V/STOL tilt rotor aircraft study. Volume 9: Piloted simulator evaluation of the Boeing Vertol model 222 tilt rotor aircraft

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NASA CR-I14602 V / STOL TILT ROTOR AIRCRAFT STUDY PILOTED SIMULATOR EVALUATION OF THE i BOEING VERTOL MODEL 222 TILT ROTOR AIPCRAFT I VOLUME IX By: H. Rosenstein P. A. Mollenkof M. A. McVeigh • _ February 1973 Distribution of this Report is provided in sponsibility f o r the contents resides in the interest of information exchange. Re- the author or org a nization that prepared it.

,A, CL _ , • • BOEING VERTOL COMPA_Y _ Prepared Under Contract No. NAS2-6598 by [ [;_.

t!_ P.O. Box 16858 _:_ " _ _ _i A Division of the Boeing C ompany I_ "- _ ' _" Philadelphia, Pennsylvania 19142 '""._: f o r _ _: . _', x • i National Aeronautics and Space Administration 8 _me..esear_ _enter _' ; and i C United States Army Air Mobility Research & Development Laboratory i" A_nes Directorate Boeing D222-I0052-I ,

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I FOREWORD This report is one of a series prepared by The Boeing Vertol Company, Philadelphia, Pennsylvania for the National Aeronaut- _ ics and Space Administration, Ames Research Center, Moffett Field, California under contract NAS2-6598. The studies re- I ported under Volumes I through IV and VIII through X were I _ jointly funded by NASA and the U S. Army Air Mobility Research _ _, _ and Development Laboratory, Ames Directorate• Volumes V i through VII were funded by the U. S. Air Force Flight Dynamics | _i Laboratory, Wright Patterson Air Force Base, Ohio.

|| [_ This contract was administered by the National Aeronautics and | Space Administration• Mr. Richard J. Abbott was the Contract f- I Administrator, Mr. Gary B. Churchill, Tilt Rotor Research Air- • _ _ craft Project Office, was the Technical Monitor, and coordina- - -, , tion and liaison with the U. S. Air Force Flight Dynamics 2 . ._ I Laboratory was through Mr D Fraga. The Boeing Vertol Company _ , i Project Engineer for the work presented in this report was _ | Mr. H. Rosenstein.

_ 2 ._ l; The complete list of reports published under this contract is _,__._ ' as follows: _ _ .... ' Volume I -- Conceptual Design of Useful Military _" I ' _'_, _ and / or Commercial Aircraft, NASA CR- _. ii_437 _'_ _ Volume II -- Preliminary Design of Research Air- _ _ i craft, NASA CR-I14438 :_ ' _ _ Volume III -- Overall Research Aircraft Project Plan, Schedules, and Estimated Cost, '_ , _' NASA CR-I14439 • , _ Volume I V -- Wind Tunnel Investigation Plan for a Full Scale Tilt Rotor Research _, _ Aircraft, CR-I14440 Volume V -- Definition of Stowed Rotor Research Aircraft, NASA CR-I14598 "_'_" _ Volume VI -- Preliminary Design of a Composite _"'_'_' Wing for Tilt Rotor Aircraft, _ASA CR-I14599 ' Volume VII -- Tilt Rotor Flight C ontrol Program ,_ _ Feedback Studies, NASA CR-I14600

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Volume VIII -- Mathematical Model for a Real Time (Boeing Vertol Model 222), NASA i Simulation of a Tilt Rotor Aircraft - CR-I14601 li Volume IX -- Piloted Simulator Ev,_luation of The Boeing Vertol Model 222 Tilt Rotor Aircraft, NASA CR-I14602 II Volume X -- Performance and Stability Test of a 1 / 4.622 Froude Scaled Boeing Vertol Model 222 Tilt _otor Air- ii craft (Phase I) , NASA CR-I14603

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TABLE OF CONTENTS PAGE F O REW O R D ............... iii f L I ST OF ILLUSTRAT I ONS ........... v i I NOMENCLATURE .............. ix t f i _ 4.0 MATH E MAT IC AL MODEL DESCRIPTION . . . . . Ii

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_ ii 5 0 DATA BASIS . • 13 6.0 FLIGHT SIMULATION FACILITIES DESCRIPTION . . . 14

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_ 9.1 FAM I LIAR I ZATI O N ......... 3 1

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" 9.2 H O VER MODE STUDIES ........ 3 2 _ I 9.3 TRANSIT IO N MOD E STUD I ES ....... 3 9 9.5 E V A L UAT IO N OF MAXIMUM NACE LL E RATE S I N - _ ;< TRANS I T I ON ........... 54 ' _ _0 • 0 CONCLUS IONS AND R E COMMENDAT IONS ...... 5S ' _ , APPEND I X A - TIM E HIST O R IE S OF S EL ECTE D P IL OTED MANEUVERS 58 } • • • • • • • • • • • • • • • f i f , LIST OF ILLUSTRATIONS F TGURE NO. TITLE PAGE 1 SUMMARY OF USES FOR PILOTED FLIGHT 2 MODEL 222 TILT ROTOR NASA RESEARCH 5 EXTERNAL VIEW OF BOZING SMALL MOTION BASE 6 INTERNAL ARRANGEMENT OF SIM U LATOR C AB . . 17 7 MODEL 222 PLOT STATION FEAT U RE S U MMARY . . 21 I 8 SIMULATOR INSTRUMENT PANEL LAYOUT . . . 22 9 POWER LEVER / COLLECTIVE CONTROL FOR I i0 MODEL 222 CONTROL FORCE GRADIENTS AND 1 .

A. 1 (a) SLOW PILOTED TRANSITION AND RECONVERSION I SAS AND LAS ON ......... 60 _" i A. 1 (b) SLOW PILOTED TRANSITION AND RECONVERSION _-- A.I (c) SLOW PILOTED TRANSITION AND RECONVERSION ! ...... _' J c i_: _

SAS AND LAS ON ......... 62 __, •

i A. 2 (a) RAPID PILOTED TRANSITION AND RECONVERSION i, _ c i,>[; ' SAS AND LAS ON 63 ,_ .,,,_- • , _ t t ,_

A2Cb_ RAPID PILOTED T_'_SITION _D RECONVERSION __

- SAS AND LAS ON ......... 64 _--_

A2C=_ RAPID PILOTED TRANSITION AND RECONWRSION

A. 3 PILOTED TIME HISTORY - RESPONSE TO LONGI- T U DINAL STICK PULSES IN HOVER , SAS AND • • • • Q • • • • LAS ON, IN = 90 ° 66 "_-_ t l

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LIST OF ILLUSTRATIONS _ FIG URE NO TITLE PAGE I • i A.4 PILOTED TIME HISTORY - RESPONSE TO i ) i LATERAL STICK AND RUDDER PEDAL 1 _ PULSES IN HOVER, SAS AND LAS ON, i iN=90 ° 67 !

