36 NS AFFECT APPROACH VARIABLEI AVERAGE ALTITVOE 119.03 FT 56.79 NS Table 2. HMD-PMD Evaluation: Touchdovn Performance Trends LANDING ZONE SIZE AND TOUCHDOWN VARIABLES I I -X DRIFT 1 +:&Rr 1 -Y DRIFT 1 +Y DRIFT 1 ZDRIFT 1 LANDING TIME LANDING ERROR

I I

( FT/S) (FT/S) (FT/S) (FT/S) :ONDSj (FEET) DISPLAY (SE( .L LARGE SMALL LARGE SMALL LARGE SMALL E 1 SMALL I LARGE I SMALL ILARGE I SMALL I LARGE I SMAL CONFIGURATION LARG I I I 1.24 1.32 4.69 4.57 233 277 36 21 0.69 1.69 3.66 1.23 1.39 2.15 PMD-PM0 Relative HMD-PMD Evaluation: Table 3.

Rankings of Touchdown Trends LANDING ZONE SIZE AND TOUCHDOWN VARIAELES CONFIGURATION T I 2 2 2 3 3 2 3.

PhlD.PMD 2 3 1 2 1 3 3 1 3 2 3 3 2 1 2 3 1 3 3 1 1 1 HMD-PMD 2 1 3 2 t 2 3 1 1 2 2 3 1 HMD.HMD 1 1 Table 5. HMD-PMD Relative Rankings of Table 4. EHD-PMD Evaluation: Approach Approach Trends Performance Trends APPROACH VARIABLES APPROACH VARIABLES PERCENT PERCENT UNDER AVERAGE AVERAGE OVERALL DISPLAY DISPLAY UNDER AVERAGE AVERAGE GROUNDSPEED ALTITUDE RANK CONFIGURATION 100 FEET CONFIGURATION 100 FEET GROUNDSPEED ALTITUDE 3 2 3’ 59.63 131.75 PMD-PM0 3 PMD-PM0 35.91 HMD-PMD 1 3 1 1.5 HMO-PMD 44.27 58.70 107.10 2 1 2 1.5 HMD.HMD - HMO-HMD 38.51 59.78 113.24 ‘3 = WORST RANKING lC2 Smoothness of Approach and Landing for Display - PMD-FMI Combination Evaluation --- - HMDPMD _ -X-X-X- HMDHMD Regression analyses were run on the distri- The radar altitude for the last nauti- butions.

cal mile before touchdown was significantly (p = 0.10) smoother for the HMD-PMD configuration than for that of the PMD-PMD. Figure 10 shows the radar altitudes approaching the landing zone PITCH ANGLE, NOSE UP - (LZ) as lower and smoother for the HMD-PMD con- DEGREES figuration. Significant differences in distance distributions were also found in pitch angle.

The pitch angle for the HMD-HMD configuration was significantly (p = 0.0028) smoother than the PMD-PMD configuration, and difficulty level 1 (small LZ) was significantly smoother than level 1.0 2 (large LZ). This difference is shown in Fig- t ures 11 and 12. Examining the time distribution indicated that the rate of descent was more con- sistent for the larger LZs. The display combi- DISTANCE TO TOUCHDOWN, DIFFICULTY LEVEL 2 - NMI nation trends, although not statistically sig- nificant, show the PMD-PMD combination to be Figure 12. Aircraft Pitch Angle during Landing more erratic across all variables than both con- Phase: Large LZ figurations in which the pilot uses the HMD.

Crash Rates for Display Combination Evaluation An examination (by chi-square analysis) of the frequency of noncrash landings per attempts showed no significant differences due to display configurations of LZ size.

Any frequency dif- ferences appeared due to chance and not experi- mental conditions.

Virtual Head Up Display (HUD) Evaluation The virtual HUD evaluation varied the co- pilot display combinations from PMD, HMD, and HMD virtual HUD while the pilot remained on the lhh HMD. The ANOVA results shown in Table 6 indi- DISTANCE TO TO”CHrmWN -NY, cate that route change as a variable has a sig- Figure 10.

Radar Altitude during Landing Phase nificant effect on percentage of time under 100 feet and average groundspeed.

Runs without route changes had a higher percentage under 100 feet with the virtual HUD configuration (73.5 percent), followed by common HMD video (57.5 percent) as shown in Table 7. In runs that contained a route change, the HMD-PMD configuration had the highest percentage of time under IUU lees \JD parcencl. tierall, the HMD configuration with the virtual HUD had the lowest average radar altitude.

However, the variability between display combinations is small, i.e., only 11 feet. Runs with a route change had faster average groundspeeds than those without (Table 8). This increase in groundspeed was predictable, since the altitudes of changed routes tended to be higher. The virtual HUD has the fastest groundspeed in runs with changes and the lowest in runs without.

the HMD-HMD combination video had the Overall, fastest average groundspeed.

Figure 11.

Pitch Angle during Landing Phase: Small Lz Table 6. Pilot Performance in Virtual HUD Evaluation: Enroute* ,NDEPENDENT VARIABLES ROUTE ROUTE STANDARD DISPLAY DEPENDENT “ARlASLES OVERALL MEAN DE”,AT,DN CONFlGURATlON DIFFICULTY CHANGE INTERACTION - 65.34 Y 14.42 N6’- NS p - 0.01 N3 PERCENT UNDER 100 FEET p = 0.02 AVERAGE GROUNDSPEED 69.10 KN 16.01 NS NS NS AVERAGE ALTITUDE 104.99 FT 17.16 NS NS NS w <O.lO *SIGNIFICANCE LEVEL LIMITED TO P **DIFFERENCES NOT SlGNlFlCANT

I

Average Table 7. Virtual HUD Evaluation: Altitude Enroute EASY ROUTE D,FF,C”LT ROUTE I I OVERALL NO DISPLAY NO ROUTE CHANGE ROUTE CHANGE MEAN COMB,NAT,ON ROUTE CHANGE ROUTE CHANGE ,,I.45 120.59 f42w 105.m lE8%) 95.04 IO2Y.l 125.16 (37.5X, i HMDPMD 112.45 157.5Y.l 108.02 64x1 107.51 “MD-HMD 93.39 162%) 119.19 (45%) ~ (COMMON VIDEO) I I 63.11 173.5%) 122.66 (37.6% Km.75 64.06 (73.57.1 ii3.m 147x1 (VIRTUAL HUD1 ‘PERCENT OF T,ME UNDER 1W FEET Average Virtual BLID Evaluation: Table 8.

Altitude Enroute EASY ROUTE D,FF,C”LT ROUTE OVERALL DISPLAY NO NO ROUTE CHANGE ROVTE CHANGE MEAN COMB,NAT,ON ROUTE CHANGE ROUTE CHANGE HMD.PMD 66.16 67.06 65.64 69.99 67.76 HMD-HMD 71.56 64.67 73.01 65.76 73.63 (COMMON VIDEO) WRTUAL HUD1 For comparison, data collected for the Army's Surrogate Trainer is shown in Table 8.

The Surrogate Trainer is a AH-1s helicopter equipped with a AN/AAQ-11 Pilot Night Vision Sensor (FLIR), IHADSS, symbol generator, and navigation system.

The groundspeed data shown for both day (27kts) and night (16kts) flights highlights the differences between the Army's tactic of nap-of-the-earth (NOE) flight and the low-level flight concepts utilized by Marine pilots in the simulator. A typical altitude plot is presented in Figure 13, which indicates the pilot flew the helicopter below 15 feet.

This altitude is significantly lower than the 107.51 mean altitude when the HMD-HMD configuration was used in the simulator.

These differences underscore the inverse relationship between clearance altitude and groundspeed and potentially reflect the difference in aircraft size.

Figure 13. Typical Night NOE Flight Profile6 for Surrogate Trainer Table 9. Line Key Errors during Route Changes Control Display Unit (CDU) Evaluation Copilot performance was evaluated as a : FREOljENCY ERROR SEDUENCE I function of display combination during low-level flights over longer routes that required a sub- DEPRESSED LINE KEY 9 INSTEAD OF 9 An enroute course stantial navigation workload.

.DEPRE.SSED LINE KEY 4 INSTEAD OF 3 change was added as a variable so that the difficulty of the copilot inserting a route DEPRESSED LINE KEY 10 INSTEAD OF 9 change into the CDU midway in a mission could be evaluated. Figure 14 contains the data matrix.

DEPRESSED LINE KEY 2 INSTEAD OF 3 Random route conditions were used so that pilots SEVERAL LINE KEY ENGAGES AFTER ONE MFK All enroute could not predict, course changes.

data runs required the copilot to manually cap- SEVERAL PAGE CHANGES AFTER ONE DIR ture the LZ.

Table 10. Uaster Function Key Errors during Route Changes FREQUENCY ERROR SEQUENCE 3 DEPRESSED FTL/PLN INSTEAD OF DIR DEPRESSED MARK INSTEAD OF FTLlPLN PMD-PMD 3 DEPRESSED STAT INSTEAD OF DIR DEPRESSED PROG INSTEAD OF DIR DISPLAY 2 DEPRESSED MARK INSTEAD OF DIR CONFIGURATION 1 DEPRESSED MAP/RTN INSTEAD OF FLTIPLN i Pilot Performance Data Summary ROUTE NO ROUTE CHANGE CHANGE The size of the landing zones affected pi- lot performance more consistently and predict- Figure 14. Experimental Matrix for Approach ably than any other factor.

The smaller zones and Landing PMD-RMD-CDIJ Evaluation required more precise maneuvering, which resulted in longer landing times, higher radar altitudes during approach, smaller radial error, Capturing the LZ required a specific etc.

To land in these zones, the pilot must three-key operation of the flight plan master have the helicopter under control.

The pilots function key (MFK) and line keys 9 and 6. Most evaluating the HMD-PMD combination generally runs had an addendum to this sequence, which was performed better using the HMD.

Ease of slewing several scale changes (line keys Il.and 12).

the sensor allowed pilots to examine terrain Discrete data was examined to determine actual features and maintain low altitude with compara- sequences. There were 2 errors in 17 operations tive ease. The pilots' landing approaches and when this sequence was performed, and both touchdowns were also smoother when using the involved parallax problems with line key 9.

HMD.

The random enroute change also involved a During the enroute portion of the mission, specific sequence of events to properly execute crew performance in flatter terrain was slightly the new route and capture checkpoints in the better while the pilot used the PMD, but perfor- The PMD-PMD configuration had the fewest old.

mance in mountainous terrain was better when the CDU errors (6), followed by the HMD-PMD configu- Copilot operation of the pilot used the HMD.

ration (7) and the HMD-PMD configuration (11).

CDU indicated that it is a useful part of the There were 11 line key errors and 13 total MFK navigation system that reduces the dead reck- errors encountered during route changes. Tables However, the oning navigation workload task.

9 and 10 display the type of errors that occurr- excessive number of copilot input errors during ed. These tables show consistent problems with route changes indicates that changes need to he parallax and misunderstanding of key functions.

made in keyboard layout and in CDU feedback The copilots depressed line keys several times cues.

in succession in trying to obtain a response or to correct an error. The copilots did not cue in on the CDU feedback (for example, the asterisk that appears with the capture func- tion). These errors indicate that the display arrangement needs to be correctd and CDU feedback must be furnished when a function key is initialized.

Pilot Opinion Data Table 11. Minimum Safe Altitude at 60 to 80 Knots and Maximum Safe Speed at After the data runs, the pilots were asked 100 to 150 Feet AGL to rate the safety and ease of display configur- RADAR ALTITUDE IFT, AND SPEED II(N) BY DISPLPIY CDNFlGURATlDN - I-- ations during an approach to the landing zone (Figures 15 and 16). While there was little variability in response, consistent trends were Pilots rated the HUD safer and easier apparent.

The variability than dual PMDs in all phases.

is small, but the pilot's HMD display is appar- ently the critical preferred feature.

EXTREMELY SAFE Pilots showed a consistent preference for

r

the HMD-PMD configuration across all aspects of VERY SAFE mission ease and safety.

The HMD-HMD virtual HUD was considered the most dangerous and diffi- SAFE cult display configuration. However, pilots expressed a preference for the copilot to use MEDIUM - the PMD for map reading and navigation.

I 'Ihey felt that the virtual HUD made it difficult to SAFETY DANGEROUS - turn their head and use the CDU, and that the time required to regain the display created a VERV DANGEROUS - dangerous situation. All pilots felt the HMD- PMD configuration was the safest, most effective EXTREMELY DANGEROUS - configuration; the HMD-HMD virtual HUD configur- ation was felt to be the least effective and safe.

PMDPMD “MD-PMD DlSPLAY CONFIGURATION Seventy-five percent of the pilots felt the CDU helped to maintain low altitude, and 50 per- Figure 15. Safety of Approach to LZ cent felt it helped to increase groundspeed.

These respondents felt the CDU simplified navi- EXTREMELY gation duties and increased orientation, which EASY allowed more time for concentration on flight VERY Copilots felt the tactical map display tasks.

EASY was useful.

Pilot Opinion Data Summary Evaluating the HMD-PMD combination resulted EASE in a definite preference for the pilot to have a DIFFICULT helmet display.

The copilot preferred an HMD VERY for mission ease and a PMD for mission safety.

DlFFlCULT The enroute evaluation indicates a consistent preference for the HMD-PMD configuration.

IMPOSSIBLE Copilots felt that the virtual HUD con- HMD.HMD figuration was more difficult and dangerous than the HMD-HMD or HMD-PMD configurations.

They DISPLAY CONFIGURATION also preferred the HMD-PMD configuration.

Figure 16. Ease of Approach to LZ The CDU was found to be an extremely useful navigation tool. It enables copilots to accur- ately assess present position, desired position, Participants were asked to indicate the and overall mission. The HMD increased copilot task loading, but operation of the CDU was still minimum safe target altitude at 60 to 80 knots possible.

and the maximum safe target groundspeed at 50 to 100 feet above ground level (AGL) that was Conclusions attainable on an actual night mission. Table 11 shows the pilot ratings of actual mission The simulation experiments have demonstra- altitudes and speeds. They believe that lower ted the ability of pilots and copilots to fly a altitudes and higher speeds are attainable when the pilot uses the HMD. night mission at low altitudes, ranging from 50 to 150 feet AGL, with the night visionics equip- ment package tested.

Further HNVS Efforts Although this experiment required no data to be generated on dead reckoning versus naviga- As a result of these experiments, a base- tion system requirements, both pilot performance line HNVS configuration has been established and and opinion data reinforced that crew station is presently being evaluated in actual flight workload was reduced with Doppler command steer- tests by the Naval Air Development Center and ing information. Incorporating the CDU naviga- Naval Air Test Center. The cockpit configura- tion capability was also instrumental in further tion is identical to that utilized in the simu- reducing navigation workload.

lation studies. Addtionally, an HNVS System Specification has been developed by the Naval Most pilots preferred flying the night Air Development Center for procuring actual pro- transport mission with the HMD instead of the duction prototype hardware for flight test. The PMD, regardless of which display configuration HNVS System Specification was developed, in the copilot was using. The precise slewing of large part, from data generated during this and the sensor with the HMD using a pilot's natural prior simulation experiments.

head movements allowed control over the sensor without changing hand position on the collective References during critical flight maneuvers, which is re- quired when operating the sensor manually.

Martin Marietta Aerospace, Final Cockpit and Software Preparation Task Report, OR 15,647-i, By contrast, most copilots preferred using May 1980.

the PMD. They found constantly moving imagery somewhat distracting when performing the CDU Martin Marietta Aerospace, Helicopter Night line key and master function tasks.

Vision System Simulation Evaluation Phase I Final Report, OR 15,986, July 1980.

The copilot group evaluating the virtual HUD mode of IHADSS did not find this mode use- Martin Marietta Aerospace, Helicopter Night - ful. Of particular concern was losing symbolic Vision System Simulation Evaluation Phase II aircraft attitude and altitude information and Final Report, OR 16,015, July 1980.

losing imagery while performing cockpit tasks using the virtual HUD.

Martin Marietta Aerospace, Helicopter Night Vision System Simulation Final Report, OR The preferred cockpit display configuration 16,026, July 1980.

was to have the pilot use the HMD and the copi- lot use the PMD. The HMD provides the pilot Martin Marietta Aerospace, Helicopter Night with precise slewing control over the sensor and Vision System Simulation Final Report, OR more visual feedback information than available 16,551, December 1981.

with the PMD. The PMD provides the copilot with sufficient aircraft position, attitude, and altitude information, yet simplifies cockpit The PMD does not introduce the workload tasks.

visual interference characteristic of the HMD or the complete loss of aircraft information char- acteristic of the virtual HUD.

Copilots found the CDU to be a useful navi- gational aid in reducing the navigation workload task. The present keyboard inputs required for enroute changes, however, are somewhat cumber- some through nonalignment of CDU symbology with the appropriate line keys. The result was co- pilot confusion and numerous copilot input errors. Through lack of an indication for posi- tive CDU line key actuation, numerous other co- pilot line key input errors resulted.

SOME PILOTING EXPERIENCES WITH MULTI FUNCTION ISOMETRIC SIDE-ARM CONTROLLERS IN A HELICOPTER J. Murray Morgan Airborne Simulator Facility Manager Flight Research Laboratory National Research Council Canada Ottawa, Ontario Abstract Experimental Hardware The installation of two side-arm mounted, isometric The NAE Airborne Simulator controllers in the NAE Airborne Simulator, a modified, variable The NAE Airborne Simulator is an extensively modified stability Bell 205’A is described, as is the development of Bell 205A-1 with the stabilizer bar removed, the standard various control systems for use with them. The results of two hydraulically boosted actuators replaced with dual mode experiments are presented indicating both the feasibility and electro-hydraulic actuators (which provide full authority acceptability of such systems for a wide variety of tasks in a electrical or fly-by-wire control from the right seat or full conventional single rotor helicopter, with a minimum of authority hydraulically boosted mechanical control from the stability augmentation. Areas of future research are indicated.

left, or safety-pilot’s seat) and extensive hybrid real-time computational capability. The safety pilot, whose controls reflect all computer inputs to the actuator system, can assume Introduction control at any time, while a safety system, which monitors the status or condition of many elements of the fly-by-wire system In the fall of 1979, the National Aeronautical can cause an automatic reversion to left seat control if a ‘fault’ Establishment (NAE), a division of the National Research or ‘out-of-limits’ condition is sensed.

Council Canada, was approached by the Sikorsky Aircraft division of United Technologies, with a proposal for a co- The on-board hybrid computation system comprises three PDP-11 processors, in mutual communication and operative project to flight test a pair of isometric side-arm operating on a computational cycle of l/64 second, supported controllers in the NAE Airborne Simulator. This was of by three fields of analog computation.

sufficient interest to the NAE, relating closely to an area of active research interest, that it was possible to agree to such a The Simulator is equipped with a wide range of motion program, the results of which were reported in Reference 1.

sensing systems which provide high quality measurements of Sikorsky provided the two controllers installed in a seat with its velocity relative both to the earth and the ambient airmass.

side-arm supports, NAE provided the interface between the A nose boom carries vanes for angle-of-attack and sideslip electrical outputs of the units and the simulator computers measurement, together with a swivelling static pressure source and developed suitable control systems, while pilots from both while dynamic pressure is taken from two wide-angle pitot organizations took part in the formal evaluations. This paper probes, nose-mounted. The usual range of inertial sensors is will describe the development process that led to the evaluated supported by a doppler radar for earth referenced velocity systems, some of the problems encountered and their solution.

measurement, while a radio altimeter provides height data, Data from the first co-operative experiment and a more when within some 750 metres of the surface.

recent NAE experiment employing similar controllers will be oresented. and intentions for future work in this area will be Figure 1 shows a simplified block diagram of a typical simulation channel as used in this series of experiments.

I) ATMOSPHERIC COMPUTER SYSTEM t STATIC AND DYNAMIC SENSOR PACKAGE PRESSURES, DOPPLER RADAR FIG. 1: A TYPICAL SIMULATION CHANNEL FOR THE ISOMETRIC SIDE - ARM CONTROLLER EXPERIMENTS Side-Arm Controllers The hand controllers (Fig. 21, standard, commercially available, ‘4-axis’ units were mounted on a standard Bell seat as shown in Figure3. Each had independant outputs in four control axes, X (fore/aft), Y (lateral), Z (vertical) and 0 (torque about the Z axis) as shown in Figure 3, while their transducing characteristics were as listed in Table 1. The controller units themselves exhibited essentially zero compliance in response to forces and moments up to the rated maximum input, but when installed the system showed slight movement due to the compliance of the supporting structure. In addition to the primary force sensing transducers, each unit carried several discrete switches, namely, a trigger, a standard aircraft ‘coolie hat’ two axis thumb switch, and either side of the latter a simple contact closure push button. The outputs from these were read and interpreted by the on-board computers, while the functions allocated to them were 1) Trigger-communication 2) Coolie hat - progressive trim in X or Y as appropriate 3) Inboard push button - datum reset trim system activation.

Table 1. Controller transducing characteristics.

Axis Max Input Sensitivity Max Output X 20 lb. F 0.5 volt/lb. 10.0 volts Y 20 lb. F 0.5 volt/lb. 10.0 volts Z 40 lb. F 0.25 volt/lb. 10.0 volts e 60 in. lb. 0.167 v/in. lb. 10.0 volts FIG. 3: CONTROLLER/SEAT INSTALLATION Experimental Software The initial proposal called for a simulation of the Blackhawk helicopter, however, since this would have required a complex, model following procedure and would have caused delays in the program and uncertainties in the validity of the model, it was not undertaken. As a compromise, the basic Bell 205 was used as the baseline model and two levels of SAS were provided, a simple rate-damping augmentation about all axes and a rate command/attitude hold mode in pitch and roll with both stiffness and damping augmentation in yaw and augmented heave damping. This approach had the advantage of being simple and certain in its implementation while presenting a ‘real’ helicopter, with all the cross-couplings and asymmetries inherent in this class of aircraft.

Control Modes The experimental software was arranged so that, prior to engagement of the fly-by-wire system, the various outputs from the hand controllers could be assigned to drive different control actuators, enabling a variety of control modes to be investigated. In all, two primary and three secondary control modes were examined (Fig. 4). For the three modes which had duplicated control functions on the two controllers, both inputs were read at all times by the computers and summed, giving the pilot the option of using either hand for the primary control FIG. 2: AN ISOMETRIC CONTROL UNIT task.

PRIMARY CONFIGURATIONS TWO HANDED/Z COLLECTlYE TWO HANDED/cONVENTIONAL SECONDARY CONFIGURATIONS FIG. 4: ISOMETRIC CONTROL CONFIGURATIONS gave the pilot the lower sensitivity he desired aroundneutral, System Development while still permitting large and rapid inputs to be made without Control Signal Shaping any disturbing discontinuities in slope.

The extent of the dead band, and the extent and Figure 5 shows three forms of signal shaping which magnitude of the linear slope segment, which were adjustable represent the progression in the development process from the in flight with the fly-by-wire system disengaged, were optimized first flight to the point at which the system was offered for for various control functions and flight conditions. For the formal evaluation. The simple dead-band and linear slope of formal evaluations a compromise set of characteristics, biased Figure 5a proved to be too sensitive for other than very limited towards operating at and near the hover were used.

hovering, and even that required a very high pilot workload.

The dual slope arrangement in Figure 5b was quite acceptable During the development flying it was noted that, due to at the hover, but still produced problems at other points in the arm/armrest/controller geometry, it was significantly easier to manoeuvring flight envelope where the ‘knee’ became obvious apply a Z force in the up rather than the down direction.

to and created difficulties for the pilot. Therefore the approach Therefore, to provide the pilot with a more subjectively even shown in Figure 5c, a small linear range blending into a response in this channel an overall asymmetry was applied to quadratic non-linear characteristic was finally evolved. This it, effectively magnifying inputs in the UP-Z sense.

MAX. OUTPUT VOLTS FORCE s -

If-

/‘/I & : d 5a r‘ 5b 5c FIG. 5: CONTROL SIGNAL SHAPING Control Position Indicators Trimming Prior to first flight, two trim systems were installed, a Late in the evaluation phase of the first experiment, progressive, constant rate trim, activated by the ‘coolie hat’ when one of the subject pilots elected to attempt off-level landings and take-offs, a significant and anticipated operational and applied to the X and Y function of the appropriate disadvantage of isometric controllers was highlighted. In a controller, and a datum reset system. This latter system was conventional helicopter the pilot has a direct bio-mechanical activated by pressing the inboard thumb button on either indication of the tip-path-plane orientation; the cyclic position controller which action disengaged that unit from the drive system, while the inputs to the control channels were held is a direct analog of the normal to that plane. This information is used by and is of great importance to the pilot during all constant by the computer. The pilot was then able to relax take-offs, but especially when lifting off from a slope. A rigid any held force and reconnect the controller by releasing the controller inherently robs the pilot of this important piece of button. Both of these proved to be unsatisfactory. While the information and, under some circumstances, visual information progressive system could be used, the force changes associated may not suffice as a replacement. Operational limitations with repositioning the hand on the controller to make contact associated with the absence of another performance related cue, with the switch were sufficient to introduce unwanted inputs tail rotor collective pitch, sensed in a conventional helicopter to the system, thereby making unacceptably large demands on the pilot in terms of the care and physical accuracy of his hand from pedal displacement, also were evident in these experi- ments. This information is used by the pilot as an indication of movements. The datum reset was somewhat easier to use, yaw control authority remaining when operating with large (except that again repositioning the hand and applying yaw rates or in the presence of large sideslip velocities. Figure 7 sufficient force to the button to overcome its spring generally caused inadvertent inputs), but suffered from a more funda- shows a rudimentary Control Position Indicator (CPI) that was mental problem. If, during the period when the controller was fitted above the instrument panel coaming to compensate for the loss of these cues and while far from ideal (the indicator was disconnected from the controlled system, the aircraft was adapted from a fixed wing auto pilot trim indicator) it sufficed externally disturbed, then on reconnection the pilot could find to expand the useable envelope in the areas indicated above.

himself with an out of trim condition even greater than that which he had been in the process of relieving initially.

To overcome these difficulties, a selectable, continuous integral trim was devised, which provided an integral-plus- proportional command signal from the hand controller to the system, with the inboard thumb button being reassigned to the activation of this system. The final configuration is shown in Figure 6. As reported in Reference 1, the handling character- istics of this type of system depend on the ratio of integral to proportional gains (Ki ), and the optimum value need not F:e entire remain constant overt flight envelope. However, for this experiment a set of constant values was used, and they are reproduced here in Table 2.

Table 2. Integral/proportional gain ratios.

1.0 Roll 0.5 Pitch FIG. 7: CPI INSTALLATION Yaw 1.9 1.5 Heave First Experimental Evaluation Description of the Experiment A series of tasks, shown in detail in Appendix A and intended to represent the greater part of the flight envelope of the 205, was selected for evaluation by a group of five pilots, two from Sikorsky and three from NAE. Cooper Harper ratings were required for each task and, subsequent to the completion of the experiment the subjects were asked to reply to a general questionnaire; their responses are reported in full in Reference 1.

The decision was made to introduce the subjects directly to the two primary control configurations rather than to train them via a force analog of a conventional displacement system. To provide an overall comparison, one of the NAE pilots, with some five years experience on the aircraft was asked to rate the tasks using the basic, unaugmented, aircraft and displacement controls. The experience levels of the evaluation FIG. 6: TYPICAL SIMULATION CONTROL CHANNEL pilots are shown in Table 3.

Table 3. Evaluation pilot experience levels.

HOVER MANoEuvRlNG Pilot Total Hours Helicopter/Fixed Wing ii TRANSITION FROM HOVER TRANSITION TO HOVEf.

ALL SUBJECTS A 3500 32501250 : PRECISION LAraING 5700 400/5300 V B FREE AIR MANCEUVRIffi x OPERATIONAL TASKS C 6900 900/6000 D 6500 4000/25OC SAME TASKS E 5600 155OI4050 : CONVENTl ONAL CONTROLS

I

Results of the First Experiment : SEE TEXT Cooper Harper Ratings Figures 8 to 10 are plots of the Cooper Harper ratings obtained during the experiment. All data points have been used and they are coded by task rather than pilot so that any effect of task on opinion can be examined.

Summary of Pilots Comments As a supplement to the numerical opinions obtained during the experiment, the following summary of the subjects’ written and verbal comments for which there was reasonable commonality is produced below: 1) When using a three axis configuration, force rather than displacement pedals were preferred. the need to use leg and foot movement when applying only forces with the hand generally being judged less natural than applying forces to all controls.

2) The assignment of collective to a twist function was not liked since it tended to be prone to inputs in the

I I I

incorrect sense, and the instinctive relationship between input RATE COMMAND ’ ’ BASIC AIRCRAFT and control response, present when collective was driven via INTEGRAL TRIM ATTITUDE HOLD the Z axis, was absent.

RATE AUGMENTED 3) All subjects felt that the more fully supported and INTEGRAL TRIM erect posture inherent in the side-arm controller installation reduced fatigue compared to the conventional helicopter FIG. 8: 4 FUNCTION, RIGHT HAND, EVALUATIONS seating position.

Pilot Adaptation With one exception, discussed in more detail below,

1------------

all subject pilots adapted easily to the multi-function configurations, to the extent that the majority of them elected to commence data runs before the allotted training period was complete.

Discussion of Results of First Experiment Four Axis System 4m ------------, Consider Figure 8 and ignore, for now, the circled data 3w points. While the degree of acceptability increased with increasing stability augmentation, as might be expected, the 2= main point of interest is that even the most primitive form of augmentation brought the peak of the rating distribution to the acceptable side of the 3.5 boundary. Note too, that the i data in the left hand column suggest that there is little

O i 4 FuNclloN 3 z CCLL) ’

difference between the basic, unaugmented, aircraft when

F OF&

PEDALS flown with displacement and isometric, four function 3 (e COLL) controllers. Also, the spread of points due to individual tasks FORCE PEDALS suggests that no particular portion of the manoeuvring flight envelope examined produced opinions radically different from any other.

FIG. 9: RATE AUGMENTED, INTEGRAL TRIM

1lllllllllll II I I lllll I I

The circled data points are of special interest, and may The Interim Period have a particular significance. They were all contributed by a single subject, who was the exception to the general pattern of From the end of the initial experiment to the summer easy adaptation to the isometric, multi function system. It is of 1981, no formal investigations were carried out, but the possible that he may represent a sub-group in the piloting body controllers were flown quite frequently, often riding ‘piggy- who will adapt only with great difficulty to such systems, and if back’ on other experiments or for the purpose of demonstration so, this could have significance in the areas of trainee selection to pilots from other organizations and countries. In this period or training washout.

too, they were flown in the IFR environment, where the ability to free one hand for ancilliary tasks, without having to abandon Effect of Control Configurations the control task met general approval. The pattern of relatively easy adaptation for the majority of pilots was maintained.

From the evidence of Figure 8, the rate damped model was selected to examine the effects of various control con- Development of an Alternate Hand Grip figurations on pilot opinion, the results being plotted in Figure 9. Of the two primary configurations there is a slight Both during the initial experiment and subsequent preference for the three-plus-pedals over the four-axis mode, flying, it had been noticed that, although the controller units with all data points for the former configuration being on the themselves had little inherent cross-talk, in use there were acceptable side of the 3.5 boundary.

several coupling tendencies, the dominant ones being a nose-up pitch with UP-Z commands and a roll into yaw. Both these The Effect of Increasing Stability effects appeared to be, if not due to, at least exacerbated by the hand grip design. Figure 1 la shows the original hand grip The final plot in this series shows the effect of supplied with the isometric controller. If a lightly cupped hand increasing stability when using the preferred control configura- applies a force to this grip in the UP-Z direction, the pressure tion. It suggests that while the tendency for acceptability to point on the handle is sufficiently displaced from the force increase with increasing stability augmentation is present, even sensing axis to result in an appreciable moment in the nose-up the ‘basic aircraft’ is within the fully acceptable boundary with sense. Similarly roll inputs, generally applied with the inside this control system.

edge of thumb or forefinger produce a moment about the Z axis, hence producing a yaw command.

lo- ------------ 9- 8- 7- m------------ 8- 5-0 4- ------------ 3-m c c c z- l- FIG. Ila --.

o BASIC AIRCRAF r ?ATE COMMAND ’ INTEGRAL TRIM 4TTlTUDE HOLD ‘RiATE AUGMENTED ’ !NTEGRAL TRIM FIG. 10: 3 FUNCTION, Z COLLECTIVE, FORCE PEDALS Biasing Factors When interpreting the above data, two factors should be considered. The possible sense of euphoria experienced by the pilots on discovering that they could not only fly, but fly well, with such a radically new control system may have introduced a favourable bias in the ratings. On the other hand their very low experience level at the time of rating (no more than some 10 hours each by the end of the flight phase, with FIG. lib the exception of the development pilot, who had about 22 hours) might have been expected to produce the reverse effect. These effects are reasonably expected to diminish as work in this area proceeds. FIG. 11: HAND CONTROLLER CONFIGURATiONS

I’

One other problem noted with the original grip was Table 4. Pilot flying experience at that the slimness and almost circular cross section of the lower the end of the second experiment portion of the design made the application of larger yaw inputs Pilot Total Helicopter/Side-Arm relatively more difficult than inputs in the other three axes.

This was one action for which it was necessary to grip the C 7200 995137 handle firmly, a most undesirable technique which leads to both rapid hand fatigue and undesired inputs, both pilot and B 8054 432170 environment induced.

G 905 313113 H 4002 4002/g To eliminate, or at least reduce the effects of these undesirable characteristics, the handle shown in Figure 1 lb Results was designed and manufactured at the NAE. Its main features are the elimination of the curvature in the X-Z plane, a some- While most of the data from this experiment still awaits what ovoid cross section and, to assist in the application of a analysis, some preliminary results are presented in Figures 12 to ‘clean’ Z force, a much larger base flange and a good sized 14, specifically, the pilots’ subjective opinions, track deviations thumb/finger support table. in the lateral translation segment, and touch-down scatter.

Pilot Opinions While no formal evaluations of this design have been made, it has found general acceptance among the project pilots As illustrated in Figure 12, the pilots generally and has been used in a recent series of tests.

considered isometric control to be more difficult and less precise, in this type of closely bounded task, than conventional control. There is also a suggestion that this judgement is less The Second Experiment severe in the case of the three axis system than in that of the fully integrated, four axis configuration. However, the greater Following the initial work with these controllers, it was number of opinions fall between the ‘same/more difficult’ and felt important that a more direct comparison between the ‘same/less precise’ responses, indicating no great difference multi-axis, isometric systems and conventional controls should from displacement controls. The relatively very short exposure be made. To this end an experiment was designed and flown of the subject pilots should also be considered when looking at in the summer and fall of 1981. these replies.

Description of the Experiment Using the marked ground course, shown in Appendix A,

pilots were required to fly, in a single run, an accelerate/stop 0

3 FuNcTloN segment, rearward, lateral and quarter translations, two ‘pedal’ turns, a precision touch-down and a lift-off. The briefing to 6 them included instructions to pay close attention to height- keeping and tracking, and to fly the course ‘briskly’.

Qualitative and quantitative data were recorded using both the aircraft data acquisition system and ground observation. The pilots were also asked to provide a subjective MllcHMmf THE SAME aai assessment of the relative ease and precision of the task when DIFFICUT yo~~ EASIER EASIER using the multi-function controllers, compared to the DIFRDUT conventional controls.

The subjects were required to fly the course alternating two runs with conventional controls and two runs with either the four-axis or the three-axis plus pedals configurations using the isometric side-arm controllers. Each flight consisted, generally, of one practice and eight data runs. Two complete I 6 sets of runs were flown with each isometric configuration and 1 6 refamiliarization was permitted for each pilot between his evaluating with the different side-arm controller systems.

The data were analyzed for precision, control activity and time as a means of investigating the relative performance of a particular subject as he moved from one control system to another.

The subject pilots for this exercise were all from NAE and Table 4 summarizes their relevant experience to the end FIG. 12: PILOT’S SUBJECTIVE ASSESSMENTS of this experiment.

CONVENTI ONAL CONVENTI ONAL 80- 0’ ’ aq .A so- 0 0 A PI LOT DES I GNATORS 2’.

“” ‘,” p 4o- ‘.‘- RUN SEQUENCES #g 30- p=’ .” .’ “” ‘“I= 20- .“’ lo ‘...“” ‘.” o ..“” I I I I I I 0 10 20 30 40 50 60 TIME (sets) CONVENTI ONAL FIG. 13: PERFORMANCE COMPARISONS, CONVENTIONAL AND SIDE-ARM CONTROLLERS LATERAL TRANSLATION TRACKING Landing Accuracy Lateral Tracking Figure 14 compares landing accuracy of systems, and a To obtain the data plotted in Figure 13 time-adjacent definite degradation in performance is noted for the isometric pairs of runs were analysed for RMS deviations and plotted one systems. (It should be emphasized that the control system being against the other, thereby eliminating, as far as possible, any flown in these tests had not been specifically optimized for the effects due to changing atmospheric conditions or pilot fatigue.

landing task.) There is an interesting difference in the pattern of landing errors between the systems; using conventional controls the errors tend to lie along the lateral axis of the aircraft, while While the general tendency is towards a more unsteady with the isometric systems, there is a definite slew towards the tracking performance with the isometric controllers, it is longitudinal. This may be due to a change in the type of visual possible that learning curve effects are still present, since there cues required by a pilot when landing with isometric systems is a consistent tendency for the RMS values for the two control compared to those he habitually uses when operating with systems to approach one another the later into each flight the displacement controls. This may demand that more of his data are taken. It is noteworthy that there is no indication of visual attention be directed towards the front of the aircraft any time penalty when using the force controllers, which may than to the side and, considered in combination with a natural suggest that even though the subjects considered the tasks to I tendency to drift the aircraft along the line of sight, may have be more difficult, and their performance to be less precise with caused this dispersion pattern. (It is worthy of note that in the the isometric than with the conventional system, the level of Simulator the evaluation pilot sits on the right, and that there degradation was not such as to cause them to proceed with are no errors either to the left or the rear with any control system.)

unusual caution.

PILOT DESICNKKIRS UOAO 8- 8- 3 FMCTION cDNvENnoNAL 4 FUNCTION CONVENTIONAL FIG. 14: TOUCHDOWN SCATTER COMPARISONS Future Intentions 3) Further direct comparisons between displacement, limited compliance and isometric controllers.

The National Aeronautical Establishment will continue 4) Investigations of integrated control/display systems its investigation of integrated, multi-axis control systems as part using multi-axis controllers and advanced electronic displays.

of the aircraft flight systems and flight mechanics research program. At the time of writing, for example, a controller, similar to the one described in this paper but with some Conclusion compliance, is being prepared for installation in the Airborne Simulator. The potential merits of limited motion will be The work at the NAE over the last two years has investigated.

demonstrated both the feasibility and acceptability of using multi-function, isometric, side-arm controllers to perform a It is expected that the main areas of interest for future wide variety of tasks in a conventional helicopter, with the study will be: minimum of stability augmentation. While these two short test programs do not provide definitive answers to all of the 1) Evaluation of the limited motion controller.

questions which the designer must ask about such radically 2) An investigation into more sophisticated control unconventional control systems, they do indicate that this systems, including mission and task level optimization, and adaptive or scheduled variations in control system characteristics. will be a fruitful area for future research efforts.

L

Acknowledgement Table 1. Task details for the first experiment.

The names of the subject pilots, in alphabetical order, Title Content Task # with their affiliations are: Hover into and across wind 1 Manoeuvring at 1.1 K. Davidson NAE 360” turn left and right Hover 1.2 S. Kereliuk NAE Lateral translation, 1.3 G. Kohler Sikorsky Aircraft moderate rate M. Morgan NAE 1.4 Accelerate and rapid stop R. Murphy Sikorsky Aircraft Transition from hover 2A E;;p from and to 2.1 D. Sattler NAE Transition to hover 2.2 A.D. Wood NAE 28 Landing 2.3 Zero speed landing from Reference hover to terminate in marked zone Report-Sinclair, M., and Morgan, M., “An Investigation of Multi-Axis Isometric Side-Arm Controllen in a Variable 3 High Speed Flight 3.1 Summetrical pull-ups Stability Helicopter”, National Research Council 3.2 Steep turns Canada, Aeronautical Report LR-666, August 1981. Roll reversals 3.3 3.4 Partial power descents 3.5 Sideslips 3.6 High power climb Pop-up and point 4 Operational 4.1 Manoeuvres 4.2 NOE course 4.3 Downwind take-off and turn APPENDIX Second Experiment Figure 1 represents the ground course marked out for SOME PILOTING EXPERIENCES WITH MULTI the second experiment. The boxes contain instructions to the FUNCTION ISOMETRIC SIDE-ARM CONTROLLERS subject, while the circled numbers indicate radio transmissions IN A HELICOPTER required for data correlation. Table 2 gives the dimensions of the various linear segments.

Task Details for the Two Experiments Table 2. Ground course dimensions.

From To Distance (ft.)

First Experiment Table 1 details the tasks required to be performed by A 8 670 the subject pilots in the first experiment. A single Cooper 8 C 445 Harper rating was requested for each task, with the exception C D 450 of Task ZA, where separate ratings for the transition to and from the hover were required. D E 500 OVER 8 THROUGH GATE NOW pOJN ACCELERATE TO 40 KNOTS STOP OVER 8 PEDAL TURN 270’ LEFT BEFORE GATE I @ B BACK TO A NOW PEDAL TURN

0 0 OVER D

360” RIGHT lzlL--

NOW CONSTANT RATE 000 LATERAL, TRANSLATION TOUCHDOWN C CONSTANT VELOCITY (10 KT), IL = 320” ON SPOT IL = 320" BEGINNING OVER C NOW TRANSLATION PEDAL TURN COMPLETE TOUCHDOWN NOW pJ FIG. 1: COURSE FOR INTEGRATED SIDE ARM CONTROLLER EVALUATIONS

I I I IllIll Ill1 I

RESULTS OF NASA/FAA GROUND- AND FLIGHT-SIMULATION EXPERIMENTS CONCERNING HELICOPTER IFR AIRWORTHINESS CRITERIA J. Victor Lebacqz, Robert T. N. Chen, Ronald M. Gerdes, and Jeanine M. Weber NASA Ames Research Center .Moffett Field, California Raymond D. Forrest Federal Aviation Administration NASA Ames Research Center Moffett Field, California Abstract As a part of their respective research pro- grams, NASA and the FAA have instituted a joint pro- A seauence of around- and fliaht-simulation gram at Ames Research Center to investigate helicop- experiments was conducted at the Ames Research Cen- ter IFR certification criteria. This series of ter as Dart of a joint NASA/FAA proaram to investi- investigations has the following two general goals: gate helicopter instrument-flight-riles (IFR) airworthiness criteria. This paper describes the 1) To provide analyses and experimental data first six of these experiments and summarizes major to ascertain the validity of the Airworthiness Cri- results. Five of the experiments were conducted on teria for Helicopter Instrument Flight,l which have large-amplitude motion base simulators at Ames been proposed as an appendix to FAR Parts 27 and 29 Research Center; the NASA-Army V/STOLAND UH-1H (Refs. 2, 3).

variable-stability helicopter was used in the flight experiment. Taken together, the results of 2) To provide analyses and experimental data the experiments indicate, among other things, that to determine the flying qualities, flight control, 1) some level of artificial stability and control and display aspects required for a good helicopter augmentation is generally required for adequate IFR capability, and to relate these aspects to flying qualities during precision instrument design parameters of the helicopter.

flight; 2) neutral longitudinal or lateral control position gradients do not result in inadequate fly- With respect to the first goal, the sections of ing qualities, given good directional characteris- the Ref. 1 criteria that deal with static and tics, but an unstable longitudinal gradient can dynamic stability attempt to prescribe quantitative prove to be inadequate for instrument operations in values of several helicopter flight characteristics turbulence; 3) pitch and roll attitude augmentation that would be required for IFR certification. To in the stability and control augmentation system the extent that these values are a carryover from (SCAS) plus directional augmentation including at fixed-wing practice or an amalgam of previous handling-qualities requirements formilitary aircraft least yaw damping is required to achieve satisfac- tory precision instrument flying qualities irres- Ref. 4), it is necessary to ascertain their (e.g., pective of the type of rotor or level of display validity for civil helicopter certification. One assistance; 4) fliqht directors provide some com- aspect of interest has to do with the requirements pensation for-poor-flying qualities in dual-pilot for stable force or position control gradients lon- situations but are of minimal assistance in this gitudinally, laterally, or directionally. Another aspect of interest is the difference in criteria regard for single-pilot operations; and 5) the SCAS level required for ratinqs of satisfactory is the for normal category rotorcraft depending on whether same (pitch and roll attitude augmentation) for the the aircraft is to be certificated single or dual pilot, particularly since most of existing substan- range of approach types considered (nonprecision versus precision, constant speed versus decelera- tiating data pertain only to dual-pilot operation.

Yet another area of concern is the influence of tion to a low speed).

displays on the instrument meteorological condi- Introduction tions (IMC) flying qualities, which is not consid- ered in Ref. 1 but has been shown in some cases to Current and projected expansion of civil heli- compensate for less-than-satisfactory inherent fly- copter operations has led to increasing efforts to ing qualities (e.g., Ref. 5).

assess problem areas in civil helicopter design, With respect to the second general goal, most certification, and operation. Of concern are the influences of the helicopter's inherent flight helicopters currently certificated for single- dynamics, flight-control system,. and display com- pilot IFR operations use advanced stability and control augmentation systems (SCAS) or displays or plement on flying qualities for instrument flight both.6 Of concern is the level of complexity of rules (IFR) flight; both in terms of design param- the SCAS required to achieve a good IMC capability eters to ensure a good IFR capability, and with because of the cost, control authority, and relia- regard to the characteristics. that should be required for certification. bility factors the SCAS introduces. Of interest also is the expansion of helicopter IMC operations to exploit the helicopter's unique capability to fly at very low airspeeds; this expansion requires

llll1l1llllllll Ill I I I Ill1 III

additional definition of the required flight dynam- directional pedals) plus control interconnects and ics, flight controls, and displays.

gearings. All feedback and control gains may be programmed as functions of flight parameters, such The various experiments discussed in this as airspeed. This structure permits the construc- paper were designed to investigate elements of tion of typical SCAS networks; it may also be used interest in achieving both goals in a consistent as a response-feedback variable stability system to fashion. SDecificallv, the objectives of each modify the basic characteristics of the simulated experiment,'listed in-chronological order, may be helicopter.