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A.5 PILOTED T I ME HISTORY - RESPONSE TO [_ LONGITUDINAL STICK PULSES IN HOVER, SAS AND LAS OFF, .....

} iN = 90° 68

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! _ A.6 PILOTED TIME HISTORY - RESPONSE TO ! LATERAL STICK AND RUDDER PEDAL "t f PULSES IN HOVER, SAS AND LAS OFF, + i * _+I )i A.7 PILOTED TIME HISTORY - RESPONSE TO . LONGITUDINAL STICK PULS E S IN TRANSI- TION, SAS AND LAS ON, iN=70 °, V= 8 0 KNOTS 70 +,f li .........

_ ' _. A.8 PILOTED TIME HISTORY - RESPONSE TO _+ IN TRANSITION, SAS AND LAS ON,

: : ,. , . . i i LATE STICK AND RUDDER PEDAL PULSES

+ " i A.9(a) PILOTED TIME HISTORY - HELICOPTER MODE : ' MANEUVERS, SAS AND LAS ON , . .

iN=90 ° 72 +_ Ii A.9(b) PILOTED TIME HISTORY - HELICOPTER MODE " ' MANEUVERS, SAS AND LAS ON, iN = 90" . . 73 • ".

I: A i0 PILOTED TIME HISTORY - PARTIA L POWER _"_+ . +_ o DESCENTS IN TRANSITION, SAS AND LAS I :_£

ON iN = 70 ° , V = 70KNOTS. . . . . . _ 4 _

A.I! PILOTED TIME HISTORY - RESPONSE TO r . ,,, A.12 PILOTED TIME H I STORY - RESPONSE TO LONG I TUD I NA L STICK PULSES AT 140 KNOTS , LAS OFF, = ........

iN 0 e 76 " i A.13 PILOTED T I M E HISTORY - RESPONSE TO L A TERAL AND RUD DE R PE D AL PULS ES.AT

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i LIST OF ILLUSTRATIONS FIGURE NO. TITLE PAGE A.14 PILOTED TIME HISTORY - RESPONSE TO LATERAL STICK AND RUDDER PEDAL PULSES AT 140 KNOTS LAS OFF, :iN = 0° . • • • . . . . 78 A.15 PILOTED TIME HISTORY - RESPONSE TO LONGITUDINAL STI C K P U LSES AT 260 KNOTS , LAS ON, iN=0° 79 ! A.16 PILOTED TIME H_STORY - RESPONSE TO LONGITUDINAL STICK P U LSES AT 260 KNOTS, h A.17 PILOTED TIME HISTORY - RESPONSE TO LATERAL & • • • • • • • • • • " '_ 1 STICK AND RUDDER PEDA L PULSES AT 260 KNOTS, LAS ON, iN=0 ° 81 _ A.18 PILOTED TIME HISTORY - RESPONSE TO LATERAL _ STICK AND RUDDER PEDAL PULSES AT 260 KNOTS, _ '_ LAS OFF , iN=0 ° 82 ;.... !

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NOMENCLATURE ' Symbol Definition Units

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h Aircraft altitude Ft.

I _ Aircraft rate of climb Ft / sec I i N Nacelle angle Rad.

p Aircraft roll rate_ positive Rad / sec when rolling clockwise (right I wing down) !

q Aircraft pitch rate, positive Rad / sec !

, / J _ when pitching nose up ' r Aircraft yaw rate, positive Rad / sec i - -_ | when yawing nose right _ !_ u Aircraft longitudinal component Ft / sec , of veloci t y _' v Aircraft lateral component Ft / sec . - _'" of velocity _" I _F Fuselage angle of attack Dt_g. .

, r , ) _F Fuselage sideslip angle Deg.

_B Longitudinal stick position , Inches positive aft _BSAS Longitudinal SAS link position Inches _TH Power lever / collective control Inches ' p ositio n _, , 6 E l eva to r angle, p os i t ive t r ailing Rad. % e e d ge do w n _ F l a p a ngle , p o s iti ve wing Deg. . ,,.

6 f _v _ ,. , _ trai l i n g e dg e do wn _ 6r Ru dd er pedal position , righ t I nches ru dd er positive _rSAS R u d d er peda l SAS link p o s iti on I nche s "- - NOMENCLAT URE S[mbol Definition Units 6 s Lateral stick position, positive Inches to the right _SSAS Lateral SAS link position Inches I (_ 0 Ai r c r a f t pi t ch a tt i t ude, Deg.

positive nose up Rotor collective pit c h at th r ee Deg.

ii 90.75 quarters radius station f - ! i_ _ Air c raft right wingroll downattitude, positive Deg.

I _i_ i _ Aircra f t yaw a t titude, pos i tive Deg.

nose right 'i_ _ ' ; Change in rotor rotational speed Rad / sec m • ¢ < i 1.0 SUMM_' This doc u ment presents the resul t _, _ : : ._i time piloted simulation conducted to investig_ _ _ :i : ndling qualities and performance of the Boeing Model 2 t _ _ otor aircraft design as described in Reference (i). Tn±. _ zed evaluation was c onducted during the period from Septe m Der 25, 1972 through uctober 28, 1972. Since this was the first piloted simulation ,-_fthe Model 222, the run program (which is shown and de- scribed in Section 8 of this document) was set up to broadly cover all regimes of flight (hover, transition, cruise , climbs, descents, etc.) and to identify potential problem areas.

During the above mentioned time period approximately 34 hours of piloted simulations were conduc t ed.

.f - The aircraft represented in this simulation was the Model 222 as described in Boeing Vertol's preliminary design study of : March 1972 (Reference 1). It differs from the aircraft of _ ; ]. ! Boeing Vertol's January 1973 proposal in several respects, the _.i I most important being: - i (a) The simulation model had outboard flaperons i and spoilers only, whereas the January 1973 _ _ aircraft has full span flaperons and spoilers.

< " (b } The simulation model had 6 rad / sec 2 control I % - ° • _ ' power in pitch and 1.0 rad / sec 2 in roll , . ' co mp ar ed t o 1.2 r a d / se c z and 2. 0 r ad / sec 2 i = respectively for the January 1973 pr o posal.

( c) The load alleviation system on the model senses nacelle pi t ching and yawing moments _ to feed back into cyclic pitch. The January 197 3 aircraf t senses pitch and yaw angle , and dynamic pressure.

(d) The longitudinal stability augmentation _ syste m in this simulation model incorporated _ {. a pitch attitude feedback loop. In addition, _ cy c lic pitch a c tuators. In the J anuary . .

1973 proposal the pitch attitude feedback . ; _ w as r e move d fr o m t he s t a b i li t y aug m e n tati o n sys t em and incorpora t ed in t o t he au t opilo t , _ -_ and the longitudinal SAS moves only the longi- we r e ma de in o r der to sim p li f y the desi g n of t udi n al cyclic p i t c h a c t ua t o r . T hese c hange s the ai rc ra f t.

I An e l even de gr e e -o f-fr ee d om m a them a ti c a l mo d e l i.e. , 6 ai r - fra m e , r otD r RP M , f i rs t w in g ve r t ical b e ndin g and t or si o n an d __ 2 n ac e ll e degr e e s o f fr ee d om w as f o r m ula t ed and i s d es cr ibed __

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in Volume VIII of this series of reports. The m a themati c al c _del was mechanized and used to drive the B_eing Small Motion L,ise Flight Simulator (SMBFS). The SMBFS provides initial I motion cues only and was modified to re[_resent the Model 2 22 .

These modifications included the addition of power lever / co]lective pitch control mounted on the left arm rest with a nacelle incidence switch located on the grip, an instrument panel deslgned to represent the tilt rotor, and an approprilte ; _rce feel system. O_ : 1_.r elements of the pil c , t's control _._- tern i.e., beep tri m , nag. brake, stick and pe l als were natl . .- factory and required no modification. The command pilot in the Model 222 is in the right seat.

_ Visual displays were computer-generated and projected onto a ]cr, en in front of the pilot. Two displays wore used for : ,! this evaluation; a ship deck for hover and a mountainc_us scene • "- _ _ l! wi t h a road and telep _ one poles for the cruise mode. Th e li _ , _ajority of the piloted simulation was however, c ondu c ted _ u sing the road only.

_'_ ! " _ The Model 222 tilt rotor aircraft was evalu a ttd at the design .ross weigh t of 12 , 000 ibs. with the nacelle-horlzontal cen- .) ty at this weight}.

; " Jnc --un prog r am c onsisted o f pil ot fa miliarization , hover m_de i _;''_ Ii }_ tar of gravity located ah 28% chord (most aft center of gravt- _ . . _ st_xiies , transition m o de studies , cruise mode studies , eva._ua- _ i : -- , . tion of maximum nacelle rates in transition and helicopter fli_,ht mode s t udies. Salien t findings and conclusions are as '_ I! foliows : ,_ i. An ef fici ent c oc kpi t d esign (ins t rument | !

panel layout and placement of primary con t rols) is required to minimize pil o t ' workload during transition.

2. There is no "bes t wa_" of t rimming the Model 222 in transition and reconversion.

! Since t he pilot has a nacelle t ilt cor.LLOI , it is possible t o t rim t he aircraft a t 1 i many different combinations of t i)t angle , bo d y i : and atti t ude.

[I 3. The Model 222 w_.s flown from hover to - .- maximum speed an d b a ck to a hover with the stability augmentation system off.

4. The aircraft is docile and eas,, to control through transition and recon- ;" version. The changes made in the longi- [ t u di n al 3AS i n t he J a nu ary 1 97 3 pr op o sa l

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i and described in item (d) above co,lid impact on the docile transition and reconversion character- should be re-ev=luated by piloted simulation.

isti c s on the i00 to 140 knot speed range. These 5. Relatively low rotor inertia requires careful design !

i and tailoring of the thrust and power management system to insure precise altitude hold capability in hover low spe ( d flight m o des.

| - , _ 6. The Model 222's longltudinal handling character- _ istics in the cruise mode {nacelle horizoDtal and SAS off) are sa t isfactory. The cruise mode lateral I directional evaluation showed a cou p led roll / spiral [ mode from the end of transition to maximum speed, and high dihedral effect. These w ere annoying to the pilot but easily controllable. These can be completely , / _ i eliminated by SAS feeding back roll rate into rudder, , and sideslip angle into aileron.

• _ 7. The descent / deceleratior / wing stall boundary in the _ helicopter mode was investigated. These preliminary evaluations indicate that if t he wing is allowed to steep _ _tall during approac}les, the rate of descent i_ builds up rapidly and recovery close t u the ground may '_... _ ,I be difficult. It is felu that insufficient c u es in _ "I the nudge base simulator (such as che _ges in noise i'_'-. ! level and no buffet onset indication) and relative [ pilot unfamiliari t y with the vehicle are complicating '_ f a c to r s and a d diti o n a l p i lot trd i nin 9 and fam i liar- _ 1 ization would obviate any problem in this area.

( Wi t h the wing leadlng edge umbrellas o r with spoilers open , descent capability is improved. It should be noted that descent rates up to 1500 f t/ min at low speed have been achieved, i 8. A preliminary eval_ation of maximum nacelle til rates in rapid acceleration t ransitions and reconversion:.- | . was cond u c t ed. One rapid transition and reconversion run was conducted with the nacelle tilt rate liriSed to 5" / sec (nomin a l maximum value is 10° / sec). _ne pilot indicated tha t lower maximum ra t es might be de- | sirable in the high speed end of t ransiti o n (100 KT conditions , while higher nacelle rates are accc ;cable I 140 KTS) to minimize pi t ch at t i t ude changes at th e se " .._,_. at lower speeds. Additional work is required _n th i s " area to evaluate the desirability o f e s tablishing a sche a ule of maximum nacelle tilt rates , an d to optim i ze contr_l schedul i ng.

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2.0 INTRODUCTION , Pilo t ed s i mula tio n i s a usefu l an d i m p ort an t t ool in t he design, developmen t and test of new flight vehicles. Figure i shows a summary of some of these uses as they could be ! applied to the Model 222 Tilt Rotor.

A:_ a part of Con t ra c t NAS2-659 8 Boeing Vertol developed a _athema t ical model of the Model 222 Tilt Rot o _ aircraft, _n- I tended primarily for use with the FSAA at Ames. As a further _,ddi t ion to t he same c ontract Boeing Ver t ol programmed this i aa_h model on its hybrid c omputer and used it to drive the I Small Motion Base Flight Simulator for preliminary pilot eva- ]ua t ion of a til t rotor aircra f t. The results of this simul a - tion are presen t ed in this r e port.