1) First experiment (ground summarized as follows.

simulation, 1978):7 develop generic models of cur- In the first experiment, the rotor design and rent helicopters having three different rotor types; helicopter geometric parameters of the mathematical explore SCAS concepts and influence of longitudinal model were selected and tuned to simulate stability static stability; and determine relative influence and control characteristics similar to those of the 2) Second exper- of IFR CornDared to VFR aDDrOaCheS. UH-lH, OH-6A, and 80-105 aircraft, which use iment (ground simulation;1979):a,g determine suit- teetering-, articulated-, and hingeless-rotor sys- ability of requirements on cockpit control position; tems, respectively.7 These same three generic examine. efficacy of several SCAS concepts; and helicopters were used as the baseline configura- 3) Third experi- explore influence of turbulence.

tions for the second experiment; only the teetering ment (qround simulation, 198O):lO determine influ- model was used in the successive experiments. Ref-.

ence of crew-loading (single pilot versus dual erence 9 lists several of the geometric and rotor pilot); determine influence of three-cue flight design parameters for them. It is emphasized that director displays; and examine suitability of addi- the resulting static and dynamic characteristics 4) Fourth experiment (flight, tional SCAS concepts.

are intended to be representative of the three 198O):ll validate selected results of ground- types of rotor systems investigated for the three simulation experiments in flight concerning static weight classes of helicopters that were simulated; longitudinal stability, level of SCAS, and flight they are not, in all respects, identical to the director displays. 5) Fifth experiment (ground characteristics of the UH-lH, OH-6A, or BO-105.7 simulation. 1980):12 examine influences of unstable static control qradients, angle-of-attack stability, Static Stability and pitch-speed-coupling; and examine influence of 6) Sixth exoeriment (qround simula- failed SCAS. One type of configuration variation carried tion, 1981):13 investigate SCAS reqiirements for across most of these experiments was changes in decelerating instrument approach; explore influence longitudinal, lateral, or directional static stabil- of electronic display format; and examine influence ity-as measured through cockpit control positions of approach geometry and deceleration profile. with speed or sideslip. For the purposes of this paper, the variations in longitudinal control posi- The remainder of this paper is organized as tion with velocitv will be emDhasized. Of the follows. The following section summarizes the three baseline helicopter models developed in the designs of the experiments with an emphasis on first experiment, the models with articulated and variations that were carried across all of them, hingeless rotors had stable control position gra- and the next section provides a review of their dients at 60 knots; the position gradient for the Fol- conduct, again emphasizing the similarities. teetering rotor was unstable. One of the SCAS con- lowing these summaries, the resultsof all the exper- cepts considered (rate damping with input decoup- iments are compared with each other, followed by ling, longitudinal cyclic to collective gearing some general conclusions.

scheduled with speed) turned out to destabilize this gradient, yielding an almost neutral gradient Experimental Design for the hingeless rotor, an unstable gradient for the articulated rotor, and a more unstable gradient Mathematical Models for the teetering rotor.7 In addition, a prelimi- nary investiqation of the influences of this gra- In the ground-simulation experiments, the dient was made in a controlled fashion for the basic mathematical model used to simulate the hingeless-rotor model by using the variable- flight dynamics of the helicopters was a nine- stability aspect of the model structure, with feed- degree-of-freedom model developed for use in nap-of- back of longitudinal velocity to longitudinal the-earth (NOE) simulations.14 The model explic- cyclic being used to vary the effective M,,.

itly includes the three-degree-of-freedomtip-path- Table 1 summarizes the gradients and the times to plane dynamic equations for the main rotorI and either half or double amplitude of the prevalent the six-degree-of-freedom rigid-body equations. low-frequency roots.

The main-rotor model includes several major rotor- system design parameters, such as flapping.-hinge This variable-stability capability was used in restraint, flaDDino-hinoe offset, blade Lock number, succeeding experiments to control the longitudinal and pitchiflap'coupling: Simulation of different control position gradient with speed, including the rotor systems (e.g., hingeless, articulated, and influences of the SCAS gearings. In the second teetering) was accomplished by appropriate combina- experiment, two levels of gradients were considered tions of those design parameters. for the hingeless rotor (stable and neutral), and neutral values were designed for the teetering and The model is structured to permit full-state In the third articulated rotor models also.asq feedback to any of the four controllers (longitudi- experiment, only the teetering-rotor model was nal and lateral cyclic, collective stick, and used, with the gradient held at neutral (to highlight influences of SCAS and displays, as will capability. Variations in the type of augmentation, be described below).1° The flight experiment con- and to some extent the level of it, were carried out across all the experiments.

sidered three levels of gradient (basic airframe, In the first experiment, these variations for each of the three increased value to roughly that of the ground exper- iments, decreased value to neutral), with the baseline aircraft consisted of 1) no augmentation; 2) pitch/roll/yaw rate damping; 3) input decoupling variable-stability capability of the aircraft being used in a fashion analogous to the ground to reduce off-axes accelerations to control inputs added to (2); and 4) pitch and roll attitude auo- simulation model to vary Mu,-and the resulting mentationadded to (3).7 The second experiment- control gradient being measured in flight.ll In the fifth experiment,-this gradient was systemati- considered again the last two of these concepts, cally varied for the teetering-rotor model from with the gains For the teetering-rotor configura- tion increased to provide response characteristic quite stable to unstable values, yielding times-to- double-amplitude down to about 6 sec.12 The values roots similar to the hingeless-rotor configuration; considered across all the experiments are summa- in addition, turn-following augmentation (increased rized in Table 1 for SCAS implementations incorpo- directional stiffness and feedbacks to reduce the rating only rate feedbacks. Dutch roll excitation) was considered, as was a rate-command-attitude-hold system in pitch and roll Other Baseline Characteristics that was implemented by adding proportional-plus- integral prefilters to the pitch and roll command As was mentioned above, ground simulation channe1s.a models of helicopters having hingeless-, articu- These SCAS types were all considered aqain in lated-, or teetering-rotor systems were used in the the third experiment, with a selectable wing-leveler first two experiments; in the remaining ground- (roll-attitude feedback) also added to the rate- simulation experiments (and of course in the flight damping and rate-damping-input-decoupled SCAS mech- experiment), emphasis was on only the teetering- rotor system. Reference 7 describes the wide range anizations to study split-axis auqmentation in a preliminary way.

of response characteristics among the three unaug- For‘this experiment, reduced levels of rate and attitude feedback were used for mented baseline models and the resulting flying- qualities deficiencies. For the hingeless and these SCAS types, to be more consistent with actual teetering-rotor capabilities. An additional teetering mqdels in particular, however, the addi- velocity-hold SCAS was designed, which augmented tion of SCAS incorporating rate damping and input decoupling effectively minimized these differences, the vertical velocity time-constant to roughly 0.5 set and used longitudinal velocity feedbacks to particularly when high-gain feedbacks were used increase the effective phugoid frequency and par- with the teetering model in the second experiment.g For this reason, only baseline configuration tially eliminate lift-change caused by speed (Zu).lO changes to the teetering-rotor model will be dis- The fourth (flight) experiment included only cussed here.

the two SCAS types of rate-damping-input-decoupling and pitch/roll attitude augmentation, with the Table 2 lists some of the stability deriva- tives at 60 knots of the baseline teetering-rotor levels desiqned to be consistent with the third ground-simulation model. These characteristics experiment.I* These same two SCAS types at the same level were also used in the longitudinal axis were held constant across all the experiments, but in the fifth experiment selected variations were for the fifth experiment, with the lateral axis held fixed at a high-gain rate-command-attitude- also considered.l* One of these variations was the For the hold type. In addition, a failed longitudinal steady-state attitude-to-speed gradient.

pitch-rate damper was also simulated by eliminating baseline model. this aradient was verv low the pitch-rate feedback in the rate-damping-input- (-O.O3"/knot at 60 knots), which cons;derably aggra- vates the difficultv of controllinq soeed at low- decoupling SCAS.l* Finally, the sixth experiment also included rate-damping-input-decoupling, rate- control gradients; 'the variation considered was to command-attitude-hold, and attitude-command SCAS increase artificially the drag damping (X,) to pro- types, with somewhat higher augmentation levels duce an attitude-to-speed gradient of -0.33"/knot considered because of the decelerating task. Addi- at 60 knots. Another variation was the angle-of- tional desiqns were a velocitv command system and attack stability, which was nearly zero for the an acceleration-command-velocyty-hold system, that baseline configuration (Table 2). This derivative incorporated high-gain feedback of longitudinal was made very stable (M, = -0.025), using the velocity to longitudinal cyclic (constant term of variable-stability system; as is discussed in hovering cubic about 1.7).

Ref. 12, this variation had a negligible influence on the longitudinal control position gradient (in Because of the consistency across most of the contrast to its effect on a fixed-wing vehicle), experiments of rate-damping-inbut-decoupled, rate- but did modify short-term response to cyclic.

command-attitude-hold, and attitude-command SCAS Again, these variations were considered in only types, these results will be emphasized in this the fifth experiment.

paper.

Stability andControl Augmentation System (SCAS) Displays As was discussed in the introduction, one of Figure 1 shows the instrument panel layout the major aspects of concern in this sequence of used in all the ground simulation experiments, experiments was the type of stability and control The instruments were arranged in except the last.

augmentation required for a good helicopter IMC was simulated. TO provide a lack of repetition, a standard "T," and were conventional, with the four different approach plates to four oil rigs exceotion of.theelectromechanical attitude indica- were devised, with different frequencies and alter- tor {ADI),-which was a 5-in. unit incorporating nates for each plate; these four possibilities heading (through longitudinal lines on the ball) as well as pitch-roll information. Turn-rate-slip were mixed randomly among the control-display com- information was presented on a separate instrument, binations. Finally. on the sinole-oilot aooroaches.

as is frequently done in helicopters, rather than the pilot did not know whether he would be'able to _ Figure 2 shows the with the attitude indicator. continue the approach or be forced to do a missed primary flight instruments for the flight experi- approach; the simulated fog was made to start clear- ment. The horizontal situation indicator (HSI) is ing at 100 ft above the decision height and then to similar to the one used in the ground experiments, either re-fog or continue clearing just below deci- but the AD1 incorporated integrated glide-slope sion height. As a result, the pilot had to make and localizer deviation data plus turn-rate-slip the decision whether to continue.

information not included in the ground simulator unit. In the last ground simulation experiment, Wind and Turbulence the AD1 was replaced with a black-and-white cathode ray tube (CRT) unit to present electronic formats.

An additional variable carried across the Figures 3 and 4 illustrate the two electronic for- experiments was the level of winds and turbulence mats considered in this experiment. As can be seen, present. For the ground-simulation experiments, a the first is a simplified analog of an electrome- simple model for atmospheric turbulence16 was used; chanical AD1 such as the one used in the flight it included three independent Gaussian gusts plus a experiment; the second is one way of integrating a mean wind which could shear in direction or magni- variety of information into one presentation, but tude.

In the first experiment, all evaluations will not be discussed in this paper. were conducted in no turbulence.

In the second experiment, the configurations were evaluated in Excluding the integrated electronic format, both no turbulence and at a representative level of therefore, the primary display variable considered turbulence (u,,.= cs,, = 3.0 ft/sec, qw = 1.5 ft/sec) across the experiments was the extent of flight with no mean wind. The third experiment added a director information provided to the pilot in addi- lo-knot mean wind that sheared rapidly in direction tion to the raw deviation data. Because the task a total of 100" at a ranae of about 1 mile out: all considered for the first two experiments was a VOR the configurations were evaluated in this wind-and approach, only course-deviation information was turbulence combination, with no zero-turbulence presented on the HSI, with the AD1 flight director evaluations. This same wind and turbulence model needles biased off scale. In the remaining ground- was again used in the fifth experiment, with evalua- simulation experiments and in the flight experiment, tions conducted both with it and in no turbulence.

'a precision MU approach task was considered; for The sixth experiment included a vertical shear of these experiments, azimuth and elevation deviation the mean wind (from 10 knots at altitude to 2 knots plus DME (range to go) information was given on at ground level) in addition to the shear in direc- the HSI. In the third experiment, one-, two-, or tion, and considered 1.5 times more turbulence three-cue fliqht directors were a display variable; = ay = 4.5 ft/sec, uw = 2.25 ft/sec); again lull evaluations were conducted in both calm air and in the flight-experiment, either no directors or three-cue directors were the variable; in the sixth with this turbulence model.

experiment, all configurations included a three-cue flight director; in the fifth experiment, no flight For the flight experiment, the level of wind directors were considered. The general philosophy and turbulence was not a controlled variable. As is discussed in Ref. 11, tower estimates of wind mag- of the flight director design is discussed in nitude and direction plus the pilots' qualitative Ref. 10.

estimate of the turbulence level were used to sepa- rate the data into two groups: one in which head- Crew-Loading Situation winds with little or no turbulence were present, All but the third experiment were conducted as and one in which there was a tailwind component or typical flying-qualities experiments; the pilot's moderate turbulence or both.

sole task was to perform the desired control task, Conduct of the Experiments with no auxiliary tasks of communications or navi- This scenario of full-attention-available- gation.

Equipment for-control is consistent with a dual-pilot crew- loading situation. In the third experiment, the configurations were evaluated assuming this situa- The first three qround-simulation experiments were conducted using the Flight Simulator.for tion but they were then also evaluated in as rea- Advanced Aircraft (FSAA) qround-based simulation listic a simulation of a sinqle-pilot situation as For the single-pilot simulations, the facility at Ames Research-Center; the last two used possible.

the Vertical Motion Simulator (VMS) facility at pilot always had to communicate with Approach Con- trol and Tower, set a transponder frequency, and Ames (Fiqs. 5 and 6). Both facilities include a complex movable structure to provide six-degree-of- switch communication frequencies; for approaches freedom motion; in the case of the VMS, a large including a missed approach, he also had to switch vertical travel (+30 ft) is available to enhance communication frequencies again, copy a clearance simulation fidelity of longitudinal motions, and from Departure Control, switch navigation and the FSAA is characterized by a large lateral travel transponder frequencies, and track a VORTAC. Radio In both facilities, a visual scene from (250 ft).

"chatter" from two other helicopters in the area a terrain board is presented throuqh the cab window summarizes the individual details of the evaluation on a color television monitor with-a collimating tasks.

lens. For the first two experiments, the Cooper-Harper pilot ratings were assigned to approaches were conducted to a model of a STOL air- each configuration on the basis of the evaluation port with helipads; the last three ground- task for each experiment, and comments made rela- simulation experiments considered approaches to a tive to comment card; task performance and control model of an off-shore oil rig.

usage data were also obtained for each.

Across all the experiments, the total number of participating Instrument conditions were simulated using an pilots by affiliation was as follows: NASA, 3; electronic fog generator which could obscure all or U.S. Army, 4; Federal Aviation Administration, 4; part of the visual scene as a function of range or NAE Canada, 2; and Civil Aviation Authority, UK, 1.

altitude. In the first two experiments, the Approximate total evaluations for Experiments 1 instrument runs were conducted entirelv in the foe through 6 were, respectively, 60, 200, 150, 50, 200, to a minimum descent altitude of 600 ft, with no - 160; taken together, therefore, over 800 evalua- breakout simulated. The third and fifth experi- tions were obtained.

ments did include a partial clearing of the fog starting at about 100 ft above the decision height, Discussion of Results which could then refog at the decision height to force a missed approach; in the sixth experiment, Influence of Longitudinal Control Gradient _ .-.

the fog always disappeared at the decision height.

In Figs. 9a and 9b the average Cooper-Harper The flight experiment was conducted on a UH-1H pilot ratings from each experiment are plotted as helicopter which had been modified as an in-flight functions of longitudinal static stability without simulator by adding an avionics system called turbulence and in turbulence, respectively. The V/STOLAND (Fig. 7). The system provides integrated data are for configurations with a rate-damping- navigation, guidance, display, and control func- input-decoupling SCAS and a dual-pilot crew-loading tions through two flight digital computers; it may situation; they include both hingeless- and be operated with or without flight-director com- teetering-rotor systems in the results for Experi- mands, in the modes of manual, control-stick steer- ments 1 and 2. To emphasize the imoortant asoects.

ing (CSS), autopilot, or research. The flight- the pilot ratings are'shown versus the gradient .

control portion of the V/STOLAND srstem uses a level (in./15 knots) for the stable cases but versus combination of a full-authority parallel servo and the inverse of the time-to-double-amplitude of the a limited authority (20% to 30%) series servo in divergent root for the unstable cases.

each control linkage. In addition, disconnect devices exist in the left c.vclic controls to allow As can be seen, the correlation among all the for a fly-by-wire mode through this research cyclic experiments is quite good. The data show a consis- stick. The riqht stick, or safety pilot side, tent trend toward a degraded capability as the sta- tic stability is reduced to neutral and then retained the standard UH-1H cyclic and cockpit instruments. This experiment was conducted in the unstable, with the trend being more obvious in tur- research mode, with the software providing a set of bulence. In terms of Cooper-Harper ratings, how- flight-control laws with variable qains and a set ever, the aircraft systems were still rated as of flight-director laws with fixed-gains.ll Instru- adequate for the tasks considered, irrespective of the static stability. Note that, with this type of ment flight was simulated with the use of an "IFR Hood." SCAS, average ratings in the satisfactory category were not attained, even at the most stable level.

Evaluation Tasks and Procedures In commenting about these configurations, the pilots noted increasing difficulties in maintaining trim and controlling speed precisely as the static sta- Althouah the evaluation tasks differed in detail among the six experiments, they were gener- bility was decreased, but also noted that the ically similar for all except the sixth. Each of instrument tracking performance was still adequate at least down to neutral stability.

the first five included a lateral guidance acqui- sition at constant altitude (about 1200 to 1600 ft AGL, depending on experiment), transition to a The IFR Appendix requires positive longitudi- nal control force stability at approach speeds for vertical descent at a constant soeed of 60 knots (1000 ft/min for the VOR approaches of Experiments both transport and normal category helicopters, 1 and 2, acauisition ofa 6" slide slooe for Exoer- reqardless of crew 1oadina.l In these exoeriments.

iments 3 through 6), constani speed tracking dur- control force and control-position stability were tied toqether throuqh the use of electrohvdraulic ing the descent (except Experiment 6), and control-loaders, and so the requirement would pro- transition to a constant-speed missed-approach hibit the neutral and unstable gradients that were maneuver consisting of a standard-rate turn at considered. Considerations for airworthiness climb rates varying from 600 to 1000 ft/min, with acceptance are likely to center on those configura- the transition occurring at the missed-approach tions whose flying qualities are assessed to fall point in the first two experiments and at the deci.- Experiment between satisfactorv and adeauate, but there is no sion height in Experiments 3 through 5.

clear correlation between acceptance and the 6 included a deceleration while on instruments Cooper-Harper pilot ratinq. All of the ratinqs according to one of three deceleration profiles, fall within the adequate category, and the di?fer- and considered two approach geometries (Fig. 8), ences between stable and neutral gradients in but a missed approach was not included. Table 3 individual experiments generally amount to about one pilot rating or 1ess.s~11*12 Taken together, Influence of the Stability and Control Augmentation therefore, the results indicate that the achieve- System ment of a clearly adequate (e.g., CHPR< 5) capabil- ity probably justifies the requirement for a stable It was noted in discussing the static gradient gradient, but a neutral gradient might be margin- results that no ratings in the satisfactory cate- ally acceptable for the dual-pilot situation. gory were achieved for the tasks considered using rate-dampinq stability auqmentation. Fiqure 12 Influence of Other Baseline Characteristics shows the ratings assigned to the three types of Ditch and roll SCAS considered most consistentlv As was discussed earlier, some modifications across all the experiments: rate damping with " to some baseline teetering-rotor model character- input decoupling, rate-command-attitude-hold, and istics were considered in the fifth experiment to attitude command. These cases are primarily for ascertain any influence of these characteristics the SCAS incorporated on a machine with neutral on the types of results discussed above. Figure 10 basic longitudinal stability; note that a rate- shows the data from this experiment for configura- damping SCAS does not alter the control position tions with a high steady-state attitude-speed rela- oradient. a rate-command-attitude-hold SCAS resuZts tionship (obtained through the introduction of in a neutral gradient, and an attitude SCAS stabil- high-drag damping X,). As can be seen, little izes the gradient because of the Ma term. As has chanqe in average rating is evident for the neutral been pointed out in the reference for each experi- or stable gradients, with a small improvement for ment,.attitude augmentation in pitch and roll.

the unstable gradient. The pilot comments for (imolemented either as rate-command-attitude-hold these configurations demonstrate mixed reactions or attitude command) is required to achieve ratings One pilot consistentlr rated the and difficulties. in the satisfactory category.7$12 The advantages high-drag configurations as better than the low- include a reduction in interaxis coupling, reduced drag ones because small speed changes resulted in turbulence excitation, and improved short-term and fairly siqnifi‘cant rate of climb changes as a lonq-term dynamics. It is interestinq to note that result of-the increased negative dy/du; hence rate the-failed longitudinal damper considered in Experi- of climb could be well controlled using pitch atti- ment 5 still had characteristics that met the cri- tude. The other pilots, however, noted that the teria of Ref. 2 (with stable gradient) and yet was requirement for large power changes with speed was rated marginal at best in turbulence.]* Because a detriment, particularly since power was still the the criteria do not directly assess short-term primary controller for rate-of-descent; therefore dynamics, acceptance of a failed state for this the required changes for speed led to apparent confiquration would rest entirely in the hands of speed-and-rate-of-descent coupling, thereby negat- the certification pilot and would likely not be granted, even though the criteria are met.

ing any advantages of more precise speed control.

As a result, therefore, in general the average ratings for the equivalent high-drag and low-drag Influence of Flight Director Displays configurations were about the same, both in no tur- bulence and in light turbulence. As a result, it Figure 13 illustrates some of the data is unlikely that the low attitude-to-speed gradient obtained concerning the influence of three-cue flight directors compared with raw-data displays.

of the baseline machine significantly influenced The Experiment 5 configurations shown were selected the ratings shown earlier.

because their stability and control characteristics Another modification to the baseline character- are virtually identical to those of the Experiment 6 configurations; these Experiment 6 data were "cali- istics was the introduction of a large increment in The data for this modi- bration" evaluations obtained with no deceleration anole-of-attack stability.

fication are shown in Fig. 11. As can be seen, the on instruments. As can be seen, some beneficial influence of the three-cue fliqht director displays influence on the oilot ratinq is high in turbulence, with the high angle-of-attack stability configura- is apparent in the Experiment 3 results, parti&-- tions being rated as inadequate for the task. As larlv with the hiqher level of SCAS (attitude auq- Considering all the experiments, in- is discussed in Ref. 12, the addition of this sta- mentation).

general the flight director assistance did improve bility did not significantly influence the longi- ratings given to the rate-damping control system tudinal control position gradient, but did lead to an "insidious" coupling between rate-of-descent sufficiently to provide a clearly adequate capabil- and speed control. Pilot comments indicated that ity, but did not improve this SCAS type sufficiently for these configurations the angle-of-attack sta- to move it into the satisfactory category. With the attitude-tvoe SCAS. however. the assistance of bility coupled through pitch attitude to large the flight directors generally pushed the ratings inadvertent speed chanqes when large changes in The rate-of-descent were made with the collective. clearlv into the satisfactorv cateaorv. This lack imoortant ooint brouqht out by these data is that of subitantial overall benefjt of the"flight direc- tors for the rate-damping SCAS type was not coupling effects have a major-influence, ahd yet exoected at the outset of the experiments, and it the criteria of Ref. 1 do not consider such effects should be cautioned that the results are likely to at all. For helicopters, other typical types of be auite sensitive to the desiqn method used.1oy13 coupling are cross-axis inputs (eliminated for most Based on these data, relaxed airframe airworthiness of the configurations investigated in the program) requirements, because of "credit" for advanced dis- and pitch-roll coupling, particularly for hingeless- plays, may be warranted in some cases, and the rotor machines; such effects should probably be absence of consideration for displays in the IFR considered quantitatively for airworthiness accept- Appendix1 may require further attention.

ance.

Influence of Task damping augmentation, even at a fairly high level and with input decouplinq, qenerally has received ratings ranging from'marginally adequate to just Because the Cooper-Harper pilot rating applies worse than satisfactory, depending on other factors. to an airframe-control-system display combination A failed rate damper was considered marginally for a specific task, and because the evaluation inadequate, even though the aircraft characteris- tasks have varied somewhat across these experiments, tics were still within the IFR Appendix criteria. it is useful as a final comparison to examine the Ratings from influence of the task on the ratings.

several of the experiments are compared in Fig. 14 4) The addition of three-cue flight directors for similar stability and control characteristics did not improve the IFR capability for rate-damping control systems to the satisfactory category, if and displays as a function of the task that was considered. It should be noted in particular that all the experiments are considered; some beneficial effect in achieving ratings in the satisfactory the difference between the dual-pilot and single- category with an attitude-augmented SCAS was appa- pilot tasks considered in Experiment 3 resulted in Inadequate flying qualities could not be a change of almost one pilot rating, justifying in rent.

improved to satisfactory with the use of flight principle the division in criteria for normal- category helicopters in the IFR Appendix, but leav- directors, but the improvement might take a marginal configuration into the clearly adequate category.

ing in question the lack of distinction for This possible improvement is not considered in the transport-category he1icopters.l It may also be seen that a decelerating instrument approach leads current criteria.

to worse ratings than even the single-pilot task Decelerating 5) Increasing the difficulty of the task with a constant-speed approach.

single-pilot or inclusion of an instrument approaches are not explicitly.considered by the IFR (e.g., deceleration) did result in degraded ratings for Appendix,l and these data intimate that more strin- equivalent configurations. A difference in gent criteria may be required for these more demanding tasks. requirements for single- and dual-pilot operations was therefore shown to be warranted. Similarly, a Concluding Remarks difference in requirements of future versions which consider decelerating instrument operations may be A sequence of ground- and flight-simulation projected.

experiments concerning helicopter IFR airworthiness References has been described in this paper. A total of over 800 piloted evaluations of several aspects of con- cern for helicooter instrument flight was obtained 1. "Rotorcraft Regulatory Review Program Notice in these experiments. Although there are varia- No. 1; Proposed Rulemaking," Federal Regis- tions in detail among the experiments, the general ter, Vol. 45, No. 245, 18 Dec. 1980.

results with respect to IFR airworthiness can be On the basis of these results, as pre- compared. 2. "Federal Aviation Regulation Part 27 - Airwor- sented here and in previous documentation of the thiness Standards: Normal Category Rotor- experiments, the following conclusions may be craft," Federal Aviation Administration, drawn, particularly concerning the proposed IFR Feb. 1965.

Appendix: "Federal Aviation Regulation Part 29 - Airwor- 3.

1) The criterion requirinq a stable longitud- thiness Standards: Transport Category inal force gradient with speed ;s probably justifi- Rotorcraft," Federal Aviation Administra- able for rate-damping types of SCAS, although tion, Feb. 1965.

little significant degradation has been shown with neutral or slightly unstable gradients; hence the 4. "General Requirements for Helicopter Flying and neutral gradient, at least, could be considered Ground HandlingQualities,l'Military Stand- It should be emphasized marainallv acceotable. ard MIL-H-8501A, Sept. 1961.

that a rate-command-attitude-hold-type of SCAS, as considered in these experiments, results in a.neu- 5. Lebacqz, J. V., "Survey of Helicopter Control/ tral longitudinal gradient; this type of configura- Display Investigations for Instrument tion was generally rated in the satisfactory Decelerating Approach," NASA TM-78656, category. Hence, this type of criterion needs to 1979.

be linked to the type of SCAS employed, which it currently is not.

6. Traybar, J. J., Green, D. L., and DeLucien, A.

G. . "Review of Airworthiness Standards for 2) Inherent characteristics of the helicopter Certification of Helicopters for Instru- lead to a varietv of tvoes of interaxis coupling. ment Flight Rules (IFR) Operations," Fed- One type explicitly considered in these experi- eral Aviation Administration Report No.

ments led to a considerable degradation in pilot FAA-RD-78-157, Feb. 1979.

ratings. The current IFR Appendix does not address off-axis coupling; perhaps future versions should.

3) In all the experiments, attitude augmenta- tion in pitch and roll has been required to achieve pilot ratings in the satisfactory category. Rate 7. Forrest, R. D., Chen, R. T. N., Gerdes, R. M., 12. Lebacqz, J. V., Forrest, R. D., and Gerdes, Alderete, T. S., and Gee, D. R., "Piloted R. M., "A Ground Simulator Investigation Simulator Investiqation of Helicopter Con- of Helicopter Longitudinal Flying Quali- trol Systems Effects on Handling Qualities ties for Instrument Approach," NASA During Instrument Flight," Preprint No. TM-84225, 1982.

79-26, 35th Annual National Forum of the American Helicopter Society, Washington, 13. Lebacqz, J. V., "Summary Report - Fifth NASA/ D. C., May 1979. FAA Helicooter Instrument Certification and Operation Simulation Experiment," and Forrest, R. D., "A Piloted informal memorandum available from FAA, 8. Lebacqz, J. V., Simulator Investigation of Static Stabil- Ames Research Center, Moffett Field, ity and Stability/Control Augmentation Calif., Dec. 1981.

Effects on Helicopter Handlinq Qualities for Instrument Approaches," Preprint No. Chen, R. T. N., Talbot, P. D., Gerdes, R. M., 14.

80-30, 36th Annual National Forum of the and Duoan. 0. C.. "A Piloted Simulator American Helicopter Society, Washington, Investigation of-Augmentation Systems to D. C., May 1980. Improve Helicopter Nap-of-the-Earth Handling Qualities," NASA TM-78541, 1979.

9. Lebacqz, J. V., Forrest, R. D., and Gerdes, R. M., "A Piloted Simulator Investigation 15. Chen, R. T. N., "A Simplified Rotor System of Static Stability and Stability/Control Mathematical Model for Piloted Flight Augmentation Effects on Helicopter Hand- Dynamics Simulation," NASA TM-78575, 1979.

ling Qualities for Instrument Approach," NASA TM-81188, Sept. 1980. 16. Aiken, E. W., "A Mathematical Representation of an Advanced Helicooter for Piloted Simulator Investigations of Control System 10. Lebacqz, J. V., Forrest, R. D., Gerdes, R. M., and Merrill, R. K., "Investigation of and Display Variations," NASA TM-81203, Control, Display, and Crew-Loading 1980.

Requirements for Helicopter Instrument Approach," AIAA Paper 81-1820, Albu- querque, N. Mex., Aug. 1981.

11. Lebacqz, J. V., Weber, J. M., and Corliss, L. D., "A Flight Investigation of Static Stability, Control Augmentation, and Flight Director Influences on Helicopter IFR Handling Qualities," Preprint No.

81-25, 37th Annual National Forum of the American Helicopter Society, New Orleans, La., May 1981.

Table 1. Summary of longitudinal control position gradients.

- -_ _ ..". -_. . . _. __. . . . _-__ ._- _..

Time-to-double Gradient, Rotor Experiment Configuration amplitude, in./15 knots set _-.--_-- _.__..... _-..-.. ".. _...... __ _.__....._ +0.06 1 Teetering 5.8 Hingeless -0.05 2 Hingeless Neutral -0 Hingeless Stable -0.63 -0.02 Teetering Neutral 3 Teetering -0.02 4 Teetering More stable --0.50 --0.25 Base UH-1H Neutral -0 5 Teetering Most stable -1.03 Stable -0.53 Neutral -0.03 Unstable +0.03 11.0 to.125 6.3 Most unstable 6 Teetering -0.41 ._ _-_ -. -. _. . -- .

.::-.‘-:':::::-:I _.- ..,.. _ ,..__ : .I: -- . -.~. . - _ Table 2. Longitudinal derivatives of baseline teetering-rotor helicopter at 60 knots.

Derivative Units Value .- rad/sec*/ft/sec -0.00022a rad/sec2/ft/sec -0.00278 : l/set -0.847b Mq l/set t0.143b MP rad/sec*/in. 0.17b Mse rad/sec*/in. 0.0223b M6C l/set -0.005a XU l/set 0.026 XW l/set -0.013a ZU l/set -1.28 ZW ft/sec*/in. -2.58b z6e ft/sec*/in. -10.00 Z% . ~_ .- ._-----~~ aBaseline, unmodified for gradient changes.

bNo SCAS.

Table 3. Task details.

.~. . -___ - - . _- .

Decision height, Experiment Guidance Speed profile Missed approach ft AGL - 1 . - - _-___-~~ ~ 1 VOR 60 knots, constant 600 Yes VOR Decelerate 80-60 knots 600 Yes before let-down, 60 knots constant thereafter 3 6" MLS Decelerate 80-60 knots 300 Yes before vertical intercept, 60 knots constant thereafter 4 6" MLS Constant 60 knots 200 Yes 5 6" MLS Decelerate 80-60 knots 300 Yes before vertical intercept, 60 knots constant thereafter 6 6" MLS Constant 60 knots until 130 No -0.5 n.mi. to go, decelerate to -15 knots on instruments ~..:.I:.I:fl::~lI=_~---~----------~ __.

I Ill Fig. 1 Instrument panel layout.

ATTITUDE FLIGHT DIRECTOR FLAG FLAG COMMAND BAR DEVIATION RATE OF TURN \ INDICATOR* INCLINOMETER NOTES *INPUT FROM ROLL/YAW RATE GYRO ASSEMBLY *‘INPUT FROM RADIO ALTIMETER (a) Attitude director indicator.

BEARING 1 HEADING COMPASS HEADING POINTER INDEX WARNING SELECT TO-FROM POINTER >ME 2 DME 1, / AIRPLAE COULE HERDING BEARING 1 SYMBOL MASK DIAL DEVIATION BEARI’NG 2 DATA SOURCE DATA SOURCE SELECT SWITCH SELECT SWITCH (b) Horizontal situation indicator.

Fig. 2 Flight director displays.

ROLL DECISION RANGE ROLL POINTER (FLASH AT DECISION COLLECTIVE INDICES RANGE) DIRECTOR I / 1. ALTITUDE TAPE 16. HORIZON BAR 2. VERTICAL SPEED 17. AIRCRAFT SYMBOL 3. THRUST MAGNITUDE (FLASH FOR DECEL) CONTROL DIRECTOR 18. SIDESLIP 4. ROLL POINTER 19. PITCH ATTITUDE 6. PITCH & ROLL STICK 20. WIND DIRECTION DIRECTOR INDEX 21. HEADING SCALE 7. LATERAL STICK 22. GROUND VELOCITY CONTROL DIRECTOR STATUS VECTOR 8. LONGITUDINAL STICK (APPEARS AT DECEL.)

DIRECTOR 23. GROUND VELOCITY 9. LANDING PAD VECTOR COMMAND / 0 (APPEARS AT (APPEARS AT DECEL.)

RAW DATA DECISION RANGE) 24. LATERAL COURSE GLIDE SCOPE 10. AIRSPEED OFFSET 11. RADAR ALTITUDE 25. GLIDE SLOPE 12. ALTITUDE INDEX (FLASHES AT 13. TORQUE INTERCEPT) LOCALIZER 26. IVSI 14. ROTOR RPM GROUND ERROR 15. RANGE TEXTURE Fig. 3 C format for Experiment 6. Fig. 4 X format for Experiment 6.

Fig. 5 Flight Simula.tor for Advanced Aircraft.

Fig. 6 Vertical Motion Simulator.

Fig. 7 UH-1H V/STOLANO helicopter.

If Y w ‘IG 4Gt-- -8500 ft --+-. -8500 f-t - RANGE DECISION RANGE, t DECEL.

- t-- -4600 ft +- -8500 ft -+-J----r 400 ft RANGE Fig. 8 Approach profile geometries. .

EXPERIMENT EXPERIMENT 0 1 23’ A 4 v 5 INADEQUATE 6 - ADEQUATE SATISFACTORY 2 SATISFACTORY 2- STABLE N = UNSTABLE t t I I I I I I I I I I I I I I 1 I -1.2 -1.0 -.8 -.6 -.4 -.2 0 .I .2 -1.0 -.8 -.6 -.4 72 0 .I .2 SE@, in/15 knot l/TD, l/set S ES/V, in/15 knots (a) No turbulence, no flight directors. (b) In turbulence, no flight directors.

Fig. 9 Pilot rating data as function of Tongitudinal stick gradient.

_-.. - . . .- , . ., AVERAGE,TURBULENCE AVERAGE, NO TURBULENCE dr9/du = -0.03 DATA FROM FIG. 6 FOR EXPERIMENT 5 I I I I J .lO .05 0 .05 .I0 1 IT%, 1 hec l/TD, ‘I/s~c Fig. 10 Influence of de/du = -0.33"/knot (from Experiment 5).

0 M, = 0 FROM FIG. 6 FOR EXPERIMENT 5 m M, = -0.025, NO TURBULENCE ti M, = -0.025, TURBULENCE 6- : n

4- -- 0

l o

2- STABLE -- UNSTABLE I t I I I -_ .I0 .05 0 .05 .I0 .15 .20 l/TX, llsec InD, l/set Fig. 11 Influence of M, = -0.025 (from Experiment 5).

EXPERIMENT 0 1 :: A4 v5 EXPERIMENT +‘3 ’ STABLE BASELINE INADEQUATE GRADIENT ADEQUATE SATISFACTORY t I I I I

IL--L-

RATE RATE 1ATTITUDE FAILED FAILED RATE RATE SCAS DAMPING COMMAND COMMAND RATE SCAS COMMAND, ATTITUDE (LONGITUDINAL) ’ INPUT (LONGITUDINAL) ATTITUDE DECOUPLING HOLD HOLD (NEUTRAL des/V) RATE ATTITUDE DAMPING, COMMAND INPUT DECOUPLING SCAS TYPE (b) In turbulence, no flight directors.

(a) No turbulence, no flight directors.

Fig. 12 Influence of SCAS.

11111111111lll IIII III II

EXPERIMENT . 3, DUAL PILOT a 3, SINGLE PILOT A4 v5 +S RAW 3-CUE RAW 3-CUE DATA FLIGHT DATA FLIGHT DIRECTOR DIRECTOR a) Rate-damping, input-decoupling SCAS.

b) Attitudecommand SCAS.

Fig. 13 Influence of three-cue flight director: in turbulence, dual pilot.

EXPERIMENT F 2 8- 2,’ FJ 1 +6 6- n F 4 RATE SCAS a 0

l

g4- 5 0 l ATTITUDE SCAS a n 0 I I I I I I J MLS MLS MLS MLS MLS VOR CONST. CONST. DECEL.

CONST. CONST. CONST.

SPEED APPROACH SPEED SPEED SPEED SPEED DUAL DUAL DUAL SINGLE DUAL SINGLE PI LOT PILOT PILOT PILOT PILOT PI LOT I \ J , Y Y RAW DATA DISPLAYS FLIGHT DIRECTOR DISPLAYS Fig. 14 Influence of task: in turbulence.

STATE-OF-THE-ART COCKPIT DESIGN FOR THE HH-65A HELICOPTER Daniel E. Castleberry HX Program Manager and Marsha Y. McElreath Avionics Pilot Instructor Collins Government Avionics Division of Rockwell International Cedar Rapids, Iowa Abstract ’ On 14 June 19’79, the USCG awarded Aerospatiale Helicopter Corporation (AHC) a contract for a Short Range Recovery The design of the US Coast Guard HH-65A “Dolphin” cock- (SRR) helicopter, the HH-65A. Teamed with AHC, Collins Gov- ernment Avionics Division of Rockwell International designed pit employs advanced integrated electronics systems technology to achieve several important goals in this multimission the avionics system for the SRR helicopter.

aircraft: Integrated Cockpit Design 1) Integrated systems operation with consistent, simplified The Rockwell solution to the Coast Guard design mandate cockpit procedures exceeds mere automation. The HH-65A cockpit design achieves 2) Mission-task-related cockpit displays and controls three additional goals: (1) integrated systems operation with consistent, simplified cockpit procedures, (2) mission-task- 3) Reduced pilot instrument scan effort with excellent related cockpit displays and controls, and (3) reduced pilot instrument scan effort with excellent outside visibility.

outside visibility In order to meet these goals, Rockwell-Collins has imple- To achieve these goals, Collins Division of Rockwell has mented the integrated avionics system to depend heavily upon designed the avionics system to rely heavily upon distributed distributed but complementary processing, multiplex digital but complementary processing, multiplex digital bus technol- ogy, and multifunction CRT controls and displays.

bus technology, and multifunction CRT controls and displays.

This avionics system has been completely flight tested and will soon enter operational service with the Coast Guard. Distributed but complementary processing is an important integration concept used in the HH-65A. Its architecture does Introduction not hinge on one centralized computer for processing all navi- gation signals, displays, control inputs, etc. Instead, distributed “On weekend duty, I always expected to be awakened in the processors perform specialized functions. The system coupler middle of the night,” recounts a veteran pilot of the USCG, unit (SCU) manages communications between the boxes and “and when flying search and rescue (SAR) off San Francisco controls radio tuning. The control display unit (CDU) provides Bay in summer, I always expected fog.” So it came as no sur- pilot access to all flight management operations. The horizon- prise when an early morning call interrupted his sleep. Tasked tal situation video display (HSVD) driver unit generates the with finding an overdue sailboat and two-man crew, the pilot navigation displays, and the mission computer (MCU) acts as and copilot strapped themselves in their HH-52 and lifted off. both navigator and flight engineer. Without pilot action, the Passing over the harbor lights, they soon disappeared into the MCU calculates a “best estimate” of present position and veloc- night. Searching amidst dense fog, the pilot contacted the har- ity, automatically tunes the navigation sensors, enables flight bor radar control and asked for position confirmation plus planning, RNAV-style (including the generation of special advisories of nearby objects or vessels. The controller quickly USCG patterns), monitors fuel consumption, and records the warned him of a ship at his twleve o’clock position and asked engine and transmission condition (Fig. 1).

if he could see it. “I don’t see anything,” responded the pilot.

These specialized processors perform distinctive tasks; yet, Suddenly, the drone of a fog horn cut through the night fog.

they cooperate as a single integrated system to accomplish Max power! Positive climb! Forward airspeed! With reflex mission objectives. A high-speed mvltiplex digital data bus action, the pilot executed an instrument take off, a Coast enables uninterrupted communication between the avionics.

Guard maneuver which transitioned the helicopter up into San Using discrete addresses, any two boxes can communicate with Francisco’s terminal control area-also, a less than desirable each other on the bus. To fly to a point, for example, the pilot flight situation, but in the pilot’s judgment, the better option.

indicates his intent on the CDU, which in turn communicates that intent to the mission computer. The MCU computes and Communication, navigation, flight control, and search sensor displays the aircraft’s navigational situation on the HSVD and management are classical avionics functions which constitute CDU, then, sends roll commands to the flight director (FD), every SAR operation. In theory, however, communication, navi- which executes the commands through the automatic flight gation, and flight control are merely handmaids to the search control system.

effort-the sole reason for the mission. Yet, how much atten- tion could the pilot in this account devote to finding the miss- Although centralized versus distributed processing does not ing sailboat? Because routine cockpit duties often monopolize necessarily alter cockpit operation, system survivability argues crew attention during SAR operations and thus impair crew for distributed processing. A mission computer failure, for effectiveness, the United States Coast Guard presented indus- instance, impacts only RNAV capability; automatic navigation try with a challenge: Build an avionics system that automates via TACAN, VOR, or localizer is not impacted. LORAN, con- the routine tasks of communication, navigation, mission man- trolled through the system coupler unit, also remains valid; agement, and flight control, and therefore, frees the crew to and since the HSVD display drivers process all VOR and TACAN signals plus generate the navigation displays, the crew focus on the mission tasks which only they can perform-the visual search and FLIR or RADAR interpretation.* retains display guidance.

I IlIlIlllllllllI llllllllllllllll I Ill Ill1 I I II II I

Fig. 1. Distributed but complementary processing.

Fig. 3. Control display unit (COMM radio control display).

Another important integration tool is using one device to do the work of many. Four multifunctional CRT devices, dual con- The CDU is a single-point control for all flight management trol display units (CDU’s) and dual horizontal situation video operations. By incorporating “function keys,” the CDU controls displays (HSVD’s), inhabit the HH-65A cockpit (Fig. 2).

numerous mission tasks: For example, pushing the COMM or NAV button dedicates the CDU to COMM or NAV radio tun- ing. Selecting FPLN dedicates the CDU to flight planning.

Likewise, pushing the PROG or STAT keys transforms the CDU into a flight progress or status reporting device. Having assigned the control display unit to a particular function, the crew uses the “line select keys” adjacent software labels to (1) tune individual radios, (2) set transponder codes and modes, (3) insert waypoints, plus a host of other functions (Fig. 3).

Because the CDU centralizes all operational inputs, it simpli- fies pilot procedures. He communicates, navigates, flight plans,

f

r ’ HSVD

HSVD etc, without having to manage dedicated controls scattered throughout the cockpit. Furthermore, CDU pilot procedures are uniform. Whether the pilot tunes a COMM/NAV radio, changes the transponder code, or enters a waypoint, he uses identical procedures.

(COPILOT) (PILOT) As the CDU is a central point of avionics control, so the horizontal situation video display (HSVD) is a central point for flight situation displays. The HSVD supplants several dedi- cated instruments: the conventional HSI, projected map, RADAR and FLIR displays, as well as a hover indicator (Fig. 4).

Nonetheless, merely replacing conventional instruments is not the purpose of the HSVD. Rather, it organizes data into “task-related“ modes which not only present the pilot informa- tion needed for specific mission phases but also eliminate extraneous information. Consider the low altitude hover over Fig. 2. HH-65A panel and console layout. water at night. Because the pilot generally faces a centrally Data Fig. 4. Horizontal situation video display modes.

positioned HSI which provides virtually no hover information, of the SAR environment?” The following scenario intends to he scans several other instruments to interpret his hover situ- demonstrate integrated system operation, specifically, as it ation. The HSVD’s hover mode integrates all hover data into impacts cockpit procedures and workload in the SRR helicop- one central display: omnidirectional airspeed, longitudinal/ ter. Assume that a pilot were flying a routine patrol when the lateral drift, radar altitude, computed wind, plus target posi- rescue coordinator calls and instructs him to proceed directly to the site of a ditched aircraft, initiate a search and rescue tion.

reported survivors.

Remaining HSVD modes, likewise, satisfy other flight phase To navigate to the downed aircraft, the pilot types in the requirements: The HSI mode is primarily an approach display.

LAT/LONG position on his CDU (the mission computer also The MAP mode serves en route navigation, where the flight recognizes LORAN TD’s, place-bearing-distance, or identifiers) plan ahead may be viewed. The RADAR and FLIR modes dis- and selects DIRECT TO. The mission computer creates a direct play the video images from these sources for searching. The course to the point. It also continuously plots present position RAD4R MAP mode relates radar returns (weather/ground) to using dual LORAN, dual VOR, TACAN, dual compass systems, the flight plan. And the DATA mode, a north-up chart presen- and precision omnidirectional airspeed sensor inputs; manages tation, facilitates impromptu flight planning.

the navigation sensors (ie, automatically selects navigation sta- tions and tunes the LORAN, VOR and TACAN receivers); and Besides suiting information to flight phases, task-related dis- flies the aircraft to the waypoint through the flight director.

plays denote “complementary formatting.” For example, The HSVD MAP mode simultaneously displays the flight plan.

because a pilot navigating cross-country uses wind information This mode combines a tactical map presentation of flight plan to plan the flight, the MAP mode incorporates a digital wind waypoints and an abbreviated HSI, which the pilot uses with readout. By contrast, the pilot in a hover does not need wind the progress and flight plan displays of the CDU to monitor en information for flight planning; he needs to visualize wind route progress.

velocity relative to the helicopter. Consequently, the hover mode incorporates a modified Beaufort wind arrow, which Meanwhile, the mission computer has already assessed the instantly pictures the changing wind velocity. Each pilot needs fuel situation. Accounting for wind, the MCU calculates the computed wind information but in a complementary for- fuel required to fly to the search point, proceed to the destina- mat-dictated by the flight situation.

tion, and leave a 30-minute reserve. If on-board fuel is insuffi- cient, the system warns the crew by announcing FUEL ALERT Typical SAR Operation on the CDU. If sufficient fuel exists, the STATUS display translates the fuel reserve (ie, fuel in excess of what’s needed Thus far, we have described technical features of the HH- to fly to the destination) into hours and minutes of flight time, 65A avionics system. At this juncture, one might ask, “How labelled BINGO. MCU fuel management gives the pilot instant does the integrated avionics system aid the pilot in the context visibility of his fuel status, and thus, how long he can search.