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I • E valuation of Tilt Rotor Handling Qualities o Stability and Control l o Control System Optimization o Evaluation of Man-in-the-Loop System Compatibility ' o Evaluation of Malfunction Effects ' • E v a l u a t i o n of Ti lt R otor P e r forman c e I r _! o Man e uver Capability J _. o VTOL and STOL Takeoff and Landing Capability

H • As a Tool to E valuate Configuration Changes

_-II o Changes in C o ckpit L ay o ut i o Cha ng es in Tai l S iz e _ [! o Changes in Geometry &' o Changes in SAS C onfiguration i ' __ o Changes in E lasti c Characteristics , I • As a Flight Test Support Tool o Development of E mergency Techniques ,i . ._.

o Familiarization of Flight Crews with Aircraft _. ? _._ , Charac t eris t i c s P rior to Flight

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O E xplo r a t i o n o f Fligh t- Dis co ve red Ph en omen a _ _i FIGU R E i. SUMMARY OF USES FOR PILOTED FLIGHT SIMULATION i Q r

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3.0 AIRCRAFT DESCRIPTION The Boeing Model 222 Tilt Rotor Research Aircraft is a three place, twin turbine engine aircraft, with two rotors displaced laterally and is designed to demonstrate "Proof of Concept" I! for follow-on military or commercial tilt rotor airplanes.

[_ F_ure 2 is a 3-view of the aircraft and provides the general ar '. -angement and salient dimensional data. It should be noted I that the air c raft simulated is th._ March 1972 version of the ii model, as described in Reference i. The aircraft incorporates two 1550 horsepower Lycoming T53-L-13B engines (modified), ea c h i' driving a three-bladed, soft-in-plane hingele_s rotor. The I '_ rotors are interconnected by cross shafts, which transfer single engine power to both rotors in the event of an engine failure. Wing leading edge umbrellas, coupled with 30% chord i , .f [ single-slotted flaps capable of 70 ° deflection, are used at _ ' hover and low forward speed to minimize vertical drag or down- load on the wing. These have been programmed to open or clo_,e I . [_ at a dynamic pressure corresponding to 50 KIAS, and a nacelle Control of t he Model 222 is accomplished utilizing rotor longi- _ tudinal c y c lic, differential cyclic, rotor thrust, and differ- _. ential collective control in conjunction with airplane control iil _ angle of 7 5 ° • surfaces. The airplane control surfaces consist of elevator, _ v l rudder and aileron / spoiler c o ntrols The rotor controls pro- - _ are phased ou t as a function of de c reasing nacelle incidence ' vide the major por t ion of the control power at low speeds bu t i i latively more effe c tive Figure 3 presents a summary of the

t i _ angle as speed inc r eases and the airplane controls become re-

primary moment-producing controls for ea c h of the three flight _ , modes.

' _ In the version of the Model 222 used for this simulation, the by providing a soft mounting for the nacelle in pitch so that the . , :._'k; _ [ hub moment generate_ by cyclic tilts the nacelles simultaneously _' : <' thrust vectoring effect of longitudinal cyclic is amplified __ . _ for longitudinal control and differentially for directional _ _i c ontrol. Note that in later versions of the aircraft the same _ effect is obtained by positive a c tua t ion of the nacelle. An .r_._.

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artificial feel system is provided which varies the control I i__- feel forces about all three axes as a function of dynamic pres- [ . .'.._ q I I sure to improve control force harmony and provide desirable -- _] levels of feel forces for handling qualities and flight safety consi d erations • stick On the Model 222, longitu d inal cyclic is connected to the for longitudinal control an d to the pedals for directional control. Both lonqitudinal and lateral cyclic are programmed I with nacelle tilt t o minimize pivot moments as par t of the load i ,

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i FIGURE 2 MOD E L 222-1 TILT ROTOR RE SEARCH A] FO L I _ UT FRA_ , I E I

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_ DEL 222-1 TILT ROTOR RESEARCH AIRCRAFT I . .

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i FLIGHT MODE PRIMARY CONTROLS [i Helicopter (Hover) - Pitch Longitudinal Cyclic ! _i - Roll Differential Collective i - Yaw Differential Lonqitudinal C, ' clic Transition - Pitch Longitudinal Cyclic and Elevator U - Roll Differential Collective, • . Differential Longitudinal Cyclic, _ Aileron and Spoiler _ - Yaw Differential Longitudinal Cyclic :: Diffe r en t ial C o llec t ive and ..... Rudder - Pitch Elevator _ - Roll Aileron and Spoiler - Yaw Rudder _._j NOT E : Airplane con t rol surfa c e s are opera t ive at all times. _<_.

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alleviation system (LAS). Roll control in hover is achieved i by differential collective pitch. Roll control in transition utilizes phased differential collec t ive pitch, differential longitudinal cy c lic and differential nacelle tilt in conjun c tion I with the spoilers and flaperons.

The rudder and elevator con t rol surfaces are conventional.

Roll control surfaces in this simulation consist of upward- I operating semi-span spoilers, and downward- o peration of the ou t board flaps. This permits use of more efficient single-slotted flaps for low speed loiter in the cruise configuration and per- I mits reduction of yaw due to roll control input be c ause of t he favorable yaw due to spoiler c ombine_ with the adverse yaw due to aileron control.

I The stability augmentation system (SAS), used for this pilote d simulation study consists of a pitch, roll and yaw SAS. The pi t ch SAS incorporates pitch rate, pitch attitude and longi- I tudinal stick pickoff feedback loops. The SAS moves the elevator -- and longi t udinal cyclic pitch actuators. Longitudinal SAS is used in hover and transition and is phased out in t he cruise lateral stick pickoff. These are phased ou t in the cruise I mode. The roll SAS consists of roll rate, roll attitude and a mode. A roll attitude hold mode is included to be use d in t he rate and yaw attitude hold. These are phased out in cruise.

i cruise configuration. The yaw SAS consists of roll rate, yaw The roll and yaw SAS's move both rotor and aerodynamic con t rol surfac e s.

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i The load alleviation sys t em (LAS) utilizes longitudinal and lateral c yclic pitch feedback loops to zero out the rotor hub I moments.

The thrus_ /c ollective pi t c h is controlle4 by throttle t y n ?

levers in the cockpi t , whi c h, in hover, command d irec_iy both i_ engine power and collective pitch. The governor adju:_&s t_._ _ ....

colle c tive pitch to maintain cons t ant rpm. Overtravel o f t .e _ cockpi t levers is provided beyond the normal maximum power _,_i_ I position. The overtravel is entered by passing through a gate _ which shutoff the governor, so tha t in t he overtravel position _, _ : the lever dire ct ly controls collective pitch only and can be , use d ju s t lik e a he li cop ter c oll e cti ve pit c h l e ve r t o perf o r m _i_.

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a c o llective flare. The mechani c al interconne c t from t hrust / collective lever t o colle c tive pitch is phase d ou t d uring t r a n- si t ion so t ha t in cruise t he pilot d eman d s power only, and pitch is governe d to main t ain rpm like a o onve_tional propeller airplan e .

A _ | It should be no t ed that the air c raft simula t ed during this

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program is not the same as describe d in Boeing Document D222- 10050, Volumes I to XII (S t u d y o f V / STOL Tilt Ro t or R e search °r' | Aircraft Program - Phase I). The air c raft geome t ry is essentially

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the same. Weights , inertias, aerodyn a mic data, load allevia- I tion and SAS configuration have b e en revised. While future piloted simulation studies may yield small differences in quan t i t a t iv e r esul t s, t he quali t ative resul t s an d tren d s s h oul d I be similar.

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4.0 MATHEMATICAL MODEL DESCRIPTION

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The mathematical model of the Model 222 Tilt Rotor aircraft is described in Volume VIII of this series of reports, and is i used to drive the Boeing Small Motion Flight Simulator (SMBFS) an eleven degree of freedom total force model. This model includes the basic six degree of freedom rigid body outer I loop equations written about the instantaneous center of gravi- ty with the inertial and aerodynamic terms included. The rotor is treated as a point source of forces and moments with , I appropriate response time lags and actuator dynamics. The ._ wing has one vertizal bending and one wing torsion degree of freedom. These structural degrees of freedom are treated on tion of the structure are much higher than the frequencies of i a "quasistatic" basis; i.e., the natural frequencies of vibra- the rigid body motion, and the coupling is in the aerodynamic terms. Each nacelle has an independent pitch degree of free- I dom about the wing pivot. The aerodynamics of the wing, tail, rotors, landing gear and fuselage are included. Wing and tail mutual interference effects and turbine engine performance and I dynamic responses are represented.

The control sys t em elements represented include pilot c ommand rate , power), three-a x is stability augmentation systems (SAS), i (longitudinal and lateral stick, pedals, nacelle position and thrust management system (includes rotor constant speed governor) and a load alleviation system (LAS). The LAS system I incorporates feedback to rotor cyclic and collective pitch for purposes of improving stability , blade load reduction, gust alleviation and increased damping of aeroelastic modes. C on- I trol system actuator dynamics are represented by appropriate second order systems. Figure 4 is a summary of the salient _ features of the mathematical model used for this study.

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(1) Full Fligh t E nvelope Capability with Total Force I R e pre se n t a t i o n I (2) 6 R igid Body Degrees of Freedom i (3) Independent Nacelle Pit c h Degree of Freedom (4) 2 E lasti c Degrees of Free d om

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(5) 1 Rotor Rotational Degree of Freedom (6) Includes the Aerodynamics of: I • Rotors • W i ngs • Rotor / Wing & Wing / Rotor Interference I • Fuselage • L anding Gear • Tail Surfa ce s I • E ngines I (7) C ontrol Sys t em E lements: • Pilot Comman d • SAS System I • L oad Allevia t ion ( L AS) '.

• Thrus t and P c w er Management Sys t em _- . _ ( 8 1 A e roe l a st i c R ep r e s en t a t i o n , , , _ • W in g V ertica l B e n d ing _ Torsion I • Wi n g • N acel l e P itc h in g De g ree of Freedo m

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I F _G U _ 4 e S ALIEN T F E ATUR E S OF _I_ T _"_ M OD E L _

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Q I 5.0 DATA BASIS The Model 222 aircraft used in this simulation study is de- I scribed in Reference i. The data basis for this aircraft was, for the most part, obtained using analytical methods.

obtained from wind tunnel tests on similar tilt rotor con- These methods have been generally substantiated by test data figurati o ns. At the time input data for the simulation was being prepared (June 1972), there existed only a limited quan- tity of wind tunnel test data on the actual Model 222 config- uration. Alth o ugh a c o mparison of these limited test results with the corresponding analytical data indicate_ favorable agreement, it must be emphasized that the data base utilized come of further wind tunnel tests. In view of this, and the for the simulati o n is subject to modification pending the out- differences in aircraft c haracteristics previously noted, it ctlara c teristics, fl_ing_ qualities, and performance of the I should be borne in mind that pilot comments on handling M o del 222, in future piloted simulation studies that incorp- orate a more complete data base may yield small differences I in quantitative results although the qualitative results and - four Boeing-developed computer programs. Hover and cruise i Rot o r data used in the mathematical model were predicted from performance (thrust-power) were obtained using a propeller performance _ analysis computer program (B-92), which uses an trends should be similar, i I explicit vortex influence technique theory (Reference 5).

Transition performance data, in-plane forces and moments and cyclic pitch effects were estimated using computer program I D88 _Refercnce b). This program uses strip theory, c o mbined with unsteady aerodynamic and non-uniform downwash to compute aeroelastic -otor loads. In-plane elastic rotor derivatives (both static and rate)in axi a l flow were estimated using _ I computer program C4i (Reference 3). Elastic rotor rate deri- ..

' vatives in transiti o n were estimated using computer program ,_ C-49 (Reference 4). Correlation with rotor test data is shown L_:' : " .I in Volume VIII, Section 7.0 of this series of reports. Wing, _"_ _ tail , f us elage and nacelle aerodynamics were estimate d us i ng _ , t _ : • DATCOM (Reference 2), combined with increments and trends f_i " I derived from References 7 and 8. Rotor ground effects also _> were obtained from Reference 7. Aircraft geometry, weigh t s and _ % inertia are as specified in Reference 1.

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6o FLIGHT SIMUm ON FACILITIES DESCRIPTIO

The Flight Simulation Facility is an integrated laboratory complex for performing unmanned and piloted real-time flight simulation studies of aircraft, control systems, and instru- mentation concepts and configurations. It is comprised of two laboratories, the Flight Simulator Laboratory and the Hybrid I Simulation Laboratory. These two laboratories are located in separate buildings and are interconnected by electrical cabling.

The Flight Simulator Laboratory contains a six degree-of- f_eedom small motion base simulator, a pilot station equipped with an adaptable instrument panel and a wide-range variable I Fli_ht Simulator Laborator_ flight control force-feel system, a cockpit-mounted out-of-the- window collimated visual-simulation display, a visual simula- tion and associated and I scene generating system interface, control and readout hardware.

i The variable flight control force-feel system incorporates actual aircraft flight controls modified to have load cells at the points of pilot applied forces, and to be positioned developed from the load cell force signals and con t rol posi- _ I by hydraulic servo-actuators controlled by computer signals tion feedback signals. Any desired relationship between pilot effort and control position can be simulated. The system

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I offers high signal-to-noise ratios, and responds to forces ranging from an ounce to more than a hundred pounds. The visual display system presents the pilot with _ _right colli- I mated out-of-the-window symbolic visual scene, the visual scene is computer generated, offering both latit u de in s c ene content and an unconstrained flight path and maneuver cap- ability. The generated scene is reproduced by a 600 line g I black and white television system for viewing by the pilot through a large collimating lens. The pilot's field of view measures 38 degrees vertically by 53 degrees horizontally, I and had a depression angle of zero (0) degrees.