Fig. 5. Sector search entry into flight plan.

While the mission computer monitors fuel consumption, the Upon reaching the target area, the aircraft automatically data link system reports en route progress to the search coor- initiates the search while the crew concentrates on the search dinator, relieving the pilot of routine position reporting. He RADAR, FLIR (forward-looking infrared), and DF radio hom- merely designates the communication radio and transmission ing, or they scan the white caps below. When the target is interval on the CDU DATA LINK display. At the specified spotted, the integrated system, with minimal crew effort, aban- time, the integrated system automatically downlinks 9 pieces dons the search and expedites the rescue operation. Over-flying of information regarding aircraft position, status, and flight the target location, the pilot pushes two buttons: MARK - to progress. mark the target’s location, and HOVER - to call up the approach-to-hover pattern. He inserts the approach-to-hover Eliminating routine flight management tasks frees the pilot pattern into the flight plan and selects APPR on the flight to concentrate on system performance and flight progress. director panel - triggering a chain of operational events. The Pushing the PROG key on the CDU calls up the computed pre- system turns the aircraft downwind to ensure a final approach sent position (LAT/LONG) and ground speed. Pushing the line into the wind, directs a minimum time procedure turn, and key adjacent any flight plan waypoint provides instant access computes a five degree descent to the hover transition point.

to “waypoint data” for that geographical point-time, distance Using the FD speed beep, the pilot. may vary the approach and course to the waypoint via the flight plan or via direct. speed. At loo-feet radar altitude, the FD APPR mode drops; T- HOV mode captures and slows the helicopter to zero ground As the aircraft nears the search area, the pilot plans his speed at 50 feet RA - just short of the target (Fig. 6). During search. He selects one of three available patterns (sector, lad- the transition to hover, the HSVD automatically displays the der, or expanding square) and then defines the pattern param- HOVER mode. The computed wind, HOVER velocity com- eters. For example, if he selects a sector search, the computer mands, omnidirectional airspeed vector, and the marked target position enable the pilot to monitor the approach-to-hover asks what track spacing is desired (Fig. 5). (NOTE: The pilot maneuver as well as modify the hover conditions. If the pilot may request search advisories by entering the sea state, visi- bility, cloud cover, and altitude; the MCU will compute the beeps either radar altitude or longitudinal/lateral airspeed, the optimum track spacing.) Selecting “INSERT -+” displays the indicators instantly verify his input.

flight plan, where inserting the pattern requires only pushing While the survivors are hoisted to safety, the pilot decides a line key at the desired datum point. The mission computer automatically plots the pattern waypoints and displays them his next course of action. Should a victim require immediate on the HSVD. medical attention, he may choose to fly to a medical center MARK TURN A A L-l) AUTOMATIC TRANSITION TO HOVER

-fh-

-3

“-=A

AT 50’ RADIO ALTITUDE NOTE: Pilot selects Approach (APPR) and Transition to Hover (T-HOV) Flight Director Modes in order to fly entire maneuver automatically f Fig. 6. Approach-to-hover maneuver.

rather than home base. With the push of the DATA mode but- ton, the HSVD displays surrounding hospital locations in a north-up, chart presentation (Fig. 7). To examine direct dis- tance, time or course to any viable alternate, the pilot simply calls up waypoint data for the respective hospital through his oHLJ59 CDU. If desired, the MCU will also compute the maximum range on that course. Once again, minimal pilot action acti- q STUB 1 024.3 vates integrated system response, to enhance crew effective- 42.5 ness.

mWPT A The technological tools of digital data bus communication, distributed but complementary processing, and multifunction q NOPHS •IMERCY CRT controls and displays have effected integrated cockpit operation in the HH-65A. Although this system has been @CID implemented for a SAR application, these techniques and this ,,A, -RO t,Y10 approach to operational cockpit integration will adapt to any /’ \ helicopter mission. A system coupler unit and CDU which cur- /’ ‘\ I’ \ rently controls radios could as easily control weapons systems.

,’ \ I A mission computer and HSVD might as easily display termi- ‘\ \ nal area approach procedures or tactical combat command and %% ‘a R1O control data. Meanwhile, the HH-65A with its integrated cock- _I’ pit operation, will benefit Coast Guard line pilots who under- take SAR despite adverse conditions.

*Cdr. David A. Young, “Avionics System Design Requirements for the United States Coast Guard HH-65A Dolphin”: Presented at the Sixth European Rotorcraft and Powered Lift Aircraft Forum, Bristol, United Kingdom, September 16-19, Fig. 7. HSVD data mode.

1980.

PERFORMANCEEVALUATION OF A KIWESTRETIC-TACTUAL DISPLAY Richard J. Jagacinski, John M. Flach, and Richard D. Gilson Associate Professor, Graduate Student, and Professor The Ohio State University Columbus, Ohio Richard S. Dunn Engineering Psychologist U. S. Army Research and Technology Laboratory Moffett Field, California 1,2,3 In a series of Abstract Engineering. studies he and his colleagues demonstrated the display's Simulator studies demonstrated the feasi- usefulness in improving the precision with which bility of using kinesthetic-tactual (KT) dis- car drivers could control the distance between plays for providing collective and cyclic command themselves and a vehicle in front of them.

information, and suggested that KT displays may An example of a single dimensional KT dis- increase pilot workload capability. A dual-axis laboratory tracking task suggested that beyond play as it might be used on a helicopter collec- reduction in visual scanning, there may be addi- tive handgrip is shown in Fig. 1. An electro- mechanical slide protrudes from the surface of tional sensory or cognitive benefits to the use of multiple sensory modalities. Single-axis the handgrip to indicate the direction and magni- laboratory tracking tasks revealed performance tude of tracking error. If there is zero error, with a quickened KT display to be equivalent to the slide is flush with the handgrip. If the performance with a quickened visual display for slide protrudes downward, the pilot moves the a low frequency sum-of-sinewaves input. The collective in the downward direction until the trackers approximated a lag in these tasks. In slide returns to the flush position.

contrast, an unquickened KT display was inferior to an unquickened visual display. The trackers A two-dimensional KT display as might be approximated a proportional element in these used on a helicopter cyclic handgrip is shown in Full scale simulator studies and/or The electromechanical slide is in the tasks. Fig. 2.

inflight testing are recommended to determine form of a ring that is flush with the control grip the generality of these results. when there is zero tracking error. The protrusion of the ring from the control grip represents a Introduction vector composite of lateral and longitudinal errors. The appropriate response is to move the The kinesthetic-tactual (KT) display has cyclic in the direction of the protrusion until been under development and evaluation since 1966. the ring is again in the flush position. 'The It provides a useful display alternative for vectoral nature of this display seems to be highly compatible with the two dimensional cyclic helicopter tasks which have high visual workload or which are incompatible with visual or auditory movement.

display devices. Examples include terrain flight Fixed Wing Aircraft Study with high demands for visual attention outside the cockpit and night flight with viewing aids which are not fully compatible with cockpit One use of the KT display has been to pro- visual displays. Numerous laboratory and simula- vide pitch commands in fixed wing aircraft.

tion studies have been conducted to develop Gilson and Fenton4 measured the performance of prototype KT displays and to measure performance novice pilots in a Cessna 172 with three dif- (1) a visual display with these displays. These studies show the ferent types of displays: concept to be feasible for helicopter application of airspeed; (2) a visual display of deviations and effective at visual workload relief. This from a desired angle of attack; (3) a KT display report first summarizes some early studies of deviations from a desired angle of attack.

oriented to workload and feasibility issues, and The KT display was mounted on the control yoke then discusses some data which provide more handle, and pilots minimized protrusion of the detailed quantification of KT display performance. display from its zero error position with fore- For controlling aft movements of the yoke.

The KT display was invented by Dr. Robert angle of attack in an approach to landing Fenton of the Ohio State Department of Electrical maneuver, the visual and tactual displays of angle of attack were comparable to each other, and both were superior to the visual display of airspeed. In a tight turn about a point at con- stant speed and constant altitude, the KT display f) Fig. 1 Control-display relationship for a one-dimensional kinesthetic-tactual display suitable for a helicopter collective. (Copyright 1979, Human Factors, Vol. 21, p. 80) Simulator (TASS) facilities at the U. S. Army's Avionics Laboratory, Fort Monmouth, New Jersey.

inves- The first study by Gilson, Dunn, and Sun5 tigated performance of an instrument flight rules decelerated landing maneuver in a simu- lated UH-1 helicopter buffeted by wind gusts.

Cyclic commands were indicated visually by hori- zontal and vertical crossbars; pedal commands were indicated visually by a rate of turn needle.

Collective commands were presented either visually by a display similar to a glide slope pointer on the left-hand side of the flight director, or tactually by a single dimensional KT display mounted on the handgrip of the collec- tive. Experimentally it was possible to make a) Leftword movement b) Rightward movement the overall task more difficult by adding a time required required delay to the cyclic roll dynamics. Adaptive circuitry adjusted this time delay so that the relationship for a Fig. 2 Control-display sum of absolute tracking errors of the four two-dimensional kinesthetic-tactual command signals reached a criterion value. The display suitable for a helicopter performance measure was the value of the time cyclic.

delay necessary to achieve this error criterion.

For all five pilots in this study, the KT dis- permitted superior performance to the two visual play permitted a longer time delay than the The superiority of the KT dis- displays in controlling angle of attack, alti- visual display.

This latter maneuver play may be due to reduced visual scanning or a tude, and airspeed.

more cognitive advantage regarding how the pilot requires considerable monitoring of visual cues outside the cockpit. The reduced need for visual processes information from multiple modalities.

scanning with the KT display may account for This issue was addressed in a later laboratory these results. study.

Helicopter Simulation Studies A second helicopter simulation study by Sun6 examined the feasibility of tactually providing Two helicopter simulation studies were con- both collective and cyclic commands while still ducted using the Tactical Avionics System providing other visual information, e.g., situational displays. The simulated helicopter output. In a critical tracking task adaptive was a UH-1. A single axis KT display was circuitry gradually shortens the time constant mounted on the collective handgrip as in the until the task becomes so difficult that the previous study. Additionally, a two dimensional subject loses control. The inverse of this KT display in the form of a ring was mounted on critical time constant at the instant control is the cyclic handgrip. A nonlinear gain was used lost is called the critical root, and is repre- to magnify the.protrusion of the ring for small sented with the symbol Xc.

tracking errors. Wearing flight gloves, pilots were able to use these KT displays to success- In this experiment, the single dimensional fully perform an instrument flight rules KT display was mounted on a control stick similar decelerated landing maneuver. With pitch and to a helicopter collective. The visual display roll rate signals used to quicken the cyclic consisted of a vertically moving line on an display, pilots were also able to maintain a oscilloscope screen. The quickened signals stable hover in the presence of simulated wind consisted of a simple addition of error and gusts, and concurrently perform a secondary error velocity with the two equally light-cancelling task. The group means of the critical roots weighted.

are shown in Fig. 3. These results replicate Recent Laboratory Studies the basically additive effects of modality and quickening previously found by Jagacinski, Single-Axis Tracking Miller, and Gilson.g The visual modality was superior, and the quickened displays were In a recent laboratory study at The Ohio the quickened KT display was superior. However, State University by Jagacinski, Flach, and approximately equivalent to the unquickened student subjects were trained on a Gilson, visual display.

critical tracking task using one-dimensional visual or KT displays with or without quickening. Following the critical tracking, subjects A critical tracking task8 requires subjects to were transferred to a stationary compensatory stabilize the output of a first-order unstable tracking task in which they used the same dis- system. Any unsteadiness in the subject's hand plays. The input was a sum of nine sinewaves movements excites the instability and in turn with the amplitudes of the three lowest frequency requires corrective stabilization by the subject. sinewaves (.35, .73, 1.08 r/s) five times greater The difficulty of this task is determined by the than the amplitudes of the other sinewaves.

time constant of the unstable system. The shorter the time constant, the more rapidly the unstable system tends to exponentially amplify small deviations from the desired constant Critical Tracking - Day 7 \” n / Visual Display - t t 6 Unquickened Quickened Stotionory Tracking - Day IO Fig. 3 Critical tracking scores for eight groups Fig. 4 Mean squared error normalized by mean of four subjects. Groups connected by squared input for thirty-one individual dashed and solid lines were respectively subjects. The symbols represent the same transferred to stationary tracking with display conditions as in Fig. 3.

system dynamics 1.5/s and 3.0/(s-1).

TACWAL OUKKENED Fig. 5 Linear transfer functions for the subjects with the lowest mean squared error in each of four quickened display conditions. The circles indicate the data points, and the solid lines represent analytic approximations consisting of a low frequency lag, a high frequency lead, a gain, and a time delay.

Fig 6 Linear transfer functions for the subjects with the lowest mean squared error in each of four unquickened display condi- tions. The circles indicate the data points, and the solid lines represent analytic approximations consisting of a low frequency lag and lead, a high frequency second-order lag, a gain, and a time delay.

Half the subjects controlled a single integrator visual-visual display conditions, the primary system (1.5/s), and half controlled a first- and secondary displays for these conditions were order unstable system (3.0/(s-1)). Mean squared respectively the vertical and horizontal posi- error scores are shown in Fig. 4. The un- tion of a single dot moving on an oscilloscope quickened visual displays were superior to the screen. For the KT display condition, a single unquickened tactual displays.

The quickened dimensional visual display was used for the visual and tactual displays produced equivalent secondary task.

error scores.

The results of this experiment for dual Describing functions were calculated for task performance are shown in Fig. 7. The each subject. For the quickened displays sub- quickened KT display permitted superior perfor- jects were well approximated by a low frequency mance on both the primary and secondary tasks.

lag, a high frequency lead, a gain, and a time In contrast to these results, the quickened KT delay. display and the two primary visual displays As shown in Fig. 5, the describing functions were very similar for the tactual and yielded equivalent performance when subjects visual displays and accounted for about 90% of performed only a single-dimensional critical the variance in the subjects' control movements tracking task alone. Therefore, there seems to be some benefit of combining KT and visual dis- (P2>.

plays in dual task performance beyond what one For the unquickened displays, subjects were might expect from single task performance. This approximated by a low frequency lag and lead, a experimental result is not due to the elimination high frequency second-order lag, a gain, and a of visual scanning because the visual displays time delay (Fig. 6). Overall the linear transfer were integrated into a single moving dot. It functions for the visual and KT displays were may be that using two sensory modalities provides very similar. Subjects using the KT display did, additional attentional resources, additional however, exhibit less low frequency phase lag. sensory buffers, and/or additional cue discrimin- About 60-70% of the variance in subjects' control ability for processing the displayed signals.

was accounted for by the linear transfer func- Further research is necessary to delimit these tions for all but the tactual condition with the possibilities.

single integrator system.

In this condition only about 40% of the variance was accounted for, and One cautionary note should be added con- there were strong peaks in the spectra at non- cerning the generality of this experimental input frequencies in the range of 3 to 7 rad/s. finding.

Preliminary data on dual task tracking Apparently some strongly nonlinear behavior of sum-of-sinewaves inputs without crosscoupling resulted in this condition.

of the two tasks has not so far revealed similar superiority of the combination of KT and visual Dual-Axis Tracking displays. However, these data are still preliminary.

A second laboratory study by Burke, Gilson, and JagacinskilO compared tracking with visual and KT displays when a secondary visual task was DUAL TASK PERFORMANCE performed concurrently.

The primary task required subjects to use their left hands to stabilize a subcritical first-order unstable system.

Three different displays were used for this primary task: (1) a one-dimensional quickened KT dis- play; (2) a one-dimensional unquickened visual display; (3) a one-dimensional quickened visual display for which the signal was additionally passed through an off-line KT display.

This last visual display condition thus had the same bene- ‘d - VISUAL- INTEGRAL DOT (VII fit of quickening and the same detriment of the servomotor lag as the KT display condition.

The secondary task required subjects to use PRIMARY DISPLAYS their right hands to stabilize a different first- order unstable system.

Adaptive circuitry similar to that of Jex, Jewell, and Allen11 adjusted the time constant of the secondary task, until subjects' time-averaged error on the primary task was 25% higher than when the pri- I I I~-! I I If I ! I mary task was performed without significant III secondary task loading. kG DAY 8 DAY 9 DAY IO The performance measures were the washout-filtered time-averaged error on the primary task and the inverse of the Mean performance on a dual tracking task.

time constant for the secondary task, X,. In (Copyright 1980, Ergonomics, Vol. 23, order to avoid the need for scanning in the p. 970) Recommendations References In summarizing the single axis tracking Fenton, R. E., "An improved Man-Machine 1.

results with the KT display, it is helpful to Interface for the Driver-Vehicle System," consider separately the quickened and unquickened IEEE Transactions on Human Factors in displays. The quickened displays may be con- Electronics, HFE-7, (4), Dec. 1966.

sidered analogous to command displays, whereas the unquickened displays are analogous to situa- Fenton, R. E., and Montana, W. B., "An 2.

tion displays used in helicopters. With the Intervehicular Spacing Display for quickened displays, the subjects approximated a Improved Car-Following Performance," IEEE lag, and tracking performance with the KT dis- Transactions on Man-Machine Systems, MMS-9, play was equivalent to that obtained with a (21, June 1968.

visual display for a low frequency sum-of- sinewaves input. On the critical tracking task, Rule, R. G., and Fenton, R. E., "On the the quickened visual display was superior to the Effects of State Information on Driver- ;zf;en:da;Tlgisplay. However, the results of Vehicle Performance in Car Following," IEEE suggest that this difference is Transactions on Systems, Man, and Cyber- due to the servomotor lag in the implementation netics, SMC-2, (5), Nov. 1972.

of the KT display.

Gilson, R. D., and Fenton, R. E., "Kines- 4.

In contrast to these results, the unquick- thetic-Tactual Information Presentations-- Inflight Studies," IEEE Transactions on ened (situation-like) visual display was superior to the unquickened KT display for both Systems, Man, and Cybernetics, SMC-4, (6), sum-of-sinewaves tracking and critical tracking. Nov. 1974.

With the unquickened displays subjects approxi- mated a proportional element or gain.

The Gilson, R. D., Dunn, R. S., and Sun, P., present results therefore suggest that the KT "A Kinesthetic-Tactual Display Concept display be used with command type displays that for Helicopter-Pilot Workload Reduction," permit the tracker to behave in a lag-like Paper No. 77.33-22, American Helicopter manner. Under these conditions the KT servo- Society 33rd Annual Forum, Washington, D.C., motor lag must be carefully designed relative May 1977.

to the anticipated task requirements.

Sun, P., "Rotor Plane Control Device," 6.

In dual task performance both the simulator FY-77 ILIR Final Report, U. S. Army Avionics and laboratory studies suggest that the combina- Research and Development Activity, Fort tion of KT and visual displays may provide Monmouth, New Jersey, 1977.

superior overall performance to the use of only 7.

visual displays. Part of the advantage of the Jagacinski, R. J., Flach, J. M., and Gilson, KT display may be due to a reduction in visual R. D., "A Comparison of Tracking with Visual Additionally, and Kinesthetic-Tactual Displays," Proceed- scanning. the use of a second sensory modality may provide some sensory and/or ings of the First Symposium on Aviation cognitive advantages over a single modality. Psychology, APL-1-81, April 1981, pp. 74-83.

However, these results need to be carefully 8.

tested for their generality beyond particular Jex, H. R., McDonnell, J. D., and Phatak, A. V., "A 'Critical' Tracking Task for laboratory tasks. Full scale simulator studies Manual Control Research," IEEE Transactions and/or inflight testing appear to be warranted in light of the promising nature of the present on Human Factors in Electronics, HFE-7, (4), findings. Dec. 1966.

Acknowledgement 9. Jagacinski, R. J., Miller, D. P., and Gilson, R. D., "A Comparison of Kinesthetic-Tactual and Visual Displays via a Critical Tracking Portions of this research were sponsored in part by the U. S. Army Research and Technology Task," Human Factors, 21, (l), Feb. 1979.

Laboratory, Moffett Field, California, and in part by the Avionics Laboratory, Electronics 10. Burke, M. W., Gilson, R. D., and Jagacinski, Command, Ft. Monmouth, New Jersey, through R. J., "Multi-modal Information Processing NASA-Ames Grant NSG-2179. for Visual Workload Relief," Ergonomics, 23, (lo), Oct. 1980.

11. Jex, H. R., Jewell, W. F., and Allen, R. W., "Development of the Dual-Axis and Cross- Coupled Critical Tasks," Proceedings of the Eighth Annual Conference on Manual Control, May 1972, pp. 529-552.

SYNTHESISOFANINTEGRATEDCOCKl'lTMANACEPEWl'sYSTEM Joseph A. Dasaro Charles T. Elliott Avionics Laboratory U.S. m Avionics Research and Development Activity Fort tinmouth, New Jersey utility helicopter manufactured by Sikorsky Air- Abstract craft. The steps taken in the development of both 'Ihis paper discusses the process used in the the system hardware and system architecture are explained in detail in reference 1. At this point synthesis of an integrated cockpit management sys- tem. Areas covered include flight displays, sub- in the process, a system design has been estab- lished and hardware is being fabricated which in- system management, checklists, and procedures tegrates the following aircraft control and/or (both normel and emergency). The process of evolv- display subsystems: ing from the unintegrated conventional system to the integrated system is examined and a brief a) flight instruments description of the results presented.

b) engine instruments Introduction c) caution/warning/advisory One way to describe the process of designing an integrated cockpit management system is as the d) comrrmnication/navigation/identification and security devices series of steps outlined below: 1) an analysis (functional and electrical) of e) aircraft survivability equipment (ASE) all signals on-board the aircraft, f) electrical system circuit breakers (67) 2) a feasibility analysis of each signalto g) secondary systems such as: cargo hook, determine if suitable for absorption into an in- lights (position, anti-collision), air source tegrated system based on safety of flight require- ments and electronic considerations. switching, environmental control unit, anti-ice (engine, windshield), pitot heaters, blade de-ice, attitude and heading reference system, gyros, 3) an analysis of the functions performed by radar altimeter, engine ignition, tail rotor servo, the operator in a standard aircraft to determine which ones must be performed by the operator, back-up hydraulic pump, and the hydraulic leak which ones the operator must bow the status of, test subsystem.

and which can be performed automatically, and In addition, it became apparent during the de- 4) following these a.nalysEs, the initial sign effort that incorporation of the checklist and mapping of aircraft control and display functions emergency procedures would be an important feature of this system.

from a conventional unintegrated cockpit to an integrated cockpit.

The Hardware Baseline This last step is the first step in an iter- ative process in which the top down system design A standard UH-60A Black Hawk cockpit is shown Figure 2 shows the instrwnent panel is continuously modified as a function of the in figure 1.

in detail, figure 3 shows the lower console, and specific detail uncovered as the process proceeds.

This paper describes the application of this proc- figure 4 the upper console. In addition, circuit breaker panels are located above each operator.

ess in the svnthesis of the intemrated cockpit for the Army JXg?tal Avionic System TADAS). The ob- An initial top down design for a digital avionic cockpit for the Black Hawk was performed by Sperry ,iective of the ADAS effort is to apply in-so-far as possible, the latest advances in digital system Flight Systems and Bell Helicopter (reference 1) :after performing the analysis described '.

technology to a current production conventionslly in steps 1, 2, and 3 in the introduction. As de- designed rotary wing aircraft. The aircraft chosen to demonstrate the application of this technology tails in the areas of the flight displays, Paging and fault tolerance schemes evolved a cockpit is the Arnly UH-60A Black Hawk, a twin engined design emerged (step 4) with an instrument panel A system block diagram isillustrated in as shown in figure 5, a lower console as shown in figure 9.

figure 6, and an upper console ss shown in figure The Flight Displsys 7.

The primary display elements on the instru- One of the major tasks of the ADAS design ment panel consist of four identical 6.8 inch by effort was to synthesize a flight display which 6.8 inch CItT*s each with eight line select would serve as a primary source of both vertical geys on a side. A mmentary @g@.e switch is and horizontal information for the pilot. In ad- located below each display. The main reasons for dition the display had to have the ability to driving all of the primary displays to be exactly display a master caution, warnings, and advisories.

alike stems from both fault tolerance requirements and also the need for line select keys on the The display shown in figure 10 meets the flight displsys for a modest interactive capabil- flight display information requirements for the A cluster of standby instruments is con- ity. ADAS. The display is interactive in that by push- tained in the center of the instrzrment panel for ing certain line select keys and entering numerics safety of flight purposes.

from the KTU the high/low bugs on the radar alti-.

tude scale can be set, baromgtric pressure or The lower console contains a control display field elevation can be set. VOR course selected, comrrmnication, unit for each operator by which all and the navigation node chosen. Some other fea- navigation, and identification equipments are con- tures worth noting are: trolled and their status displayed, an intercom control for each operator, and a keyboard terminal a) VOR radial information is continuously unit (KTU) for each operator. In the center ofthe displsyed (if VOR tuned to a station), lowerconsole for use-by both operators are the stabilator controls, the automatic flight controls, b) wetic bearing to an ADF station is and a miscellaneous control panel.

continuously displayed (if ADF tuned), The upper control contains a number of switch c) ground track angle is continuously dis- functions which were not integrated for various played on the heading scale when the doppler is reasons and a small (10 litits) caution/advisory operating, panel which will be used prior to onboardauxiliary power unit (APU) start. After the APU is started, d) both magnetic bearing to destination from the ADAS system provides these caution/advisories.

current position and course information (either doppler or VOR) are displs+d.

The four CRT's are configured such that either a full screen display or half screen display can In all cases information is presented only if be exhibited. In normal operation the outer CRT's the specific subsystem is operating and the func- are reserved for the full screen flight displsy or tional mode selected.

full screen'wsypoint map. These displays can be called up by either operator or depressing the FLT A full page Way-point Map (figure 11) can be DIS or MAP buttons in the OUTER column of the KTU selected for viewing by either operator on either (see figure 8). CRT (see KTU dedicated buttons, figure 8). A single heading up display on a 1:l M scale with The inner CRT's can be used either for two doppler wsypo?nts shown is depicted. Wsypoint 0~ half screen displays or a full screen displsy.

dering can be selected by depressing the line For interactive paging routines the lower half of All line connections are select key (lower left).

the inner CRT will be the primary display. The erased and a new waypoint order can be entered via two columns of buttons on the tou left of the the KTLJ.

KTU call up the functions which will appear in the lower half (viz CAU - caution. EMGY - Ergine Displays emergency procedures, ASE - airborne survivability equipment, CHK LST - checklist, SEC SYS - second- A full page engine display (figure 12) will be ary systems). The top half of the inner CRT will used during the engine start mode and is available normally be devoted to the ENG IVK)N(engine moni- for call up at anytime by either operator on the tor) half page. The inner CRT can also be used to inner CRT. Analogous to the "yellow" and "red" display a full screen engine page (FULL ENG) or indications used in the standard Black Hawk, re- the wsypoint map (MAP).

verse video and flashing reverse video are used..

Allowable time remainim in an "out of limits" The momentary toggle located below each.CRT condition is displayed as shown in figure 13.

is used for slewing through pages displayed on the lower half of the CRT. A down motion causes slew- During normal flight operation an engine moni- ing through a set of pages (e.g., 1 of 4, 2 of 4, tor half page (figure 14) will be displayed on the etc.). An up motion returns the display to the too half of the inner CRT. RPM and ermine torque branch page one level higher.

ark displayed continuously in both ana?og and digital form (see reference 2). Allowable time those functions which are controlled through ADAS, remaining in an "out of limits" condition is dis- interactive control provided. If a detailed pro- plsyed as shown in figure 15. cedure may be required, access to the detailed procedure is also provided. Figure 16 illustrates The Paging Scheme a caution message where both the operator proce- dure and access to a detailed procedure are pro- The primary areas for interactive functions vided. If more than one caution occurs at the same are the lower half of the inner CRT and the inte- time, the messages are arranged in priority order grated avionics control display unit on the lower and accessed through the paging switch. Depressing console. This latter control display is used for the caution button when the master caution is not all functions associated with the comrmni cation, exhibited will bring up in priority order all out- navigation, identification, and security equip- standing cautions (if any) or the message "NO ments. Five dedicated buttons (plus one spare) CAUTIONS."

on the left side of the KTU are used to access the major branches of the interactive pages. 'Ihe In addition to the master reverse video caution dedicated buttons are: on the flight displsy, there are five reverse video warninp-;;.(ENGlOUT, ENG 2 OUT, LOW R'IR, FIRE,RYD).

1) CAU - caution In the case of a warning situation, a half page warning message is automatically displayed on the 2) CRK LST - check list lower half of the inner CRT. There are twelve (12) warning pages to cover the various warnings and 3) EDGY - emergency procedures aombinations of warnings. An exsmple of a warning message with interactive capability and access to 4) SEC SYS - secondary systems a detailed procedure is shovm in figure 17.

5) ASE - aircraft survivability equipment 2) Checklist A total of 370 interactive pages were genera- Cnecblist automation is accomplished in the ted to map the functions associated with the above ABAS by a series of messages which require yes (or into the ADAS architecture. The process used in OK) or no as operator response. The yes response synthesizing the pages required a detailed know- (the looked for response) can always be given by ledge of each of the subsystems, the current op- depressing the left top line select (lower half erator functions as described in the operator's inner CRT) or when on the ground a cyclic button.

manual, and the level of integration of the TheAGAS checklist is divided into eight (8) electricsl/control function into ADAS. branches as shown in figure 18. At system initia- tion (APU start), ADAS is automatically initialized Rules of operation and the hardware confiw to the first line of the Before Start sequence. As ation were first postulated. 'Ihen as the detailed shown in figure lo, if the checklist requires an pages were synthesized, these rules and the hard- action normally accomplished through ADAS, the spe- ware configuration were refined to provide for cific interactive page on which the action is ac- cases which were exceptions. Each significant complished is automatically displayed on the top The checklist continues to change required a complete page analyses itera- half of the inner CRT.

tion. During this process the system input/output sequence through to the end of the Before Takeoff was modified several times (both by additions and sequence automatically. If the operator desires a deletions). Only after this process was completed detailed procedure (see figure 20), branching to could realistic flow charting for the functional the specific detailed procedure can be accomplished portion of the operational software begin. by depressing the line select opposite the > symbol.

A brief description of some of the features of 3) Emergency procedures the ADAS paging scheme are: Ninety-nine (99) pages of emergency procedures 1) Caution All can be accessed by are contained in the AIMS.

depressing the emergency button and paging to the When the word "CAUTION" appears in reverse The primary purpose for this specific procedure.

video on the flight display (outer CRT), the actual major branch is to allow the operators to gain caution message(s) appear on the bottom of the familiarity with the ADAS version of the emergency inner CRT (in priority order if Tix3re than one). ., procedures. In all cases, these procedures will If the operator then depresses the caution button appear either automatically in the case of warnings In the case of a on the KTU or a button on the cyclic (during or with the caution message.

flight), the reverse video master caution message failure in the caution/warning sensors, this branch on the flight display is extinguished and the does provide a means for accessing the necessary inner CRT displays a complete half page caution emergency procedure if the operators ascertain bhe message. This message will include the emergency fault condition in another manner.

'procedure associated with the caution and, for

I IIIIIII

4) Secondary systems Conclusions Depressing the SEC SYS button on the KTG Synthesis of an integrated cockpit management system requires sn iterative multidisciplined pmc- br&ngs up a menu page which provides access to the ess in which initial conceptual system design is control Ifunctions of eighteen (18) subsystems. An example of a secondary system page is shown in continuously modified by the effects of the speci- fits of the system being addressed. This process figure 21.

does not end with fabrication of the system hard- 5) Aircraft survivability equipment ware but must be continued through carefully de- signed simulation and flight experiments. In Depressing the ASE button,lthe KTU brings up addition, the system must possess the flexibility a menu page which provides access to the surviva- to incorporate changes in procedures (e.g., updated emergency procedures) which will odcur during the bility equipments on board the aircraft.

system life cycle.

Future Plans References Functional verification of the ADAS cockpit is scheduled to take place during the first half 1. Dasaro, J. A., Elliott, C. T., "Integration of of next year in the Tactical Avionics System Simu- Controls and Displays in U.S. Army HeZicopter Cock- Proceedings of 32nd NATO Guidance and Control lator (TASS) at Fort Mornruth. A complete dynamic pits, " system simulation will be performed by tying the Panel Symposium on The Impact of New Guidance and ADAS 1553 data bus to a data bus port on a PDP Control Systems on Military Aircraft Cockpit De- 1145. Support software which simulates all the sign, Stuttgart Bad-mtatt, Germany, May 1981.

multiplex remote terminal units will be used.

2. '%Zectronic Pfaaster Monitor and Display System, After functional verification in the TASS, Human Engineering Sumnary Report," AVRADCOM Tech- the system will be flight tested in the Avionics nical Report 79-0270-3, Aircraft EQuipment'Mvision, Laboratory System Test Bed for Avionics Research General Electric Company, Binghamton, New York, (STAR), a UP-6OA. June 1981.

UR-60A cocknit.

Fig. 1.

Fig. 2.

UH-60A instrument panel.

Fig. 3. UH-60~ lower console.

Fig. 4.

Ub6OA upper (overhead) console.

ADAS instrument panel.

Fig. 5.

i. _..: )I ;;.’ .;.,::. : : ..: !.‘.. .n ., : : : :r: : : ::: : ::::;i j xi : ii::.; : :: : : : : .: : :: :.: i i: ,i i;‘; : : ::: .i I .y ..:: : . ..

ADAS lower console.

fig. 6. :. ; :: : .:: : . . . ..

I - :: ,. ; :...:.:.:I .m. -.;:. .!y;,..‘;..

,:.:$5 /: ‘y;:$..:\ ADAS upper (overhead) console.

IQ?. 7.

OUTER ( AT!

OlSl cl cl ----- .n WI” .*o T.,L .,I mwu I Fig. 9. ADAS system block diagram.

I

bOG- 500-H NAV ?AOiii- DOPPLER::$

I

VOR/ILS

xX(

SE1CRS--- Fig.. 10. ADAS flight display.

SCALE ADAS wayp3int map.

Fig. 11.

1_1 I--I 1_1 IEEl I--I 1_1 [LI I-I I-I I--I I--I 1_1 (_I I--I I--I I--I

1_1

I--I

I--I

Fig. 12. Engine display (Ml Page).

mgine display tithout of limits I Fig* 13* condition .

r( NG 27~42 L--I I--I I_( I--I 1_1 B 1_1 (_I Fig. 14. Enghemmitor (half page). mg. 15. Ek&ne mmitor without of limits condition.

Figure 16. Caution page example.

IcigL7.r~ 17. Warning page example.

Fig. 20. Checklist example 2.

Secondary system page example.

Fig. 21.

THE ROLE OF VOICE TECHNOLOGY IN ADVANCED HELICOPTER COCKPITS Howard P. Harper Senior Human Factors Engineer Sikorsky Aircraft Division, United Technologies Corporation Stratford, Connecticut Abstract Voice Output This paper describes the status of voice Techniques for producing voice output and voice recognition technology in output range from electro-mechanical recorders to digital relation to helicopter cockpit applications. The SSlllDLinfC and storane to more sophisticated this many maturing of technology provides dig&l Storage techni&es such as linear predic- opportunities for new approaches to crew work- Each of these methods has tive coding (LPC).

The paper covers the helicopter load reduction. advantages and . disadvantages which will be operating environment, potential application areas reviewed in the following paragraphs.

and the impact on advanced cockpit design.

The best example of the use of a recorder Introduction based voice output system is the ASH-19 voice warning: svstem used in a number of military the CH-54 Flying Crane-.

Utilizing increasingly more sophisticated and aircrafr &luding complex on-board systems, helicopter crews will Feedback from operational units has indicated be required during the conduct of missions to that the system has functioned well over its perform multiple tasks which include monitoring twenty year life. It does, however, suffer from aircraft systems, monitoring and initiating com- some -of- the reliability problems which one would munications, navigation, target detection, air- expect from a complex electro-mechanical system designed in the 1950’s. One drawback in an to-air attack/coordination, active/passive defense against radar, laser and infared air- and electro-mechanical system is the variation in access time to words due to positioning the ground-based detectors and designators, obstacle detection and avoidance, and monitoring mission- playback head to the location of the next desired specific subsystems. word.

In many cases the crews will be required to The second technique for voice output is perform such tasks in an all-weather, night, the use of digitally sampled and recorded voice environment that demands signals. The method is simply analog to digital nap-of-the-earth out-of-the-cockpit visual attention and hands-on- conversion of the speech signal and usually stick control readiness. As a result of this involves storage in read only memory (ROM).

loading crew work load is The resulting voice quality can be excellent, but increased task In cockpit approaching its maximum limit. depends largely on the sampling rate and encod- A minimum of about 15,000 bits concepts where a single man crew is envisioned, ing precision.

this limit clearly will be exceeded unless a new of storage is typically required per second of technological approach is found. Computer voice speech. The access time to words is extremely This paper interaction is one such approach. fast and as a result messages made up of strings reviews that technology and considers how it of individually recorded words can be put might be applied to solving some of the workload together in a satisfactory manner.

problems.

Linear predictive coding was developed After reviewing the progress in this area it primarily to reduce the data storage requirements is apparent that now is an opportune time to for voice output systems. This is the technique seriously investigate cockpit applications. There used by Texas Instruments in their “Speak ‘N voice are two aspects to voice technology: Spell” teaching system and in a series of chips Both are cur- output and voice recognition. designed to be incorporated into a variety of rently being applied in aviation and elsewhere. other applications. This technique allows stor- In the voice output area applications range from age of one second of speech with about 3,000 toys and home appliances to sophisticated text- The result of the data bits of digital memory.

The uses of voice recog- to-speech processors. compression is some loss in intelligibility when nition systems are not yet as wide spread but digitally sampled system.

compared with a manv aoolications are currently in full opera- Standard vocabularies are available but special For example, they are in use for tion2 Use. vocabularies must be processed by the manufac- assemblv line quality control and in post office turer.

of a voice input mail sorting. - Development typewriter is the subject of major research There are other systems available which are efforts at several companies. even more economical in terms of data storage requirements. These may truly be called speech Current Research synthesizers because there is -no recording and playback of a human voice. The speech is built Voice technology has generated a great deal ;p -of phonemes which are the basic elements of of interest both commercially .and in the govern- speech sound. Using around 40 of these basic ment. Many companies are carrying out research sounds along with the capability to vary pitch, and development activities directed toward mili- intensity and timing, a synthesizer can produce tary applications of both voice input and output understandable speech. The resulting speech tehnology .

All branches of the military as well has a robotic quality, but storage requirements as NASA and the FAA have research programs in this area.

are only about 80 bits per second of speech. There have been several conferences dealing with coordination of this work, the most While initially not as intelligible as the speech recent sponsored by the Naval Air Development produced by other systems, it improves greatly with training and continued exposure. Center in Warminster , Pennsylvania.

The technology does exist now for the use The Navy has, perhaps, the longest history of voice output in the cockpit environment. of military applications of this technology.

They Considering this it would now be difficult to have demonstrated its usefulness in performing justify continuing the use of tone combinations cockpit switching functions and in the more as the primary auditory warning system. Defin- complex man-machine interactions of an airborne ing a system will require a choice among the anti-submarine warfare system.

NADC currently voice output has a study under way to understand and define technologies described above.

Reliability considerations will probably rule out the problems of the Navy aircraft cockpit opera- the electro-mechanical recorder. If voice quality ting environment. This includes the effects of jet aircraft cockpit noise and the effects of G is the primary criterion the pure digital sample and store system will likely be judged best. If, loading on the physiology of speech. This study however, a large vocabulary is required one of relates primarily to the fixed-wing environment.

the data compression techniques may be neces- sary . If a virtually unlimited vocabulary is The Air Force is currently sponsoring a required, as might be the case if the system study directed toward flying a prototype voice were called on to output the emergency proced- interactive system in the F-16.

This program is ures now found in the flight manual, then a being conducted jointly by Lear Siegler Inc. and phoneme based system is the only practical General Dynamics and is expected to fly this choice. year. In the development program, progress has been made toward accommodating the unit to the jet aircraft cockpit environment.

Voice Recognition This has included dealing with problems such as the Computer recognition of a speech input is a effects of the oxygen mask on speech recognition.

much more challenging problem than the produc- tion of a voice output. A variety of techniques Helicopter Research have been used. The specific method depends on a number of variables: There are may differences in the mission the size of the vocabulary, the necessary level of recognition and the operating environment of helicopters and accuracy, the number of users, the need for fixed-wing aircraft that will have an important isolated word or continuous recognition, the time effect on the usefulness of voice interactive available for training of the system, and the technology in the cockpit. First the missions environment in which the recognizer must be are markedly different. The helicopter night, operated. nap-of-the-earth, all-weather scenario imposes long duration, high workload conditions on the crew. Attention A typical isolated word recognizer works by must be fixed outside the having the user say all of the words in the cockpit and for long periods hands cannot be vocabulary one or more times to train the system. taken off of primary flight controls. These During this process the voice signal is analyzed conditions are often sustained for the major by a bank of filters which measure the amount of portion of the mission. On the other hand, energy in a number of frequency bands. Each fixed-wing aircraft have periods during the word is broken down into a number of equal mission where workload is very high but these temporal parts and the filter bank outputs for are generally of a much shorter duration.

each are stored. This creates a template against which incoming words are tested. The computer Another factor differentiating the helicopter finds the best match for the incoming word and from the fixed-wing aircraft is the crew station environment. There are primarily two character- carries out the appropriate action assigned to that word. There are at least ten recognizers istics which contribute to this difference.

The on the commercial market today. Each claims 99% first is cockpit noise. Figure 1 shows typical plus recognition accuracy and it probably is true spectra for the two aircraft types. This clearly that under some specific set of conditions that shows the difference in frequency content.

claim can be met. It is unlikely, however, that Much more energy occurs in the speech frequen- cies in the helicopter.

any of them will approach that accuracy in a The second aspect is the military helicopter cockpit. modulation of the voice due to cockpit vibration.

This effect is shown in Figure 2 by noting the l- HELICOPTER 110 / i-

/‘.’

d’

JET FIGHTER ‘\ ‘.

-.

-5~ -.

-4 -.

/a-------- SPEECH FREQUENCIES -4 1 2doo t .-. .{ FREQUENCY - HZ Figure 1. Cockpit Noise Spectral Differences for a Typical Military Helicopter, and a Jet Fighter B I ii ?

i UNMODULATED FREQUENCY - HZ Figure 2. Modulation of a Speech Sound Due to Whole Body Vibration difference between the voice spectra under to specify the requirements for a voice inter- conditions of vibration and no vibration. As active system for a rotary wing aircraft. This might be expected, a system trained under one process is needed to determine whether the time and expense of doing a complete and detailed set of noise and vibration conditions and asked to recognize under other conditions may perform systems and human engineering analysis is unreliably. warranted.

Helicopter cockpit-related voice technology First it is necessary to list the assumptions research is currently going on at NASA Ames on which the system design will be based: 1) Research Center and at the II. S. Army Avionics the availability of a speech recognizer with 100 Research Facility at Ft. Monmouth, N. J. The word vocabulary with the capability of training Ames facility has a long record of voice related by two users and having a demonstrated accu- research work. References 1 and 2 are Ames- racy of 95 to 99.9 percent under all flight sponsored studies relating to the cockpit use of conditions ; 2) a voice output device with a synthetic voice warning concepts. More recent demonstrated intelligibility at least as good as work has addressed problems of the voice recog- current inter-communications systems.

nizer in the helicopter’s noise and vibration environment. The results are extremely encour- Ideally this preliminary design effort would Even under the most adverse conditions, take place after the completion of a detailed aging.

voice data entry compares favorably with key- analytical study of all the man-machine inter- board entry. Accuracy differences never actions. The results would allow evaluation of exceeded 2%. These results clearly establish the the workload reduction quantitatively and allow feasibility of using voice recognition in rotary- the desginer to investigate the effects of design wing aircraft. In a second study currently variables on the performance and usability of the underway, a commercial voice recognizer is being system. The time to do such an analysis is used to control an aircraft performance computer. before starting a design effort for a specific This study is beginning to contribute information application. In the heat of a design effort the on the problems of using this equipment to system designers cannot wait for the results of perform a real function in a present day helicop- such an effort.

ter. The Army at Ft. Monmouth has taken the lead in military helicopter voice-related research. One of the design concepts planned is the At present they are studying the noise environ- use of the “intelligent copilot” model. All can- .

ment of their inventory of helicopters to define didate voice interactive functions are evaluated the effects on the performance of currently in terms of whether they are consistent with the available recognizers . Their plans call for behavior of a hypothetical copilot who knows implementation of a voice interactive system to when to talk, when to listen and who prioritizes become part of the advanced digital avionics information in a logical way that is appropriate system to be flown on a UH-6OA. to the mission phase. A second design concept is that the system will provide feedback on all Pragmatically we have to recognize that a inputs and will require secondary verification of military helicopter is far from the ideal location the more critical items. If, for example, the for a voice recognizer but, because of the work pilot were to say “Jettison Tank” the system done at Ames and elsewhere, we can be reason- might respond visually or orally: “Tank Jettison ably certain that the problems can be solved. Requested” and the pilot would be required to Therefore it should become our purpose to let confirm the request by giving an action command.

the manufacturers of this equipment know that Thirdly, all voice inputs are backed up with a we are interested, that we can see many poten- manual entry mode which would be considered a tial applications, and that there is a market in secondary operational mode and, therefore, might the helicopter industry. Furthermore we should require a deeper level of paging. The fourth define the operating environment so that they concept is the use of a switch on the pilot and can do the necessary development to make equip- copilot cyclic grip which he will press to indicate ment that will function adequately in our cock- that he is talking to the recognizer. Lastly the pits. Alternatively they may tell us what has to training of the recognizer will not be done on be done to our environment to make the equip- the aircraft, it will have been done earlier and ment work. We will then have to address the stored on a cassette or in a ROM cartridge which problem of whether the value of a voice inter- can be plugged into the aircraft for a rapid data active system warrants the cost of an improved transfer.

cockpit environment. This will provide parallel pathways for the solution of the operating envi- The voice output must be unusual enough ronment problem and development of applications to be easily distinguished from other crewmen or which make maximum use of the technology to air traffic controllers. This is not meant to imply that a robotic voice is required, however reduce cockpit work.

the voice must stand out clearly from the routine Helicopter Applications voice communication traffic. The major difficulty with robotic quality voice is that people have The following are some of the thoughts troubIe taking it seriously and this effects its which must go into the preliminary design effort acceptability to pilots.

system (AFCS) there are a number of functions Next we will look at each of the various which can be considered for integration with a systems on the aircraft and try to understand voice interactive system. These include system where voice input and output technology might fit into operation of that system. turn-on, function selection, monitoring of per- formance, response to problems, and system Communication shut-down. The AFCS initiate and shut-down functions are best reserved for manual action In the area of communications we will in- since they generally occur before and after the clude radios for air-to-ground, air-to-air, and crew workload is at its highest. The selection data links; and systems for communication within of AFCS functions is a good candidate for voice the aircraft. The functions which must be actuation. Here such functions as airspeed performed with this equipment include tuning, hold, altitude hold, heading hold, or approach to selection of the system, and volume/ keyiw, hover might be selected through inputs to the squelch control. Tuning is a function which is voice recognizer . This is one case where a very narticularlv adaotable to a voice recognition positive feedback system would be required. A ‘system. The pilot might say “Tune VHF 322.7” secondarv command would be reauired nrior ro or “Tune VHF Channel 5”. Selection also fits in the initiation of any of these functio&. The well with a recognizer system. The pilot would pilot would say “Hold Heading” and the system say “Select UHF” and subsequent transmissions would respond “Heading Hold Requested”. The would be made on the UHF radio. The use of pilot, after seeing that the system understood voice to control volume, squelch, or keying does his input, would give an action command such as not seem to be practical because the voice com- “Do It”. Had the feedback been incorrect the mand would interfere with the material being pilot would cancel the input and try again verb- sent. On the voice output side, it seems poss- Voice ally or, at his option, engage it manually.

ible that voice synthesis may be used to recon- output could be used effectively to provide the struct messages encoded digitally and sent to the pilot with information on the status of the sys- aircraft from the ground via a data link.

tem .