Small Motion Base Fli_ht Simulator I Facility Description: The small motion base simulator provides six-degree-of-freedom motion employing the relatively rigid strut actuator concept. The small travels of the actuators m • _ result in approximately uncoupled motion and deliver nudge- type acceleration cues to the pilot ot satisfactory validity.

Three of the six electro-hydraulic actuator struts are vertical I and t hree are horizontal. The Moog valves of the s t ruts re- spond t o command signals genera t ed from t he ma t hema t ical model progr am med on the hybrid computing system. The cockpit is I equippe d wi t h a v ari ab l e fli gh t c on t r o ls forc e -feel s y ste m a n d 1 4

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a c o c kpit-mounted out - of-the-wi, , dow c ollimated visual display.

I Figure 5 shows the external structural details of the motion the simulator cab.

' resting C apabilities: This facility permits a wide variety of studies and pilot evaluations of aircraft handling and flying qualities, automatic and manual flight cont ol systems , I [ base ana cab, and Figure 6 shows the internal arrangement of I and design criteria definition for the praid, economlcal, manageable, and safe development of aircraft and aircraft of V / STOL aircraft because of its capability for providing c ockpit motion in six-degrees-of-freedom.

I I sy s tems. The facility is particularly adapted to the study Motion System Performance Payload (includi n g pilot ) 770 ib I Travel Limits (stop-to-stop t o tal ) : Vertical 5 in.

Lateral 5 in.

I Longi t udinal 5 in.

Pitch 13 deg .

Roll 19 deg I Yaw 19 deg Pitch Tilt . 26 deg I Rate Limits with Zero Acceleration: !

Vertical + 26 in / sac !

Longit u dinal _ 41 in / sac I Lateral 26 in / sac Pitch _ 69 deg / sec Roll _ 97 deg / sec I Yaw _155 deg / se c Acceleration Limits for Zero Rates (incremental values): I Ver t i c al + 64.4 f t / se c _ Longitudinal _ 35.4 f t / sec _ La t eral _ 28.9 ft / sec 2 deg / sec 2 I P it c h 248 Roll ; 414 de g / sec 2 : Yaw _ 745 deg / se c 2 I During the piloted simulation e f fort , it was ne c essary to tailor the mo t ion sys t em o f t he nudge ba s e s imula t or. The _ m m o ti o n s y s tem r equ i re d tail or ing for the t il t ro t or i n the verti c al an d longitudinal axes. The vertical a cc eleration

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c ap a b il i ty of t h e t i lt r otor a i rcraft in cru i s e fli g ht w a s - _ 8 si gn ificantly highe r than th a t o f the o the r airc r af t us e d t o

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14 '36 4 1 5. EXTERNAL FIGURE i VIEW OF BOEING SMAIL MOTION BASE FLIGI{T SIMULATCR ah.

Q t w, • , I 1 57 9Q _ ._,-----_ I FI G URE 6. INTERNAL ARRANGEMENT OF s IMULATOR CAB I 17 Q . i

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establish the motion system dynamic characteristics• In order I to keep the simulator from hitting its motion limits, the ver- tical axis gain was reduced• aircraft was also higher than the acceleration experiences in i The longitudinal acceleratio n capability of the tilt rotor helicopters. It was found that the long term cockpit tilt used to represent longitudinal acceleration was very dis- I orienting to the pilot. In order to eliminate this disorienta- tion, the cockpit longitudinal tilt due to acceleration was attenuated by a factor o f 4.

I Visual Simulation System the Im a ge Generating System and the Visual Display System.

i The Visual Simulation System comprises two main subsystems: The image generation of a landing zone and horizon line are provided by a high-speed repetitive operating analog computer• purposc I This computer is part of the general Hybrid Simulation laboratory. The capability, therefore, exists for expansion or adjustment of the visual scene to suit the customer's i ' | simulation task requirements. A number of different visual

I scenes have been used to date. One is the symbolic representa-

tion of a helicopter landing pad on a destroyer afterdeck.

This particular display is capable of handling up to four surf- I l aces (i.e. upper deck, lower deck, etc.) plus horizon line _ with perturbation of ship and aircraft motion. _ .

I The Visual Display Systems contains an Image Transfer Unit and a Visual Display Unit.

looking at the face of a 5-inch os c illoscope through a beam t The Image Transfer Unit has a cl o sed-circuit television camera splitter. The display computed Dy the Image Generating System as time-varying X and Y signals produces an animated pictograph _ 1 i on the scope which the camera converts into a video signal. _ The video drives an 9-inch monitor at the camera station for E ;_ focusing and alinement reference, and a 14-inch monitor at the • _( test observer station It also works the Visual Display U nit |_ ....

consisting of a 23-inch television monitor attached to the |_ , _ collimating lens in the front window position. The pilot thus Flight Simul_tor cab behind a 16-1 / 2 by 22-1 / 2 inch plastic li _ " views a bright, enlarged, infinity-focused picture thr o ugh a % 38-degree by 53-degree sighting aperture. Head-position paral- '_ lax is eliminated by the lens which lends an apperance of _ real-world depth to the scene. To assist the pilot in per- _ ceiving the picture as representing the outside world, a model _;_ of an aircraft nose boom is mounted between the collimating ___ , lens and the face of the 23-inch monitor. It appears in _ realistic 3-D outside his window. A beam-splltter permits the insertJ.on of various instrument indications as a heads-up - - .

display presentation in the window. Lateral acceleration, 1 8 [

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and position velocity and position, and longitudinal velocity _ indications have been provided in this , ,dnner.

H_.ybrid Simulati o n Laborator Z The Hybrid Simulation Lab o ratory consists of an IBM 360 / 44 digital computer system connected to five Applied Dynamics , Inc. analog computers, providing capability for solving time- critical pr o blems. The liybrid system is connected to the limited-motion base flight simulator and includes a disc system, m agnetic tape units, data adapter units, processing unit, card _i read / punch, printer, display stations and digital function generators.

the analog and digital computers, thereby permitting system The Hybrid system combines the best operational features of simulation involving the interaction of several technologies, such as flight control, aerodynamic performan c e, and vibratory i analysis. Hybrid simulations have the ability to run in real time and include system parameters in a voltage analogy. This permits inclusion of flight hardware and actual loop-closure lations are realized on the ground, materially reducing in- I effects into system analysis. More sophisticated flight simu- flight development programs and their attendant expenses.

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7.0 P ILOT STATION DEFINITION i

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represent the Model 222 aircraft. The configuration changes included I instrument panel modifications, design and fabrication of the power lever / colle c tive control and nacelle incidence I _ The cab of the Small Motion Base Simulator was configured to control, modification of the pilots force feel system to pro- I vide the proper breakout forces and gradients as a function of dyn a mic pressure, and electro-mechanical limits placed on stick and pedal travel to properly simulate maximum control stick and pedals and "back drives" on the primary controls to I travels. Additional features include a magnetic brake on provide initial control position trim in the cab. Although there is only one seat in the cab, instruments and primary controls such that flew I were positioned the pilot as if from the right seat, as in conventional helicopters• A summary of Model 222 pilot station features are shown in Figure 7.