Navigation Subsystems Control and operation of navigation equip- The engine, fuel, APU, hydraulics, elec- ment offer opportunities where both voice input and transmission subsystems trical, anti-ice and voice output would be very effective in might make use of voice. The possible crew workload reduction. The systems which might functions would include system start, condition be controlled are the doppler/inertial navigation monitoring, system control, malfunction response The func- system, Tacan, VOR/DME and ADF.

and system shutdown. A specific engine para- tions of this equipment are to provide: current meter which is a very possible candidate for position, steering information in X, Y, and Z voice monitoring is power available. Information coordinates, the map situation in terms of the about power margin has a high priority at times relationship of current position to other geo- when the pilot’s attention is outside the aircraft graphical information, system updates and accep- The pilot and both hands are on the controls.

tance of flight planning inputs such as way point might say “Power” and the system would respond locations. These functions for the most part, with a voice message “10% Torque Remaining”.

are adaptable to voice interactive techniques.

Contingency power selection is a mode which For example, current position might be called up allows pulling additional power from one engine with the voice input “Position”. The system This when the other experiences a power loss.

might respond in map coordinates or in terms of selection must be set up quickly at a time when bearing and distance to a known point. Steering the pilot would be very reluctant to remove information could be requested and provided In the fuel either hand from the controls.

verbally. For example the pilot might ask for system there are a number of possibilities.

“Directions Waypoint 3” and the system would Voice requests could be made for fuel status respond “325 Degrees, 2 Miles”. Map situational with the system responding in pounds of fuel information could be of the following types: remaining or in terms of flight time remaining at request for nearest fuel or request for height the current flight condition. In addition to the and location of highest terrain in the area.

low fuel warning normally provided, a program- Navigational system- updates could easily be mable voice system could be used to provide a accomplished verbally; the pilot saying “Update warning at any fuel state or time remaining Waypoint 3. . .Mark” when directly over the selected by the pilot. The APU could be started DOiIlt. In addition the flight could be planned and shut down by voice command but since this using a verbally prompted-waypoint entry rou- is generally a ground function where workload is tine.

not critical it would not be worth implementing in Aircraft lighting is an area the voice system.

Flight Controls where a recognizer could be particularly effec- tive . Lighting controls are numerous and fre- The primary flight control system would not system could quently accessed. The voice be directly interfaced with the voice recognizer, select, actuate and control both interior and but system faults would trigger appropriate exterior lighting systems. In addition the voice verbal messages. In the automatic flight control recognizer could be used to select various sub- __._ .- .- -.~ system status monitor modes such as engine Cockpit Impact instruments, electrical or hydraulic parameters, or emergency procedures as suggested in Ref- Table I summarizes the possible voice erence 3. Voice interaction with the remaining system applications discussed in the last para- subsystems would be limited to voice messages graphs. This is an exercise to identify what related to malfunctions.

could be done. It is important to emphasize that the next logical step would be a thorough analy- sis of the functions reuqired by the mission to Caution, Warning, Advisory determine a reasonable design solution.

The information provided to the crew by the caution, warning and advisory systems is The single place cockpit is the application where the need for an “intelligent copilot” is potentially convertible to a voice output system.

greatest. The recognizer/synthesizer will be Those messages which are currently supple- mented with an alerting tone pattern should be required to take over many of the functions renlaced with a voice message. With voice normally assigned to the second cockpit crewman.

A single place helicopter cockpit which includes technology available pilots should not have to a voice interactive system is shown in Figure 3.

identifv a failure by the pattern of tones in the alerting signal. It seems apparent that voice The physical impact of the voice system is not might become the primary alerting system for all dramatic. The only special control is the switch of the warning messages and for the more critical on the cyclic to key the recognizer. In addi- tion , of the caution messages. This would allow the ROM cartridge with the pilot’s voice characteristics is replacement of the current matrix of dedicated inserted in a slot. The caution lights with a three or four line priori- remainder of the displays and controls will only This type of alerting system will differ slightly from a non-voice cockpit since tized display.

require some new thought because of the single manual and visual backups will probably be dimensional quality of the auditory channel. provided for the voice functions.

Two messages cannot be presented simultan- eously ; all inputs are sequential rather than The major improvements will be in the All possible messages must have a pilot’s ability to keep his hands on the controls parallel.

during critical flight phases, and in his capa- priority value which determines the order of their presentation, To complicate matters fur- bility for being fully informed on aircraft system ther these priorities may have to change with status without bringing his eyes inside the mission and phase within the mission. cockpit. The concept that the recognition system will respond to simple commands will eliminate the component workload associated with Two recent studies (References 4 and 5) have presented conflicting data on the value of finding and actuating a manual control. The use using voice warning to supplement the visual of voice actuation facilitates the use of multi- Reference 4 found no important function manual controls and thus reduces cock- alerting system.

difference in the time required to respond var- pit space requirements to some extent.

ious combinations of voice, tones and visual signals in a jet transport simulator. The author It should be further emphasized that voice explains that this is because the pilots always cannot be successfully introfuced to cockpits on a piecemeal basis.

checked the voice message against the visual We are beginning to see caution panel before responding. The study various individual systems such as ground reported in Reference 5 investigated the pilot proximity warning systems and altimeters with reaction times from the presentation of a voice or voice output capability. This is manageable light warning while flying nap-of-the-earth in a now, but further poliferiation of voice systems In this case there was a dramatic helicopter . could become chaotic. The full benefits will only in response time with the voice improvement be achieved by an integrated approach.

It was found in this study that the system.

pilots were willing to respond without confirming The design of a voice interactive cockpit the malfunction on the caution panel because it system requires an appreciation of the single took approximately 3 seconds to stabilize the channel nature of the auditory system. With flight path of the helicopter sufficiently to look visual displays the designer can put up a great This is further indication that the inside. deal of information at one time in the hope that helicopter and fixed-wing aircraft may require the pilot can pick out what he needs for a significantly different approaches to integration particular task. With a voice system, sequenc- of cockpit voice technology. ing and prioritizing of inputs and outputs is necessary since only one thing can be going on There are several possible uses for voice at any time.

input to the alerting system. One would be to acknowledge messages instead of pressing the master caution capsule to indicate recognition of Another function might be to the message.

change caution priorities. If, for example, a particular system was operating marginally the pilot might want to raise its caution priority to the top of the list.

References Conclusions 1. Simpson, C. A. and Williams, D. H., The following conclusions can be drawn “Human Factors Research Problems in Electronic from this investigation of helicopter cockpit voice Voice Warning System Design”, NASA TM X interactive technology : 62,464, 1975.

1) Voice output technology is available for 2. Simpson, C. A., “Synthesized Approach use now.

Callouts for Air Transport Operations”, NASA Contractor Report 3300, 1980.

2) Research results look very favorable for the development of an accurate, reliable voice 3. McGee, J. and Harper, H. P., “Advanced recognition system for helicopters.

Subsystem Status Monitor, Sikorsky Aircraft”, USAAVRADCOM Technical Report 80-D-5, U. S .

There are many possible voice inter- 3) ,Army Applied Technology Laboratory, Ft. Eustis, action applications which will result in workload Va., 1980.

reduction.

4. Wheale, J., “The Speed of Response to 4) A thorough systems and function analy- Synthesized Voice Messages”, AGARD-CP-311, sis is required to maximize benefits and to be 1981.

sure that the system is acceptable to crewmen.

5. Reineck, M., “Voice Warning Systems : Some Experimental Evidence Concerning Applica- tions”, AGARD-CP-311, 1981.

TABLE I.

POSSIBLE COCKPIT VOICE APPLICATIONS - SYSTEM m OUTPUT COMMUNICATION TUNING DIGITAL MESSAGE RADIO SELECTION RECONSTRUCTION NAVIGATION POSITION REQUEST POSITION REPORT STEERING REQUEST STEERING INFORMATION MAP INFORMATION REQUEST MAP INFORMATION POSITION UPDATE I FLIGHT FUNCTION SELECTION FEEDBACK I CONTROLS ACTION COMMAND , SUBSYSTEMS POWERINFORMATION REQRS.

POWER INFORMATION SELECTlCONTINGENCY POWER REQUEST FUEL STATUS FUEL STATUS LIGHTING CONTROL SELECT DISPLAY MODE C :AUTION, PRIORITY SELECTION PRIMARY ALERTING SYSTEM k IARNING, A iDVISORY - h Figure 3. Single-Place Helicopter Cockpit Incorporating a Voice Interactive System COCKPIT INTEGRATION FROM A PILOT'S POINT OF VIEW David L. Green Vice President PACER Systems, Inc.

Arlington, Virginia Abstract Some of the reasons for this degraded capa- bility are readily apparent when IMC operations are studied. Navigation must be conducted via reference to electronic aids and this means a new level of air traffic control is required, with a Extensive experience in both operational and concomitant increase in the time consumed by engineering test flight is used to suggest straightforward changes to helicopter cockpit and radio communications. These procedural changes increase cockpit workload and reduce the time control system design that would improve pilot and instrument flight available for flight control. With less time performance in marginal conditions. Basic differences from airplane fight available to allocate to flight control and characteristics justify distinct treatment of dramatically degraded visual cues, the pilot flies with less vigor. Everything happens a bit helicopter cockpit flight control configurations.

Helicopter use of collective for direct lift slower, while the pilot tries to fly with greater control and collective to coupling are precision.

yaw emphasized in drawing these distinctions. Need for good downward peripheral visibility and truly One obvious solution to the IMC case is to horizontal glare shield profile are cited for incorporate sensor-display concepts which return natural visual cues during marginal VMC and the real world visual cues to the cockpit (FLIR).

approach transition. Needed control system Under certain circumstances today's technology 1) separation of yaw from makes this type of visual augmentation possible, improvements include: but even the best of these concepts still have cyclic force trim; 2) pedal force .proportional to serious displacement rate; and 3) integration of engine shortcomings in truly bad weather. In stick. Needed display event, this type of visual augmentation is controls in collective any improvements include: 1) natural cuing of yaw considered heroic for many military applications rate in attitude indicator; 2) collective position and all civil applications.

indication and radar altimeter placed within primary scan; and 3) omnidirectional display of From a pilot's perspective, this inability full range airspeed data.

to fully exploit the unique capabilities of the helicopter is a problem which is common to Introduction helicopters of all manufacturers. That is, there common man-machine interface are a number of The helicopter has one very unique capabi- characteristics, which as they stand, detract the ability to hover efficiently and pre- from the pilot's ability to accomplish the pilot- lity 9 cisely for extended periods. And when compared ing task. And although pilots may desire change, to the airplane, it has the advantage of being they aren't always able to articulate a winning and land able to fly into coneined areas, hover, argument for the features they feel they need.

vertically without any concern for the stall-spin They may not even understand the source problems phenomena. they are experiencing. So without such convinc- argument, many worthwhile improvements go ing The trained pilot has no problem exploiting unidentified or deferred.

the capabilities of a helicopter during VMC, but when the task is proposed in the MC environment, What follows then are observations, explana- the pilot often appears to fall short. This same tions, and suggested requirements for change during marginal trend in performance also exists whiCh do not require heroic efforts. These VMC and during transition from IMC to VMC for the author's comments are principally based upon landing. That is, the pilot-machine combination personal experience and observations as an opera- is less capable when external visual cuing is tional helicopter pilot, an engineering test marginal or non-existent.

pilot, a research pilot, an experimental test pilot, and a flight test engineer. The intent is Approaching a hover spot the pilot must to provide insight into the factors which may be flare to stop. When he flares the view over the nose is often inadequate. When it is, or when confusing to non-pilots and revisit a number of helicopter cockpit design features which have there is an obstruction in the over-run, the suffered at the hands of the "accepted conven- pilot will often approach with a crab angle. A t ion" . sideward flare will be used, or the helicopter will be stopped short and air-taxied so that the The scope of this paper will not allow an pilot can see the spot out the right side. When in depth treatment of all applicable characteris- this type of approach is flown in U.S. heli- tics which are candidates for change. Nor is it copters, left pedal is required to sideslip to possible to consider all phases of flight, or the right. More left pedal means that more tail helicopter applications. Instead, this effort is rotor power is required. If the pilot were to generally focused on the high workload or high sit on the left side he would hold right pedal stress situations where pilots are routinely and less power would be required to maintain unable to accomplish the transition to hover, or hover altitude. Seems like the U.S. helicopter to conduct other slow speed tasks safely. The pilot is on the wrong side or the U.S. main rotor purpose of -the paper is to persuade the reader is turning the wrong way.

that there are many reasons to revisit cockpit design and question the conventional wisdom which Collective has been handed down for generations. The pre- mise is that given a bit more design considera- The importance. of the collective and its tion, a helicopter pilot can generally achieve control characteristics are substantially under- more than is currently expected of him. He appreciated by the helicopter community. This cannot achieve more, he can do it more only -device should be recognized as a direct lift safely. control, that permits precise and quick control of the vertical degree of freedom. The stored angular momentum permits small inputs to be The Cockpit accomplished without the need to trade airspeed for altitude and without concern for the But. before we deal with the tough questions engine(s') ability to accelerate or decelerate.

related to IMC flight, I would like to first Figure 1 further illustrates how the collective conduct a walk through of the basic cockpit to can be used to climb even while the nose is pilot interfaces--the controls, external visi- pushed over to accelerate and allow the pilot to bility, the seats, and some of the things that keep the trees in visual contact.

differentiate helicopter control from airplane control.

Seat Assignment -.

When a pilot gets into most helicopters (there is always an exception) he flies from the right side. This is true even though there is really no rational reason for such a choice. In fact, when you consider the need to work with airplanes in left-hand traffic patterns, it makes little sense at all. For when the helicopter is in a left bank, the pilot in the right seat generally can't see where he is going during the turn. The pilot's line of sight is blocked by the overhead of the cockpit cabin which normally supports circuit breakers, switches, and engine controls.

The real reason helicopter pilot's are in the right seat has nothing to do with any great engineering logic. It just happened to come out that way. Mr. Sikorsky meant for the pilot to be in the left seat, but because of early vehicle training problems, the first operational pilots learned to fly in the right seat. The point here is that there is nothing sacred about the pilot Figure 1. Comparison of Airplane and Helicopter being in the right seat. But is there any reason Pitch Attitude Characteristics During to consider changes? There may be..

Climb Over Obstruction.

When compared to the airplane task, the Linear Force Cues collective but it simplifies height control; makes horizontal speed control comparatively more To further understand helicopter control difficult. For example, when an airplane pilot techniques and the need for enhanced visual cues, it can be useful to consider the nature of the climbs back to glideslope, power is added and.the nose is raised to increase angle-of-attack. In linear accelerations which are felt by the pilot (airplane vs helicopter) during a level decelera- contrast, the helicopter pilot simply increases tion. Consider the case where the throttle of an collective to climb.. To change speed at constant airplane is rapidly retarded. The linear force altitude, the airplane pilot reduces simply along the X axis causes the pilot to move thrust and the aircraft decelerates. Altitude is maintained via the elevator. The helicopter forward, and is restrained by a seat belt and pilot pitches nose up and commands deceleration shoulder harness. In contrast, the pilot of a while maintaining height via the helicopter decelerates by lowering the collective with attitude collective. This last technique requires more as he pulls the nose up. If the pilot doesn't control coordination. lean forward (so as to see out or keep his vertical orientation), the result can be no Some analysts tend to believe that airplanes forces or an increase in the forces on the and helicopters are controlled in the same basic pilot's hack as gravity pulls him against the way. Engineers with this viewpoint will read the seat.

two sets of descriptions above, and conclude that Seats-Controls the airplane techniques and the helicopter tech- niques in forward flight are essentially the Engineers underestimate the need for ad- sane. And this is where many of the helicopter pilot's problems begin. justable seats and pedals. The pilot must be able to comfortably locate himself around the On a more positive note, there has been one controls. This includes the pedals which need to noteworthy innovation in physical design of the be adjustable as well. A pilot who is uncom- collective control. This new design was first fortable or must sit on an angle, is probably installed on the Bell-222 and later on the Bell more susceptible to spatial disorientation.

214ST. The collective grip tends to move aft and upward as collective pitch is increased. I had Looking a little deeper, we find many pilots no problem with this motion. The hand grip and fly with their right forearm resting on their arm motion were comfortable. But more control with right leg, manipulating the cyclic very important, this design permits the installation their fingers. This is a method which is The I fly of two engine controls on the collective. particularly appropriate for IMC flight.

left side of the split grip is for the No. 1 this way and often feel like I have to adjust the just to engine and the right side is for No. 2. With seat too high relative to the pedals, obtain a satisfactory grip on the cyclic. I this design, one can readily advance or retard an engine in an emergency without releasing the don't really see anything which one might do to collective. An admirable solution to a difficult improve pedal positioning but a cyclic which problem (see Figure 2). could be adjusted in height an inch or two might enhance many pilots' abilities to fit into their machine. Again it's important to be comfortable to avoid disorientation during high stress or high workload situations.

I I

Force Trim Systeuis No.1 Eng.-, rNo.2Eng.

I I There should be a cyclic force trim system

in all IMC capable helicopters. This

system should always incorporate an instantaneous Force Trim Release (FTR) switch, even if the system uses a Four Way Trim Switch (FWTS Coolie Hat) to trim fore and aft, and laterally (see Figure 3).

A simple force trim system is required to hold the cyclic control where the pilot puts it.

And when such a spring system is added, the' designer should not become confused as to its Figure 2. Characterization of the Collective purpose. In a helicopter the Force Trim system Control Incorporated in the Bell-222 holds the control at some pre-selected point.

and Bell-214ST Nelicopters.

That is, the longitudinal and lateral-directional For example, the need to see down through the feet seems to be one of the least appreciated needs for visibility. Yet the pilot receives much visual data through peripheral vision when he can see the ground down through or near the

FTR\

feet. In slow speed flight or hover, horizontal motion is best controlled via this cue source.

It is even possible to receive a beneficial cue of pitch rate through this window when in a.hover or even at altitude when operating without a horizon.

Some helicopters have little or no downward visibility, and experience has shown that they are clearly more difficult to land with equal precision.

And flares from steep approaches are Figure 3. Typical Cyclic Control. much more readily accomplished when the downward visual path is available. When forward visi- bility is poor, as it is during heavy haze, and static stability of a helicopter are generally so at the bottom of an IMC approach, the pilot may weak that the force characteristics which can be actually acquire initial visual contact through developed (via a simple force feel system) do not the lower panel. This can happen even when he is handling quali- substantially enhance helicopter heads up looking for the landing area. And when ties. But when augmentation is incorporated, it you depart vertically out of a confined area, does become very important for the control to there is no substitute for this downward stay precisely where the pilot puts it.

visibility.

In contrast, putting a force feel spring in the yaw control is totally counterproductive.

Friction is more than adequate to hold the controls in place. And if you put a spring in FTR switch the pilot is forever pushing the yaw, If the so that the pedals can be repositioned.

FTR releases cyclic trim at the same time the yaw control is released, you have defeated the reason for the cyclic force gradient. The best design, from a pilot's point of view, incorporates fric- tion to hold the pedals in place, with the possible addition of a hysteresis damper that to the provides an opposing force, proportional rate of application. For VMC type maneuvers, the That pedal rate damper is even more appropriate.

the best feel for the the pilot obtains is, maneuver he is conducting if he feels a control force which is proportional to the rate a given control is deflected.

Typical Sources of Pilots Primary This type of control rate damper can also be Figure 4.

Visual Cues.

incorporated in the cyclic control with advan- tage. This is true hecause many pilots depress centering forces the FTR to release the stick Another problem occurs during attempts to rate feedback during rapid maneuvering. If the conduct steep approaches. This is illustrated in when the FTR is depressed, Figure 5. Here the pilot visually acquires the forces remain, even landing pad with his eye on a 20 degree approach the pilot's reaction is very positive.

angle. This is his limit of downward vision over Visibility the nose so that when he pitches up to de- celerate, he loses visual contact with the pad.

There is a great deal of variation between So he doesn't pitch up first. First he lowers designs when it comes to cockpit visibility. The the collective and flies down. As he descends he tmportance of external visibility is hard to becomes able to pitch up for the deceleration Yet, it is an aspect of design which while still keeping the overstate. target in sight. This seems to receive insufficient weight when cock- may explain how tail rotors get involved in trees pits are configured (see Figure 4). and fences on final approaches to confined areas.

Figure 5. Flight Path Flown when Pilot Attempts to Keep Landing Site in View After Having Started a Steep Approach.

Horizon Reference When the sun goes down, and the horizon reference weakens, the pilot trades the outside visual cues for the information available from the cockpit displays, aircraft sounds, cockpit control positions (and forces) and the force cues (vertical accelerations, etc.) Impressed upon his person. In many cases there is a period of transition where the pilot is flying via primary reference to his instruments even through some of the outside cues are still there. During such periods the pilot can experience an unexplained uneasiness, and for some reason there is a problem keeping the ball centered on the inclino- meter.

I believe this situation also develops dur- ing transitions from IMC to VMC on final approach and during certain other slow-speed hover tasks.

This uneasiness is also related to the aircraft where a pilot has a much different feel when flying from the left.

right seat as compared to A probable explanation is illustrated in Figure 6.

When the glass shield is curved, or it is sloped down to the outside, the pilot is pre- sented a very strong erroneous attitude reference. And under certain circumstances, I believe there is an unconscious tendency to match the horizon and the glare shield line. This causes the ball to he out to the left when the pilot flies and out to the right when the This glare shield line needs to co-pilot flies.

he truly level. When the aircraft visual reference is level, a weak horizon line, can be a powerful cue even when no conscious positive reference is made to it.

During night hover operation in the SH-3, it was not uncommon to work with no horizon. (Even if there is a horizon you still fly the machine on instruments). But there would be nights when Figure 6. Impact of Glare Shield Design on Crew just a faint hint of a horizon line was avail- Visualization of the Wings Level Attitude.

ahle. One never looks at it, but somehow you would know it was there. But the running lights Heading hold is not required to compensate for (navigation lights) would backscatter light into poor static directional stability; nor is con- the cockpit from the mist over the ocean and this trol quality or control power of the directional When would often mask the faintest of horizons. control system at fault. The problem clearly I was tired and I was uncomfortable, I would turn stems from the fact that it is difficult to find the running lights out (leaving the tail light the new directional control trim point when an on) while I hovered. I could then faintly make input is required to compensate for the out the horizon reference and it made all the collective-to-yaw couple.

difference.

The desired yaw control position, which the The IMC Problem pilot cannot easily locate, is the position which will yield a zero yaw rate.

He does fine when In a historical sense, the design of today's visual cues are available, but during IMC he has conventional IMC cockpit, was derived from a trouble because the yaw rate cues available in marriage of airplane instrumentation to the cock- the cockpit are totally inadequate. To under- which were originally pits of helicopters stand why, lets review the fundamentals and designed for visual flight only. And when heli- actual experience.

copter pilots were unable to accomplish an IMC on instruments, When the collective of a U.S. helicopter is hover, or an approach to hover increased during hovering flight, the pilot must vehicle stability took a large share of the blame. Automatic Flight Control Systems (AFCS) move the left pedal forward to compensate for an were subsequently incorporated to solve the pro- increase in main rotor torque. Right pedal is blem. The result was a dramatic .improvement in required under similar circumstances in a French man-machine performance, but the man had a new helicopter where the main rotor turns in a role. The pilot was now a manager, no longer in direction which is opposite to that of the U.S.

machine. One might expect a pilot to have the direct control of the helicopter. The pilot He was able to fine tune trouble switching from the U.S. to the European became a safety pilot.

the AFCS while it operated normally, while also convention. But generally there are no problems being there to recognize failures so as to at all when the yaw rate cues are sufficiently strong. But some piloting errors do occur when extract the aircraft from an approach or hover the strength of the heading-rate cue decreases.

task if safe limits were exceeded.

the Flight Director Indicator In reviewing my own experience, I can report Next came (FDI) which in many cases could double as an Auto that I have had no problem associated with take- off or hovering flight; but at high altitude or Pilot Computer (APC). The FDI brought the pilot while operating in heavy haze, I have found my back into the direct control of the aircraft, but The pilot was in- left foot moving forward with up collective.

this time he was a servo.

to follow commands on an ADI, matching That is, when the visual cues were powerful, I structed Keeping had no problem. But when the cues were weak, my pointers to their respective indices.

all the pointers in their proper place would keep learned response (which was nurtured for 22 years the aircraft on glide slope or in a hover. The in U.S. helicopters) took over, even in the European machine.

two big advantages of this display format were that the pilot didn't have to think much, and all centrally located on a single This experience illustrates the importance commands were of yaw rate cues. Although I had no problem display. I might add that the basic flying adding right pedal with up collective during my qualities of the helicopter were improved so that Auto Pilots could control the outer loop. When first takeoff, I experienced confusion at alti- this happened, it became possible for the pilot tude where the yaw rate cues were not lost, but But even when to fly almost as well with an FDI. distant and subdued. I didn't even have to enter trouble with directional pilots are allowed to fly with reference to an IMC to start having control coordination.

FDI, they are typically required to operate above unless features like heading hold say 60 knots, are incorporated. Obviously static directional stahility was not at fault. This parameter is obviously of Collective to Yaw Couple greater aid during forward flight than in the hover where I had no problem at all. One can now The conclude that the static directional stability This last point is very important.

and the yaw control system are adequate all the heading hold feature is required to accomplish an So what is missing?

way to a hover.

approach to hover (IMC) because the collective to of the single rotor helicopter is so yaw couple For the answer, compare the function of the powerful and interactive that it dramatically display which is provided for pitch and roll, to increases pilot workload when it goes unchecked.

the function of the display(s) provided for yaw control (or heading). The AD1 is a fine analog A nominal one-for-one match.

of the real world.

But look what has been provided for yaw. The most obvious instrument is the RMI (or HSI). The indicator which cue is a dial that rotates in an YOU would have to look is mounted below the ADI.

down through a hole in the floor of the heli- real world motion which would copter to see The RMI and HSI are relate to this display.

navigation and heading management indicators, not yaw rate displays.

needle of the Then there is the vertical Today this indicator is turn-and-bank inchlcator.

so small and underdamped that it is virtually useless during IMC hover or approach flight.

Electronic HSI's generally provide an en- hanced heading cue, but again, the cue is dis- and roll placed from the primary cues of pitch found on the ADI.

Figure 7. Pitch Roll Yaw Attitude Indicator (Characterized from ADI-811).

In the most modern military helicopters, we find the Electronic Vert‘ical Situation Display A heading reference strip is normally (EVSD). The yaw rate needle of the turn and bank This is presented across the top of the display. indicator may still he in this compressed scan, a step in the right direction, but it is clearly but this indicator does not readily transfer the not conventional equipment. In any event, it is It does not exhibit any characteristics message.

civil machines and not currently offered for analogous to yaw rate and therefore it must be generally beyond the scope of this paper. interpreted. Experience would suggest that con- trol logic in the mind gives priority to control Now revisit the piloting task for a moment. of the most life threatening parameter(s) and During hover and approach to hover, speed changes shuts down data inputs which either relate to low The collective require collective adjustments. priority'control or need interpretative process- and the The turn needle fits both of these criteria couples to yaw, yaw produces a sideslip, ing.

helicopter subsequently rolls and pitches as a for deferred priority.

But if the cue is so result. So in a conventional cockpit, when the strong that it works through peripheral viewing, makes an adjustment to the collective to pilot the mind accepts and acts on the data. The stay on glide slope, he excites a chain reaction, explanation may not be entirely correct, but the a chain of couples that impact the equilibrium of observations of pilot response are absolutely the aircrft as though they were gust upsets. So accurate.

couple upsets an unattended collective-to-yaw attendant deviation roll and pitch, with an This brings us to consider one possible yaw, from the desired flight path. solution. Why not present heading on the atti- tude indicator? Rotate the attitude ball of the Consider a helicopter in an ILS approach, on ADI when the aircraft turns. The cue will be so If the pilot tries strong that it can be treated peripherally; as it speed but below glide slope.

to control glide slope with collective (as we is during VMC operations. The transfer is more pitch and real world. When you turn left the face of the teach him to do) he exacerbates the And pitch attitude indicator moves from left to right. This is not roll attitude control task.

Years a new idea, it has been incorporated for years in is the primary means of airspeed control.

combat aircraft (see Figure 7).

of observation leave no doubt that as the atti- tude control task becomes more difficult, the pilot becomes highly stressed. And when the Height Control pilot operates under a sufficiently high level of Another problem control task in slow speed stress, the feet stop working.

flight involves altitude control and maintenance The feet are relatively dumb control of glide slope during IMC operation. As in the the cues elements. They work well when are case of yaw control we find a quick, precise and strong, but when the workload goes up, and the powerful control in the collective. It is a visual cues are poor, this is the first control direct lift device which has no lags to confuse Under the stress of path which fails the pilot. its application. During VMC hover operations, maintaining altitude (or glideslope) and pitch one can hold hover height within inches of the attitude, the pilot's scan breaks down and the desired value, even during turns and speed changes. But when the visual cues are gone, so displays which are not directly in his compressed scan are ineffective. is precision performance.

I

The pilot does feel vertical acceleration in The power curve is essentially flat or neutral in the bucket, stable on the front side the cockpit when a collective input is initiated.

and unstable on the back side. Finally, the But during IMC maneuvers these forces are quite gradient is typically much steeper in the slow small and they often get masked by vibrations.

regime than in high speed speed forward flight.

Sometimes the vertical forces which are produced The stability of the curve and the magnitude of via (pitch) angular accelerations similarly mask the gradient all influence the pilot's ability to collective inputs. So, as in the case of the yaw cope with the vertical degree of freedom. That control, there are really no reliable natural is, all of these characteristics contribute to cues which remain, once the external visuals are define the task. the task is.least In summary, gone.

difficult on the front side and most difficult on the back side of the power curve. And it is And as in the case of yaw control, we have easier to control the vertical degree of freedom another classic control trim problem. The pilot with the collective when airspeed is held con- has a difficult time finding the control position There are several stant vs. control in conjunction with horizontal for zero vertical rate.

reasons for this problem. speed changes.

This doesn't mean to infer that the task is First, the trim point moves around anytime ever easy under IMC, slow-speed operations.

the horizontal speed changes. For example, Because the pilot must still observe the error starting from a stabilized constant altitude near hover, a small increase or and know how to precisely respond with the situation collective.

decrease in horizontal airspeed will cause a in the power required to maintain level change Take the easy case first.

flight. The aircraft then starts to climb or Flying an ILS approach, the glide slope descend, requiring a collective adjustment to signal is conven- tionally presented quite adequately.

cancel this unwanted rate. so visualizing "above" or "below" glide slope is not Reviewing the power required characteristic a problem. And in the real world we find pilots tend we find that the power for level flight decreases to lock down the collective with friction as speed decreases below VWR in much the same way and use the cyclic control to fly up or down to It bottoms out in a achieve glide slope. This works during operation as it does for the airplane.

on the front side of the power curve which is typical bucket, then increases again to peak at zero airspeed. When speed is increased to the where all civil IMC is flown.

right, left, or rearward from zero, the power required by the rotor decreases in a way similar Pilots probably use this technique for two reasons. They know they have trouble making to forward flight (see Figure 8).

accurate adjustments to the collective setting and that several adjustments will be required before they get it right. They also know that any collective change will require a directional control "pedal" input.

We've already covered the last problem under the discussion of collective- to-yaw coupling. So why are there problems setting power (Collective)?

The pilot has no precise cue of collective position. When the pilot adjusts the collective he observes the results via a cockpit display of engine torque.

But this indicated value of torque is subject to all sorts of masking.

Changes in tail rotor thrust (pedal position), a nose up control input, a roll control input, a vertical gust, and a commanded change in rotor RPM will all cause the indicated torque to change more than the amount that the pilot typically needs to input to accomplish for a climb back to glide slope. So these miscellaneous inputs mask the pilot's collective input.

Another problem involves display locat,ion.

Typically the torque indicator is displaced too far from the primary viewing area to be included Figure 8. Helicopter Power Required for Level in a high gain scan. This seems to be a very Flight.

serious problem in the civil commmunity. In this Radar Altitude When we complete an IMC approach to a hover, and hold an IMC hover, another cue deficiency becomes evident. Arriving at the Decision Height (DH) altitude the pilot becomes more aware of his absolute altitude above the ground. And another instrument becomes important, the radar alti- meter. And where is it located? In civil helicopters it is typically found in the lower right hand corner (see Figure 9).

This is a totally unsatisfactory location for such important data. The standard pressure altimeter is simply not adequate during tran- the low-speed sitions to, and operations in, regime.' Both altitude and altitude rate are unreliable to the degree generally required for Figure 9. Typical Location of Radar Altitude controlling height during an IMC hover. Since Indicator and Engine Torque Meter (Q) radar altitude and visually derived "radar alti- Shown Relative to the Primary Attitude are the best cues available in the tude rate" and Heading References.

cockpit, these data need to be presented with higher priority in the cockpit. The display group more priority is given to adding an addi- should be given higher priority, but I have tional attitude display than properly locating another solution which seems to work very well.

the torque indicator. (See "Q" in Figure 9).

This solution is illustrated in Figure 11. Here Altitude are the Decision Height (DH) and Radar Finally most torque indicators appear to be presented digitally on the lower edge of the ADI.

underdamped. I'm not sure why they are under- This is an excellent format for the final phase damped, but I believe this damping characteristic The display I evaluated is of the ILS approach.

contributes to problem of selecting the the It has an update rate which appears by Sperry.

desired power setting. For in one case, a to be well suited to the task of interpreting properly damped indication of main rotor torque altitude rate as well. If this is an radar (Bell-222) provided excellent results (as accurate assessment, such a display clearly would illustrated in Figure 10). Yet would anyone enhance a pilot's ability to hover and maneuver seriously consider asking an Auto Pilot to close in the slow speed regime.

The Sperry AD1 the vertical control loop on torque information?

evaluated also has a rising runway indicator to display absolute height. I agree that this is a Designers of FDI'S for helicopters were proper approach hut not the total answer.

faced by the same problem, a problem which they solved by including a small edge mounted pointer to indicate collective control inputs. It is not the ultimate device but its presence lends credi- bility to the need. From personal experience I a clear indication of collective can say that position allows pilots to find trim very quickly with an absolute minimum of effort. So the solution is a full range collective position indicator.

LDH

L RADALT

Figure 10. Characterization of Torque Meter Used Figure 11. AD1 Including Digital Presentation of in the Bell-222 Model Helicopter.

Decision Height and Radar Altitude.

vice a versa. The first concept is' true only if Back-Side Speed Control the second is true. That is, the variation of Then there is the problem of speed control pitch attitude must be stable or at least neutral before one can on the back side of the power curve. The power use pitch attitude to reliably attain and hold airspeed. 'In most required.curve typically has a steep gradient on cases the And the slope stable attitude characteristic required and de- the back side of the curve.

represents an unstable situation when considered sired does not exist.

in the context of man-machine control. That is, if the aircraft slows down it will descend. This Since pitot-static airspeed indicators be- means that to climb back on glide slope the pilot come totally inoperative below about 40 knots trade airspeed cannot simply pull up (flare) and anyway (depending on the aircraft and the flight act causes to regain the glide slope. Such an profile), some sort of reliable speed cue is the aircraft to lose speed, with a momentary required so that a pilot can separate the speed positive.response but then the aircraft settles control task from the vertical control task.

further below glide slope. Ground speed can be derived from many current equipments, so that is one possibility. But the So during a slow speed aproach, the aircraft is actually responding to the airmass, collective is clearly the control which the pilot not ground speed. So it seems obvious that an must use to modulate descent rate or angle. And airspeed system which operates down to zero to simplify the control activity, it is necessary airspeed is clearly required.

to hold airspeed constant. But again there is a Again we are faced with a question of where problem with the data available to the pilot.

to locate this additional data display. Collins Airspeed Display in cooperation with PACER Systems, Inc. is developing such a display for the U.S. Navy (see When power is added, most pitot-static air- Figure 12). This display is multi-mode, allowing speed systems reflect an apparent change in both pitot-static and omnidirectional low range airspeed. A change in slideslip angle or angle- airspeed to be presented on a single indicator.

of-attack will have a similar result. So even at Thus it can one-for-one replace the current airspeeds where the pitot-static airspeed system airspeed indicator.

is still supposed to function (above 40 knots) the pilot can- find himself chasing changes in Experience has shown that this type of "position error-. Thus, he actually causes speed airspeed data is not subject to the problems that changes to occur in a needless attempt to hold the pitot-static system.

plague Now the pilot speed constant. The real speed changes are can use the longitudinal control to directly therefore confused with the changes produced by regulate airspeed. He no longer must try to hold changing values of position error, and the pre- a constant attitude to determine, after some cision of the entire approach task deteriorates. several seconds, what might happen to airspeed.

Thus, we have decreased the amount of time that To avoid reliance on pitot-static airspeed, the pilot must allocate to the ADI to accomplish pilots are told to maintain a constant pitch the airspeed control task. This reduction in attitude to hold airspeed. They are also told workload further releases the pilot's control that more and more nose down trim attitude will logic to handle the heading and attitude control produce a faster and faster trim airspeed, and tasks discussed earlier.

Figure 12. Collins ASI-800, Omnidirectional Airspeed Indicator (OAI).

Conclusions the pilot is in a pOSitiOn to Use NOW get ahead airspeed and all-the other new cues to In conclusion, let's review the highlights.

He no longer must put in of the aircraft.

see a response. In Reducing the workload associated with horizontal control inputs and wait to to antici- the pilot can learn speed control reduces the pilot's workload so actual practice, of collective change which is other display- pate the amount that he is better able to use the to maintain level flight (or glide control enhancements such as: required slope) as speed is gradually changed.

o Pitch-Roll-Yaw Attitude Indicator Since the Omnidirectional Airspeed Indicator o Collective Position Indicator (OAI) of Figure 12 is able to present airspeed for flight in all directions from zero, it is o Radar Altitude Indicator also now possible for pilots to observe the cross wind component as just that, a component of o Excellent downward visability airspeed. They can learn what 5, 10, and 20 knot components will mean as he decelerates and The net result is to bring the pilot's achieves a hover.

And when the component is too could be control task in line with what.a human high for safe operations, the pilot will be able achieve. The more' the pilot can expected to to anticipate the situation before an unmanage- achieve with the basic helicopter the more viable able hover is attempted near obstructions, etc.

the helicopter will become in civil and military The Landing applications.

To land under IMC conditions, I would expect that for some long number of years into the future, pilots will be required to have visual contact with the ground.

I really see no reason why there should ever be any reason for a CAT III type flight control system in a helicopter. What the industry needs is a system which the pilot can use to get into close proximity to the landing surface. I have hovered in some really dense fogs, but I can never remember a case where I couldn't see the ground at 20 feet. Here is where downward visibility re-enters the picture.

Nothing is as accurate and reliable as the pilot when it comes to accomplishing a vertical land- ing. So mostly the problem is stopping the helicopter over the landing pad, at an altitude of 50 feet or less. With excellent downward visibility the pilot continues to fly the air- craft down to a touchdown.

I

INTEGRATED COCKPIT FOR A-129 Dott. Ing. Filippo Reina Director, Helicopter Systems Engineering Costruzioni Aeronautiche Giovanni Agusta S.p.A Cascina Costa, Samarate, Italy James A. Gracia Systems Opezp Engineer Bryce W. Koth Human Factors Engineer Harris Government Information Systems Division Melbourne, Florida Abstract and 3) the prototype field trials. The ground The Agusta A-129 is a compact and lightweight simulator will validate the svstem for the most tandem cockpit combat helicopter under development effective flyable configuration. The first flight for multimission usage with a full complement of will initiate fine tuninq of this conceot to the electronic aircraft control and ASE equipment.

flight environment. Throughout the flight test of Weight. size, and mission requirements for the the four prototype helicopters, a large amount of A-129 mandated an integrated -system approach for experimental data will be amassed and used for the the crew/cockpit interface design. Instead of the final definition of the operational cockpit usual multitude of cockpit controls, indicators, system. The incorporation of these experimental and lights, the primary crew interface is wges, test results will confirm the operational a single multifunction keyboard and one or more flexibility of the software intensive man-machine multifunction CRT disolav units. This cockpit interface.

design approach imposed unusual constraints.upon the system architecture to overcome the inherent A-129 Definition* information access limitations of a data input/output window that was restricted by the The Agusta A-129 (Fig. 1) is a light, twin available space. This paper describes the turboshaft powered, combat helicopter under conceptual approach and resulting design of the development for the Italian Army to serve A-129 cockpit with the intent to enhance the primarily in an anti-tank role. It has a single development of cockpit standardization.

four-blade articulated main rotor and a two-blade semirigid tail rotor. The helicopter design, Introduction presently completed and frozen, reflects the results of extensive trade-offs down to the The A-129 integrated cockpit is the crew's component level in order to satisfy the Italian interface with a sophisticated weapon system. Its Army's requirement for an agile, small size, particularly small size and mission scenario limited cost aircraft, which retains the flexibility make the A-129 man-machine interface a advantages of state-of-the-art technology.

challenge to optimize. Its crew compartments must accommodate pilot and gunner comfortably, provide excellent forward quadrant visibility for the pilot to safely fly in the Nap-of-the-Earth (NOE) environment, and provide command and control capability for aircraft and mission equipment. A highly integrated helicopter system and a highly integrated cockpit are demanded to satisfy the A-129 requirements.

Integrated systems are presently emerging but the design technique has not. There are isolated examples of integrated cockpit design techniques, but a validated design practice that is fully accepted by the industry and the government-user community does not exist. The A-129 demanded a highly integrated cockpit, therefore a conceptual system design approach had to be evolved. This paper addresses the approach followed in the Fig. 1. A-129 Helicopter definition and development of this innovative hiqhlv inteqrated cockpit. However further rerinements-are planned during 1) the ground simulator test phase (October 1982 - August 1983), l Lovera, Bruno. “The Awrta A-129”. Vertiflite, 26 (61, November/December, 2) the early flight test of the A-129 (late 1983), 1980, pp 69.

UHF-AM, VHF-AM, VHF-FM and HF communications The crew of two is seated in tandem with the radios with standard NATO encrypting devices are aircraft comnander/pilot located aft and above the integrated and controlled though the A-129 IMS.

Primary armament in the present copilot/gunner.

Primary armament consists of the TOW M-65 missile plan is the TOW system with 8 missiles carried on system, but space, weight, and power provisions 4 pylons mounted to the stub wings. Rockets, and external fuel tanks can be have been incorporated into the A-129 design for machine gun pods, FLIR Augmented Cobra Tow System (FACTS), Laser interchanged in any combination with the TOW missiles. Augmented Aerial Tow (LAAT), and second generation missile systems.

The A-129 has some unique design features which The A-129 will also be fully provisioned with give it unprecedented capability and flexibility The cockpit configuration a full suite of Aircraft Survivability Equipment as a combat air vehicle.

(ASE). Candidates for inclusion are Radar and provides both crew members with identical and Laser Warning Receivers, Radar and Infrared unequaled flight visibility. The main transmission,-main rotor shaft, rotor head, and Jamners, and Flare and Chaff Dispensers.

primwy flight controls are designed to provide protection against icing and ballistic and/or wire The A-129 baseline system capabilities are listed in Table 1.

line damage as well as a stable mounting platform for a mast-mounted sight (MMS). Finally, the A-129 incorporates a modular, expandable, Configuration Flexibility multiprocessor-based data bus system which is presently unequaled in comprehensiveness and Due to its size and performance, the A-129 is flexibility in a rotary winged aircraft.

an attractive candidate for a variety of weapons options as well as other combat helicopter roles Extensive Italian Army and Agusta study and and missions. The possible armament options range experimentation, including the use of full-scale from the heavier more potent HELLFIRE to the ultra- cockpit mockups, were used to arrive at the A-129 lightweight air-to-air Multi-Launch Missile System cockpit configuration. The "camel" configuration, (MLMS). With respect to differing roles and as shown in Fig. 2, was selected to maximize missions, there is strong interest in several NATO visibility for both crew members in NOE mission countries in a capable, survivable, multi-role environments.

attack/scout helicopter. Consequently, the A-129 design had to be versatile and flexible. And the design of the inteqrated cockpit had to be sufficiently flexible to acco&nodate conversion to either HOT or HELLFIRE and for MLMS or other air- to-air missile systems.

In the visionics arena, the initial A-129 prototypes will incorporate the nose mounted, day only, M-65 TOW Siqht Unit (TSU). with orovisions for-LAAT and FACTS. However, several other visionics options are either in production or R&D which will enhance the capabilities of the A-129 as a survivable attack or scout helicopter. In particular, provisions have been incorporated into the A-129 cockpit design for conversion to a Mast- Mounted Sight/Target Acquisition and Designation

sy5.m (FMS/TADS). This equipment. oresentlv

under development, will retain co&onality with the TADS develooed bv the U.S. Anav. The visionics package wiil be completed with PNVS and IHADSS (Integrated Helmet and Display Sight Fig. 2. A-129 Cockpit configuration System).

Integrated System' The Integrated Multiplex System (IMS) provides Baseline Configuration this small helicopter with unprecedented flexibility, and extends its basic caoabilitv via The A-129 is equipped to fly and attack in day integration and automation. The heart of the IMS and night under instrument and visual flight conditions. In addition to full instrumentation, is a redundant MIL-STD-15538 data bus communication and centralized data processing system. The stability augmentation, and auto-pilot capability, multitude of functions performed by the IMS fall the Integrated Multiplex System (IMS), the heart into four major categories: of the integrated system, incorporates a navigation computer capability. Inputs to this system come Mission Electronics - Optimized integration of from the installed Doppler radar system, the radar altimeter, and standard navigational radios. For all mission equipment, including radios and navigation equipment, which provides for change and night visual flight, a Pilot's Night Vision System (PNVS) will be incorporated. growth as well as efficiency. Performs navigation, weapons delivery, and provides performance monitor Schedule and Cost Objective capabilities.

Agusta's objective is to enter production in Basic Aircraft Systems - Handling of electrical the mid-1980's with a modern, versatile combat power distribution. Dower slant monitorinq. helicopter which is (1) capable of performing caution and warning presentations, and integration effectively and surviving in a NATO threat of other environment; (2) capable of being adapted to a general airframe related electronics. Provides variety of attack, scouting/reconnaissance, and checklists and status of integrated equipment. battlefield management roles; and (3) reliable, maintainable and affordable.