I Instrument Panel Layout the Model 222 configuration. The stan d ard "tee arrangement i The ins t rument pan e l of t h e simulator was mo d i ' ied to r e p r esen t of the primary flight instrument was retained an d the loca t ion of instruments unique to the Model 222 were defined af t er con- I sul t a t i o n wi t h t he pr o je ct t es t pilo t .

Figure 8 is a photograph of the slmulator control panel. In- i I struments not labeled were not used for this simulation.

Nac e lle angl e , si de ward speed an d "g" me t er were l o ca t e d on t he left side of the panel. Engine condition , rotor RPM , angle of panel. It should be noted, however , that as the test program i attack and flap position were grouped on the right si d e of the progressed, i t becam e evi d ent t ha t t he engine t orquemeters a r e more properly placed on the left side of the panel. Space _ I constraints preclu d ed making this c hange in t he simul at or I _ ,W Transm i s s ion li mi t s a n d normal rotor RPM positi o ns were m ar ke d _ - : ; o ** th e a pprop_ [a t e ins t rumen t s. Dual eng i ne an d single e ngin e _!_ I . , , % * transmission limits are indicated by re d in d ex marks on the engine to rq ue m et e r d ials. The d ual engine t ransmissi o n limi t I ' , - _ torque. Re d index marks were placed at 100% and 70% rotor _ i _ ....

i was pl aced a t 74% t orque an d t he single engine li mi t s at 97% _ _ RPM, th e n or m a l h o ve r an d cruis e value s . S in ce ro tor RP M is _m_o automatically s che d uled as a function of nacelle angle, t he I index marks merely provi d e an indi c ation that the automati c system is operational.

_" _ P rimar_ Controls ' _

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Th e co n t r o l sti ck a n d p ed als in t h e s imu l at o r r e qui r e d n o mo di- fication f or the M odel 2 22 si mul at i o n. Lon gi t u di n a l st i ck

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I CAB INSTRUMENTATION : I Instrument Range Vertical Situation Indicator +90 ° Pitch and Roll • Airspeed 0 _ 520 KIAS I Horizontal Situation Indicator +120 ° Heading Pressure Altimeter 0 _ 10,000 Ft Radar Altimeter 0 + 1000 Ft I Rate of C limb + 6000 FT / MIN Tur n and Bank +3 Needle Widths +i 1 / 2 Ball Widths , fl Nacelle Angle 0 _ 120 ° I "g" Meter -i, +3 "g" Clock i <. |_ Sideward Velocity +_ 40 Knots _ Angle of Attack +_ 20 ° Wing Flap Position 0 _ i00 ° !_ Rotor Speed 0 _ 125% _ % _ Engine Torque Meters(2) 0 _ 125% ,, PRIMARY FLIGHT CONTROLS :_e- Stick (+6 " Long.; +5" Lateral) ;i:i. , Pedals -{'+ 2 .5") ,- * ,, Power Le_er (0_8" Normal ; 0- . I0" Emergency) / _-_ _ Nacelle Position Thumb Switch MISCELLANEOUS EQUIPMENT AND FEATURES _ * Back Drives to Trim Stick and Pedals while in Initial Condition (I.C. ) Landing Gear U p - Down Switch with Indicator Light "' "" I! SAS ON - OFF Switch Detent Switches on Spring Cartridges (Pedals & Lateral Stick - Magnetic Brake on Pedals, Long. and Lateral Controls Long . and Lateral Beep Force Trim on Stick Power Lever Null Meter Toe Brakes X Specified Force Feel System

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t l_ FIGURE 7. MODEL 222 PILOT STATION FEATUR E SUMMARY

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i[ t ovol was mechanically limited to 12 inches lateral stick

travel was mechanically limited to _5 inches and pedal travel was mechani c ally limited to +2.5 inches. A beep force trim "hat" switch was mounted on {he stick. This enabled the pilot was used for precise trimming of the aircraft at cruise speeds.

! t o zero out longitudinal and lateral stick forces and also A compromise beep trim rate of 1 / 2 inch / sec was used for most of the test program. A magnetic brake, operated by a button _j o n the stick, was used to zero the stick and pedal forces simultaneously. This was used primarily in the low conversion speed range. Detents on the lateral stick and pedals were set at _.050 in c hes.

The Model 222 Tilt Rotor uses a single lever to command the power of both engines and to provide collective pitch lead in hover and transition with rotor speed con t rolled through a ill governor. Rot o r speed is programmed as a functi o n o f nacelle f _ _ incidence angle. Rotor speed is maintained at 551 RPM to a nacelle angle of 45 ° . From 45 ° to 0 °, the rotor speed is

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linearly decreased to 386 RPM. A p_oportional thumb switch with detent, breakout and gradient mounted in the hand grip, . _ controls nacelle tilt. The power lever arrangement, shown in _ _ Figure 9, is mounted on the left arm rest. This can be rotated °_ up for easier pilot entrance and exit The power lever / collec- ' horizontally at approximately the center of the hand grip) an d I tive control has a normal travel of eight inches (measured simulates the range of engine powers from flight idle to maxi- _ .

mum power. For single or dual engine failures, direct pilot l control of collective pitch for a flare is provided by sliding I the power lever through a detent on the arm rest. This turns off the rotor governor and transforms the power lever into a I collective lever. Two inches o f overtravel (measured horizon- tally) were provided in the simulator. There was essentially no breakout or gradient (except that provided by fric t ion) in the power lever.

S t ick and pedal breakout forces and gradients were develope d _ to meet the stick force per "g" requirements for satisfactory _ I flying qualities as specified in MIL-F-8785B(ASG); an d to im- _1• prove control harmony among. _xes. Stick and pedal force gra- _': dients are specified as a func i o n of dynamic pressure, with use d in the Model 222 simulation are shown in Figure l0 for l the breakout forces constant, fhe breakout forces an d gradients " / the longitudinal and lateral sticks and rudder pedals.

computer I Two generated visual displays were available for use during the test program. One is a symbolic representation of a helic opter la n di ng pa d o n a d e stro yer a f terde c k. I n t h is _ d i s pl a y th e b od y axis geome t ry i s tr an s f or me d t hrough a n ear th

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axi s system to a point in the a i r c raft , with t h e resul t t ha t ship an d air c r a f t m ot i o n are i n de pe ndent ly p oss i b l e . T h i s i c a n b e u s e d for t he ho v er an d n ea r -h ov er mo de . T he ot her i • '3

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NACELLE UP I NA CE LLE POSITI O N SW I T C H I N CRE ASING POW E R

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NA C EL L E D O WN HAND GRIP AND R E ST a_ " 2 ,., DETE NT

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[, - / 1579 0 2 _ . . F I G U _ 9 . P_ E R LE VE_CO L LE CTI VE C ON T RO L FOR M222 S IM_A TI ON ! .