Flight Control and Stability - Including motion sensors, digital stability augmentation systems, RequirementsSummary and flight director functions, plus a complete redundant backup fly-by-wire control system which In essence the basic requirements for the A-129 can be engaged and disengaged at will. integrated cockpit design can be distilled as follarrs: Cockpit Control and Display - Including integrated flight management, equipment control, 1) Control a comprehensive complement of instrumentation, and workload reducing automation. aircraft/mission equipment 2) Flexibility to modify mission requirements The cockpit is where the benefits of the IMS with ease of fleet retrofit are most apparent. The need for a multitude of 3) Canaonality of tandem cockpits controls, switches, indicators, displays, and 4) Limited cockpit real estate lights is eliminated. In their places are a single 5) Acceptability of crew workload multifunction keyboard (MFK) unit and one or more 6) Integrated system approach to meet video-type multifunction display (MFD) units for mission/weight requirements each crew position. 7) Program cost and schedule objectives Table 1. A-129 Baseline system capability EQUIPMENT INTEGRATED FUNCTIONS .-- COMMUNICATIONS COMMUNICATION/IOENTIFlCATlON CONTROL AND OISPLAYS UHF/VHF TRANSCEIVERS PRESET SELECTION HF TRANSCEIVER ICC T”l.,bm”.,“CD NAVIGATION/FLIGHT CONTROLS FLY-BY-WIRE NAVIGATION OOPPLERlAlR OATA NAVIGATION AIR DATA SYSTEM . . .

STABILITY/CONTROL AUGMENTATION ‘PLER RADAR OOF ATTITUOE HOLO DIRECTIONAL GYRO HEADING HOLO RADAR ALTIMETER ALTITUDE HOLD AUTO OIRECTION FINDER GROUND SPEED/AIR SPEEO HOLO VERTICAL SPEED HOLD FLIGHT CONTROLS VERTICAL GYROS (THREE) HOVER ACCELEROMETER TRIAD COURSE HnLO --- STANDBY INCTRIIMFNTS I WAYPOlNl -rENTRY . ..I..._..._._._ COUPLE0 FLIGHT PLAN AFCS ACT IVATORS I SURVIVABILITY EQUIPMENT WEAPONS STORES MANAGEMENT RADAR AN0 LASER WARNING RECEIVERS WEAPONS DELIVERY FLIGHT CONTROLS FLARE AN0 CHAFF DISPENSERS POWER TRAIN WEAPONS ENGINE MONITOR SYSTEM M-65 TOW ROTOR TRANSMISSION MONITOR TOW SIGHT UNIT HYDRAULICSMONITOR AND CONTROL 2.75 ROCKETS VIBRATION MONITOR FUEL CONTROL ENGINE FUEL MONITOR ROLLS ROYCE GEM-Z ENGINES (TWO) ENGINE MONITOR SENSORS COCKPIT INTERFACE ELECTRICAL POWER CONTROL FUEL SUBSYSTEM CONTROL ROTOR TRAN~ISSION PERFORMANCE MONITOR HYDRAULICS STATUS ELECTRICAL POWER DISTRIBUTION CHECKLISTS INTEGRATED COMPUTER/MULTIPLEX BUSSYSTEM Cockpit System Design 5) Performance Monitor Subsystem computes information required by the crew for mission planning.

The A-129 cockpit system design was constrained by requirements for duplicate capability in each 6) Navigation Subsystem provides graphic and cockpit and the limited real estate in each interrelated alphanumeric displays which allow the cockpit. A perspective view of the pilot's crew to use, control, or update the aircraft cockpit is shown in Fig. 3. Since all A-129 subsystems are not incorporated within the navigation.

integrated system, some space was allocated to conventional controls and displays for these 7) Checklist Subsystem provides a semi-automated sequence for crew execution of subsystems. The remainder of the cockpit was procedures normally listed in flight manuals.

dedicated to the integrated cockpit system which functions as the crew/IMS interface. Programmable These seven subsystem categories are interfaced by displays and controls for the integrated cockpit system were located to optimize visibility and the crew through a comnon basic dialogue design.

accessibility. The balance of the cockpit real estate was budgeted by priority-criticality of Basic Design information and control and/or frequency of usage.

The cockpit system consists of two interface devices - a multifunction disolav unit (MFD).

which is basically an interactive video‘dispiay terminal, and a multifunction keyboard (MFK), which provides the control interface for the crew to select displays, create IMS entries, and manage the Automatic Fliqht Control Svstem (AFCS). Cockpit panel space restricted the-CRT for the MFD to'a usable viewing area of only 4.80 inches square.

Accounting for U.S. DOD and other recognized requirements on minimum character size, and the expected viewing distance, a maximum MFD data window (page size) of 15 lines and 28 characters per line was provided. A sumnary of character requirements and design parameters of the A-129 alphabet is shown in Table 2. The MFK was similarlv constrained bv the limited soace in the Fig. 3. Pilot cockpit perspective upper left cockpit console area: the baseline A-129 allocation was 5.75 in. wide bv 7.50 in.

long, of which only about 50 percent"could be The crew interface aspects of the A-129 IMS utilized each for the AFCS controls and indicators modular subsystems can be grouped into seven area and the MFK keypad area. Despite the use of general categories: a maximum density U.S. DOD compliant keypad configuration, the requirement.for full-.

1) Automatic Flight Control System (AFCS) is a alphanumeric pilot entrv capability limited the high priority module that demands dedicated number of special functjon display-call keys to 15.

controls and displays.

Access to all alphanumeric MFD data and controls representative of all subsystems integrated within 2) Basic Equipment Controls and Displays are the A-129 IMS therefore had to be prioritized under represented by callable display pages that have a 15 key hierarchy. The graphic displays are been translated from conventional control/display called by three dedicated MFD keys.

devices. Connnunications. Weapons Stores Manaqe- ment, Engine Monitor, Fuel, Rotor/Transmission, Hvdraulics. Vibration. Electrical Power Control.

The A-129 control/display access hierarchy is IMS test and configuration I aid Utility (e.g., shown in Fig. 4. At the far left are listed all control) are subsystems that fall within this dedicated control and display devices: these category.

devices are available at all times for the A-129 crew usage. The second column lists all IMS 3) Caution/Warning Subsystem consists primarily subsystem displays and controls which are directly of dedicated programmable display space for instant accessed by depressing one MFK or one MFD button.

presentation of alerts and a means of storing The third column lists all secondary subsystems alerts for later retrieval.

and functions accessed through menus called by single button pressings, while the fourth column lists displays accessed only from the secondary 4) Status Monitor Subsystem is a branching structure which collects and reports on-line BIT menu displays or fran other secondary display pages. The directly callable subsystem displays fault indications from throughout the IMS.

Starting with a top level display, the crew can were selected with a sensitivity to pilot needs in access lower level status surmnaries to isolate the presently defined Agusta mission scenarios.

faults. Appropriate priority was given to basic aircraft Table 2. A-129 Alphanumeric character font requirements and design (3) REllllIREMENT(1) DESIGN PARAMETER 0.192 in.

20’(2) Vertical Height (20’ @ 33 in.)

62.5% 60-100% Width to Height Ratio Stroke Width to Character Height 12.5% 12.5-14% Ratio 21 Stroke 2 Stroke Vertical Spsing (Between Characters) Width Widths Horizontal Spacing (Between Lines, 62.5% 50% Relative to Character Height) NOTES: (1) Applicable requirements from “Human Engineering Design Criteria for Militaw Systems, Equipment, and Facilities, “MI L-STO-14728. Note 1.

10 May 1976; Woodson, W.E., Human Factors Design Handbook, 1981; and Shurtleff, D.A., How to Make Displays Legible, 1980.

(2) Symbol size must additionally be corrected for off-axis viewing angles per Reinwald (as published in Shurtleff, 1980).

(3) The font is 10x16 pixels with 2 pixel stroke width.

SUB-MEN” OR PAGE CALLABLE r.tENu CALLABLE ISECONDARY) DEDICATED DISPLAYS AND CONTROLS DIRECTLY CALLABLE (PRIMARY) (TERTIARY) / FUEL DATA AFCS INOICATORS CAUTION/WARNING ALERT TEMPS WEAPONS e NAY DATA ICS PANEL MASTER CAUTION/MASTER WARN ANNUNC STATUS PANEL REMOTE OISPLAV “NIT - POWER CONTROL PANEL COMPASS ALIGN IFFIAOF ENGINE REVERSION SWITCH UHF/VHF 1 ------- COMPASS CALIBRATE ENGINE START SWITCHES VERTICALSCALE ,NSTR”MENTS “HFNHF 2 LIGHT SWITCH PANEL W,NDS”,ELD NEATER SWlTCHES TOW,TSU LASER RANGE FINDER ROCKET TRIGGER CYCLIC, COLLECTIVE. RUODER CONTROLS ENGlNE FIRE ANNUNCiSWlTCHES ’ LlGHTS,WlNOSHlELDiPlTOTPOWER STANOBV INSTRUMENTS ERG PRESTART CHRONOMETER CHECK LIST ____\ ;,“;F;;;; AND RUN UP AAOAR WARNlNG -SHUT DOWN - HOVER CAPABILITY POWER CHECK PERFORMANCE MONITOR s EgF POWER AVAILABLE FUEL DATA CRUISE PRFY Fig. 4. Awsta A-129 control/display system hierarchy Status indicators present operational status and subsystems (such as the Basic Aircraft subsystem, passage of self-tests. The FBW switches control which accesses the Engine, Transmission, and Fuel as well as Electric Power Control main rotor and tail rotor electronic control subsystems, etc., systems and arm the automatic sensing of control that provides access to all remote control circuit breakers throughout the aircraft), the special linkage severances. The Stabilization section of graphics displays (which when coupled to the AFCS AFCS provides pilot control over A-129 augmentation systems in pitch, roll, yaw and collective, provide flight guidance cues), emergency and fault provides automatic hold in altitude, attitude, condition monitoring (Status, Caution/Warning), heading, and speed (vertical, ground, and air) and Navigation, and mission critical Weapons.

provides automatic trim and coimiand to wing level Priority was also given to special requirements attitude. All Stabilization modes have fly-through (such as the full complement of readily-accessed communications units) and automated pilot capability.

information aids (Performance Monitor and Checklists).

The Flight Director functions, in the display mode, provide flight cues to the pilot on the Enhancement of the pilot/IMS cockpit system graphic situation displays. In the engage mode, interface was addressed through automation, the Flight Director operates as a full autopilot systematic standardization of visual coding for with fly-through capability. The Flight Director control options and display types, and a friendly indicators, supplemented by aural alarms, are used dialogue style requiring simple keyboard entries for visual conrnand cues while in the Flight Plan fran the pilot and providing fixed visual prompts.

mode. The Flight Plan indicator flashes as Further automation is a subject for future growth.

waypoints are attained: this button must be It has been implemented in the present IMS where depressed to commit to the next flight plan leg, tradeoffs with system complexity were acceptable.

otherwise the flight plan mode disengages.

Automation is most apparent in the caution/warning alert generation and recording, in the collection of all on-line BIT type status information for Basic Equipment Control/Display Design easily accessed pilot viewing, in the coupled aircraft control and weapons delivery subsystem, A generic example of pilot usage of the A-129 in the aircraft performance monitor calculus, and cockpit system for basic aircraft equipment control in the manner in which sequences of displays were and monitoring is shown in Fig. 5. The hydraulics automated to follow a logical mission task subsystem is first accessed by depressing the MFK A standardized method of information orientation.

"BSC ACFT" key. The MFD displays the basic coding was implemented to help the pilot aircraft menu list and the line address key differentiate between data, control requests, and adjacent to "HYD" is pressed to call the HYDRAULIC control status information displayed simultaneously display page. This page contains switches for the Sane of these coding techniques on the same page.

three hydraulic power supplies and presents status included the use of character attributes: control information on the position of each switch, and status positions are designated by boxes, special present values from three pressure sensors and timely or cautionary information is displayed in three temperature sensors. The HPS switches are inverse video, and keypad entry control prompts controlled as shown. Pushing the line address keys are underlined. Coding conveys contextual shown operates the switches in rotating fashion information to the operator; standardization with the present switch position always indicated reduces training and learning requirements; and by a box.

both simplify the task.

Pressure and temperature data is conveyed in digital format and in an analog gauge format. The AFCS Design gauge displays are analogous to dedicated cockpit instruments and provide dynamic range information The Automatic Flight Control System is managed as well as a "quick look" capability. Gauges may by the pilot using the upper portion of the MFK.

contain up to one "green", two "red", and two The AFCS interface includes IMS redundancy control, "yellow" zones as indicated by the steps shown.

Fly-By-Wire (FBW) control, Stabilization functions (attitude hold, headinq hold, vertical speed hold, The boxed double arrow in the lower left hand altitude hold,.airspeed and groundspeed hold, corner of the display indicates paging options to the pilot. In this case, the pilot can slew the autotrim. and winss level) and Fliqht Director display "up" (allcwing a return to the basic modes (hover, attack, course hold,-and flight aircraft menu) or "down" (to the second page of plan). The interface was implemented using 0.75 in hydraulics information). The box indicates that sauare liqhted leqend pushbuttons for an integrated control/indicator-design. Switches have visible these options are enabled and available to the legends (backlighted at night) and dead face pilot via the MFD rocker switch.

indicators; legends are visible in full 10,000 Caution/Warning Subsystem Design footcandles ambient sunlight; indicators are color coded for daytime cue enhancement; legend lighting Caution and warning alerts to the pilot are is compatible with present (second generation) Night Vision Goggle systems (NVGS). critical to success of the mission and the safe operation of the A-129 aircraft. Accordingly, the IMS redundancy control switches allow the pilot caution/warning subsystem has a dedicated display: the 15th (bottom) line of all MFD graphic and to manually select one of two redundant processors alphanumeric displays. Other dedicated cockpit or to engage the automated self-selection mode.

interface devices for this subsystem include the MASTER WARN and MASTER CAUTION annunciators and switches. Sane critical warning alerts are accompanied by an audible alarm fran the ICS.

The caution and warning subsystem for A-129 continuously compares data from various sensors and BASIC AIRCRAFT other subsystems to predetermined thresholds and When a caution or warning failure criteria.

FUEL condition occurs, the alert name is presented on VIBRATION ENGMONITORSYSTEM the MFD (in a priority queue with all warnings presented first) and the MASTER WARN or MASTER TRANS CAUTION annunciator is lighted. When the pilot acknowledges the alert by depressing the ROTOR appropriate annunciator, the alert name and other HYD information are entered into a quasi-LIFO list which is retained in nonvolatile memory for retrieval by the pilot or by ground support personnel. The list is accessed by depressing the.

MFK PCAUT" key and a sample display as shown in Fig. 6 is presented to the pilot. In addition to the alert name, information is also presented to indicate whether the alert was a caution or warning, whether it is a historical alert or is presently active, and whether the alert sensing device is inhibited or active. Line address keys adjacent to each caution/warning alert provide the capability of inhibiting (or re-activating) a sensor which may be faulty or to suppress repetitive presentations of the same alert. For HYDRAULIC switch type alerts, the number of occurrences is also reported. For analog sensor alerts, the PIPS 1ONm value reported is always the highest value HPSZm OFF HPS3ONlOFFj attained.

PRES 1 1200 PSI1 67 r- Status Monitor Subsystem Design PRES 2 1155 PSI1 m r PRES 3 1870 PSI1 TEMP 1 105 'X 2, The status monitor subsystem continuously TEMP 2 54 PC 2~ monitors and collates the status of on-line BIT TEMP 3 32 PC m-- results throughout ttie A-129 IMS system and presents this information to the pilot when the

B

"STAT" key on the MFK is depressed. Two of the HYDOILT1HIGH 105% unique status monitor displays presently implemented on the A-129 IMS are shown in Fig. 7.

Three information display conventions are used. A subsystem fault is indicated as a "NO GO" in inverse video. Absence of a fault indication for any subsystem is indicated simply by a blank field.

Where it is possible to track down a fault condition fran a subsystem to a lower equipment level, the inverse video NO GO is supplemented by a downward arrow symbol. Depressing the adjacent line address key calls up a l@fer-level display HYORAULlC which may contain the fault. In the example shown, the IMS status is NO GO because of a failed Master HPS 1mOFF HPS2m OFF Unit (MU 2) and a failed Data Bus (D/B) Test A.

HPS 3mOFF Lower level access to status information is available for the IMS equipment, the Navigation and PRES 1 1200 PSI- PRES 2 1155 PSI- AFCS equipment, the Electrical Power Control PRES 3 1570 PSI l&'- remote circuit breaker trips), the subsystem (e.g., TEMP 1 II ‘X Basic Aircraft subsystems, and the radios.

TEMP2 54oC A TEMP 3 32 OC &w Performance Monitor Subsystem B HYDOILTlHlGHi05PC The A-129 performance monitor subsystem provides information to the pilot for making preflight and inflight plans and decisions based on aircraft performance predictions. By viewing Y CAUT

P

STATUS MONITOR CAUTlON/%VARNlNG #0k&wlMS C 10 HFS TEMP 1 105oC UHF/VHF 1 NAVIAFCS

C q TRANS OIL HOT 12

UHF/VHF 2 ELEC POWER CONT Cm0TRANS BRG 1 95oC IFF BASIC AIRCRAFT W 0 HP-9 TEMP 2 145OC HF Wff#JOATT HOT 1

El

‘UEL 1 LOW IMS STATUS CAUTION/WARNING MU 1 Cfjn HFS TEMP 1 105oC ?@@MU 2 ERT 1

c q TRANS OIL HOT 12

ERT 2 RU 1 C lm TRANS GRG 1 95X RU 2 SYM GEN 1 W 0 HPS TEMP 2 145% SYM GEN 2 MFK 1 WI0 OATT HOT 1 MFK 2 MFD 1

III

Ei FUEL 1 LOW Fig. 6. Caution/Warning inhibit function Fig. 7. Status function operation Navigation Subsystem the MFD, the pilot has available all information The navigation subsystem integrated within the normally accessed in the typical pilot's flight A-129 IMS provides a highly complex, highly pilot- manual. Data such as flight range, flight time, interactive, and highly automated capability that autorotation envelope, etc. are presented using combines on-board sensors to calculate current either actual environmental conditions (e.g., OAT, aircraft position and progress against a prestored ALT, etc.) from aircraft sensors or pilot input flight plan. When coupled with the AFCS in the values for either preflight or inflight flight director mode, a prestored flight plan can calculations. One of the most critical variables be executed automatically.

to these calculations, aircraft weight, is derived in real time by interaction with the weapons and The principal sensors providing inputs to the fuel subsystems. Engine performance data A-129 navigation subsystem include three orthogonal calculated in the Engine Monitor Subsystem is accelerometers, Doppler velocity sensor, vertical utilized to provide actual aircraft performance gyros, TSU range and bearing data, magnetic characteristics.

Predictions and calculations heading, air data (air speed, barometric altitude, obtained through usage of the performance monitor rate of climb), and radar altimeter. Interfaces subsystem are thus greatly improved over results with directional gyros and OF radios are also formerly obtained through conventional pilot usage integrated within the navigation subsystem, of manuals or nonintegrated standalone performance allwing automatic tuning, compass/gyro systems. synchronization, and compass calibration. Other navigation subsystem functions include computation Checklist Subsystem of track, cross track, and track angle errors; position update over a known waypoint, and target The A-129 checklists subsystem provides a semi- position acquisition by pilot input or TSU; automated mechanism for interactive control and prompting of pilot procedures normally performed computations of corrected heading considering with flight manuals. Engine Pre-Start, Engine magnetic heading variations and wind speed and Start and Run-Up, Preflight, and Shut Down direction.

procedures are implemented as sequences of alphanumeric checklist pages with quick access to The highly pilot interactive navigation other subsystem status pages where required (e.g., subsystem accepts manual keyboard entry of up to to monitor aircraft oil pressures and temperatures 100 waypoints (specified by latitude/longitude or UTM coordinates, elevation and altitude, and a during run up).

target/enemy/friendly designation) which may be Future Activities arranged into as many as ten different flight plans. Full editing capability of flight plans The definition and design phases of the A-129 and waypoints is provided. Flight plans can be integrated cockpit system are complete. The created prior to flight or while in flight and implementation phase is currently under way. The allow a fly-to-waypoint capability. Other pilot next steps in the system evolution are the interactive capabilities include the selection of navigation mode, input of meteorological and integration of the many IMS subsystems, checkout of the A-129 prototype, field trials, and eventually magnetic data, waypoint fix in offset or flyover production.

modes, and navigation update in flyover or manual position input modes.

Parallel with the cockpit system implementation phase a simulation is being constructed which Besides an alphanumeric navigation data page, the primary display outputs of the navigation mechanizes the display hierarchy. This computerized simulator will serve a dual function: subsystem are the graphic situation displays - the 1) to familiarize the test pilots with the Horizontal Situation Display (HSD), Forward Situation Display (FSD), and Vertical Situation integrated system and gain their acceptance, and 2) to provide pilot feedback for system fine Display (VSD). The A-129 HSD is shown in Fig. 8.

tuning.

The HSD is a graphic map display with compass rose, a variety of aircraft control and guidance The system integration and checkout phases of cues, several aircraft performance indications, and flight plan related data. Aircraft control the A-129 IMS will provide the first real measure cues include aircraft heading, aircraft track, a of pilot workload. The overall goal of the IMS cockpit design is obviously to enhance pilot track-offset based steering cue, a digital aircraft course readout, and a crosstrack error mission performance through reduction of the scale. Aircraft flight performance indicators workload associated with normal aircraft control and mission operations. The pi1 ot acceptability present digital values for air speed, ground speed, and vertical speed and direction. of the workload reduction afforded by the IMS Flight cockpit system is ultimately the final measure of plan related data that are presented on the HSD success.

include a connected waypoint sequence with leg distances, interest and avoidance area waypoints, Integration and checkout phase activities will waypoint type, number, and alpha identifiers, be followed by the A-129 operation and field trial selected flight plan number, next waypoint, time and distance to next waypoint, and time remaining phase. This will be the final test of the IMS on target. cockpit design concept, and the outcome of this Other display related information included on the HSD are the display scale and map orientation, Y ,

both of which are pilot-controllable using the MFD \ q

line address keys and rocker switch. As in all MFD display formats, the HSD has dedicated fields for a pilot entry scratch pad and for presentation of caution/warning alerts.

The FSD is a subset of the HSD with the aircraft in the larer center of the screen. A linear course prediction line is also provided.

The VSD is similar in concept but it displays a perspective ground plan pictorial relative to aircraft position, aircraft to ground and aircraft to tree top distance cues, and aircraft velocity CAUTION + WARNING MSG

/

vectors in all dimensions.

%-il~

Fig. 8. Horizontal situation display phase will be the production IMS system baseline.

Flight tests of the A-129 prototypes will The test trials will provide an environment for provide data to verify the integrated cockpit further evaluation of the design in meeting multi- desiqn approach, which will then be useful in mission sortie requirements, another opportunity to establishing pilot accepted standards that may evaluate the pilot acceptability of the workload.

become as popular as round dials have been in the To improve system orientation to a particular past.

mission. field trial results mav indicate the need to retrofit the A-129 cockpit design. Any subsequent redesign activity will be easily accommodated by the overall flexibility and modularity of the cockpit system design.

During the production phase and remainder of the life cvcle of the A-129 IMS. manv oooortunities to enhance-the cockpit system design-anb'aircraft mission capability will occur. The new emerging technologies in pilot interactive devices (e.g., visionics, supplemental helmet mounted displays, voice recognition command and synthesized voice alerts, color CRTs for additional information coding and workload reduction, digital moving maps and perspective terrain presentations, etc.) are all software-driven. At the center of the A-129 integrated cockpit system is a separate operator interface software subsystem. This system architecture isolates and thereby simplifies the process of updating the cockpit interface design to accommodate emerging technologies.

There are of course many different approaches to pilot interaction in an integrated cockpit design. The approach we have followed incorporates a set of concepts which represent a quantum advance to integrated cockpit technology. We anticipate that this design evolution will provide information to foster further development of standards for cockpit interface systems, thereby freeing future creative effort to concentrate on other technology aspects of the operator interface.

Concluding Remarks The A-129 integrated cockpit design is driven by a stringent requirement for handling a sophisticated system within a small cockpit area.

This has resulted in a highly integrated system.

Its conceptual design has been completed and it is in the implementation stage. In a short time the prototype models will be inteqrated with the A-129 helicopter systems and then-they will be field tested. At this point the desiqn will have matured to a state that'it will be ready for production and deployment.

The modularity of the integrated system's software has separated the subsystems to permit easy retrofitting of new or other subsystems required to support new mission requirements.

This modular desiqn was carried into the integrated cockpit design so that the controls/ disolavs of additional subsvstems can be easilr incorporated. This modular-approach was also - planned to permit the incorporation of new developments in the design of cockpit interface equipments. Hopefully this degree of technology independence will permit cost-effective optimization of the A-129 during its fielded life while maximizing operational capability.

NEW DEVELOPMENTS IN FLYING QUALITIES CRITERIA WITH APPLICATION TO ROTARY WING AIRCRAFT Roger H. Hoh Principal Research Engineer Systems Technology, Inc.

Hawthorne, California Abstract precise point on the runway.

In a paper presented to the AGARD Flight Mechanics Panel in 1981 the Some recent considerations and developments in authors of Ref. 4 cited a case where a pilot gave handling quality criteria are reviewed with empha- a surprisingly good rating to what should have been sis on using fixed wing experience gained in devel- a particularly poor configuration. However, the oping MIL-F-8785C and the more recent Mil Standard landings were not in the prescribed touchdown area and Handbook. Particular emphasis is placed on the and the author (who was also the safety pilot) tasks and environmental conditions used to. develop insisted that the evaluation pilot improve his the criterion boundaries, SAS failures, and poten- performance. On the very next run, in an attempt' tial fixed wing criteria that are applicable to to achieve the required precision, a severe PI0 was rotary wing aircraft.

encountered near touchdown. Needless to say, the Introduction Historically, the handling qualities of rotary wing aircraft have been vastly inferior to their Level 2 Mil-F-8785C For example, the pitch fixed wing counterparts.

//,,,,//,/,11/,,,/1/111/1~ attitude control of many operational helicopters Level 3 Mil-F-8785C - will not even meet the Level 3 requirements of MIL- F-878X. (Level 3 is defined as a Cooper-Harper rating of worse than 6-l/2 or "Flying qualities such that the airplane can be controlled safely but pilot workload is excessive or mission effective- ness is inadequate or both."). An example is 1 where it is shown that the illustrated in Fig.

time to double amplitude for several operational helicopters is in the extreme Level 3 region. The major deficiencies of rotary wing aircraft are nearly always associated with: excessive cross- axis coupling; inadequate dynamic stability; and unacceptable stick force gradients. Interestingly, the Cooper-Harper pilot ratings from many heli- copter handling quality studies (for example, Refs. 1 and 2) indicate that rotary wing pilots are AIrspeed ikt) willing to accept much less than their fixed wing This is shown in Fig. 2 where pilot counterparts.

Fig. 1 Illustration that conventional unaugmented ratings of 2 to 3-l/2 are found well into the helicopters fall well below fixed wing Level 2 region defined for pitch control in MIL-F- standards even for a failed SAS 8785C. (Level 2 corresponds to pilot ratings of (data from Ref. 3) 3-l/2 to 6-l/2 in MIL-F-8785C.) This is felt to occur for two reasons: 1) helicopter pilots are trained to cope with, and expect as "normal," severe instabilities and cross-axis coupling; and 2) the tasks used in the evaluations were not sufficiently demanding.

Consideration of Handling Note : to meet Level2 ffying Ouality Evaluation Tasks quo/i,ies in hfd - f -878X, I>0 In recent years the task used in experiments 0 to obtain handling quality pilot ratings has been 0 found to have a profound effect on the results.

For example, in the landing approach experiments of 0 0 0 I- 1 I I I 0, Ref. 4 the pilots were required to touch down at a -.3 :2 :I 0 .I .2 .3 4 .5 “Phugoid” Damping Ratio, 5 Fig. 2 Cooper-Harper pilot ratings vs. damping ratio in hover; w < 0.5 rad/sec (data from Ref. 1) tified in terms of a scale as shown in Fig. 3a.

evaluation pilot revised his rating downward con- Certain specific closed-loop considerations, which siderably. The point here is that only by insist- were considered in formulating the scale, are ing on a precision task was the experimenter able summarized below and by the generic closed-loop to expose deficient handling qualities that would have otherwise gone unnoticed. In using existing structure in Fig. 3b.

1) A requirement for closure of the attitude data to develop boundaries for the helicopter loop implies VKC conditions and must prevail for handling qualities specification, we must criti- Some suggested evaluation adequate control.

cally evaluate the task.

2) If the equivalent system dynamics require factors might be: 1) Does the task require the same precision closure of position and position rate, but not minimum set of operating conditions as required by operational missions? attitude, a quantified as OVC = 3 is defined.

2) Does the task require the same degree of 3) OVC = 4 quantifies the operating condition aggressive maneuvering as the proposed operational missions? where velocity and attitude cues are not available; 3) Are the tasks well defined, or does the that is, only the outer loop in Fig. 3b can be closed by the pilot.

task encompass a series of subtasks such as an 4) OVC = 5 indicates that no outside visual entire approach, hover, and vertical descent? If can we identify what subtask cues are available.

the latter is true, has the most impact?

Pilot workload can also be reduced via im- 4) Are the data being used as a compromise proved displays. Recent work in the control/ because no better data are available?

display tradeoff area includes the Calspan X-22 5) Are the atmospheric disturbances of low flight tests (Ref. 8) and the CH-46 variable- enough frequency and large enough magnitude to stability helicopter (Ref. 9).

displace the aircraft from its path?

Based on the above considerations, the re- 6) Are the available outside visual cues quired level of augmentation and cockpit displays consistent with the proposed mission?

Unfortunately, these factors may well elimi- can be related to the visibility levels associated nate most existing data. The last factor was found with the missions defined for the helicopter. An initial attempt was made to establish a format for to be especially important for low speed and hover specifying the augmentation and displays required in Refs. 5 and 6 and is briefly reviewed in the following section. for various levels of outside visual cues in Refs. 1. 4, and 5 and is repeated in Table 1 for _ .

convenience.

Effect of Outside Visual Cues on Required Level of Augmentation and Display Most of the available data for low-speed and criteria have been obtained with hover handling good visual outside references and with no require- ment for unattended operation. The real-life existence of secondary tasks, and intermittent to total loss of visual references, places increased demands on the pilot -- an effect which is not For example, pilot discernible from such data.

ratings for an unaugmented helicopter (Ref. 2) and a highly augmented translational rate command (TRC) system (Ref. 7) all fall within the acceptable region (pilot rating better than 3.5). This result is a consequence of experimental scenarios that tend to be tailored toward the systems being inves- al Ouonfification of Oufside Visual Cues /OVCl That is, with pure rate systems the tigated.

scenario is usually benign, thereby usually allow- ing intense, full-time attention; whereas with a translational rate command system the task tends to Pilot The most critical contributor Posttim be more demanding. Perceived Position cues, x to be the to the total pilot workload appears quality of out-the-window cues for detecting air- craft attitudes, and, to a lesser extent, position and velocity. Currently, these cues are cate- gorized in a very gross way by designating the A more discrimi- environment as either VMC or INC.

nating approach is to classify visibility in terms Requires OVC 5 2 (VMC) of the detailed attitude and position cues avail- able during the experiment (or proposed mission), Requires OVC 5 3 and to associate handling qualities requirements Reauires OVC C 4 with these finer-grained classifications.

61 Required Outside Visual Cues for Confro/ The need for certain specific outside visual cues has been inferred from closed-loop considera- tions. These OVC levels have been logically quan- Fig 3 Development of outside visual cue scale Table 1. Augmentation and displays required for various levels of outside visual cues Pilot display Integrated display- MIL-F-8785C Llight director plus flying Mechanical aircraft velocity Augmentation quality level flight director information Rate Level 1 2 3 Level 2 5 Rate command/ Level 1 3 3 attitude hold Level 2 5 5 Attitude Level 1 3 (response feedback) Level 2 5 Attitude Level 1 4 (model following) Level 2 5 Level 1 5 Translational rate with attitude Level 2 5 Translational rate Level 1 5 with direct force control Level 2 5 rotary wing aircraft are reviewed in the following SAS Failures paragraphs.

The concept of "Levels" is used in MIL-F-878X to specify the allowable degradation in handling Lower-Order Equivalent Systems qualities in the presence of failures.

The speci- fication of Level 2 and 3 handling qualities will The basic intent of lower-order equivalent tend to be more critical in rotary wing aircraft in systems is to define a very high-order system in terms of driving the cost and complexity of the terms of a few variables that describe the funda- SAS. characteristics important to the This is a result of the relatively poor mental response handling qualities of the unaugmented helicopter (see Ref. IO). This can be done in the pilot and hence the large change in dynamics before and time domain or in the frequency domain, although after a failure of the SAS. This is illustrated in all work done to date has been in the frequency Fig. 4, which shows a dramatic shift in the charac- domain. Equivalent systems are a viable concept for teristic modes after a SAS failure in the CH-53D. defining Level 1 flying qualities for helicopters.

Clearly, the specification of Level 2 handling qualities that are better than most unaugmented helicopters would have significant implications on SAS complexity and cost.

On c-l Potential Fixed Wing Criteria Applicable to Rotary Wing Aircraft The mission requirements for rotary wing air- craft have become increasingly severe to the point where marginal handling qualities can no longer be tolerated. In most cases satisfactory inherent stability and coupling cannot be obtained without some level of stability augmentation.

Indeed, many modern helicopters employ a stability augmentation SAS system. It is therefore not unreasonable to expect Off the same quality of response (to control inputs and turbulence) in helicopters that is currently enjoyed by fixed wing pilots. In fact, the rapid and precise maneuvering required in some NOE mis- sions may make it necessary to impose _more strin- gent requirements than are necessary for fixed wing aircraft.

Fig 4 Effect of SAS failure on key The applicability of some requirements cur- response modes of CR-53D rently proposed for the fixed wing MU-Standard to (Data from Ref. 3) From a pilot's point of view, a high-bandwidth However, the complexity of the responses of unaug- response would be described as "crisp" or perhaps due to inter-axis coupling, mented helicopters, "rapid and well damped." Typical commentary for a makes it unlikely that useful equivalent system low-bandwidth response might be "sluggish response forms of sufficient generality can be defined for to control input" or "tends to wallow." There the Level 2 and 3 boundaries.

is a long history of correlating such commentary with basic aircraft stability derivatives and/ Bandwidth Criterion mete s made up of such derivatives (e.g., ; 'e'&Mqr- M, , etc.). The term bandwidth comes The bandwidth criterion was developed origi- m%e naturally into play when feedbacks and cross- nally for fixed wing aircraft with direct force feeds are combined to produce aircraft responses control. Because of the almost infinite variety that are unconventional in that the classical modes due to inter-axis of responses that can occur are no longer appropriate definitions.

it was difficult to define a lower-order coupling, equivalent system form for aircraft with direct The advantage of this approach is that it does In looking for an alternative force control.

not assume a particular form of response.

Hence it solution it was hypothesized that the coupling may be suited for helicopters, where coupling tends itself was incidental, and mattered only to the to mask the classical response forms.

The defi- extent that it interfered with the pilot's ability ciency of the bandwidth criterion in its present to adequately perform tight closed-loop tracking.

form is that it does not directly account for the This of course is directly related to the band- "The band- pilot's ability to supply crossfeeds to counteract width, which was defined in Ref. 11: coupling. It seems intuitively obvious that re- width of the specified response to a particular sponses requiring only a control input is defined as the lowest frequency simple crossfeed (such as pure gain) would be more acceptable than those for which the (open-loop) phase margin is at least requiring complex shaping.

45 deg and the gain margin is at least 6 dB." (See This concept was inves- tigated in Refs. 13 and 14 for the turn coordina- Fig. 5 for a graphical description.)

tion problem in fixed wing aircraft and is reviewed briefly in the following section.

The Ref. 11 variable-stability in-flight simu- lation results indicated that the Bandwidth Hypo- Inter-Axis Coupling thesis was indeed valid, i.e., the coupling itself mattered only to the extent that it affected band- Inter-axis coupling is well recognized as one width. These results were extended to pitch atti- of the most severe handling quality problems with tude control in Ref. 12.

unaugmented rotary wing aircraft.

While fixed wing aircraft tend to be much less affected by such coupling, a significant amount of yaw response to roll control inputs is not uncommon at high angles ODen Loop Transfer Function of attack. In such cases the pilot must use rudder coordinated with aileron inputs to eliminate the (S + I/T)e-= -= undesirable heading excursions that occur.

It was 6 se+ 2Q.lJs + 2 hypothesized in Ref.

14 that the pilot opinion of roll-yaw coupling would be directly related to the magnitude and shaping of the rudder control re- quired. Such an approach is expected to be directly applicable to inter-axis the coupling characteristics of helicopters. Because of its possible direct application to helicopter coupling, the results of Ref.

13 are briefly described below.

While the use of "coordinated" aileron and rudder is accepted as common piloting technique, a w (rod /set) + I quantitative measure of what exactly is acceptable or desirable is not known.

The purpose of this study was to provide a quantitative measure of the aileron-rudder sequencing required to eliminate roll-yaw coupling and thereby achieve coordinated turns, and to correlate this with pilot opinion ratings from available data. To achieve this end Ref. 13 considered the aileron-to-rudder crossfeed i necessary to exactly cancel the inter-axis cou- ’ c&=45” pling.

This idealized crossfeed provides a measure -l80”-- of pilot acceptability of heading control because -200 .t i 4 it is indicative of: the complexity of the rudder activity necessary to achieve perfectly coordinated \ turns; and the heading excursions that occur when the pilot does not use rudder.

Note that these considerations equally well to the known apply coupling between pedal, Fig. 5 Effect of using gain and phase margins power, cyclic and collec- tive in an unaugmented helicopter.

to define bandwidth Table 2.

Parameters defining the Two parameters are defined in Ref. 13: u, aileron-rudder crossfeed which defines the shaping of the rudder crossfeed; and Ns /L's , which defines the magnitude. The freque%y r%ponse characteristics of the aileron- to-rudder shaping as a function of the sign of u Analytical Pilot-centered are shown in Fig. 6 in terms of literal expressions Parameter function function for the Bode asymptotes. These asymptotes indicate that the magnitude of the rudder required to coor- IJ Defines shape Determines complexity of dinate is a function of Ns /N& at all frequen- rudder activity necessary cies and that the shaping oy thd?udder response is Of 'CF for ideally coordinated determined by u.

These parameters are summarized turns; also defines phas- in terms of their analytical and pilot-centered ing of heading response functions in Table 2.

when rudder is not used.

The details of the criterion are presented in Defines magni- Determines magnitude of Refs. 13 and 14. The criterion boundaries and the required and/or high- tude of YCF data used to support these boundaries are given in frequency yawing induced Fig. 7. It is interesting to note that the ideal by aileron inputs.

crossfeed was not a pure gain (u = 0). Actually, a little proverse yaw (u = -1) is seen to be desir- able. Similar results could be expected with helicopters, i.e., the coupling can actually be favorable.

Conclusions A great deal of the experience gained in developing handling quality criteria for fixed wing aircraft is directly applicable to rotary wing air- craft as well. In this we have reviewed a par== for/d > 0 Lag Lead Compensation

I Nk,c

/ L’s,,~TR (I+‘) i

fLwp< 0 Lead Lag Compensation = normalized rudder control %C normalized aileron control Tat = Fig. 7 Pilot rating correlation with crossfeed parameters Fig. 6 Asymptotes of aileron-rudder crossfeed Hoh, R. H., and Ashkenas, I. "Handling summar- 6. L*, few areas that seem particularly salient.

izing, these are: Quality and Display Requirements for Low Speed and Hover in Reduced Flight Visibility," Jour- 1) The piloting task and environment are nal of the American Helicopter Society, 26, overriding considerations in developing and using (l), Jan. 1981.

handling quality criterion boundaries.

2) Helicopter pilots have historically been Bryant, W. B., Cattel, J. C., et al., "VTOL willing to put up with considerably more degraded 7.

Advanced Flight Control System Studies for handling qualities than have fixed wing pilots.

The increasing severity of helicopter missions is All-Weather Flight. Vol. I: Task I Report," reversing this trend. USAAMRDL-TR-75-13A, July 1975.

3) Outside visual cues and cockpit displays Lebacqs, J. V., and Aiken, E.W., "A Flight must be considered when structuring a helicopter 8.

handling quality specification. Investigation of Control, Display, and Gui- 4) The poor inherent handling qualities of dance Requirements for Decelerating Descending rotary wing aircraft make SAS failures more criti- VTOL Instrument Transitions Using the X-22A cal than for fixed wing aircraft. Attempting to Variable Stability Aircraft. Vol. I: Techni- cal Discussion and Results," Calspan Corp., impose fixed wing requirements for Levels 2 and 3 Buffalo, NY, Rept. AK-5336-F-1, Sept. 1975.

is probably not practical in terms of cost and complexity.

5) Many handling qualities criteria developed 9. Niessen, F. R., Kelly, J. R., Garren, J. F., for fixed wing aircraft should be directly appli- et al., "The Effect of Variations in Controls cable to helicopters with appropriate revisions in and Displays on Helicopter Instrument Approach the numerical limits and boundaries. Capability," NASA TN D-8385, Feb. 1977.

References 10. Hodgkinson, J., and LaManna, W. .I., "Equiva- lent System Approaches to Handling Qualities 1. Hoh, R. A., and Ashkenas, I. L., 'Development Analysis and Design Problems of Augmented of VIOL Flying Qualities for Low Speed and Aircraft," AIAA Atmospheric Flight Mechanics Hover," NADC-77052-30, Naval Air Development Conf., Hollywood, FL, 8-10 Aug. 1977.

Center, Warminster, PA, Dec. 1979.

11.

Hoh, R. H., Myers, T. T., Ashkenas, I. L.

2. Seckel, E., Traybar, .I. J., and Miller, G. E., Ringland, R. F., and Craig, S. J., "Develop- "Longitudinal Handling Qualities for Hover- ment of Handling Quality Criteria for Aircraft Princeton Department of ing," University, with Independent Control of Six Degrees of Aeronautical Engineering, Rept. 594, Dec. Freedom," AFWAL-TR-81-3027, Air Force Wright 1961. Aeronautical Laboratories, Wright-Patterson AFB, OH, Apr. 1981.

3. Heffley, R. K., Jewell, W. F., Lehman, .I. M., "A Compilation and 12. Hoh, R. H., Mitchell, D. G., and Hodgkinson, and Van Winkle, R. A., Analysis of Helicopter Handling Qualities "Bandwidth -- A Criterion for Highly Aug- J..

Data. Volume One: Data Compilation," NASA CR- mented Airplanes," AIAA-81-1890, AIAA Atmos- 3144, Aug. 1979.

pheric Flight Mechanics Conference, Albuquer- que, NM, 19-21 Aug. 1981.

4. Smith, R. E., "Effects of Control System Dyna- mics on Fighter Approach and Landing Longitu- 13. Ashkenas, I. L., Hoh, R. H., and Craig, S. J., dinal Flying Qualities (Volume I)," AFFDL-TR- "Recommended Revisions to Selected Portions of 78-122, Air Force Flight Dynamics Laboratory, MIL-F-8785B(ASG) and Background Data," AFFDL- Wright-Patterson AFB, OH, Mar. 1978. TR-73-76, Air Force Flight Dynamics Labora- tory, Wright-Patterson AFB, OH, Aug. 1973.

5. Hoh, R. H., and Ashkenas, I. L., "Effect of Reduced Visibility on VTOL Handling Quality 14. Hoh, R. H., and Ashkenas, I. "Handling L-9 and Display Requirements," Journal of Guidance Quality Criterion for Heading Control," Jour- and Control, 4, (2). Mar.-Apr. 1981. nal of Aircraft, 14, (2), Feb. 1977.

HELICOPTER SIMULATION TECHNOLOGY: AN AMES RESEARCH CENTER PERSPECTIVE Richard S. Bray Ames Research Center, NASA, Moffett Field, California Abstract It is the objective of this paper to review recent helicopter simulation experience at Ames Helicopter handling qualities have been the Research Center for evidence relating fidelity of subject of many simulator programs at Ames Research motion and visual cueing to subjective fidelity and Center over the past decade. The earlier experi- confidence in research results. The scope of this ences, in fixed-cockpit simulators, demonstrated experience in terms of objectives, facilities, and the basic difficulties of simulating the inherently simulated flight tasks is briefly described.

complex control tasks of helicopter flight to the Approaches to optimization of the utilization of level of subjective fidelity required for confident unique cockpit motion- and visual-simulation capa- evaluation.

It became recognized that deprivations bilities are discussed, and several experiences in visual and motion cueing were probably major that offer hints regarding the role of vertical factors in the problem. More recent simulations acceleration in hovering tasks are described. Con- have utilized large-amplitude cockpit motion sys- cluding remarks address the need for a firmer tems, and efforts have been made to optimize the understanding of the effects of cueing deprivations effectiveness of the visual simulations. This and suggest a program of directed research on the paper reviews the total experience for evidence subject.

regarding the levels of motion- and visual-cueing fidelity required for handling-qualities research Scope of Research Activities in ground-based simulators.

Positive contributions of cockpit motion were identified, but much remains Objectives and Tasks to be learned regarding the sensitivities of indi- vidual control modes to cueing attenuation. A Several papers presented at this meeting dis- firmer understanding of the pilot's utilization of cuss recent helicopter research conducted in Ames visual and motion cues is the key to more efficient simulators. A series of handling-qualities use of simulation in helicopter control-systems studies, in the context of a "nap-of-the-Earth" research. flight task, is discussed in Ref. 1. That paper touches on the relationships of simulation facili- Introduction ties and procedures to the'interpretation of results. The results of tests to guide the devel- Flight-simulation technology is especially opment of helicopter IMC flight certification challenged by the helicopter. Mathematical model- criteria are presented in Ref. 2. Control systems ing and verification procedures are difficult. and guidance displays were evaluated in an ILS-like Flight modes include those often characterized by approach that included deceleration to hover on low stability and cross-axis control coupling, instruments. The study of Ref. 3 closely examined conditions that tend to produce unrealistically variations in enqine and control-system response in high workloads in simulation. A sense of realism, critical height-control maneuvers.- This simulation better termed subjective fidelity, in the simulated reauired optimized visual cues and, like the study of'Ref. 2,'anticipated benefits from the utiliza- flight task is essential for its use in research; and, depending on the research objective, some tion of a simulator with a large-amplitude cockpit Motion- and visual-cueing consider- moderate to high level of objective, or engineer- motion system.

ing, similarity to the flight task is required to ations in that study will be expanded upon in a create that realism. There is no basic obstacle later section of this paper.

to the attainment of objective fidelity in the simulation of an aircraft except in the areas of Although these are typical of helicopter simu- cockpit motion and the outside visual scene. At lation studies beino conducted at Ames, others must best, simulation can provide only part of the cues be mentioned to indicate the broad scope of objec- available in the aircraft. The effects of these The XV-15 Tilt Rotor aircraft has tives oursued.

deprivations, their individual contributions to the been the subject of simulation exercises for the 9 diminution of subjective fidelity, is not clearly years since concept proposals were evaluated. This understood; they have not been subjected to ade- program, conducted in support of the vehicle devel- quate study. In the absence of better information, opment and flight tests, used a variety of Ames research simulations are configured and used in the facilities; it is documented in Ref. 4. Another manner that experience indicates to be probably example of support of a research aircraft is the effective. recent simulator studies defining optimum operating procedures for the winged, or "compound" version of the Rotor Systems Research Aircraft (RSRA). A dedicated simulator cab is being constructed for continued support of the two flight vehicles.

.99 Individual simulator exercises in these flight- than +0.5 g, but are generally satisfactory for support programs might have one or several specific helicopter simulation. The larqe transport-type objectives, but collectively they involve tasks cockpit has two pilot stations,-and is equipped with hvdraulic control loaders. visual simulation covering the operational envelope of the aircraft.

They represent a specific challenge and opportu- TV mon;tors, and head-up display equipment. As in nity. The requirements for fidelity are severe, all simulators (except the several "dedicated cock- but since the aircraft exists in a flight-test con- pit" simulators), this cab is reconfigured for each figuration, the opportunities for verification are new simulation. Over the past decade, this facil- Reference 4 reports several illuminat- ity has been used in simulation of a wide range of excellent.

ing exercises comparing simulation responses with Currently, helicopter simulations make aircraft.

those of the Tilt Rotor aircraft, and relating them up about 25% of its operation.

to the pilot's subjective impressions.

The newest facility, the Vertical Motion Simu- Another helicopter research effort in its lator (VMS). is shown in Fiq. 3. The oresent cab early stages is one that joins VTOL studies in is of the same specifications as the FSAA, but is addressing the special guidance and control prob- driven in angular motion by a small, six-actuator lems of approachins and landinq on a destroyer in hydraulic system. This is mounted on a laterally very adverse weather, perhaps among the most dif- driven carriage with 13 m of travel atop a beam ficult tasks to simulate adeauatelv. This task which can be moved vertically in a 19 m envelope.

will also be the subject of further discussion.

These latter two drives are electric, and are capa- ble of nearly l-g accelerations. A second hori- A number of helicopter simulations have been zontal motion component is not provided; however, used in terminal-area traffic control studies, and the cab can be rotated to substitute fore-and-aft a current program is assessing airborne radar con- motion for lateral motion. A later section of this cepts of guidance to offshore oil platforms. These paper will discuss the capabilities of those large are IMC flight tasks with very modest maneuvering motion systems to reproduce the motion cues of requirements.

maneuvering flight.

This overview has not touched on all of the Visual Simulation Systems. Ames operates two Ames helicopter simulation activity, but perhaps it Redifon TV model-board visual scene generators.

has described those efforts in which the quality of These systems can provide a 34" by 48" visual field visual or motion cueing, or the effects of their on a 525-line color televison raster format. The absence, should have been a consideration. model-boards have accumulated a variety of features modeled at scales from 1:300 to 1:1200. Half of Facilities one of the model-boards is devoted to hilly terrain appropriate for helicopter NOE flight tasks. A Cockpit and Motion Systems. This discussion variety of aviation ship models, mechanized to pro- of facilities is limited to those factors defining vide deck motion, are provided. An oil drilling the pilot's immediate environment: displays, con- platform is also available. A recent acquisition trols, and, most importantly, cockpit motion- and at Ames is a Singer-Link computer-generated-image visual-cueing systems.

The-simulator cab illus- (CGI) visual simulation system. This device can trated on Fiq. 1. desiqnated Chair 6. is oooular produce four independent 34" by 48" visual fields with experimenters who-are in the preliminary on 1024-line raster formats. The scenes, which are phases of a research program, or who are studying in color, can present simulations of day, dusk, and navigation or display questions unrelated to the night conditions. Scenes presently available higher frequency dynamics of the helicopter. It is include an airfield and surrounds, a destroyer with a box on wheels that can be located handily in the helicopter landing facilities, and a small carrier.

computer laboratory, but like most Ames simula- A new simulator cab, shown in Fiq. 4, is confiqured tions, it is equipped with a collimated TV monitor for a helicopter pilot's station-and is equipped for displaying a scene generated by a model-board with four collimated CRT "windows" for display of system. It also has provisions for a collimated the CGI scenes. It has operated as a fixed-cockpit head-up display. To avoid the complications of simulator. Within the year, this cab, which is the hydraulics, control loaders are simple electro- first of a series of "interchanqeable" cabs, will mechanical devices.

Another fixed-cockpit simu- be installed on the VMS motion system to combine lator, which utilizes a salvaged UH-1 cab and the increased viewing area with the large-motion control hardware, is used primarily in the devel- capability.

opment of software for a helicopter avionics flight program. Cueing Effectiveness The Flight Simulator for Advanced Aircraft The preceding descriptions of the motion and (FSAA), illustrated in Fig. 2, features a lateral visual systems fall short of defininq the extent to motion envelope of 30 m, together with 3 m of which those systems can reproduce the cues sensed vertical travel and 2.5 m of fore-and-aft movement.

bv the nilot in flioht. This definition can be Three independent drives provide generous ampli- obtained only through examination of the specific tudes of anqular motion. All drives are electric.

simulated flight task -- the accelerations of Linear acceleration capabilities are modest, less flight compared with the limited spectrum available in the simulator -- and the visual information of the simulated helicopter tasks involve less vital to the task in flight compared with what is lateral maneuvering than provided for in this case, available in the simulator. The followinq oara- although one series of experiments at Ames, conduc- graphs initiate this process by establishing gen- ted in a fixed-cockpit simulator,, utilized a high- eralized maneuver-cueing relationships for the VMS speed NOE task that included 60"'to 60" roll motion system and the visual simulation devices. reversals.

The high-frequency dynamic response capabilities of these systems will not be addressed here. The gen- Body-axis lateral accelerations are produced eral topic of allowable lags in motion and visual essentially undistorted, the short-term components systems is well covered in the recent literature. provided by the lateral drive system and the low- frequency components generated by easing a "tilt" Motion-Cueing Capabilities component into the cockpit roll attitude.

The motion commands to the VMS drives are com- Vertical Motion. Two response curves are posed of 1) the computed motions of the modeled shown for vertical motion. They describe the rela- aircraft subjected to second-order high-pass fil- tive capabilities of the VMS and the FSAA to repro- tering and possibly attenuated; and 2) discrete duce the vertical motions seen in a flight task limiting logic that arrests the motion at the involving maximum lower-frequency vertical accel- excursion limits, if the primary mode of confine- erations of about to.3 g. Helicopter low-speed ment is overpowered. The characteristic fre- tasks and hovering tasks usually fall in this cate- quencies of these "washout" filters are directly It can be seen that the VMS, with a washout gory.

related to the maximum amplitudes of the lower- frequency of 0.4 rad/sec, provides an unattenuated, frequency accelerations anticipated in the simu- effective (less than 30" phase error) band-pass lated maneuvers, the degree to which direct between 1 and 6 rad/sec. The relatively limited attenuation of the accelerations is acceptable or excursion capability of the FSAA defines a washout necessary, and the excursion envelope of the frequency of 1.4 rad/sec and an effective band-pass related motion-system mode. The roll and pitch between 3.0 and 6 rad/sec. Such increases in ver- \ modes are not usually constrained by their own tical washout frequency have not produced the angular excursion limits, but rather by the con- strong conscious motion-visual disparity disturb- sequences of logic that attempts to minimize spuri- ance seen in roll.

ous longitudinal or lateral accelerations owing to cockpit tilting. Thus, roll excursions are iim- Visual-Cueing Capabilities ited by the capability of the lateral drive to retain the specific force vector in its proper ori- The comparison of visual cues provided in entation as the cab is rolled. Gains and washout simulation with those present in flight is not the frequencies typical of those used in the VMS in straightforward process demonstrated for motion helicopter simulation are indicated in the Bode A visual scene has many measures, and the cues.

diagrams of Fig. 5. This diagram describes the significance of each to the pilot's perception of "band-pass" of the system -- those portions of the his position and velocity remains ill-defined.

maneuvering spectrum that can be reproduced accu- However, some of the obvious capabilities and rately. It also illustrates that motions at fre- limitations of the model-board and CGI systems can quencies near the washout frequency will be highly be noted.

distorted in phase. The roll-off in dynamic response shown at the high frequencies is typical Field-of-View. In Fig. 6, the extent of four of the drive system, not the motion constraint visual fields, as might be generated by the CGI logic. system, is compared with the pilot's outside visual field in a typical helicopter. The model-board Lateral Motion. The curves labeled "roll" in systems are capable of supplying only the single Fig. 5 represent the combined mode of cockpit roll forward scene. The fourfold increase in field and lateral motion mentioned earlier. The indi- offered by the CGI system still falls short of cated gain of 0.5 and the washout frequency of 0.7 matching the flight condition, though it adds side- rad/sec are appropriate for the simulation of ward and downward scenes that are assumed to be of lateral maneuvers'involving angles-of-bank seldom prime importance for position and velocity cues in which in this case would result precision hovering. Also, the argument is made qreater than 30", in a lateral excursion of about 5 m. Accommodation that the more generous lateral field improves the pilot's perception of rates-of-change of aircraft of higher-amplitude lateral maneuvering would require more attenuation or an increase of the attitude.

washout characteristic frequency. Experience has quality of the Scene. Visual systems are most indicated that for active iateral maneuvering, the former is the preferred option. In simulated severely tested in simulations of flight in proxi- visual flight tasks, motion-vision phase dis- mity to-the terrain or structures, exactly the parities can be consciously disturbing, as washout tasks usually chosen for critical helicopter There is no inherent frequencies are increased above 0.7 rad/sec. It is control-systems evaluations.

seen in this case that motions in the frequency limit to the extent that real-world textures and range from 0.7 to 1.5 rad/sec are transmitted with detail can be reproduced on the model-board; how- large leading-phase distortions. Fortunately, many ever, models are only seldom detailed to match the because of the loss of virtually all visual infor- limited resolution capability of the camera optics mation at large nose-up pitch attitudes. Precision and video system. Even without the deliberate hover is made difficult because of the lack of detail, objects on the model-board usually possess By comparison, translational velocity cues that normally are some level-of resolvable texture.

obtained from sidewarh and downward views. The the CGI svstem, because of the limited number of extent to which these factors limit the validity of lines it can draw, is severely limited in presenta- the simulation varies'with the maneuvers of the- tion of detail. The CGI scene, with its compara- simulated flight task. Validity also depends to tively high resolution, is excellent at medium to some degree on the pilot's sense of subjective large distances but tends to lose its realism as fidelity. In the fixed-cockpit simulations, the terrain is approached. The concentration of the system's image-producing capacity on a limited visual-scene limitations were often assessed as a In the more scene feature, as in a ship model, offers at least major cause of performance difficulty.

recent programs using cockpit motion, these criti- a partial solution to the difficulty. More will be cisms have been less strident. The visual con- said about simulated scene content in discussions straints on the task are recognized, but perform- of several specific simulation applications.

ance difficulties within that constrained task are Observations and Discussion not so often attributed to a lack of visual cues.

This latter assessment more closely agrees with the results of flight tests5ys in which limitations of Validity of Simulation the pilot's field of view affected performance to a lesser degree than anticipated in view of early The foregoing has considered a variety of simulator experiences.

research objectives, and the varying limitations of From the facilities used in the investigations.

this experience, observations can be made relating Cueing Optimization obiective. cueinq caoabilities, and validity of the The experimenter has the opportunity, and the simulation. As defined here, validity is the obligation, to shape the simulated flight tasks to effectiveness of the simulation as a means of .'

achieving the research objective, and thus does not take best advantage of motion- and visual-cueing imply a specified level of subjective or objective capabilities in the pursuit of his research objec- fidelity in the vehicle simulation itself. tive. A standardized procedure is not offered here; instead, the simulation of Ref. 3, which is Fixed-Cockpit Simulations. In the earlier considered a particularly effective example of is discussed in detail. The experience at Ames, handling-qualities issues were cueing optimization, objective of those experiments was the evaluation addressed in fixed-base simulation, with limited results. The simulations of light, agile vehicles of variations in height-control parameters. The drew strong adverse comment from the pilots who critical maneuvers were determined to be climbs experienced exaggerated, unrealistic workloads in over obstacles at low forward speeds while minimiz- conventional helicopter maneuvers. Pilots required ing exposure time above the obstacles, deceleration considerable practice to reach a stable level of to hover under cover of obstacles, and a "bob-up" to a momentary surveillance position above the performance, and performance differences between pilots tended to be large. Subjectively, the obstacles before a return to hover. A particular pilots considered the aircraft model suspect, and arrangement of simulated obstacles, identifying a judged the limitations of the model-board visual course on the model-board, minimized the signifi- system to be another prime source of their diffi- cance of visual limitations while defining flight culties. The experimenters, recognizing that maneuvers that optimized the cueing potential of motion-cue deprivation might be a major part of the the VMS.

problem, began to seek the use of the FSAA and the Visual Simulation. The pilot's view of the VMS for their stability and control studies. How- ever, lack of motion did not appear to present experimental course, as seen at the instant of serious problems to all experimenters. Simulations passing over one of the obstacles, is shown in of larger stabilized vehicles, used in studies of Fig. 7. Obstacles are laid out between two rows of navigation and display systems, were generally trees that define the straight-line course. This accepted by the pilots. The summary observation is avenue is terminated in the distance by a crossing made that if the character and workload of the row of trees. The obstacles were arbitrary in vehicle control task is not severely distorted, and form, and made no contribution to a sense of real- if the pilot is not asked to pass critical judg- ism in the scene. Rows of trees might have been ment on the vehicle's dynamic responses, the fixed- more aesthetically pleasing. Models of ground cockpit simulation appears to be adequate. vehicles were included to help establish a sense of scale, and the level surface between obstacles Adequacy of the Visual System. The single included scattered shrubbery to aid in the sense of forward window provided by the model-board visual The avenue of high trees proximity to the ground.

system places a limit on the fidelity of heli- did more than identify a course; it served to opti- copter simulation in visual tasks. In turninq mize visual perception of height and height-rate flight near the terrain, the inability to see-what from the limited forward field-of-view. This cue- lies ahead in the predicted path is disconcerting ing augmentation was vitally important during the and unrealistic. Quick stops are almost prohibited deceleration to hover. The pilot's view during development of more sophisticated computer- this maneuver is seen in Fig. 8. Even though the pitched-up attitude severely constrains the view generated scenes, but little research is under way of the surface, the trees offer an effective set of to address the question of how to use current capabilities most effectively.

references for the perception of vehicle veloci- ties. Some of this effectiveness is attributed to the fact that the trees did not completely obscure Benefits of Improved Cueing -- the more distant scene.

It has been noted that in restricted viewing fields, in which close objects What benefits are seen as a result of such or surfaces completely predominate, the visual efforts to increase the cueing fidelity of simula- cues of angular and linear motions can become tion? Like the cues themselves, the benefits tend confused. to be subtle, thouqh, as will be seen shortly, startling effects can be demonstrated if the appro- Even with the motion cueing At several points in the task, field-of-view priate tests are made.

provided by the FSAA or the VMS, there remain many limitations were especially noted. It was diffi- reminders to the pilot that he is operating a simu- cult for the pilot to assess his clearance dis- The motion system contributes its own tance when passing over the obstacles; and during lation.

reminders if the motion logic is improperly condi- his bob-up maneuver, it was very important to retain sight of some tree tops over the nose in tioned for the simulated flight task. The intro- duction of large-amplitude cockpit motion to heli- order to maintain position reference.

copter simulations does lead to these general Cockpit Motion. Because no lateral maneuvers observations: 1) the pilot's initial assessment of were required other than to maintain position subjective fidelity is somewhat improved; 2) his or time to a performance plateau between the rows of trees, the lateral motion con- "transition time," straints of the VMS were minimized. with an unfamiliar vehicle and task is shortened; Altitudes in the task did not exceed 80 ft, a height that is 3) maneuver amplitudes and control "style" compare only slightly greater than the vertical excursion more favorably with those of flight; 4) less vari- capability of the VMS. Thus, vertical accelera- ation in performance and assessment is seen across tions were reproduced with unusual fidelity for a group of pilots; and 5) ratings and commentary ground-based flight simulation. The vertical regarding handling-qualities issues appear to be acceleration band-pass noted for the VMS in Fig. 5 offered with greater ease and confidence.

was realized; and moreover, because the task was so limited in altitude, vertical accelerations to the Two peripheral observations are worth noting: limits of the machine could be utilized. there is noticeably less criticism of the visual system's limitations, and comments regarding motion Another Optimization Opportunity. Recent are limited almost exclusively to those inspired by limited experience with the four-window CGI display anomalies, such as limit encounters, or by audible Again, the reader is suggests that the radically increased field of ;iew noise from the motion drives.

does not relieve the experimenter of the need to reminded that even with excellent motion cues, the seek optimization of the visual information. If, oilot is dealinq with a simulated fliqht task; he for example, aircraft systems are to be evaluated will have reseriations regarding the fidelity-and for their adequacy in landing on a moving ship validity of the simulation until he has accommo- deck, the visual simulation must approach the real- dated to the remaining artificialities, especially world scene in the orovision of attitude and Dosi- those of the visual simulation.

tion cues. As mentioned earlier, the four-window CGI system falls short of Dresentina the in-fliaht Some effects of improving the visual cues are fieldlof-view. In the example discussed earlier, a In Fig. 9 are shown-the four " more obvious.

scenes, as presently configured, representing the particular flight task was enabled by configuring pilot's view near touchdown. His only significant model-board elements to optimize the information in view of the deck is in the lower right window, and the single forward field. The increased field of this view is notably separated from the other view offered by the CGI system enables a simulated visual information sources. What we see is a landing on a ship or a drilling platform. These problem of limited (or perhaps non-optimally ori- additions are consciously appreciated by the ented) field of view compounded by the geometry of pilots; they see an increased validity of the the deck and superstructure. simulated mission, but contributions to a sense of subjective vehicle fidelity are unclear.

The argument is made that neither the window Some Observations Regarding Vertical-Motion Cueing placement nor the simulated ship geometry should ---_-- be constrained by real-world measures, if as a result of either constraint the task is made The most uniaue asoect of the Ames simulation experience has be.& the'availability of vertical unrealistically difficult. Window placement should Thouqh this facility has been be optimized and the scene elements desianed to motion in the VMS.

operational for nearly P-years, no formalized provide attitude and position cues of maiimum effectiveness. Unfortunately, there is little in investigation has been conducted in an attempt to the literature to guide the experimenter in this identify the contribution of the vertical motion cues to the validity of simulation. Other than quest. Great effort is being expended on the helicopter studies such as those discussed earlier, is somewhat pessimistic. In the brief "back-to- the use of the facilities has been limited to Space back" tests in the VMS, the pilot had no oppor- Shuttle control-systems verification studies. The tunity to accommodate to the altered visual-motion It is probable that if the entire oilots recoqnized the Shuttle simulations to be of relationship.

unique quality, particularly in their examinations program had been conducted with reduced vertical of the Ditch-control modes for PI0 tendencies, but motion, ratings would have been degraded less than perhaps'their most significant specific comment demonstrated here; however, it remains for some relative to motion cues was, "This is the first directed studies to consider this question in the time we have experienced realistic turbulence in a detail it deserves.

simulator." The turbulence model was conventional, the same as thev had experienced in FSAA Shuttle Another example of evaluations differing with simulations. For the first time, they were variations in vertical motion cues was seen in a physically sensing the lower frequency vertical fixed-base simulator investigation of the use of a gusts. integrated side-stick controller (SSC) multiaxis, in lieu of conventional helicopter controllers for Another limited, but striking, item of evi- nap-of-the-Earth flight. It was discovered that dence of vertical-motion effects was obtained with sufficient levels of stability and control durina the exoeriments of Ref. 3. The objectives augmentation, up to three axes of control (pitch, and the task of that study were described-earlier roll, and yaw) on the SSC provided handling- in this paper. The pilots were asked to evaluate qualities equivalent to those achieved with the a number of collective-control and engine-response However, the addition of conventional controller.

configurations in terms of a formalized handling- the fourth controlled axis (vertical) to the SSC qualities rating scale and subjective commentary. yielded significant degradation in pilot rating.

As in many experiments of this kind, the evalua- In contrast, in a follow-on moving-base simulation tions were "blind"; that is, the pilot was not made on the Vertical Motion Simulator, the same four- axis SSC configuration was given pilot ratings aware of the specific variations as his evaluations equivalent to those achieved with conventional con- progressed from one configuration to another.

During the latter part of his participation in the trollers.

tests, one pilot, in several instances, was sub- jected to a variation in vertical motion instead The sensitivity of this height-control problem of a variation in the vehicle model. He was not to cockpit motion brings to mind the difficulty of achieving subjective fidelity and flight-like per- informed of this change during the tests, nor did formance in the simulated airplane landing maneu- he consciously sense that the simulator motion had been changed. He assumed he was evaluating modifi- ver. The hypothesis is offered that the visual cations to aircraft parameters. The change cues of linear motion are often very weak, espe- effected was an increase of the vertical-motion cially in the case of vertical motion; thus, ver- washout frequency from 0.4 to 1.4 rad/sec, con- tical acceleration cues are heavily relied upon in straining the cockpit motion to that experienced in the conduct of precise control of height rate.

the FSAA (see Fiq. 5). The effects of this change This dependency might extend to the lower ranges of maneuvering frequency (near 1 rad/sec). Visual on the piiot's subjective ratings of two helicop- ter confiaurations is shown in Fig. 10. Subse- cues of angular motions are much stronger. Sensi- quently, the pilot was informed of the experiment tivities to angular-motion-cue deficiencies are usually manifested at the higher frequencies typi- and asked to repeat the evaluations in the absence cally seen with high-response control systems of cockpit motion. His commentary accompanying (3-6 rad/sec).

the ratings of those cases with attenuated motion cited insufficient vertical rate damping in the vehicle. Concluding Remarks A review of helicopter simulation experience The two helicopter configurations differed With the at Ames Research Center indicates that experi- only in their values of vertical damping.

menters seeking sound pilot evaluation of vehicle full VMS vertical motion, they were given the same handling qualities have developed an appreciation rating. The descriptor associated with the 4.5 for, if not an understanding of, cockpit motion.

rating is "minor to moderate annoying deficiencies With reduced cock- It is observed that low-order, well-damped, requiring pilot compensation."

pit motion, one configuration displayed "very uncoupled control modes, in the presence of strong objectionable but tolerable deficiencies, requir- visual cues, are not sensitive to motion-cue depri- ing extensive pilot compensation," and the other vation. As these descriptions -- order, damping, coupling -- move toward the other end of their was assessed as having "major deficiencies requir- scales, or if visual cues are weakened, sensitivity ing improvement." In the fixed-cockpit evalua- to motion-cue deprivation is increased. There are tions, the ratings were further degraded.

indications that helicopter height control, with It might be inferred from these results that its collective and cyclic contributions, benefits if the research program had been conducted in the strongly from large-amplitude simulator motion.

FSAA, the degraded evaluations would have pre- All significant experience at Ames with simu- vailed, leading to quite erroneous experimental It is likely that such an inference lation of visual flight tasks has been obtained conclusions.

with a limited forward field of view. The advan- References tages of a fourfold increase of viewing area are anticipated, though they remain undefined. With 1; Chen, R. T. N., "Unified Results of Several Analytical and Experimental Studies on optimization of scene elements, the single-window model-board view has demonstrated surprising ade- Helicopter Handling Qualities in Visual quacy in a number of simulator studies. Terrain Flight," presented at AHS/NASA Specialists' Meetinq on Helicopter Han- These are very generalized observations, and dling Qualities, Ames Research' Center, they do quide the utilization of simulation facil- NASA, Apr. 1982.

ities at-Ames; but still lacking are the well- documented demonstrations of effects of cue 2. Lebacqz, J. V., Chen, R. T. N., Gerdes, R. M., deprivation that are required in the development of Weber, J.-M., and Forrest- R. D -"Results an understandinq of the motion- and visual-cueing of NASA/FAA Ground and Fliqht S;&lation processes. The-experiences with the VMS verticai- Experiments Concerning Helicopter IFR Air- motion capabilities suqqest their use in carefully worthiness Criteria," presented at conditioned studies of-the roles played by vertical AHS/NASA Specialists' Meeting on Helicop- To be acceleration in helicopter piloting tasks. ter Handlinq Qualities, Ames Research "carefully conditioned," such experiments should Center, NASA,.Apr. 1982.

employ the most promising human performance measurement and modeling techniques, and should be 3.

Corliss, L. Il., "The Effects of Helicopter designed in recognition of the probably influences Engine Response Dynamics and Excess Power of learning and task complexity on the pilot's on-NOE Handling Qualities," presented at utilization of motion cues. The visual simulations AHS/NASA Specialists' Meeting on Helicop- used during these experiments, in combination with ter Handling Qualities, Ames Research the piloting tasks, must be of the highest achiev- Center, NASA, Apr. 1982.

able fidelity to minimize contamination of the results as a result of visual deficiencies. 4. Churchill, G. B. and Dugan, D. C., "Simulation of the XV-15 Tilt Rotor Research Air- i Another attractive objective is the further craft," presented at SES-SFTE Conference development and evaluation of substitute cueing on Simulation -- Aircraft Test and Evalua- mechanisms (variable geometry seats, torso/helmet ation, Naval Air Test Center, Patuxent pullers) in the context of helicopter flight tasks; River, Md., Mar. 1982.

the VMS offers the opportunity for direct compari- son of the effectiveness of such devices with that 5.

Milleli, R. J., Keane, W. A., and Keneally, of essentially unattenuated motion. Also, the VMS w. u., "A Fliqht Program to Define V/STOL offers some opportunity to study visual fidelity Visual Simulator Requirements," presented factors in the presence of high-quality motion at the Second Flight Simulation Symposium, cues. The studies suggested here should produce the Royal Aeronautical Society, May 1973.

results facilitating more graceful and intelligent accommodation to simulations with limited cueing 6. Carico, G. D. and Corliss, L. D., "Effect of capabilities, and providing a firmer basis than Field of View on Performing a Low Altitude presently exists for further simulation technology Maneuvering Task," proceedings of the development. Third Interservice Industry Training Equipment Conference, Orlando, Fla., Nov. 1981.

Typical "fixed-base" simulation cab incorporating visual Fig. 1.

simulation and head-up display.

i Fig. 2. The Ames Flight Simulator for Fig. 3. The Ames Vertical Motion Simulator (VMS).

Advanced Aircraft (FSAA).

- ROLL - - VERTICAL, VMS - - - - VERTICAL, FSAA LEAD I I I I I ci LAG /I I I I , I .5 1 2 5 10 20 FREQUENCY, rad/sec Fig. 5. Simulator motion response Fig. 4. "Interchangeable Cab," with four- window CGI visual simulation display. relative to that of the modeled aircraft.

DESTROYER CGI DATA BASE

VIEWED FROM SH-SF ICAB I I I I I I -60 -30 0 30 60 90 AZIMUTH, deg Windows provided by visual simulation systems compared Fig. 6.

with typical helicopter fields of view.

,. “?

:, .

Fig. 7. View,of helicopter longitudinal Pilot's view while decelerating Fig. 8.

maneuverinq course on visual simulation to hover.

model-board.

207’ D’ 30' CGI DECK LANDING SCENE 5 ft ABOVE TOUCHDOWN POSITION I I I I I I

60" L . I I I

I 60" 30" 0 30" 60" 9o" # Pilot's views just before touchdown on deck as provided by CGI visual Fig. 9.

system.

COOPER-HARPER PILOT-RATING CONF. 1 lo- 0 Cl CONF.2 UNACCEPTABLE g 8- .- IMPROVEMENT 5 7! I- 0.

WARRENTS 3- SATISFACTORY 2 _ I I I (“.iS, 1.5 ( FSAA) (FIX;D, WASHOUT FREQUENCY, radkec Fig. 10. The effects of changes in cockpit motion on pilot ratings of helicopter handling qualities.

PAST APPLICATIONS AND FUTURE POTENTIAL OF VARIABLE STABILITY RESEARCH HELICOPTERS William S. Hindson Senior Research Associate Joint Institute of Aeronautics and Acoustics Stanford University Stanford, California Summary development. Nevertheless, variable- stability research aircraft have continued Variable-stability research helicop- to be used throughout this period, during ters began to be used more than 25 years which their capabilities have improved and ago to investigate flying qualities crite- their applications have diversified.

ria for helicopters.

However, their . .

application was soon diverted to investi- It is the purpose of this presentation gate mainly the problems of V/STOL air- to review briefly the evolution of vari- craft. This emphasis prevailed until the able-stability research helicopters, with past decade when the greatly increased use a view to emphasizing that these facilities of helicopters for a wide variety of more are general-purpose in nature and represent demanding applications resulted in renewed major long-term investments similar to a use of these facilities for rotary-wing large wind-tunnel or a sophisticatedground- research. based simulator installation.

The historical development of Some past variable-stability research helicopters and and recent applications of severalvariable- some of their previous applications are stability research helicopters are reviewed presented as a guide for assessing their as a means towards understanding the role future potential. The features of three that these facilities may have in the general-purpose rotary-wing flightresearch future. The features of three different aircraft that provide complementary capa- variable-stability helicopters having com- bilities are described briefly, and a num- plementary capabilities, and some of the ber of future applications are proposed. considerations involved inairbornesimula- tion technology are summarized to provide a basis for determining their future poten- Introduction tial.

To conclude, a number of applications to future rotorcraft research are noted.

In the past 25 years, variable-stabil- ity aircraft have made major contributions The three variable-stability research to the formulation of flying qualities helicopters which are given particular criteria, guidance, control, and display attention in this paper are the NASA/Army systems requirements. They have also been CH-47B and the NASA/Army UH-1H helicopters used as development tools for particular operated at the Ames Research Center, and designs, and as training vehicles to demon- the Bell model 205A-1 airborne simulator strate a wide range of generic control operated by the National Aeronautical characteristics or to provide pilot famil- Establishment (NAE) in Ottawa, Canada. A iarization prior to flying a new aircraft. BO 105 rigid-rotor variable-stability heli- In the early days, use of these airborne copter operated in GermanylReference l), simulation facilities was fairly extensive, the NASA/Army Rotor Systems Research Air- since ground-based simulation equipment, if craft (RSRA) described in Reference 2, and it existed, had extremely limited capabil- various other rotary-wing aircraft having ities. Until recently, this was particu- a variable-stability capability, but which larly true for helicopters and V/STOL are in the class of technology demonstra- aircraft, for which the motion and visual tionor training vehicles, are not dis- requirements in hover and in low-speed cussed.

maneuvering flight placed severe demands on simulation fidelity. Today, however, major advances in the capabilities of ground- Evolution and Past Applications --__ based simulators, exemplified by the Vertical Motion Simulator at NASA-Ames and Historically, variable-stability by multi-window computer-generated imagery research aircraft have usually evolved with systems, have tended to cause much greater the appearance of new classes of vehicles confidence to be placed in this means of that have exhibited unsatisfactory or aircraft and systems design, and criteria unusual flying qualities in their basic lllll!111! !lllllllllIlllllllllll Ill1 or in their subsequent application design, Ryan X-14A operated at NASA-Ames, the Bell to more demanding flight tasks. For heli- X-22A operated by Calspan, and the Short copters, one of the first developments of SC-1 at RAE Bedford.

a variable-stability research vehicle that involved an electromechanical control sys- In addition to the use of these facil- tem and electrical response-feedback tech- ities for the development of general V/STOL niques was at the NACA Langley Aeronautical criteria, some of them were also used as Laboratory in 1955 (Reference 3). With development tools for specific designs.

this helicopter, various levels of control The use of the NAE airborne simulator in power, control sensitivity, and rate damp- separate development programs for the ing augmentation were evaluated in an Canadair CL-84, the Hawker-Siddeley P-1127 effort to provide a data base for handling (Kestrel), the Vereinigte Flugtechnische qualities criteria during instrument ap- Werke VAK 191B, and the DeHavilland DHC-7 proach. For this control task, themilitary is summarized in Reference 14.

helicopter flying qualities criteria, MIL H 8501A (Nov 521,developed forvisualflight, Although some of these researchefforts were found to be inadequate (Reference 4). were also applicable to helicopters, such However, it seems that the profusion of as in the areas of lateral-directional fly- V/STOL configurations that also began to ing qualities and steep low-speed instru- appear at this time re-oriented the appli- ment approaches, it was not until the early cation of variable-stability helicopters to mid-seventies that rotary-wing applica- largely toward this field of research, an tions began to be emphasized by the NASA emphasis that persisted until about 1970. and NAE variable-stability helicopters. By then, these facilities consisted of second During the period 1955-1970, more and third generation research vehicles than 25 different VTOL and V/STOL non- equipped with much higher capacity hybrid rotary-wing configurations were test-flown computing equipment (References 15-17). In by NATO countries. Although this number response to an emphasis on all-weather of V/STOL concepts was probably signifi- capability, and in recognition of thepoten- cantly smaller than the number of new tial for trading-off vehicle control system helicopters introduced in the same period, complexity for sophistication in cockpit attention was focussed on them because of displays, these aircraft also began to their novel operational capabilities and acquire the capability for flight-systems the diversity in the design of their pro- research involving advanced navigation pulsive-lift and flight control systems, equipment and programmable CRT displays.

not to mention their often notorious safety This change in emphasis away from V/STOL In response to the need for crite- applications was perhaps partly due to the record.

ria which could more efficiently lead to a failure of any V/STOL aircraft (with the successful and capable V/STOL design, a notable exception of the Harrier) toachieve succession of efforts was undertaken by operational application, from which would NASA (Reference 5), AGARD (References 6-9), have emerged the much needed experiencewith which to validate, revise, or extend the and the U.S. military (Reference 10). To create an additional source of data for V/STOL criteria. However, it was precisely these efforts, the U.S. Army provided heli- this stage of development that was taking copters to NASA Langley and to p?AE for place for the helicopter. New and more modification as V/STOL research vehicles. demanding mission requirements were creat- These flight research aircraft, described ing the need for improved flying qualities beyond those which had been adequate in the in References 11 and 12, were the first to The nature of these requirements, use the model-controlled method of airborne past.

simulation which was originally proposed in and some recent applications of variable- Reference 13. As shown in Fig. 1, this stability aircraft in addressing them, are method is virtually identical to a ground- discussed briefly in the following section.

simulation implementation except,in this case,the moving "cab" can follow the model- generated motion commands without restric- Recent Applications tion, and with fidelity determined by the bandwidth of response and the degree of Perhaps the most significant factor (uncoupled) control achievable in influencing the recent use of variable- pure that axis. Because only four relatively stability research helicopters has been the uncoupled controls are normally available strong civil demand for dual or single in a helicopter, the motion can be accu- pilot instrument flight capability to sup- rately controlled only in four degrees-of- port natural resource operations, or to freedom, hence creating some limitations allow effective commercial use of helicop- for the simulation of V/STOL and STOL air- ters in weather conditions at least equiv- Nevertheless, significant contribu- alent to CTOL operations under Instrument craft.

tions to the V/STOL criteria were made by Flight Rules (IFR). Although the first these helicopters, and by other variable- instrument flight certification of a civil helicopter occurred in 1960, the strong stability research vehicles such as the

-

I

demand of the seventies suddenly emerged Unlike the need for helicopter all- at a time when the National Airspace Sys- weather capability that has persisted to tem was ill-equipped to allow the unique varying degrees for the past 20 years, the capabilities of the helicopter to be used military requirement for Nap-of-the-Earth efficiently. This led the Federal Aviation (NOE) operations has more recently created Administration (FAA) to institute a broad genuinely new needs for flying qualities program in cooperation with industry and and agility criteria, and cockpit engineer- NASA (Reference 18), parts of which were ing advances to include both displays and to better define the minimum requirements This mission requirement is controllers.

for helicopter IFR certification, and to so severe that it can only be partially investigate systems for improving the oper- addressed in even the most advanced moving- ational efficiency of rotary-wing aircraft base ground simulator, or in airborne simu- in instrument flight conditions. Among lators, where well-designed task elements research facilities that have been used in perhaps could be separately developed this program are the extensive ground-based towards a satisfactory solution. One simulation facilities at the Ames Research research effort using the NASA/Army UH-1H Center, and the UH-1H and B205A-1 variable- variable-stability helicopter to evaluate stability helicopters operated by Ames and different flying qualities during a simu- NAE respectively. Three flight programs lated NOE mission is described in Reference that were carried out in support of this Another investigation, carried out in 25.

requirement are described in Referenceslg- the NAE B205A-1, evaluated various multi- 21. As indicated, this requirement for side-arm controller con- axis, isometric, flying qualities criteria arose differently figurations as alternatives to the conven- than had been the earlier case for V/STOL tional helicopter cyclic stick, pedals and aircraft, the helicopter since in general, collective controls (Reference 26).

manufacturers and the avionics companies were able to provide a satisfactory capa- Variable-stability helicopters have bility,without the need for detailed guide- not been widely used by the rotorcraft lines. Rather, the motivation for this industry as development tools for particular effort was more to assess the validity and, designs. Kowever, some recent examples where necessary, extend the scope of a set where specific flight programs have been of "interim standards" which previously had conducted are described in Reference 27, been employed in the certification process. pertaining to the RSRA, and in Reference 28, which describes the role of airborne simu- Although there do not appear to be lation during part of the development pro- any major flying quality problems in cur- gram for the Sikorsky S-76.

rent helicopters for the relatively conven- tional instrument approach task, there has The broad capabilities of these been general agreement that their very low- general-purpose research facilities have speed capabilities have not yet been been characterized by referring to their In the fol- exploited for all-weather operations. As past and recent applications.

reviewed in Reference 22, a considerable lowing sections, the principal features of amount of research in this area has already three current variable-stability research been carried out, much of it at the NASA helicopters having complementary capabili- Langley Research Center using the CH-46C ties are summarized, along with a short and CH-47B variable-stability helicopters. discussion of some of the considerations Control and display requirements for carry- involved in implementing the technology of ing out decelerating approaches to hover in airborne simulation.

instrument conditions were investigated for both manual and automatic control as described, for example, in References 17 Principal Capabilities of Three Variable- and 23. A more recent investigation in Stability Research Helicopters this area using the NASA/Army UH-1H heli- copter, combined an automatic decelerating General Features approach with a helical let-down trajec- tory designed to perhaps permit helicopter The NASA/Army CH-47B is a twin-engine instrument approaches to a busy airport tandem-rotor cargo helicopter capable of without impacting existing CTOL operations lifting an internal or external payload of The investigation of (Reference 24).

The aircraft approximately 10,000 pounds.

means to improve the operationalefficiency is specially equipped with high bandwidth of rotary-wing aircraft in the National parallel electrohydraulic actuators that Airspace System is well-suited to these are able to drive the basic helicopter con- variable-stability helicopters. Their trol system over its full range through broad sensor complement and programmable electrohydraulically-operated clutches.

navigation, control and display systems During variable-stability operation, the allow for fairly rapid implementation of evaluation pilot's electrical control inputs system conepts, followed by comprehensive drive these actuators through the flight evaluation in the real environment.

computer and the engaged clutches, thereby operating the basic CH-47B controls. The computers. Other modifications include parallel control mechanization permits the removal of the stabilizer bar to improve safety pilot's controls to follow thebasic rotor responsiveness to cyclic inputs, and helicopter controls at all times, although installation of a separate electrohydraulic in general, the action of the safety actuator to drive the horizontal stabilizer pilot's controls will be quite different which was disconnected from the longitudi- than that of the evaluation pilot's. Sev- nal cyclic. The latter feature is gener- eral mechanical safety features are incor- ally not used except to trim fuselage atti- porated to insure that the safety pilot tudes in forward flight. This airborne can control the aircraft in the event that simulator is the third such general-purpose a clutch fails to disengage following a research facility that has been developed system disconnect. This helicopter had by NAE. It has been carrying out various originally been used in the technology research programs since 1974.

demonstration program described in Refer- ence 29. After its completion, the air- craft was acquired by NASA-Langley where Simulation Envelopes it was modified for use as their third- generation variable-stability research The capability of a variable-stability The aircraft helicopter to simulate the flight regime helicopter (Reference 17).

was transferred to the Ames Research Center and dynamic response characteristics of in 1979. other aircraft is limited in part by its own flight envelope, the control power The NASA/Army UH-1H is equipped with available in each axis, and the bandwidth the V/STOLAND avionics systems described and authority limits of the electromechan- Its variable-stability icalorelectrohydraulic actuators used in in Reference 30.

control system consists of high bandwidth the variable-stability system. As mentioned limited authority electrohydraulic series in a previous section, when only the four servos as well as lower bandwidth limited conventional helicopter controls are avail- : rate but full authority parallel electro- able, then motion can be controlled accu- The parallel actuators rately in only four degrees-of-freedom.

mechanical servos.

This has greater implications than just are used to off-load or to assist the series servos during sustained or aggres- precluding accurate simulation of V/STOL or sive maneuvers commanded by the evalua- compound helicopter designs with their tion pilot, or for following the lower special longitudinal force-generating fea- frequency components of automatic control tures, since sideforce and turbulence re- laws implemented in the flight computer. sponce characteristics are also compromised.

Although several airborne simulators for Although the action of the parallel servos can be isolated from the evaluation pilot's conventional aircraft have been operatinq longitudinal and lateral cyclic controls, for several years now with additional con- any action of the parallel servos in the trol devices installed to provide control main or tail rotor collective channels is over all six degrees-of-freedom, only NAE has undertaken serious study of possible reflected to the collective and pedal con- trols of both pilots. (The evaluation configurations that could provide this capa- pilot can momentarily disable these paral- bility in a helicopter.

lel servos if their action interferes with his own control inputs; however, the series An important consideration that can actuators may saturate during this time.) also strongly influence the available simu- Despite these limitations, this aircraft lation envelope is the method used tomoni- can be an extremely effective research tool tor the acceptability of maneuvers generated since it is supported by a dedicated fixed- during the in-flight simulation. Automatic base simulation facility that permits effi- monitoring of control rate and position is cient development of flight software. The usually incorporated, particularly in the aircraft has been in operation at the Ames case where series servos are used in the Research Center since 1977. variable-stability system. If only parallel or dual-mode actuators, such as those in The NAE B205A-1 (Reference 161, essen- the NAE B205A-1, are employed, then the tially the civilian equivalent of the safety pilot can be relied upon to a much UH-lH, has been extensively modified to greater extent for monitoring the remaining maximize its capabilities as an airborne control margins. This usually permits simulator. It is equipped with full au- simulations of more aggressive maneuvers thority dual-mode actuators that were such as may be encountered following simu- specially designed to replace the boost lated engine or stability augmentation sys- actuators of the basic production helicop- tem (SAS) failures. Flight programs where The actuator servo valves are mechan- these considerations influenced the simu- ter.

ically signalled when the safety pilot has lated evaluation task in contrasting ways control; in the variable-stability mode are described in References 25 and 31.

they are commanded electrically from the evaluation pilot's control via the flight A summary of factors influencing the be devoted to in-flight steady-state and available simulation envelopes of the three dynamic calibrations, especially for air- aircraft described here is provided in speed measurements, to obtain the degree Table 1.

.of precision required of a general-purpose research facility. For example, the iner- tial and air mass velocity measurements in Modeling Techniques the NAE airborne simulator were developed to sufficient accuracy to warrant its use A central issue in the use of varia- for several-atmospheric wind and turbulence ble-stability aircraft, and one which also measurement programs (Reference 35). Of influences the simulation envelope, is the additional benefit, the frequent availa- fidelity of motion response that can be bility of redundant measurements from a achieved during the in-flight simulation. variety of sensors, combined with the For some investigations, such as those recent remarkable advances in digital com- now make it possible to involving only generic flying qualities, puting equipment, the importance of accurately representing implement modern state estimation and fil- specific dynamic response characteristics tering algorithms to achieve improved may not be of great concern. However, the accuracy and noise suppression.

accurate simulation of a specific design, the investigation of higher frequency While navigation equipment usually modes of motion, or the representation of plays a supporting role for pilot in-the- turbulence response characteristics may loop flying qualities investigations, it require a level of performance from the can assume a more central role for inves- variable-stability control system that is tigations of a systems nature, such as very difficult to achieve. curved decelerating approaches. These may be carried out using either manual or auto- In general, control of the dynamic matic control.

response characteristics is accomplished either using response-feedback and control- The motion and navigation sensor com- feedforward techniques, effectively equiv- plements of the NASA and MAE research alent to most conventional stability and helicopters are summarized in Table 3.

control augmentation system implementa- tions, or with model-following systems such as that shown in Fig. 1. Some of the Evaluation Pilot Controls and Displays considerations involved with each method are summarized in Table 2, which identifies An important requirement in any that there are major advantages, at least piloted simulation is the representation of Similar to theoretically, in using model-following control force characteristics.

techniques. Although model-following auto- sophisticated ground-based research simu- pilots with quite good performance (i.e. lators, nearly all variable-stability air- moderate bandwidth) were relatively easy craft today have the capability to model a wide range of force-deflection character- to develop for some of the earlier light single-rotor variable-stability helicop- istics, including breakout, hysteresis, viscous and coulomb friction, and non- ters (Reference 12), the larger facilities presently in use have presented difficul- linear spring gradients. These character- ties that tend to be associated with con- istics may also be influenced by themotion trol crosscoupling and higher frequency of the simulated aircraft being "flown" by structural modes, which are in addition to the evaluation pilot. The flexibility that the usual difficult aerodynamic and vibra- the three variable-stability research heli- tional environments. The use of modern copters have for varying the evaluation multivariable control system design tech- pilot's control characteristics is summa- niques (e.g. References 32,33), or methods rized in Table 4.

involving the inverse solution of the equations of motion of the basic platform Rapid advances have also taken place (Reference 34), are possible means for in cockpit display system hardware that now it possible to consider more diffi- improving the motion fidelity of variable- make stability helicopters which could benefit cult control tasks such as curved or decel- The programmable dis- the three facilities described here. erating approaches.

play equipment available in the NASA and NAE helicopters is also noted in Table 4.

Platform Instrumentation The in-flight simulation objective Computational Capacity imposes severe accuracy requirements on the motion measurements of the helicopter It is usually not possible for vari- which, able-stability aircraft today to employ the in the final analysis, are used to validate the dynamic response character- full potential of current computer technol- To take advantage of increasingly istics. Particular attention must usually ogy - compact and more powerful computing equip- motions can be a source of difficulty). In ment would compromise the availability of addition, simulations of large heavy-lift the aircraft for conducting research pro- helicopters and airships might be possible.

grams. As a notable exception to this statement, the NAE B205A-1 has recently Associated with an expanded simulation been equipped with a locally-designed envelope is the continuous need for im- multi-microprocessor digital computing sys- proved simulation fidelity. Greatly tem that replaced the original minicomputer enhanced computational capacity combined installation. This development has pro- with modern multivariahle control system vided the ability to carry full laboratory design methods should ultimately result in operating-system software,and to implement improved variable-stability system perform- on-line data handling programs in addition ance. If model-following methods are to the necessary flight programs. employed, an associated area to benefit is the simulation of wind and turbulence, While the NAE capability is exception- including windshear. Also requiring the computational capacity of the improvement is the simulation of instrument al, CH-47B and the UH-lH(V/STOLAND) research flight conditions, particularly the transi- listed in Table 5, is adequate helicopters, tion to visual flight at instrumentapproach to meet requirements at their current minimums. Technoloqv to permit more real- stages of development. istic representation-of this critial area would be of major benefit to all in-flight The objectives of presenting these simulators, and possibly ground-based simu- brief descriptions have been to illustrate lators as well.

the differing yet complementary capabili- ties of these research helicopters, to The three variable-stability helicop- identify areas where further development ters that are described hriefly in this could be warranted, and to provide a basis presentation could indeed benefit from for assessing the potential of these facil- these and other improvements. However.

each has distinctly different and ities to carry out future rotorcraft research. These considerations are dis- complementary capabilities that tend to cussed briefly in the following section. focus its applications. The UH-1H (V/STOLAND) helicopter's navigation sensors make it ideally suited to investigations of such as terminal-area Future Potential of Variable Stability a systems nature, or the implementation Research Belicopters approach procedures, and testing of new automatic guidance and There is little doubt that the appli- control concepts. In this regard, a full- cation of variable-stability helicopters to flight-envelope autopilot designed using various general or specific research and the inverse model techniques described in development problems would be broadened Reference 34,is under development and is Although the capabil- significantly if their capabilities were nearing flight test.

improved. Such indeed turned out to be the ity of this aircraft to simulate a wide range of flying qualities or to perform case when several of the conventional vari- aggressive NOE-type maneuvers even with low able-stability aircraft developed five or six degree-of-freedom simulation capabili- levels of stability augmentation is severely ties in the past decade. Some of their limited by its variable-stability system the facility is considered ade- applications to new classes of aircraft actuators, and to basic research in the field of quate for representing the generic flying human response studies, for example, are qualities of most current SAS-equipped hel- noted in Reference 36. However, achieving icopters during conventional instrument full six degree-of-freedom controlled approach tasks.

motion capability in a helicopter is admit- (The additional longi- Alternatively, the NAE B205A-1 is tedly more complex.

undoubtedly the superior vehicle for gener- tudinal and lateral force-generating capa- bility in hover needs to be provided by an al flying qualities research, including the auxiliary reactive propulsion system.) simulation of specific designs. Limited Increased application to V/STOL vehicles only by the inherent control power avail- is the usual justification given for pro- able from its teetering rotor, and its four posing this capability, but may also be controlled degrees-of-freedom, it is able, one reason why it has not yet been real- among other attributes, to accommodate aggressive maneuvers such as might arise ized. A fairly large amount of longitudi- from simulated systems failures, even when nal-force control power is usually con- sidered necessary for this application: close to the ground. However, it has a whereas, a considerably smaller amount limited cockpit display and navigation sys- could still permit investigation of impor- tem capability with which to conduct tant rotorcraft problems such as instru- advanced integrated systems investigations.

ment flying qualities criteria with exter- nal loads (where oscillatory longitudinal The CH-47B, also intended primarily of boundaries defining minimum acceptable for pilot in-the-loop flying qualities standards for FAA criteria, or to meet investigations, is distinctive for its Level II and Level III military flying ability to address problem areas associated qualities and performance specifications.

with external load control. In addition, This usually involves the systematic varia- its greater amount of control power in tion of configurations in realistic mission pitch, which arises from the use of differ- simulations for which general-purpose ential collective for this purpose, permits ground-based or airborne simulators are simulation of the response characteristics well-suited. Rarely, however, can a single that may be associated with different rotor facility provide all of the required data system designs. The aircraft is also with the level of confidence necessary to equipped with a programmable symbol gener- establdsh criteria. Instead, a number of ator and associated electronic CRT cockpit carefully planned investigations using displays, giving it the greatest capabil- facilities with complementary capabilities ity in this area of the research helicop- are usually conducted. The unique features However, the ters discussed in this paper. of variable-stability research helicopters, CH-47B is presently at a considerably lower such as those described in this presenta- level of development than the other facil- tion, offer important capabilities with ities. which to address these issues.

are capable of making These aircraft significant contributions to the develop- Concluding Remarks ment of flying qualities criteria and sys- tems requirements for a variety of mission The application of variable-stability requirements applications, such as the FAA research helicopters to support new devel- certification and military NOE programs opments in the rotorcraft industry has mentioned earlier. In addition to the increased significantly in the past several research programs that have already been years. This has been associated mainly undertaken, a number of other applications with developing criteria to support the also within these general areas are as recent widespread use of helicopters in proposed: more demanding missions, and to a lesser The development of sensors and extent, with the development of new designs.

1) control laws for automatic hover control, Still, recognition of the potential of these including precision control of an external facilities has been overshadowed by the load and hover relative to a moving plat- confidence, much of it yet to be substan- form; and the development of stability tiated, that has been growing in the new augmentation systems and displays to sup- capabilities of modern ground-simulation This presentation has called port the manual execution of these tasks. technology.

The investigation of stabilization attention to the historical development of 2) systems for external loads in hover and in these aircraft that places them in the class of long-term general-purpose flight forward flight, along with associated fly- research facilities. The review of their ing qualities in instrument flight condi- tions. previous applications, and the summary of The development of navigation, their current and potential capabilities 3) control and display system that have been presented, suggest the nature guidance, requirements necessary to exploit the very of the applications that could emerge for these vehicles in the future. At a time low-speed capabilities of the helicopter in when rotorcraft and flight control system instrument flight conditions in both remote and congested areas. technologies are making rapid advances, and The evaluation of new man-machine the use of helicopters for a variety of new 4) tasks is becoming more widespread, it is interface technology, such as voice actua- tion, tactile controllers, and multi-axis probable that variable-stability reseach side-arm controllers, that requires devel- helicopters will continue to serve an increasingly useful role.

opment to exploit new electronic flight control systems.

The investigation of energy man- 5) References agement techniques and associated control and display requirements applied to engine- (1) Attfellner, S., and Rade, ??.,"~0105 In- failure situations in single or twin- Flight Simulator for Flight Control and engine helicopters. Guidance Systems," First EuropeanRotor- The evaluation of advanced theo- craft and Powered-Lift Aircraft Forum, 6) retical control system concepts for which Southampton, September 1975.

modeling errors and sensor noise and accu- (2) White, S., Jr., andcondon, G.W., "Flight racy may represent major limitations. Research Capabilities of the NASA/Army Rotor Systems Research Aircraft," Pre- An important aspect of these criteria sented at the Fourth EuropeanRotorcraft and system development efforts that is and Powered-Lift Aircraft Forum, Stresa, sometimes overlooked is the determination September 1978.

Whitten, J.B., Reeder, J-P., and Crim, (3 (18 Nelson, J.R., "The FAA Helicopter A.D., "Helicopter Instrument Flight Operations Development Program," and Precision Maneuvers as Affected Presented at the Fifth European Rotor- by Changes in Damping in Roll, Pitch, craft and Powered-Lift Aircraft Forum, November 1955.

and Yaw," NACA TN 3537, Amsterdam, September 1979.

Salmirs, S., and Tapscott, R.J., "The (4 (19 Lebacqz, J.V., Weher, J.M., and Effects of Various Combinations of Corliss, L.D., "A Flight Investiga- Damping and Control Power on Helicop- tion of Static Stability, Control ter Handling Qualities During Both Augmentation, and Flight Director Instrument and Visual Flight," NASA Influences on Helicopter IFR Handling TN D-58, October 1959.

Qualities," Presented at the 37th AHS (5) Anderson, S.B., "An Examination of Forum, New Orleans, May 1981.

Handling Qualities Criteria for V/STOL (20) Kereliuk, S., and Sinclair, M., "Eval- Aircraft,; NASA TN D-331, July 1960. uation of IFR Handling Qualities of (6) Anon., "Recommendations for V/S!?OL Helicopters Using the NAE Airborne Handling Qualities," NATO,AGARD Rep. V/STOL Simulator," Atlantic Aeronau- October 1962. tical Conference, Williamsburg, March 408, (7) Anon., "Recommendations for V/STOL 1979.

Handling Qualities with an Addendum (21) Peach, L.L., Jr., et al., "NASA/FAA Containing Comment on the Recommenda- Flight-Test Investigation of Helicop- tions," NATO, AGARD Rep. 408A, October ter Microwave Landing System 1964. Approaches," Presented at the 36th "V/STOL Handling," NATO, AGARD (8) Anon., Forum, Washington, May 1980.

577, Part I-Criteria and Discus- (22) Lebacqz, J.V., "Survey of Helicopter Rep.

December 1970, Part II-Documen- Control/Display Investigations for sion, tation, June 1973. Instrument Decelerating Approach," (9) Anon., "V/STOL Display for Approach NASA TM 78565, March 1979.

and Landing," NATO, AGARD Rep. 594, (23) Niessen, F.R., et al., "The Effect of July 1972. Variations in Controls and Displays "Military Specification-Flying (10) Anon., on Helicopter Instrument Approach Qualities of Piloted V/STOL Aircraft," Capability," NASA TN D-8385, February December 1970. 1977.

(11) Garren, J.F., Jr., and Kelly, J-R., (24) McGee, L.A., et al., "Automatic "Description of an Analog Computer Helical Rotorcraft Descent and Landinq Approach to V/STOL Simulation Employ- - Using a Microwave Landing System," ing a Variable-Stability Helicopter," Presented at AIAA Atmospheric Fliqht NASA TN D-1970, January 1964. Mechanics Conference, Aibuquerque; (12) Daw, D.F., and McGregor, D-M., "Devel- August 1981.

opment of a Model-Controlled V/STOL (25) Corliss, L.D., and Carico, D.G., "A Airborne Simulator," National Aero- Preliminary Flight Investigation of Cross-Coupling and Lateral Damping for nautical Establishment Rep. LR-352, Nap-of-the-Earth Helicopter Opera- August 1962.

"The Model-Controlled tions," Presented at the 37th AHS (13) Gould, D.G., Method for Development of Variable Forum, New Orleans, May 1981.

Stability Aircraft," National Aero- (26) Sinclair, M., and Morgan, M., "An nautical Establishment Rep. LR-345, Investigation of Multi-axis Isometric June 1962. Side-arm Controllers in a Variable (14) Hindson, W-S., "The NAE Airborne Stability Helicopter," National Aero- V/STOL Simulator as a Design and nautical Rep. LP.-606, August 1981.

(27) Sinclair, S.R.M., Xereliuk, S., and Development Tool for V/STOL Aircraft," Canadian Aeronautics and Space Jour- FTood, A-D., "Simulation of Hovering nal, December 1970. Flight Characteristics of the Rotor (15) Daw, D.F., Lum, K., and McGregor, Systems Research Aircraft Using the D.M., "Description of a Four Degree NAE Airborne V/STOL Simulator," of Freedom V/STOL Aircraft Airborne National Aeronautical Establishment Simulator," National Aeronautical Rep. LTR-FR-53, July 1977.

Establishment Rep. LR-499, February (28) Wright, G.P., and Lappos, EJ., "Spirit Handling Qualities Design and Devel- 1968.

16) Sinclair, S.R.M., Roderick, W.E.B.. onment,ll Presented at the 35th AHS and Lum, K., "The NAE Airborne V/STOL Forum,. Washington, May 1979.

Simulator," Presented at AGARD Flight (29) Anon., ~ "Tactical Aircraft Guidance Mechanics Panel on Rotorcraft Design, System Advanced Development Program May 1977. Flight Test Phase Report Vols I and 1 7) Kelly, J.R., et al., "Descrition of II," USAAMRDL TR 73-89A and B, April the VTOL Approach and Landing Tech- 1974.

nology (VALT) CH-47 Research System," NASA TP 1436, August 1979.

(30) Baker, F.A., et al., "V/STOLAND Avionics System Flight-Test Data on a UH-1H Helicopter," NASA TM 78591/ AVRADCOM TR 79-23, February 1980.

(31) Sattler, D.E., et al., "An Investiga- tion of the Recovery from an Engine Failure in a Twin Engine Augmentor Wing Aircraft Using the NAE Airborne Simulator," Canadian Aeronautics and Space Journal, First Quarter 1981.

(32) Rynoski, E.G., "Adaptive Multivariable Model-Following for Aircraft," Pre- sented at the Joint Automatic Control Conference, San Francisco, August 1980.

(33) Stengel, R.F., et al., "The Design of Digital Adaptive Controllers for VTOL Aircraft," NASA CR 144912, March 1976.

(34) Meyer, G., "The Design of Exact Non- Linear Model Followers," Presented at the 1980 Joint Automatic Control Con- ference, San Francisco, 1980.

(35) Sinclair, M., and Hindson, W.S., "The Wind and Turbulence Measuring System of the NAE Airborne Simulator," DME/ NAE Quarterly Bulletin No. 1977(4), National Research Council, Canada.

"V/STOL Flight Simulation," (36) Anon., NASA TM 81156, November 1979.

_____ -__---_---------- REAL WIND B I SIMULATED TURBULENCE WINDS a DISTURBANCES TURBULENCE 1 1 , ,____ -_--- -----.

-I ” HIGH BAND- SIMULATED WIDTH YODEL +HELICOPTER AIRCRAFT FOLLOWING DYNAMICS AUTOPILOT INCLUDING NAVIGATION COCKPIT EOUATIONS DISPkAYS I VEHICLE cmTRoLa l TRMdSDUCERS NAVIGATIDN ,INFORYATION I ; i I REAL WORLD VISUAL I I I / I AND MOTION CUES I SIMULATED I I I SERYOhHXWNISY ENVIRONMENT A 1 I ___________- -_------ _-----------.

Figure 1. Model-Following Method for Airborne Simulation Table 1.

Comparison of Variable Stability Eelicopter Simulation Envelopes NASA NhE NASA uH-lew/sToLnND) B 205A-1 OH-478

T

-30 tc 120 kts longitudinally -30 tc 120 kts longitudinally Flight envelope of basic -30 to 160 kts longitudinally 35 kts in lateral flight. 35 kts in lateral flight. 35 kts in lateral flight.

production aircraft.

Maximum bank angle 40 deg below 145 kts, 20 deg at 160 kts.

I

Fully articulated tandem Single teetering rctcr. Single teetering rctcr, sta- Rotor systfm.

counter-rotating rctcrs. bilizer bar removed.

I .- -------~_.__

Longitudinal cyclic Differential collective Longitudinal cyclic Basic pitch Lateral cyclic Lateral cyclic Lateral cyclic Cc"trclS : roll Differential lateral cyclic Tail rctcr collective Tail rotor collective available . yaw Main rotor collective Main rctce collectives Main rotor collective . heave Independent ele".tcr trim Independent longitudinal Mechanical elevator inter- . lcngitudi- cyclic trim connect nal pitch trim 1.0 r/se=2 1.9 r/.ec2 1.0 r/set; 2.5 r/sec2 1.9 r/see2 2.5 r,sec2 1.8 r/sec2 0.9 r/sec2 1.8 r/6.x power in . yaw 1.2 g 1.1 g 1.15 g hover . heave'

_ -. - ..- _

Parallel electramechanical Variable stability system Parallel ectrchydraulic actu- Dual mode electrchydraulic actuators 4-axes 100% actuators in d-axes with actuator characteristics. atcrsin 4-axes with 100% nearly 100% authority, lover authority high banduidth2.

authority high bandvidth2.

bandwidth3, and rate limits Step-tc-step travels achieved Stop-tc-stop travels achieved giving stop-to-stop control within 0.75 sec. in approximately I.5 sec.

travels between 5x4 set (collective) tc 9.3 set [tail rotor). series electrchydraulic Notes actuatcrs in J-axes with high 2 vicinity 50hz bandwidth2 and authcrity lim- 3 vicinity 4Obz ited between 19% (collective) and 308 (tail rotor) of full travel. Rate limits 7 times faster than parallel ser"c8.

-, ~~~~ Safety pilot mcnitcrs ccn- Hardware and software mcni- Basic helicopter ccntrcl Control system trcl rate and position tcrs with trip thresholds rates mcnitcred by hardware monitoring except for trips near max based on persisting serve system with adjustable trip command-response errors. swash plate angles sensed by thresholds.

flapping angle transducers.

Extension tc 5 or 6 degrees Lcngitudinal cyclic trim of freedom with auxiliary Remarks gives very limited 5 degree- thrusting engines under inves- of-freedom control.

tigation.

I

-

I

Note 1 T/H L max wt for hcver c.g.e. at 6.1. with max cont. power/normal operating weight.

Table 2. Comparison of In-Flight Simulation Methods Consideration ___ Response-Feedback Model-Following Implementation of simulated Desired dynamic response of Standard equations-of-motion dynamics. each simulated configuration model structure with aerc- must be separately ccnstruc- dynamics of simulated "ehiclc ted from basic vehicle char- incorporated directly for acteristics plus scme ccmbi- each program.

nation of feedforwards, feedbacks.

Requirement for precision Not necessarily required Required far the degrees of on-line mcticn estimation.

on-line. freedom in which motion is controlled.

Requirement for in-flight Typically necessary tc Desirable but not generally confirm characteristics. required. Dynamic calibrations far each simulated ccnfigura- ticn.

GKwledge required cf Low precision except a* High precision.

>asic vehicle response needed for basic autopilot :haracteristics. design.

Real turbulence effects Zapability tc control and Real turbulence effects net suppressed. simulate turbulence suppressed.

=eSp"Se.

Simulated turbulence easily Simulated turbulence response introduced, including wind- difficult tc effect without influencing maneuver response shears in the degrees of freedom that are controlled.

Summary of Principal Instrumentation Table 3.

NAE NASA UH-rHW/sToLand) B 205A-1 . INS linear velocities a-axis body-firer lotion Sensors *INS linear velcc- a iMS gimbal angles ~cppler radar in addition to . j-axis body-fixed Instrumented boor 3-axis linear ocppler radar with a,0 vanes accelerometers and swivelling .u,ras and J-TEC low X-axis rate gyros airspeed sensors static port Vertical and direc- . Laser dcpplee velo- tional gyrds Cimeter low airspeed (3-axis) sensor planW3 with a,B vanes (remcvahlel and static ports . ~"strumented boom with a,B vanes and static ports _. ._~ -~ . "on,Loc, ILS MIA Vaviqaticn and .MLs Radar altimeter . TACAN. "HE related sensors . INS . INS input to flight m~adar altimeter . MIS :cmputers Cubic WE-based trian- gulation system . Radar altimeter Table 4. Summary of Evaluation Pilot Control and Display Hardware NAE B 205*-1 Z-axis (pitch-roll1 Spring cartridge X-axis programmable programmable fcrce- hungees with magnetic force-feel system.

brake release and feel system in prc- Electric power lever curement. fixed gradients.

or a ccllecti"e can Evaluation Pilot Magnetic brake on be installed.

Control* collective lever and Side-arm controllers pedals.

can he installed.

Adjustable spring Programmable electrc- Prcgrammable e1ectrc- Programmable electrc- mechanical flight mechanical flight mechanical flight director (AD-350). director (HZ-6FI and director (FD-109).

Cockpit Displays Table 5. Summary of Computational Capacity P!ASA CH-478 1 Sperry 1519A mini- ccmnuter 321 1S bit Digital Canputers w&s of RAM 50hr-- frame rate.

120 operational amplifiers

IA”+Computers 1EA1 TR-4g

60 integrators 120 manual pots 30 servo-set pots.

A PIID&IN-THLI.LXXANALYSIS OF SESXR4LKINDSOFIELI~PlERA~~~~~ Ibbert K. Wffley Principal E&search Ehgineer systeu!s lMlnology, Inc.

Mxmtain View, Glifomia 94043 &stract Pilot's integral gain in normal speed KI change maneuver It is becoming increasingly convenient to measure and analyze directly the control strategy Pilot's speed loop gain %J of pilots involved in performing authentic tasks -both in simulators and in flight. As a Pilot's position gain in dash/quickstep KR it is new possible to begin conpiling a result, catalog of engineering descriptions of various Pilot's closure rate gain in dash/ Kk flight tasks, the associated piloting technique, quickstop and the perceptual pathways involved. This paper describes hm a certain class of helicopter Range (actual) R flight tasks, namely acceleration/ deceleration Ftxition comnsnd manewers. can be quantified and put to use in Rc the fields of handling qualities, flight train- perceived range and evaluation of sirmlator fidelity. 'Ihe ing.

Rp three specific cases include the norms1 speed Closure rate B maneuver, the nap-of-the-earth dash/ change quickstop, and the decelerating approach to Maximm~closure rate tnax hover. All of these manewers share conrson gen- eric features in terms of pilot adaptation and ii Ikceleration mathatical description; yet each differs in U Psrturbation forward speed terms of the essential feedback loop structure, implications for handling qualities requirwents.

Forward speed u and simulator fidelity criteria.

Speed conmnnd "C btatim x Fore-and-aft displacement Speed damping stability derivative A Gilinsky's perceived range constant EL Controlled element transfer function Gravity constant g yco Pilot control strategy transfer function h lkight %I A0 Perturbation pitch attitude h Vertical velocity I&aping ratio Maximnn sink rate during terminal landing r %k maneuver Closed-loop damping ratio of () task .

ro 'Ibuchdawn sink rate for landing manewer htd r Pitch attitude Pilot's effective gain in approach task K, Pitch attitude ccmnmnd % EIX~~~JIpitch attitie during quickstop %k manewer Ihase margin Q Ehase margin of () task 9) w Mzural frequency individual task or maneuver description (i.e., specific flight tasks and then individual axes of Closed-loop natural frequency of () task .control for each task). With regard to the Y> rotary-wing specification, MIiK8501A, there is &ossover frequency % the mention of specific flight tasks in connec- tion with various power and speed conditions but, Effective crossover frequency of () task % again, no quantitative definition. &nce. as sJbscripts specialized environments such as NOEhave entered the scene, it has been necessary to consider a Approach to hover task significantly more stringent response standards such as those suggested by Fdenboroxh and f landing flare task &rnicke8. A-I example of the level of task breakdawn which should be considered is shown in r I&h/quickstop task mble lg. 'Ihis is based, in part, on careful tabulation of Army training objectives.

U NXUB~ speed change task X Fore-and-aft position regulation task Table 1. Army Flight TJ&k.sand hewers cl Pitch attitude regulation task (btary- and Fixed-Wing) Introduction WEAPONDELIVERY HOVERFIRE 'Ihe purpose of this paper is to describe, RUWIING FIRE DIVIIIG FIRE using a set of examples, certain elements of an Au!

approach to handling qualities which can quanti- JNSTRUMENT FLIGHT tatively account for the pilotqehicle response TAKEOFF needs in performing specific flight tasks or maneuvers. 'Ihis is accomplished by modeling the lLcwwx TAKEOFF TO Hovm flight task or maneuver in a way which permits HOVER HOVER CHECKS the inference of the pilot's loop structure and HOVERTuans the relative dependence of task perfonrrance on Fowrao HOVER various essential and supporting loops. lhis cqlements and is fully coupatible with the equivalent systems approach to describing the vehicle dynsmic~'*~ and, in fact, provides the needed context for applying bandwidth criteria3.

If handling qualities are "those stability and dynamic response characteristics of an air- craft and its control systen which iupact' the pilot's ,$Zliq to complete some useful task or mission, then wz rmst be prepared to quantify not only the vehicle but also the task. '0sk quantification is the real subject of this paper; and wz illustrate the concept using examples of several kinds of helicopter acceleration/ deceleration msnewers.

Historically, handling qualities requirements have not been very closely tied to specific 7his holds for fixed-wing5, flight tasks.

and rotary-wing aircraft7. &rhaps the V/SToL6, closest that existing specifications ccnmz to dealing with individual flight tasks is the fixed-wing handling qualities specification, MIL-F-878X, and its three "flight phase categor- The aim of this paper, then, is to show how a ies;" hwever, we shall- be dealing with at least more thorqh treatment of individual flight one or tw, additional tiers of detail in the tasks and lnaneuvers can result in better understanding of the piloting technique, the &he iuportsnt tool in the X-10 analysis was perceptual pathways, crucial vehicle characteris- the use of a phase plane plot of the "cormrand tics, and the role of supporting pilot loops. loop" (extrane outer loop) -in that case height 'Jhe hope is to arrive at a more rational and versus height rate-of+hange. Based on the phase selective approach to handling qualities which it WIS observed that the plane trajectory, looks after the key ingredients of any particular landing flare was equivalent to an unforced piloting task. lhis approach can also be useful second-order response beginning with a set of in judging the validity of simulator investiga- state initial conditions and a set of state tions of handling qualities. ccrmmnds appropriate to touchdawn. lhis is shcwn in the sketch in Pig. 1.

In order to illustrate the above concepts we shall consider one class of helicopter flight tasks, namely speed changes. E&presenting this class are three rather specificmanewers: General second-order, 1) Nxmal spee$ change maneuvers phase plane w trajectory 2) NOEdash/quickstep \ described by f 3) Decelerating approach to hover.

As wa shall see, each involves a unique combina- tion of abruptness, pilot compensation, essential loop structure, and crucial vehicle features. In effect, each maneuver represents a particular context for judging handling qualities.

Technical Approach Flare Segment The approach to analyzing the speed change maneuvers listed above is adapted from a Flare Initiation Constraint particularly successful and insightful analysis of the ii f 2&L+ + ub2p = 0 transportlO?di&~ed%rea d?~ct"es~~~onjedf closed-loop flight path response for the flare Figure 1. Phase Plane Depiction maneuver, pilot control strategy was quantified in considerable detail. This resulted, in turn, of Landing Flare in identifying differences between landings per- formed in flight and in a simulator, the effects of training pilots in flight as opposed to on a 'Ihe closed-loop danping and natural frequency simulator, and the key features in the pilot or parameters, can be found using cf and y, identification procedures, aircraft responsible for landing rigorous parameter any simple phase plane estimation difficulties. although even methods wrk well. The sketch in Pig. 2 outlines 'Ihe analysis procedure applied to the X-10 all that we shall need in order to address the landing flare consisted of identifying the effec- speed change manewers of interest here.

tive second-order closed-loop response parameters-l (e.g., frequency and damping) and subtracting the Ebr the landing flare, it was found that a open-loop aircraft response in order to infer the fairly large ssnple of pilots preferred a closed- Each of these compo- pilot's control strategy. loop damping ratio of about 0.7 f 0.1 and a nents, of course, has value, i.e., natural frequency of about closed-loop 0.4 f 0.1 radlsec. In terms of an effective 1) Qosed-loop pilot-vehicle respamez banckLdth (crossover frequency) and phase margin, abruptness or urgency of the task and specific the DC-10 flare was found to have: context for supporting loops or pilot actions.

2) cp~loop aircraft re9pmse: specific roles or influences of vehicle stability, con- Crossover Frequency, w = 0.2 to 0.33 rad/sec trol, and perforumnce characteristics.

Cf 3) Pilot control &rat=: availability of cues, ease of compensation, and level of skill. l&se Margin, +% ='70 to go deg _ . _._ .-- a helicopter the normal technique for slawing dam is to simultaneously pitch up and lower the collective. Ihe relative amount of collective control change tends to be in direct proportion to the airspeed; but, collective control is a separate issue which can be handled apart from the pitch attitude control, per se.

?he main determinant of a helicopter speed change is the use of pitch attitude since it can shown that to a be first-order good approximation": To this we can add the pilot's closed-loop con- trol of attitude in terms of a first-order lag approximation involving pitch crossover fre- htd quency, wc , i.e., as 0.83 - 0.6 7 e Cf hpk (2) \ Thus a pilot control law can be expressed in +!d terms of a pitch attitude conrman d rather than a cyclic pitch control conmand, per se.

pk lhe basic control strategy for either regu- lating or changing speed will involve a speed Ebrmalized Phase Plane and Figure 2.

Relationships for Extracting Closed-Loop feedback in the W wnnnsnd loop," i.e., as shorn in Dmnping and Natural Frequency Fig. 3. 'Ihe job of the pilot is to adopt a speed 'Ihese values therefore establish a highly qusnti- tative context by which to judge basic airplane response characteristics and the degree of pre- cision and control of pitch attitude required for support of the lauding maneuver. As an example we might apply a factor-of-five rule of thunb for setting the necessary inner-loop pitch response &nce the equivalent-system pitch bandwidth.

attitude bandwidth requirement for landing in the Figure 3. Control Strategy for the DC-10 should be at least 1 to 1.7 rad/sec - a Ibmal Speed Lange hewer reasonable range of values.

control strategy, Yp , which will result in an effective management%f speed, and wz can obtain strong clues of the pilot's control strategy by lhe norm1 speed change maneuver in a heli- observing a phase plane plot of speed versus copter might include takeoff as ~11 as up-and- acceleration. In several available flight cases, away flight. It is not unlike the corresponding it can be observed that the phase plane maneuver in a fixed-wing aircraft. Cyclic pitch trajectory of a speed change is essentially (or elevator) and collective (or throttle) are second order. Figure 4 shows sme examples.

coordinated so as to effect an x-axis accelera- tion with minimal disturbance to flight path. In

I

Initial Condition at a Terminal Steady Forward ,Condition Velocity\ u-u, Figure 5. Typical Phase Plane of a Ebrmal SpeedChange Figure 4. TypicalPlight&amples of Normal Speed &anges (5) The kind of data shawn in Fig. 4 can be re- plotted in conventional phase plane terms as shm in Pig. 5. even thqh good definition of the terminal condition is lacking. mere it is Airframe Closed-loop so ill-defined, we nust estimate or assu0e a Pitch Speed closed-loop damping ratio, 5,. A value of 0.7 to Rssponse Wspmse 0.9 is probably reasonable in view of the desire to avoid significant overshoot in sny discrete manewer. (Recall that for the landing flare a and, assuning an integral-plus-proportional speed damping ratio of 0.7 was measured.) 'Ihe ratio of control, peak pitch attitude change (or x-acceleration) to total speed change is directly related to the closed-loop natural frequency. According to the relationships shorn in Fig. 2.

then * w = 2.4 g (3) U AU S3 - s +gsKl=O +(1+$2+k+)jl) Using the predominant closed-loop response “C e e and the essential helicopter dynamics, it is thus (7) possible to solve directly for the pilot's con- trol law, Y pu' It can be show that for w ce >> t+ the s3 term is negligible and the s2 coefficient is nearly unity. (Also Ku is often negligible.)

i.e., 0 = 1 + Y l Yc - s2 + 2~93 + 4 (4) pu u

lllllllllllllllllll Ill Illll III I III I I II I

Thus since position is of ultimate importance. A phase plane portrait of the dash/q+kstop is therefore correctly depicted in the R - R plane Note that ua can handle either the of Fig. 7.

K 4 2wil + %l md (8) s (9) dash-quickstop combination or the quickstop alone I = 25uwu + 't 5Jp g depending upon how uz pick initial conditions, but the family of phase plane trajectories would be the same.

Typical flight data may show a 10 deg pitch change for an 80 kt speed change which therefore corresponds to an q, of 0.1 rad/sec according to of 0.7, this muld yield a Fiqn. (3). For a 5, crossover frequency of 0.07 rad/sec and a phase It should also be noted margin of about 85 deg.

that only a pitch attitude cue and a speed cue indicated airspeed) are needed to accom- (i.e., plish this task. The integral term implies a trimming function in parallel with the basic pitch attitude connmnd. Thus the basic pilot gains (assuning a typically negligible X,, for helicopters) would be Figure 6.

Corsnand Ioop for the NOEspeed (Ibsition) Change Kb = 0.42 and KI = 0.07lsec ('0) ,U') Initial Condition Terminal at Rest (dash In retrospect it can be seen that the usual plus quickstop) closed-loop pitch attitude bandwidth (mc ) of about 1 rad/sec is not critical to the p&for- mance of a the normal speed change umneuver; in it could be as low as 0.35 rad/sec and fact, still provide adequate support to the task.

Takeoff time histories for a UR-6O12 seem to substantiate these estimates well in that an airspeed inverse time constant of about O.l/sec and sn attitude inverse time constant of about &.ickstopsegment Dash Segment 0.33/set can be observed.

\- Initial Condition at a Steady Forward Velocity (quickstop NOEtbsh/@ickstopl%amuver alone) This is a far more aggressive variety of speed change maneuver than that considered above. The NOEspeed change-really a position change _ also involves use of collective pitch Figure 7. Range Fhase Plane Assuning to offset height changes and prevent ground con- Second-&der Closed-Loop Behavior tact. As before, though, wa shall treat only the the pilot's control law for effect- x-axis, i.e., ing a speed change through use of pitch attitude If the NOEspeed change is assuned to involve and set aside the important collective control, both a range and a velocity feedback, then (At the same time, wz are es- control aspects.

tablishing the context of the collective control task.)

= KR + KfiS (12) 'pr The basic control strategy for the NOEmsn- euver involves a range couamnd-loop (Fig. 6) The controlled elanent is the same as before agrees well with the pitch damping (essentially except for an additional integration, i.e., pitch attitude bandwidth) swested by Eidenborqh and Wernicke8 for the NOE regime.

'Ihis bandwidth requirement, of course, is at .- yC r * - G+q (' 'It, (13) great variance with the pitch dwing specified in MIL-K8501A (see Bf. 11).

-Ihcc- Airframe CLosed-Icop IkceleratingApproachtoEbwr x-Position Pitch Response F&sponse 'Ihis is a flight task for which the estima- tion of a simple pilot control strategy is obscured by the effects of visual perception of thus O=Y&Y+ +1 - s2+2~+lrs + or2 (14) range. Wen, et a1..13 collected numerous approach profiles, such as those shm in Fig. 8; but it is not possible to fit simple linear, and constant-coefficient models as in the previous tw cases.

S3 xu2 + (gKR - Xu)s + gKR = 0 (15) o+(l -w,)s ‘e e and with the same simplifying conditions as be- fore for the s3 and s2 terms, 25rUr I$"7 snd 'k-s (16),(17) Cbservations made for a UH-1H performing quickstops in flight9 were that

8or

‘pk - n (18) .

R max

i5 oE

starting fran 40 kt, the peak pitch-up .5 e.g., 35 during the deceleration was about 40 deg. Based R (nm) on these observations, Figure 8. Typical Approach Profiles KR-4$and I$=12 (1% Masured by &en, et a1.13 This corresponds to I+ = 0.8 rad/sec and, for It was found, hcwever. that if the "perceived range" function of Gilinsky'4 was assuned to be = 0.7, the effective crossover frequency is % operating, i.e., 0.5 rad/sec and the phase margin is 85 deg. 'Ihis is an extraordinarily high bandwidth for an Perceived x-axis task! Again applying a factor-of-five cm Range, bandwidth requirement for pitch attitude, an NOE dash/quickstep should require about 2.5 rad/sec where A is an arpirically obtained perceived nearly sn order of magnitude higher thsn Wee- range constant and R is the actual range, then the normal speed change task. Also, this value the pilot control strategy for the entire approach follcwd by hover is a sinple, Note that a value of 0.25 for K, and 500 ft stationary form suchas shok;n in Fig. 9.

for A wuld give a crossover equal to about 0.035 rad/sec at 0.5 nm, 0.065 rad/sec at 0.25 nm, and 0.25 rad/sec at hover, i.e., a steadily increasing bantiidth. It is particu- larly interesting that the model applies to a steady hover as ~11 as to the entire speed Furthermore, the above estimated transition.

value of 0 at hover agrees ~11 with the 'a made by Ringland, et simulator measurements a1.,16 using an open cockpit on the NASA knes Figure 9. tkcelerating Approach-to-Ibver Research Center s.01 six-degrees-of-freedom Control Strategy siuulator. Those data shawed hover position bandwidth wc = 0.2 rad/sec for three pilots.

A closed-form solution of the approach pro- Che 1a.s: observation for this case is that in terms of deceleration or file can be derived15 pitch attitude versus range: the supporting pitch attitude bandwidth require- ,. ment wuld be about 1.3 rad/sec, and crucial only Ki R ..

during the very last portion of the manewer.

gA0 - R = (21) the Ringland data16 (the This agrees with (1 + R/A)3 was about 1.4 rad/sec) and other measured w where Ka is sn effective pilot control strategy multiloop klytical approaches as exanplified by gain and the effective crossover frequency can be Ckaig, et a1.,17.

expressed as a function of range by Ihniliqg Qmlities hplications w =K C a'I&7iX a As a result of the above analysis, we have defined the x-axis control for three basic heli- The goodness of this model is shawn in In each case copter speed change smnewers.

Fig. 10 along with two fittings to a set of there wsre variations in cues used and in the flight data -one slightly better at long range abruptness and, therefore, the quickness required and the other at short range.

in the attitude response. This is sumnarized in Table 2.

Smm-ary of Helicopter Speed Table 2.

Change Characteristics I 0’ .125 .25 .5 R (4 Figure 10. Comparison of Deceleration Profiles Betwen Analytical Meland Flight Test lkta It shaild be noted that certain handling limit just haw aggressively the pilot performs in qualities requirements having fair agreement with a critical situation. It should be further noted present standards have been derived fran a that the pilot model arising fran the flight task direct, simple analysis of basic discrete- analysis can also be used as a tool for urunanned maneuver flight tasks. Furthermore, the parsmeters used to characterize the outer-loop corrputer sirmlation in early design very discrete umneuvers are identical in form to the stages. Ihus realistic closed-loop investiga- inner-loop regulatory or tracking hctions such tions can be conducted into "stability and as attitude control. For exauple we can deal control" and "performance" interactions.

with pilot control strategy gains, pilot ccupensation. crossover frequencies, N-e The main handling-qualities-related objective margins, etc.

of the analysis approach presented hss been to errphasize the rational, direct relationship be- Ihe very limited depth of the foregoing an- twsen a task and its supporting handling alysis nLlSt be recognized, however.

The amount qualities features.

and quality of flight data supporting the nuneri- cal results presented is grossly inadequate for setting design standards. bta for individual Siuulator Fidelity flight tasks mJst be gathered systematically for reasonably large populations of skilled pilots Simulator fidelity is a basic issue in the and various vehicle types. Ps shown, analysis field of handling qualities when flight simula- methods do not require large arrays of vehicle tion is the main source of pilot and performance state records, therefore extensive flight test data. Pbrmally simulator fidelity is established insttunentation is not really needed.

To an by focusing on the correctness of dynamic re- extent, existing flight and simulator data could sponse of the simulator motion and visual systems be reanalyzed. Useful data can also be obtained and the vehicle mathenmtical model. The result nonintrusively frm flight and simulator investi- is frequently great simulator system sophistica- gations having other priaary objectives.

tion and model complexity.

A thorough quantitative definition of heli- Che criterion for simulator fidelity is the copter flight tasks and manewers should include extent to which the sinulator induces the same those listed in ?$ble 1 with special emphasis on piloting technique or control strategy for a the critical mission segments such as NOE or given task as does the actualaircraftg. Thus we air-to-air combat or difficult operating might measure pilot control strategy in the simu- enviromnts such as nighttime, instrunent lator in the manner segested here and conpare it meteorological conditions, or extreme atmospheric to flight. This was done in the case of the disturbances.

K-10 landing maneuver" and found to reveal sig- Handling qualities are not solely tied to nificant differences accounting for landing performance problems. In addition, certain ad- "stability and control" but csn also impact "per- verse training effects were spotted in term3 of fornnnce" aspects, especially in extreme pilot control strategy.

maneuvers. Fbr example, in normal speed change maneuvers (including takeoff) or in an approach A sinulator fidelity effect which relates to the speed change maneuvers analyzed here was to hover, large torque transients due to the found in a recent set of unpublished data ob- pilot's use of collective pitch are not likely.

tained fran sn Army UK60 training simulator.

Performance of a very abrupt quickstop, on the These data, shm in Fig. 11, describe a quick- other hand, requires collective pitch applied stop maneuver as performed by an instructor flying at low altitude over a runway.

with connnensurate quickness to avoid ground-tail contact or excessive increase in altitude. Ihe specific amount of maneuver abruptness (in terms of + or w ) implied by the quickstop analysis cr presented here is likely to lead to the rotor drive-system/fuel-control coupling discussed in Ref. 18. The result may be significant rotor underspeed/overspeed transients which, in effect, sinulator specificationI this was 18 deg, and the maximus pitch attitude recorded during the maneuver was 13 deg.

Using, as an example, three specific kinds of helicopter speed change maneuvers, we have danon- strated haw each of the maneuvers can be modeled and interpreted in terns of its ok individual pilot control strategy. 'Ihe normal speed change maneuver relies only on a speed feedback loop with sane proportional-plus-integral conpensa- tion. 'Ihe maneuver is mild and requires minimal response bandwidth in the supporting pitch atti- tude regulation.

The NOEdash/quickstep contrasts greatly with the normal speed change maneuver in terms of abruptness and requires both range and closure- rate feedbacks. The pilot's aggressiveness in Figure 11. C&i&stop Phase Plane Data the maneuver calls for a very large pitch atti- Fran W-60 Training Simulator tude bandwidth in order to adequately control the vehicle. In addition. the collective pitch con- trol response required to support the msnewer in Direct inspection of the phase plane of k terms of height regulation may precipitate versus R reveals a constant slope of 0.071 engine/fuel-control deficiencies in adequately ft/sec/ft with no apparent preference for controlling rotor rpn.

range. 'lhe approximate closed-loop roots are therefore (s + I+ )(s + 0.06s)~. Equation (15) The third manewer, the decelerating approach can thus be used t8 estimate KR and KR, i.e., to hover, is intermediate to the other tm in terms of abruptness but involves pilot perception in a special way. It is shorn that the pilot 0 = s3 + S2 + g KRs + gKR - control strategy can remain relatively invariant “C e throughout the approach and ensuing hover and that the main source of closed-loop variation arises from the nonlinear effect of range ~1 c + (1 +y) s2 + 0.065s w-9 perception.

e e Ihndling qualities implications can be drawn hence in each case by inspecting the role of vehicle dynamics either in the direct response (in these cases, speed response) or in the response of (23) ,(24) supporting axes or controls (e.g., pitch attitude due to cyclic pitch change). This was denon- strated for the simple cases considered here by Comparing these values to the 4 deg/lct and applying a "factor+f-five" inner-loop/outer-loop 1 deg/ft, respectively, estimated from flight, we bandwidth criterion. Amore thorough, systematic see that in the sinulator the closure-rate feed- treatment muld, of course, be required to set back was more than an order of magnitude smaller firm handling qualities requirements.

and that the range feedback was essentially non- Simulator fidelity was also addressed in existent. Having such a disparity should, of the analysis approach illustrated terms of course, discourage any use of the sirmlator for here. The main fidelity criterion used was the that particular maneuver, but it also can help to direct, quantitative comparison of the pilot diagnose the source of stilator fidelity prob- control strategy induced in a particular simula- lems. in the case cited above, it is likely that tor versus that induced by an actual aircraft the main limiting feature was the d-ward field counterpart. Discrepancies in control strategy of view over the nose. According to the can then be used to aid in searching for specific 7. hon., 'Military Specification -- Helicopter sources of deficiencies in the simulator motion Flying and Qound Ikndling wlities; Gsn- MIGH-8501A.

or visual systars or in the computer mathematical era1 Requiremmts For," models of the vehicle and envirmnt.

/me&sent 1, April 3, 1962.

It is smested that the general approach 8. Edenborough, H. K.; and Wsmicke, K G., illustrated here be applied in a broader, more "Control and Manewer Wquirements for Armed thorqh manner to the field of handling quali- AHS Twentieth Annual National Helicopters," ties. The approach provides a rational way to Fonnn, Wsshington, D. C., May 13-15, 1964.

account for the handling-qualities needs in sup porting a given flight task. It also offers a I-kffley, &bert K.; Clement, Warren F.; 9.

means for evaluating the validity and effective- Ringland, Ibbert F.; Jewzll, Wayne F.; Jex, ness of flight simulation tools which nust be I&n-y R.; McRuer, bane T.; and firter, handl'ing qualities used in establishing Vernon E., "Determination of I&ion and requirements.

Visual System Requirements for Flight Train- qFa2ms l?2chnology, Inc., ing Simulators," Technical Report No. 1162-1, August 1981 R2fe?xmces (forthcaning AR1 report).

1. Stapleford, R. L.; PbRuer, D. T.; I-bh, R.

mbert Id; Schulmsn, Ted M.; 10. Heffley, H.; Johnston, D. E.; and Eeffley, R. K., Randle, Ibbert J., Jr.; and Clement, Warren "(Xltanarting MILF-8788(ASG), the Military F "An halysis of Airline Ianding Flare systems Flying Qalities Specification," I& Based on Flight and Training Simulator Technical Report Technology, Inc. , Measurenzents." systgas lkchnolofg, Inc. , No. 190-l. August 1971.

Technical Report No. 1172-lR, March 1982 (forthcoming NASATM).

"Bquivalent System Approach 2. Ibdgkinson, J., for Flying @lities Specification," SAE 11. &ffley, Robert K, "A Compilation and An- Control and Olidance Systems Cbnnsittee of I-klicopter Handling @alities alysis (MCAIR Meting, Dznver, CO, March 1979, Volune Two: IF&a Analysis," D&a.

Paper 79-017).

NASACR-3145, August 1979.

3. Ibh, R. H.; Mitchell, D. G.; and Ibdgkinson, 12. Nagata, John I.; sinner. Gary L.; Buckanin, M "Bsndwidth - A Criterion for Highly Robert M.; Robbins, Robert D.; and Williams, A&nted Airplsnes," AIAA Paper Ibbert A., "Airworthiness and Flight &arac- No. 81-1890, AIAA Atmospheric Flight P&h- teristics tialuation, UH-6oA (Black Ihwk) Conference. Albuquerque, NM, anics Final Report on USAAEFA Project Helicopter," hgust 19-21, 1981.

No. 77-17. September 1981.

4. George, Rrank L.; and worhouse, &vid J., 13. I%en, Gene C.; DiCarlo, l&niel J.; and "Relationship of the Flying wlities @eci- Yenni, Kenneth R., "A Parametric Analysis of fication to Task Performance," AIAAW3rkshop Ihlicopters," Visual Approaches for on Flight Testing to Identify Pilot &rkload NASA TN b8275, December 1976.

and Pilot I@zmics, AE'FTC. Edwards AFB, CA, Jsnuary 19-21, 1982.

14. Gilinsky, Alberta S., "Perceived Size and Distance in Visual Space," Fsychological 5. hon.. 'Military Specification -- Flying Review, Vol. 58, 1951. pp. 460-482.

Airplanes ," Qalities of Piloted MIFF-878%. November 5. 1980.

15. Heffley, &bert R., "A tide1 for Manual Dzcelerating Pgproaches to l-bver," Proceed- 6. Chalk, Charles R.; Key, bvid L.; Kroll, ings of the Fifteenth Annual Conference on John, Jr.; Wasserman, Richard; and Radford, &nual Control, AFEDL-TR-79-3134. mvember Itbert C., "Background Information and User 1979, pp. 545-554.

Olide for MIIrF-83300Military Specification -- Flying C&alities of Piloted V/STOL Aircraft," AFFDLTF&70-88, March 1971.

16. Ringland, R. F.; Stapleford, R. L.; and Bgdaleno, R. E., V&ion Effects on an IFR tivering '0sk -- /malytical Predictions and erimental Fk33iks," NASA CR1933, bvmber 1971.

17. &aig, Samuel J.; and Campbell, hthony, "!nalysis of VIOL Handling @alities Requirements. Part I: Iongitiinal ever and YBxnsition," AFFDL-TR-67-179, Part I, October 1968.

8. Kuczynski, W. A.; Cooper, D. E.; 'Ikomey.

W. J.; and Howlett, J. J., "The Influence of kgine/fiel Control &sign on I-klicopter Qnamics and %ndling Qalities," Journal of the American Helicopter Society, ml. 25, No. 2, April 1980, pp. 26-34.

19. schalm, P. s., "Specification for the UlTAS I-klicopter Synthetic Flight Training System, tkvice B38." NIEC 2222-l 152, &tober 30.

1975.

E -

Applications of System Identification Methods tp the Prediction of Helicopter Stability, Control and Handling Characteristics G. D. Padfield RAE Bedford, UK R. W. DuVal NASA- Ames Research Center Abstract -solidity ;1 -time to half amplitude (s) The paper describes a set of results -trim values of aircraft forward u;, we from the first phase of an RAE/NASA colla- and normal velocity components borative programme on rotorcraft system (ft/s, m/s) identification that has the main objective -velocity perturbations along ua v*w of improving prediction methods. Flight body x,y and z axes (ft/s, m/s) measurements collected at RAE Bedford on -sideways and vertical velocity Vo 'wo an experimental Puma heliCODter are components (ft/s, m/s) reviewed and some notable characteristics -control vector (equation (7)) u highlighted. Following a brief review of -state vector (equation (7)) x previo& work in rotorcraft system identi- -sideslip angle (rad) % fication, the results of state estimation -rotor longitudinal and lateral 61c' e1s and model structure estimation processes disc tilts respectively (equa- applied to the Puma data are presented. tion(1)) The results, which were obtained using -Lock number (equation (1)) Y NASA developed software, are compared with ri -relative damoing theoretical predictions of roll, yaw and -longitudinal-and lateral stick n1s' nlc pitching moment derivatives for a 6 degree position (% aft, std) of freedom model structure. Anomalies -pedal position (% left) reported in other investigations have -longitudinal and lateral cyclic reappeared in this study. The theoretical pitch angles (equation (1)) methods used are described in the Appendix -aircraft pitch angle where a framework for reduced order model- -system eigenvalue ling is outlined.

-rotor downwash (normalised) velocity (equation (10)) Notation -flap frequency ratio -normal velocity component at A -state matrix (equation (7)) rotor disc (equation (10)) B -control matrix (equation (7)) -air density -rotor induced torque coef- -relative air density CQi ficient (equation (10)) -aircraft roll angle -rotor thrust coefficient -rotorspeed CT (equation (3)) -dutch rollfrequency (equation -fin sideforce coefficient (9)) 'YFN (Fig 5) F -force and moment vector Introduction (equation (7)) -gravitational acceleration A collaborative programme between -height of aircraft cg below :R the Royal Aircraft Establishment and the rotor (ft,m) (equation (7)) NASA - Ames Research Center to develop and I I I -moments of inertia of air- xx' yy' zz exchange information on rotorcraft system craft in roll,pitch and yaw identification is now underway with the (slugs ft2. kg m2) main objective of improving prediction L v, -rolling moment concise deri- Lpetc methods. In the first exercise of this vatives (normalised by I,,) kind at RAE, flight tests have been made Mu, -pitching moment derivatives Mwetc on a Puma helicopter and the results have (normalised by I been analysed at NASA - Ames, using soft- -pitching moment &ivative Nslc ware developed in support of the Rotor wrt flapping (normalised by Systems Research Aircraft programme.

IYY) This paper reports on the first phase of Nv, -yawing moment derivatives Nr etc The paper begins by des- this activity.

(normalised by I,,) cribing the aircraft and data processing -yawing moment from main rotor NRi system at RAE Bedford and attempts a pre- (normalised byI,,) liminary interpretation of selected flight -roll, pitch and yaw rates P¶ q,r data based on comparison with theory.

(rad/s) Aspects of rotorcraft system identifica- R -rotor radius (ft,m) tion are then reviewed and the techniques R2 -(multiple correlation coef- ~. . currently in use are described.

ficient)d Derivatives, estimated by these techniques Data pertaining to the trim condi- for the Puma at a nominal 100 kn trim con- tions for flight 325 are given in Table 1.

The nominal t%m IAS was TOO kn from which dition, are presented and compared with a series of shaped control inputs were made results predicted by a linear, 6 degree of and the ensuing response measured and re- freedom theoretical model developed at RAE1. Some of the major anomalies are dis- corded. The inputs-included steps, doub- cussed. Finally the use of reduced order lets and "3211" multistens and aenerallv linear models for handling studies is produced repeatable response patterns; " reviewed, and approximations to the natural atmospheric conditions at the test points were very smooth.

modes discussed in some detail.

Time to recovery varied with input type and Flight Mechanics Investigations with size but typically step inputs gave about Puma XW241 at RAE Bedford 10 s of data and 'return-to-tri.m' multi- steps gave about 20-30 s duration.

The Aircraft and Data Processing System Typical results from cyclic and pedal in- Puma XW 241 (Fig 1) is a multi- puts are reproduced in Figs 3 and 4. Before purpose experimental aircraft operated and introducing the system identification pro- managed by the Helicopter Section of the cess a preliminary interpretation will be attempted on these results together with a Flight Research Division at RAE Bedford.

limited comparison with theoretical Since its procurement in 1974, the Puma predictions-.

has been used in a variety of research pro- grammes and has recently been fitted with a digital Pulse Code Modulated (PCM) Preliminary Interpretation of Flight recording system; the results described in zsults and Comparison with Theory this paper were obtained with this equip- ment. A block diagram highlighting the An initial assessment of the flight results and comparison with theory provides features of the current data acquisition a suitable background to judge the effect- and processing system is shown in Fig 2.

iveness of both the theory itself and the system identification method, described The instrumentation system includes three packs of linear accelerometers of later. The theoretical results presented both ac pickoff and force feedback type, in this paper are based on a 6 degree of two packs of rate and attitude gyros and a freedom linearisation of the nonlinear simulation model described in Ref 1.

heading gyro. Air data sensors give air- During the development of this nonlinear speed measurement from a conventional model, a degree of validation was achieved pitot static and incidence and sideslip through flight/theory comparisons of Puma angles from vanes located on a nose boom.

data; these results are reported in Refs All four control displacements, the three 1 and 2.

swash plate jacks and the tail rotor pitch jack are sensed by potentiometers. In addition, one of the four blades is cur- The short term pitch and roll rate response to longitudinal and lateral cyclic rently instrumented to measure flap, lag steps inputs respectively are shown in and feather angles at the bearings. Samp- Fig 3. The comparisons with theory are ling rates vary from 32 to 256 per second, encouraging and indicate that the principal the lower rate restricted to slowly vary- ing quantities, eg airspeed. The current damping, control sensitivity and 'static overall rate for the multiplexed system is stability' parameters ought to be predicted 8024 words (12 bit) per second but the with reasonable accuracy by theory, There is some evidence that the 'initial' angular pattern of signals in each data field is acceleration is sharper in theory but this flexible.

is to be expected with a quasi-steady rotor During the Spring of 1981 a trial model. This phenomenon and the difficul- was conducted with the object of data ties it can present to a derivative estima- gathering for a system identification tion process are discussed in more detail exercise. Trims and responses to pilot in the next section.

shaped control inputs over a range-of Turning now to the flight results flight conditions were recorded. Data tapes from one of these flights, designa- shown in Fig 4, we see the coupled response to a doublet input applied to the pedal.

ted 325, were sent to the Ames Research Centre for analysis using NASA software This set of data reveals the presence of a during the first round of the RAE/NASA lightly damped lateral/directional oscilla- collaborative programme in Rotorcraft tionthatwas very apparent to the pilot and Flight Mechanics. Results described in observer performing these tests. This this paper are drawn from this exercise oscillatory Idutch roll' mode is also pre- dicted by theory and a comparison of the and used‘in subsequent interpretation, measured and predicted characteristics is analysis and comparison with theory.

It can be seen that the reductions impose a positive yawing moment shown in Table 2.

on the fuselage when the yaw rate isamaxi- lateral/directional characteristics are mum, giving a negative damping contribu- roughly in agreement whereas the longitu- dinal couplings, particularly the phase tion. This effect is discussed in more - _ _ The relationships are significantly different. quantitative terms in the Appendix.

above description of the increased insta- Having establjshed possible basic ingred- ients of this mode it is useful to examine bility of the lateral oscillation is de- the theory to determine the origins. The pendent on the correct phase relationship between longitudinal and lateral variables.

derivative matrix used in the analysis is given in Table 3 and the system eigenval- As highlighted earlier this is the area of ues in Table 4. Included in Table 4 are most serious discrepancy between flight and theory. It should perhaps be emphasised the longitudinal and lateral subsystem eigenvalues derived with coupling terms that engine/rotor dynamics are not included set to zero. Conventional fixed-wing air- in the t eoretical model but that there is k craft terminology is used to describe the evidence that including this degree of different modes. The frequency of the freedom can reduce the effective yaw Included in Fig 4 is the main dutch roll oscillation is determined pri- damping.

marily by the directional stability, ie rotor torque variation (via shaft strain gauge measurement) during the oscillation.

- q q 1.011 rad/s.

Although fairly noisy, the fluctuations Wdutch roll are seen to be-phased-relative to yaw rate, so that as the aircraft is yawing to star- An important effect displayed in board, the engine is applying a clockwise Table 4 is the considerable reduction in torque (minimum) to the fuselage. This is damping of this mode when the coupling in the same sense as predicted-by theory terms are included. The damping is rather anditmay be surmisedthat the phase shift low even for the lateral subsystem and relative to pitch rateand vaneoscillations this can largely be attributed to the low is accounted for in the engine dynamics.

value of the yawing moment derivative NV.

This derivative contributes to the mode As a final note in this exploratory damping when the effective centre of the section some measurements were made on the oscillation is a condition of non-zero Puma following a pedal doublet input with sideslip. Wind tunnel data3 for the Puma the pitch axis stability augmentation fin in sideslip, as used in the theoreti- switched on, to give some indication of cal model.

is shown in Fig 5. The side- the effect of reducing the longitudinal force is strongly nonlinear-with sideslip, coupling. A comparison of yaw and pitch but over the range -5O < 13 < t8O. nracti- vane responses with and without augmenta- tally no lift is-produced by the-fin. It tion is shown in Fig 8. Clearly the damp- is believed that this effect is due to the ing of the oscillation has increased with suction on the rear of the 'lower surface' the reduced longitudinal coupling. In at small angles of attack; a characteris- fact, the time to half amplitude-has now tic of thick aerofoi sections with large reduced to about one cycle, ie approxi- i trailing edge angles . This aspect is mately the same as predicted by thelateral discussed further in Ref 5, in relation to subset calculation. The frequency is prac- the Puma and how design improvements tically unchanged. Longitudinal coupling obviate the effect on the AS 332 Super obviously does play an important role in Puma.

the damping of this oscillation.

Returning to the coupled system, a Rotorcraft System Identification - sensitivity analysis reveals that the main coupling derivatives affecting the dutch The state of the art in the field of roll damping are Mp and NW . For a system identification has largely been perfectly governed rotorspeed the deriva- developed in the fixed wing community, the tive NW reflects the torque changes pro- most recent comprehensive review of which duced by the powerplant, following rotor- can be found in Ref 7. There have been speed variations due to incidence perturba- several attempts to apply the various tions. Theeffectof this derivative on techniques to helicopters, a review of the dutch roll and longitudinal short per- which also appears in Ref 7. Before des- iod eigenvalues is illustrated in Fig 6.

cribing the methods used in the RAE/NASA is destabilising for the dutch roll NW collaboration it is perhaps worth making a mode and the roots are close to the longi- few observations on these past efforts and tudinal and lateral subset approximations highlighting some of the lessons learned.

when NW is zero. Fig 7 illustrates how the phase relationship between w and r Observations on Previous Work (incidence and yaw rate) perturbationsvary with NW for the dutch roll mode. They The ground rules for helicopter are seen to converge as NW increases and system identification were laid down in the clearly when they are in phase, torque pioneering work by Molusis in the early and lateral cyclic pitch respectively.

seventies8-lo. In his papers, Molusis Non-uniform inflow effects have been neg- emphasised the need for long data records lected for this example.

and combined manoeuvres in order toprovide Constraining the sufficient information for reliable esti- body to rotate only in pitch and neglecting mation. He also stressed the importance forward speed effects, the above equations of providing satisfactory initial esti- can be augmented by the body pitching moment equation mates of derivatives in maximum likelihood algorithms by using an optimal filter/ smoother in conjunction with a least ;1 q +M (2) squares estimator. Of particular rele- f3p YY vance to the helicopter problem were Molusis'sobservations on the effects of is the pitching moment of inertia and IYY rotor/fuselage coupling on derivative esti- the pitching moment derivative can mation. These were clearly demonstrated MBlC by attempting an identification of a 6 be written in terms of the rotor thrust coefficient degree of freedom (dof) model from simula- 'T ' tion data that included rotor flapping modeslo. - h Q(QR)~vR~C The identified derivatives were MBlc q R T ' substantially different from quasi-steady predictions even though time histories Here hR is the cg distance below the showed reasonable correlation. Including rotor, p the air density and R the the three principal multi-blade coordinates rotor radius. For present purposes we in the assumed model and performing the have assumed in equation (3) that the reduction to a 6 dof model after the iden- rotor thrust remains normal to the disc tification, resulted in very good agree- during pitching motions.

The quasi-steady ment with quasi-steady values. The con- form of equation (2) can be written in the clusion was not that 6 dof quasi-steady form theoretical models were necessarily inade- quate but that profound difficulties could be expected in trying to estimate these M from flight measurements of fuselage ME Bls as1c lc aBlc motion alone. Molusis suggested an empiri- where M =----as> M =-- 9 I cal correction based on computed theoreti- els Iyy aels * YY IC\ cal data but his own results showed that . . . . . \,I this was not very satisfactorylo. A From equation (1) the quasi-steady flapping perusal of the simulation results in derivatives can be written as Ref 10 reveals that perhaps the most ser- ious discrepancy is the underestimation of

aOlc _

--

the primary-rate damping derivatives by aO1c - 16/ya --

- 1. (6)

,

aq

the 6 dof identification. These usually aels produce dominant effects about all axes" and relatively simple theory should be To demonstrate the phenomenon in question we A gross underestimation fairly reliable.

now attempt to estimate the derivatives Mq of this effect will obviously lead to a and in a model structure given by “I@& further corruption of secondary effects in equation (4) from data generated by equations each equation when the time histories are (1) and (2), following a step input in els .

forced to agree. A possible cause of this anomaly can be demonstrated by a fairly Configuration data used for the numerical simple example.

study are y q 8.0 , Q q 27.8 rad/s and q -6.36.

Using a least squares Consider a hovering rotor with zero flap- MQc'lyy ping hinge offset and centre of gravity estimator, the effect of data length on the (cg) on the rotor shaft. The first order estimated derivatives is shown in Fig 9. A flapping equations in multi blade coordi- sampling interval of 0.04 s was used for nates and that represent this case. The effect is dramatic, particu- Blc Bls ' larly on the estimated damping M, .

longitudinal and lateral disc tilt respect- Using only-a short data length a-positi;e Mq is ively, can be written as predicted and even after 2 s the estimation is still only 70% of the quasi-steady value.

For a 2 s data run, the effect of sampling interval on the estimation is shown in Fig 10. Increasing the sampling interval to 0.2 s results in an accurate estimation - . . . . . ..\I.

of the quasi-steady derivatives.

Obviously where y is the rotor Lock number, sl the the model structure given by equation (4)is rotor speed, p and q the body roll and inadequate for portraying the short term the longitudinal pitch rates and elsJ % dynamic effects of the flapping equation low as 2 Hz. Some exploratory work was done on this theme by one of the present Cl), which in this case produces a mode with frequency 5.5 rad/s and time to half authors in Ref 14, where the estimation amplitude of about 0.1 s. A comparison of process was transferred to the frequency pitch rate response using the full equa- domain. The data can thenbe filtered by tions and the quasi-steady representation excluding data beyond the frequency range of interest. This idea will now be exploi- is illustrated in Fig 11. Clearly the ted for the analysis of the Puma data.

angular acceleration in the former case is zero at the origin and builds up to the Current Methods quasi-steady value after about 0.2 s.

Using data in this interval for estimating quasi-steady derivatives tends, therefore, The techniques used in this paper are based on the system identification approach to underestimate the damping in the motion to the extent that for, very short data described in Ref 15, where the data analy- lengths, an unstable system is predicted sis is divided into three main stages.

as shown in Fig 9. state estimation, model structure estima- tion and parameter identification. The When analysing flight data one would, state estimation stage involves the recon- of course, struction of all states and controls from use a much longer data record than used in the previous example and the the available sensor me surements using a step input is rather limited in its excita- Kalman filter/smoother 15 This process tion spectrum. also eliminates the biases and scale factor However, a multi-step input would typically extend over about 30% of errors from the measured data and reduces the data record and the contamination the level of measurement noise. The second effects described above might persist well stage utilises a ste wise least squares P into the record. regression technique 7 to derive an ade- quate model structure from a range of pos- Several other attempts to identify sible candidates. The method uses certain 6 dof helicopter models have resulted in statistical criteria to determine which underestimation of primary damping deriva- states should be included in the model and tives relative to the quasi-steady values estimates the corresponding coefficients.

The third stage adopts these estimates as (Refs 11-13). All these references report using advanced statistical methods for starting values for a maximum likelihood estimating derivatives but the phenomenon estimation process that generates unbiased, efficient, parameter estimates.

is still present in the results. The obvious solution to the problems high- The analysis of results described in lighted by Molusis is to measure individual the next section will be confined to the blade motions and to use higher order model structures in the estimation process. There first two stages described above. The equations used in the state estimation pro- will obviously be flight cases when such a cess are essentially kinematic relation- solution is mandatory to the identification of certain phenomena, particularly when 6 ships. For the present data set the recon- dof theoretical models are known or sus- structed states include the three velocity pected to be inadequate. components and the translational and rota- For a wide range of conditions within the normal flight tional accelerations of the aircraft centre of gravity. The external forces and moments envelope, however, it is hoped that 6 dof are then directly related to the accelera- models can be used to predict handling characteristics adequately and it is worth- tion estimates and for the model structure estimation we choose the 6 dof formulation while exploring other potential solutions to the data processing problems before of the form resorting to the measurement of blade E q AX + BE motions. One approach is to use an input (7) signal with a frequency content that excludes the rotor modes. A possible F x and u are the external force and ;;;d,eEt vecto& , vehicle state vector and shortcoming of this method is that the data could be starved of information necessary control vector respectively. The regres- sion process currently used treats each to predict some of the derivatives, par- ticularly for hingeless rotor helicopters equation separately, but rather than find where the lower frequency rotor modes are the coefficients of this limited model that are valid for the entire frequency range of not far removed from some of the body modes. Nevertheless, the use of smooth the data, it is more appropriate to deter- inputs rather than sharp steps or ramps mine the best set for the frequency range of interest. This is readily accomplished has an obvious appeal in this context. A similar and perhaps complementary solution by transforming the data to the frequency domain using a Fast Fourier Transformation involves filtering the data to exclude the 'higher' frequency content but again this (FFT) and then truncating the data beyond ' could mean coming down to frequencies as the bandwidth of interest.

If it is assumed that the coefficients do not vary be seen that only two cycles of the lat- with frequency then the regression can be eral oscillation are available.

Two fur- performed in the frequency domain with the ther points are worth noting; a bias has vectors in equation (7) interpreted as the been detected and corrected for in the corresponding Fourier transforms. An pitch rate gyro and the initial slope of additional advantage of estimation in the the normal velocity component (w) indicates frequency domain is that the resolution an initial climbing and decelerating flight can be improved by padding the data with condition. The ef?ect of this unsteady - zeros. This effectively forces the spec- initial condition has not been explored.

tral lines to be more closely spaced with- In discussing the results of the model out affecting the frequency content of the structure estimation we will refer to the data. The higher resolution should result multiple correlation coefficient (R) for in more accurate estimation of the low the fit. In essence, the closer R2 is to frequency modes, which are difficult to unity:, the better the overall fit of the from a limited duration data data in the frequency domain.

Table 5 summarises the estimated rolling moment coefficients (derivatives) for the cases studied, along with the current quasi- Model Structure Estimation with Puma Data steady theoretical predictions.

Deriva- tive estimates with a structure containing The results presented in this sec- only lateral variables are compared for tion were produced at the Ames Research thre; tffe;u;;cy ranges, 0 to 0.5 Hz, 0 to 1 Center and forwarded to RAE Bedford for . In all cases the high R2 Hz, The inherent communica- values indicate a reasonably good fit with interpretation.

this limited model.

tion difficulties in this process limited The roll damping Lp the scope of this first phase of the col- is seen to increase by about 50% as the laboration. The analysis software is cur- bandwidth is reduced but is still markedly ' rently being implemented on a VAX computer lower than the theoretical prediction. The at Bedford when these difficulties should Fourier transform of the rolling moment is compared with the estimated fit in Fig 13, be largely overcome.

for the two lower frequency ranges.

Most of the data is contained in the range 0 to From the assortment of data collec- ted during Puma flight 325, three man- 0.5 Hz, and the dominant peak at 0.25 HZ oeuvres were selected for analysis, namely corresponds to the lightly damped lateral doublet inputs in lateral cyclic, pedal oscillation. The fairly close fit at and and longitudinal cyclic. The data used above this frequency appears to be achieved consisted of measurements from one of the at the expense of the fit at lower frequen- cies. Referrin packs of inertial instruments (accelero- again to Table 5, it can RS be seen that meters, rate gyros, attitude gyros), the rises to 0.96 when the longitudinal variables w airspeed sensor and the pitch and yaw and q are vanes. The instrument pack referred to added, with some associated modification to included a 2-Dole Butterworth filter the lateral variable coefficients.

Compar- - 10.6 Hz- 5 = 0.73) that effectively ing the estimated and theoretically predic- (tin removed the do;inant b/rev noise in the ted derivatives in Table 5 we can see that the major anomalies are for the rate der- measurements. The results of the model ivatives structure estimation process reveal a and Lr . The magnitude of L? these effec s seems to have been reversed marked variation in estimated coefficients, particularly the contributions from coup- for the flight results but it is difficult ling effects. The reasons for this are to rationalise the physical significance not yet understood and we therefore choose of this; indeed the result seems rather dubious; to concentrate on a synthesis of the prim- Further exploration into the details of the regression analysis is ary moments for each input, ie rolling clearly required.

moment for lateral cyclic doublet, yawing moment for pedal doublet and pitching moment for the longitudinal cyclic doub- Pedal Doublet/Yawing Moment Synthesis - The sample rate for all-channels let.

was 64/s, a rate that limited the record The smoothed state estimates for this manoeuvre are illustrated in Fig 14 and lengths to.about 15 s. This limitation the yawing moment derivatives derived from will be removed in future analysis. The the lower frequency bandwidth of data are results from each manoeuvre will be dis- compared with theoretical predictions in cussed in turn.

Table 6. Results from two different model Lateral Cyclic Doublet/Rolling Moment structures are included, one with lateral Synthesis variables only and the other with the addi- tion of the longitudinal variables w and The multiple correlation coefficient Results from the filter/smoother 9 - is hardly affected by these additions.

process are shown in Fig 12, where it can

I -

The least squares in the frequency domain.

It can be seen that the directional stiff- regression analysis does of course have its ness (NV) and damping (Nr) compare very shortcomings, particularly when noise is well with theory but the control power nresent in the datal8, but also when there (N ) is somewhat larger according to nP How- is strong correlation between states.

The most striking differences are theory.

ever the technique is appealing in that it in the coupled rate derivatives Np and offers a simple and systematic approach to The small value of Np predicted by model structure estimation, and investiga- Nq - theory is the result of the cancellation tions of the type described above will be of two larger effects from the main (nega- pursued in the continuing collaborative tive NP) and tail-rotor (positive N ).

When fully exploited and under- programme.

It is-interesting to note that the P va ue stood, the results of the model structure of R2 was raised to 0.91 by the first two estimation stage will then be used to variables drawn into the regression, v initiate the more complex maximum likeli- and np , the remaining variables then hood process.

being added to account essentially for the The model structure estimation stage out of phase component of the response.

serves another useful purpose in the vali- For the lightly damped yawing mode the sum dation of reduced order approximate models of these contributions should obviously be Simplified of helicopter flight mechanics.

small. Once again, a detailed breakdown which still indicate trends models, of the in-phase and quadrature components accurately, have obvious advantages and of these 'damping' terms is required to various schemes are suggested in the aid further interpretation of the results.

Appendix for the 6 degree of freedom heli- Longitudinal Cyclic Doublet/Pitching copter as a framework for analytic model Moment Synthesis The results described are based reduction.

on the theoretical predictions of Puma The state estimates derived for this characteristics discussed earlier in the manoeuvre are shown in Fig 15, where the Clearly, however, we have failed to paper.

bias error in the pitch rate gyro is again validate the theoretical arguments put for- apparent; also the normal and longitudinal ward to explain the low damping of the velocity estimates again indicate an initial lateral oscillation. In particular, the decelerating climb. The least squares and Nq in yawing moment derivatives NW estimates in the frequency domain are com- to their Table 6 bear little resemblance pared, in Table 7, with the quasi-steady theoretical counterparts, and, unfortunately theoretical predictions for the derivatives.

for this study no estimate was obtained for Results are presented for the longitudinal (another important the derivative Mp model structure for the three frequency effect in the theory). There is no reason ranges. Both the pitch damping (Mq) and to believe that any major physical effect control power (Mnls) are seen to increase has been neglected in the theoretical quasi- as the bandwidth is reduced and both rise steady derivatives, but it is possible that to nearly 80% of the theoretical predic- the omission of the rotorspeed degree of tions. The remarkable agreement for the freedom has distorted the model structure speed stability derivative (Mu) is somewhat estimation. With this possibility in mind, overshadowed by the incidence stability future estimations will include this addi- derivative (M,) comparison.

The variable tional state.

W actually just managed to become part of Concluding Remarks the fit, accounting for on1 the last few per cent in the value of R3 . It is pos- Results from the first phase of an sible that the estimation of an unstable RAE/NASA collaborative programme in rotor- is related to the unsteady initial con- MW craft system identification have been des- ditions for this manoeuvre, as the effects cribed. Flight data from three manoeuvres of the initial deceleration persist for the with an RAE Puma helicopter have been pro- duration of the record as shown in Fig 15.

cessed by the state estimation and model It is intended to explore this topic structure estimation processes of a NASA further.

system identification software package.

A comparison of moment derivative estimates The results from the three manoeuvres with theoretical predictions has been used discussed above are both encouraging and as a guide to the likely accuracy involved.

perplexing. Unfortunately, time was not No firm conclusions can be drawn from this available to explore further the uncovered first exploration but a number of features anomalies, for this paper.

The underesti- are worth highlighting.

mation of roll damping, relative to theory, a feature common to several earlier Encouraging results have been obtained (1) attempts at rotorcraft parameter estima- by performing the regression analysis in tion, is perhaps of greatest concern.

The the frequency domain. Reducing the band- reduction effect of the lower frequency width of data used resulted in an increase flapping modes, described in an earlier in roll and pitch damping and control power, section, should be minimized by truncation as expected from time domain considerations, performing such a 'partitioning' analysis but, in some cases, the results still fall can often stem from the enhanced physical well short of expected quasi-steady values.

appreciation it inspires. In this Primary yawing moment derivatives (2) Appendix we concentrate on the theoretical estimated from a pedal doublet manoeuvre results discussed earlier and outline a show good agreement with theory but the framework for reduced order analysis. One adverse yaw (Np) and coupling from longi- can clearly imagine the 6 dof model as tudinal motion (Nw,Nq) show significant already being a reduced order approximation differences. The contribution of these of a higher order system containing rotor coupling terms to the 'dutch roll' mode as well as engine/fuel control system damping, so elegantly expressed by simple degrees of freedom.

theory, cannot, therefore, be The method of analysis is described substantiated.

in Ref 19 where the concept of weak coup- Roll damping (Lp) and control power ling is introduced and conditions of I? ) estimated from a lateral cyclic application are quantified. The technique 'Ilc has been applied to strongly controlled doublet manoeuvre are considerably lower aircraft motions20 and more recently to than theory predicts - a result in keeping describe the range of application of the with previous reports.

This anomaly in the longitudinal short period approximation for rolling moment is accompanied by a very helicopters21. In the present paper, a high estimate for the derivative Lr and form of approximation for lateral/direc- the temptation at this stage is to question tional motion is sought. For the Puma the structure estimation, rather than results described the strong coupling from theory, for these effects.

longitudinal motion renders the search in The closest agreement with quasi- (4) vain. In the following, aspects of open steady predictions were obtained for pitch- loop stability characteristics only will be ing moment derivatives from a longitudinal addressed.

cyclic doublet manoeuvre, except for the The reduction process is based on a static stability derivative M, which is partitioning of the system matrix into estimated to be very small and positive by lower order subsystems that are weakly the least squares regression.

coupled. Details of the method can be A theoretical framework for exploring (5) found in any of the references cited above reduced order model structures has been and will not be elaborated upon here. In outlined in an Appendix to the paper.

order to achieve the correct partitioning A coupled longitudinal/lateral fourth order for rigid body modes of motion we need.to system is required to describe the current introduce the vertical velocity w0 and theoretical predictions of the Puma 'dutch sideways velocity v0 as new variables, roll' oscillation.

replacing pitch attitude and yaw rate in the equations of motion.

The system The collaborative programme is still matrices, in partitioned form, for decoupled at an early stage and more detailed longitudinal and lateral motions are shown investigations using the model structure in Table 8, along with the approximating estimation process are planned for differ- characteristic polynomials. One would not ent manoeuvres of longer duration and poss- expect these formulae to give very accurate ibly improved model structures.

results in general but they do serve as a Appendix yardstick against which the effects of coup- ling terms can be measured. For the Puma The Use of Reduced Order Theoretical Nodels derivative data given in Table 3 the weakly The complex nature of helicopter coupled approximate results are shown in flight mechanics make it a prime candidate Table 9.

for treatment as a sum of interacting sub- A comparison with the decoupled longi- systems, with the attraction that phenomena tudinal/lateral results given in Table 4 may be described by considering a series of indicates again that the dutch roll damping lower order problems and their interactions.

is badly o&predicted. This poor compari- Conditions under which this type of approx- son will result when the oscillatory side- imation is valid are often based on intui- slip motion has a significant component of tive reasoning but they can be formulated sideways motion superimposed on the side- more precisely using notions of subsystem slip due to yawing motion. The overpredic- dynamic separation (eg widely separated tion of the phugoid damping can be attri- characteristic times) and interaction buted to a similar effect in the longi- strength. For aircraft flight dynamics tudinal plane. For the other three modes such opportunities often arise for des- the approximations are clearly satisfactory.

cribing rigid body/aeroelastic interaction The importance of the directional stiffness OP in the description of the individual on the dutch roll damping is not, of modes using familiar arrangements of motion NV course, predicted at all by this type of variables, eg longitudinal short period mode approximation. The root loci in Fig 16 made up of incidence and pitch rate excur- illustrates this effect; the two curves sions. Once again, the attractions of above system are, shown are for the coupled longitudinal/ lateral system and for the lateral system x - 1.0 + 1.079i alone. Both loci are approaching the short period q asymptotic value given in Table 9 as Nv - 0.114 + 1.115i ie dutch roll damping z (Nr +Y,).

increases, 'dutch roll q The point being made is that as directional giving the right order of damping reduction stability increases the formulae given in for the dutch roll mode. If we assume a Table 8 are not only becoming better neutrally stable oscillation exists for the approximations to the lateral subset damp- above system thenagood approximation to ing but that both are improving relative this damping decrement can be obtained by In these con- to the fully coupled result.

deriving the steady state frequency res- ditions the other modes are also approxi- ponse of equation (8i. For this case w mated fairly accurately, eg spiral mode is approximately 180 out of phase with marginal stability is well predicted, i (hence in phase with yaw rate) and the although these results are not shown here.

effective damping becomes, The results discussed above indicate that the degree of coupling between longi- damping(dutch roll) * tudinal and lateral motion can be strongly L N M influenced by the directional stiffness or, in other words, by the frequency of the + $ (g - NpUe) + p Nr, + Yv lateral oscillation. This result is some- P PWO what intuitive and for the present coupled system the coincidental similarity between . . . . . . (9) the longitudinal short period and dutch = 0.234 , roll frequencies is bound to result in hence strong longitudinal/lateral interactions.

R,(X) * - 0.117 As discussed in an earlier section these interactions are brought about mainly (where w. is the dutch roll frequency) through the derivatives Mn and NW ; the main rotor contribution torthese deriva- which agrees with the result given by tives being negative and positive res- equation (8) above.

pectively for 'clockwise' rotating rotors.

The strong coupling present leads to a The derivative NW , as stated further reduction in dutch roll damping earlier, represents the quasi-steady torque through the mechanism discussed earlier.

variation produced by the engine, in res- This effect can be quantified by consider- ponse to rotorspeed variations. The ing a reduced order system made up of both validity of this implicit weak coupling the longitudinal and lateral 'fast' assumption can only be assessed when an oscillatory modes.

engine/rotorspeed control system model structure, representative of Puma, is Assuming on the one hand that trans- itself incorporated. The derivative NW lational velocity excursions (u,vO,WO) are is produced mainlv by the 'induced torque', much slower and weakly coupled, and, on written in coefficient form, the other, that the roll subsidence approximation interacts in a quasi-steady cQi = manner, the 4th order approximate system - CThz - x0) . (10) takes the form, W 2 0 0 zW + "e q dq -M L /L MpLr/U L q Q.

MW Mq Pv P ep ZE - V 0 0 0 0 + (LV/Lp)(g-NpUe) > NY + yv + (L,'Lp)(g-NpU L.+l kNwUe - ‘eNv ( . . . ...(8) is the rotor thrust coefficient and Here we have neglected all coupling terms CT X0) is the upwash, normal to the disc.

except Mp and NW , but with the strong (uz - coupling remaining, The variation in the semi-normalised yawing no further reduction moment (yaw acceleration), from this source, would be reliable. Examination of the with normal velocity w is shown in Fig17.

Routhian for this 4th order system indi- The variation is seen to be moderately non- cates that the oscillatory stability bound- linear particularly in the normal helicopter ary is crossed when NW - 0.02 , which is working state (w < 0). The derivative NW in accordance with the results presented is seen to increase as autorotation is earlier in Fig 6. The eigenvalues for the - 0) but the effect 10 Molusis, J.A. "Rotorcraft Derivative approached (u - x0 disappear once the engine Identification from Analytical will, of courge, It is known that the direc- Models and Flight Test Data."

disengages.

AGARD CP 172 (Methods for Air- tional stiffness can also reduce in this craft State and Parameter region due to fin shielding effects; the Identification)." Nov. 197'1.

loss of 'dutch roll' damping for these flight conditions should, therefore, be 11 Kaletka, J., Rix, 0. "Aspects of System fairly severe if the current predictions Identification of Helicopters."

are correct.

Proceedings of the 3rd European Rotorcraft and Powered Lift Air- Acknowledgements craft Forum, Aix-En-Provence, Grateful thanks are extended to France, Sept. 1977.

colleagues by both authors for the help 12 Kloster, M., Kaletka, J., Schaufele, H.

given in the preparation of material for "Parameter Identification of this paper; in particular to .Mustafa a Hingeless Rotor Helicopter in Demiroz and Susan Chu at NASA Ames and to Flight Conditions With Increased Tony James and Jane Whitbread at RAE Instability." Proceedings of the Bedford.

6th European Rotorcraft and References Powered Lift Aircraft Forum, Bristol, UK, Sept. 1980.

Padfield, G.D. "A Theoretical Model of 1.

13 Hodge, Ward, "Comparison of Analytical Helicopter Flight Mechanics for and Flight Test Identified Application to-Piloted Simula- Aerodynamic Derivatives for a tion." RAE Technical Report Tandam Rotor Transport 81048, Apr. 1981.

Helicopter." NASA TP-1581, Feb. 1980.

Padfield, G.D., Tomlinson, B.N., 2.

"Simulation Studies Wells. P.M.

14 DuVal, R.W. "The Use of Frequency of Helicopter Agility and other Methods in Rotorcraft System topics." RAE Technical Memorandum Identification." First AIAA "Simulation Helicopter+ Flight Testing Conference, Sarconi, G.

3.

caracteristioue du SA 330.” Paper 81-2386, Las Vegas, Aerospatiale-Note Technique, Nov. 1981.

330.05.0080, 1975.

Hall, W. Earl, Jr., Gupta, Narendra K., "Rotorcraft Fluid Hansen, Raymond S., Hoerner, S.F., Borst, H.V.

4.

System Identification Techniques Dynamic Lift, Hoerner Fluid for Handling Qualities and Dynamics, N.J., 1975.

Stability and Control Evaluation."

Roesh, P., Vuillet, A. "New Designs 5.

Proceedings of the 34th AHS for Improved Aerodynamic National Forum, Washington, Stability in Recent Aerospatiale May 1978.

Proceedings of the Helicopters."

16 Bryson, A.E., Jr., Ho, Y.C. "Applied 37th AHS Annual Forum, New May 1981. Optimal Control, Blaisdell Orleans, Publishing Co. Waltham, Mass, Kuczynski, W.A., et al. "The Influence 6.

1969.

of Engine/Fuel Control Design on Jennrith, R.I. "Stepwise Regression."

Helicopter Dynamics and Handling 17 Proceedings of the Statistical Methods for Digital Qualities."

35th AHS Annual Forum, Washington, Computers, Wiley, NY, 1966.

May 1979.

18 Fiske, Philip H., Price, Charles F.

Parameter Identification, "A new Approach to Model Struc- 7.

ture Identification." AIAA Paper AGARD Lecture Series 104, Nov. 1979. 77-1171, 1977.

"The Analysis of Weakly "Helicopter Stability Milne, R.D.

8. Molusis, J.A. 19 Coupled Dynamical Systems."

Derivative Extraction and Data Int. J. Control, 2, No.2, 1965.

Processing Using Kalman Filter- 28th AHS ing Techniques."

20 Milne, R.D., Padfield, G.D. "The National Forum, Washington, Strongly Controlled Aircraft."

May 1972.

Aeronautical Quarterly, May 1971.

is, J.A. "Helicopter Stability Molus 9.

"On the Use of Approxi- 21 Padfield, G.D.

Derivative Extraction from mate Models in Helicopter Flight Flight Data Using the Bayesian Mechanics." Vertica, Vol 5, Approach to Estimation."

243-259, 1981.

PP.

Journal of the American Heli- Copyright 0, Controller HMSO London, 1982 copter Society, Jul. 1973 - Dccouplcd long/lne Coupled system Nominal* IAS 100 kn Hode subsystems Relative densiry, ~1 0.818-0.837 Roll subsidence -2.242 -2.209 Rotorspeed. C 26.6-27.2 radls -0.0451 tl.047i -0.193 fl.079i ‘Dutch roll’ ROfO3z torque - 11500 ft lb (564 hp. 421 kU) Spiral -0.1166 -0.1194 *VW ,3018-12169 lb (S¶O4-5519 kg) -0.9054 ?l.lS6i -0.7645 f0.9354i ‘Short period’ I** 6380 slug fez (8650 kg m2) ‘Phugoid’ -0.00833 fO.1764i -0.0168 ?0.203Si I** 25483 slug ft2 (34550 kg m2) ______ -_____-- YY I’* 20283 slug fc2 (27500 kg m2) zz Table 4. Comparison of coupled and uncoupled longitudinal/lateral system eigenvalues - Cg locecion below rotor hub 7.05 fc (2.157 m) Puma, 100 kn Cg location forward of rotor hub 0.086 ft (0.026 m) Rocar Lock number 7.86 A2 1.0516 Solidity. s 0.0917 . -~_~._- * Accunl IAS varied from 95-100 kn giving an EAS 1 105-110 kn.

+ Flight estimates ** Manufacturer’. estimates (gef 3).

lateral variables only w and q -. .-.-_I Theory Table 1. Puma flight 325 - nominal trim conditions O-O.5 Hz O-1 Hz o-4 Hz o-o.5 Hz lx2 0.92 0.91 0.9 0.96 L” -0.013 -0.013 -0.013 -0.0123 -0.022 L -0.4 -0.32 -0.28 P -0.35 -2.05 L’r I .? I .82 I .9 3.03 0.294 Flight Parameter Theory L ll1.c 0.022 0.021 0.02 0.019 0.044 (approx) L w 0.0054 -0.00027 I.4 I .39 0.839 Prx-iod (s) 4.5 6.0 ---------- 14.0 15.36 Tj (~1 Table 5. Rolling moment derivatives - estimates from lp/vl 0.02 0.019 flight and theory LPlV I35O I58O 0.012 0.0096 Ir/vl Lrlv Fliaht estimates -95O -8S” x0-0.5 Hz) 0.006 0.0027 Theory Idvl + Lateral Lq/v 98’ -48O variables w and q only

I+Jl 0.336 0.354

LWlV 3o” -107O

R2 0.97 0.98

Table 2. Comparison of measured and predicted ‘dutch roll’ oscillation characteristics 0.0053 0.0071 0.00605 NV -0.53 -0.0009 N -0.6 P Nr -0.636 -0.572 -0.528 N -0.027 -0.028 -0.043 T)P -0.003 0.0127 Nw N 0.322 -0.328 q Table 6. Yawing moment derivatives -estimates from flight and theory Flight estimates Theory Roll subsidence -2.05 'Dutch roll' -0.346 f1.116i O-O.5 Hz O-I Hz O-4 Hz Fpiral -0.1124 'Short period' -0.768 kO.968i 0.73 R2 0.802 0.76 'Phugoid' -0.0349 +O.l969i 0.00246 0.00243 0.00243 0.00245 MU Table 9. Approximate eigenvalues for longitudinal and lateral subsystems 0.00074 -0.0056 0.00092 0.00085 Mw M -0.683 -0.667 -0.648 -0.831 M 0.0316 0.0302 0.029 0.038 Ills Table 7. Pitching moment derivatives - estimates from flight and theory -__ (Note E Ue) = Mw(Z + lJe) - nqzw) ‘1 Longitudinal subset X = fu,vo,w,qli Y” = w-u0 e Fig 1 Puma XW241 A3 4 (short Period); A2 - (2” + Mq)A - tx+ Ue) = 0 .

- .- Lateral subset 5 = l;“.“,;.P); ;, = ; + Uer (up = Y, = 0) +(Nv+$(t-NJ)“e = 0 (toll subsidence); A = L %.

P -.-._ ._-.- __ _... - ..__.._ __.. __ -__-.-.. . .- -.. .._-. -.-..-~.--__ Fig 2 Puma flight mechanics data acquisition and Table 8. Approximate formulae for longitudinal and processing at RAE Bedford lateral eigenvalues Fig 5 Puma fin sideforce coefficient vs sideslip angle ‘Dutch Im IA1 n\ 1.

(rrdlsl A Fig 3 Flight/theory comparison of rate response to cyclic inputs . Subset poles

;/ *p-e-.-------I.-

Re(Al 111~1

Ik+t--t--t~ t+-I-t+t+tt I ,t$ii

- ti-tt-

-1.0 -0.5 l O.S I Fig 6 Root loci for fast oscillatory modes with varying yawing moment due to incidence (NW) u >- .- yk?

m-0 5- L a .- u- - ul c”a mu .- -0.01 0 0.01 NW 0ssa-l Fig 7 Variation of the phase angle in incidence and yaw Fig 4 Puma flight 325, Run 12 - response to padal doublet rate excursions during ‘dutch roll’ oscillation with NW -theoretical predictions (a) 32512 pitch AFCS OFF I 0.25 P (radlsl I 0 -.M.-- -I3 ’ (b) 35202 pitcyAFCS ON I2 1 Puma -yaw and pitch vane response to pedal doublet Fig 9 Fig 9 Variation of derivative estimates with record length

MeIs L 0.5 MBlS

& .o 2.0 % -o.s- , 1 I I I 1 0.1 AIlSI 0.2 2.0 40 6.0 6.0 10.0 tw 12.0 Fig 12 Comparison of measured (-) and estimated (- - -) states (Kalman filter/smoother) for lateral cyclic doublet manoeuvre x lo-’ Fig 19 Variation of derivative estimates with sampling interval a Fig 11 Comparison of pitch rate response with and without flapping dynamics Fig 13 Fourier transform of measured f-) and estimated (- - -) rolling moment - lateral cyclic input -1 0.5 r lradhl 0.0 0.25 q imdlsl - /-----_\ ____. _.---.

000 _ *-___-' TJ- \ \.

=___*- -025 ,,,, ,,,, ,,,, ,,,, ,,,, ,,,, 25 50 75 10.0 I2 5 ,,5) 150 1&t - Fig 15 Comparison of measured (A and estimated (- - -) states (Kalman filter/smoother) for longitudinal cyclic doublet manoeuvre -20 ,,,, ,,,, ,,,, ,,(, ,,,( ,,,( 1151 15.0 25 50 75 IO 0 12 5 Fig 14 Comparison of measured (-1 and estimated (- - -1 : states (Kalman filter/smoother) for pedal doublet -NRi manoeuvre (radls')) (m(A) (radlr) I I , I I -10 0 -30 -20 10 wfftls) I I I , -0.2 -0.1 0.1 0.2 Re(hl(l/sl Fig 16 Root loci for ‘dutch roll’ mode with varying Fig 17 Variation of ‘induced’ rotor torque reaction NV - comparison of coupled and lateral/ with normal velocity perturbations directional subset modes 3. Recipient’s Catalog No.

2. Gownmutt Acaasion No.

I. Report No.

NASA CP-2219

5. Report Date 4. Title and Subtitle

April 1982

6. Performing Organization code

HELICOPTER HANDLING QUALITIES

8. Perfor ‘n Organization Report No.

7. Author(S)

A-8?&

10. Work Unit No.

T-4007Y

9. Performing Organization Name and Address

San Francisco Bay Area Chapter of the American

11. Contract or Grant No.

Helicopter Society, Washington, D.C. 20546 and

NASA Ames Research Center, Moffett Field, CA 94035

13. Type of Report and Period Covered

Conference Publication

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

14. Sponsoring Agency Code

Washington, D.C. 20546 and the American Helicopter

Society, 15. Supplementary Notes Chairman and Organizer: David L. Key, Aeromechanics Laboratory (AVRADCOM),

Ames Research Center, Moffett Field, CA 94035

.

.

Phone: (415) '965-5839: FTS- 4h8=&319, All+ovon-

359-5839

16. Abstract

This conference publication contains the formal papers of a specialists'

meeting on helicopter handling qualities.

The conference was co-sponsored

by the American Helicopter Society and the NASA Ames Research Center, and

was held in April 1982.

Helicopters are being called upon by the military and civilian commun-

ities to perform more and more tasks, and extend operations into poor weather

and at night. Accompanying this increased use is a significant increase in

pilot workload and a need for better handling qualities.

The ability to

define handling qualities required to perform,such missions has not kept

pace with the actual uses.

The objective of this specialists' meeting was

therefore to develop an overview of the status and problems in the development

and specification of helicopter handling-qualities criteria, and highlight

topics for future research efforts by government and industry.

r

__--- ..---..

18. Distribution Statement 17. Kay Words (Suggested by Authorlsl)

Helicopter handling qualities

Unclassified - Unlimited

Pilot workload

Integrated cockpits

All-weather and night operations

Subject Category 03

Agility and maneuverability

22. Price* 21. No. of Pages 20. Security Classif. (of this pagal fg. Security Uassif. (of this report)

Unclassified Unclassified 254 Al2

*For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA-Lang1ey , 1g@ Postage and Fees Paid SPECIAL FOURTH CLASS MAIL

National Aeronautics and

National Aeronautics and BOOK

Space Administration

Space Administration NASA-451 Washingtbn, l?.C.

20546

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