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I Gr ad i ent _ L B / IN L at e ra l S t ick F o rce ! [

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Sideward Flt. out Force Constan t - l.0 LB Ba c kwa rd a nd --- /// _Br e ak

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t _t',_ _ , i - 20 .... / f B ac k ward a n d / !

l i : / I Breakou t Force Cons t an t =6 ,.0 LB I S ld e w a r d Flt.- - / - T 1 i . Q '_ .... J I . 40[ ................. , Gradient L B / IN + - _ : I L o n git ud i_l Stick Fo r ce _ ' :

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Bac kw ard a n d ......... . ' Si de wa r d Fit I _ B r e ak o ut F o rc . C o n ,ra nt 1 0 LB

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I I | , l ' , l i0 0 0 I Q 0 3 00 3 00 4 00 50 0 DYN_ I C P RESSU R E _q_LB / F T _ I FI GU R E I O. MODE L 222 C ONTR OL F O RC E G _D I ENT S AND BREAKOUT FORC ES [ 2 5

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display is a symbolic representation of a road with telephone I p o l e s on th e s i d e for r e fe r e n c e . Mou n tai n s are pr o vi ded i n t he background. This is used primarily for cruise mo d e stu d ies, al t hough i t was used almos t exclusively for t he Model 222 I simul ato r prog r am.

Additional fea t ures of the cab included a landing gear I p-down pedals an d SAS on-off swi t c hes. The simulator's primary con- I lever wi t h indica t or light, toe brakes on the directional trols i.e., stick and pedals are back driven to an ini t ial trim po s i t ion from signals computed in the mathema t ical mo d el.

I Ini t ial power lever / collective control trim is accomplished by the pilot by moving the lever until t he power lever t rim i ndica t or (null me t er) is z ero. The s i mulator can t hen be I "flown" by depressing the operate switch. While the simulator is in opera t e, the back d rives are inoperativ e .

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i 8 . 0 SIMULATOR RUN PROGRAM

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. The run program developed for the Model 222 piloted simulation approximately equal emphasis on all flight modes. Since t he i was predicate d on evaluating the full flight envelope, with _ scope of this program was large, and the objective was to evalua t e t he f ull flight envelope, t here were some conditions _ I that were not evaluate d in dep t h. These areas are not_ i in • Sec t ion I0 and are recommended as subjects for additional "_ work. The simulator run plan was as follows.

I i. Familiarization b) "Learn t o fly" i a) General comments on cockpit layout 2. Hover Mode S tudies I a) H eight control c apabili t y - adequacy of control I - precision of c ontrol b) Longitudinal and lateral stick and pedal i I pulses (from trimmed flight conditions) c) SAS evaluation (evaluate on, off; rate , } .

attitude , LAS) I d) Control sensitivity - adequacy of respo.,se ?

I e) Control response to large inputs - con t rol co u pling f) Rcspons_ in gus t s < / ' _ g} E ngine out operation :_ 3 . T ransition Mode Studies a) Slow acc e leration an d d eceleration through _.'_ ._ I t ran s i t ion _ _'_ , b) Rapi d accelera t ion and decelera t ion through i transition c) Con t rol se nsitiv it y (roll , pi tc h a n d yaw) d) I L ongitu d inal and la t era l s t i c k and p ed al pulses (from t rimmed f ligh t conditions) U - _ 2 7 • 0 . _ , L

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_ e ) SAS evalua t io n ( e valu a t e o n, off; r ate , _ I '_ att itude, L AS) • f) Cont rol resp o nse to l a r ge inpu ts - c o nt r o l c o upling d _ I - ad equa c y o f re s pon se g) R e sp o nse in gusts

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h ) F l i gh t p a t h co n tro l (i n clu din g l o w sp eed c li mb s an d d e s cents) I - ad equacy of contro l - p rec isi o n of con t r ol ' i i) Contro l sens it iv it y - - j) E ngin e ou t o per at i on I 4. Cruise Mode Studies a) Lo ngi t u d inal, l a t e r al st ic k a n d pe da l pu lses I (fr o m tr imme d fligh t c o n d i tio n s ) b ) SAS e va l ua t i o n ( e v a lua te on , o ff; rat e, I a ttitud e , LAS) i c ) M a ximum ac c e leration a n d d ecele r ati on .

B d) Co n t rol r esp o n se t o lar g e inpu ts _ - a d equacy - c on t rol c oupling _ _ e l R es p o ns e t o gu sts l f) Cl imbs and d es ce n t s _j•_ .

g) En gin e o u t op eratio n " i I 5. E v a luatio n of M ax im u m N ace l l e R at es i n T ran siti on _*_ 6 . Eva lu a te He lic o p t e r F ligh t M od e .4.y Vo ic e r e c o rd e r , us ed p a rt o f t he ti m e , a n d 36 channe l s o f . _- b rush recorder data w ere obtain e d f o r e ac h m a neuver exc e p t f o r th e familia r i z ation runs where o nly bru s h re c orded data w e re o btained. These will be retained for fu t ur e r e fer e n ce .

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i 9. 0 PILOTED RESULTS Th is sect i o n co nta i ns th e pi lot comm e nt s th at w e re o bt a i n e d I during the piloted simulation p o rti o n of this program. The primary experien c e of the Tilt Rotor p r oject p ilot is in flyin g helico p ter s . His ex per ien c e in f l ight testi n g o f V / S TOL ai r sh o u l d b e interpreted accordingly. Du r ing di s cus s i o ns held i c r aft _s l i mited. Therefor e , t he c o m m e n t s pre s en ted h ere i n p rior t o the start of this p ha s e, t h e pilot wa s inst r ucted t o be as critical a s pos s ible to e nable the Tilt Rotor project I to def in e t he sign if i ca n t p r obl e m a re as. Al tho ug h the M od el 2 22 as si m u l ated had g e n e ra l ly acceptabl e s t ability an d hand- l ing c h ar a c teristi c s, the following modificati o ns have been I incorporated as the result of the pilo t s c o mm ents presented in this secti o n.

increased a s follo ws : - I • H o ver control pow e r and sensitivity have been I Axis _o ntr o l Sensi t ivity R ad / Sec 2 Control S e nsitivi t_ Rad / S e c 2 !

This Si m ulati o n ! Current This Simulation [ Curren t • Pit c_ .6 i 1.2 .i .2

u Roll i 1 . 0 20 2 L .4

.5 . 5 .2 I .2 Y a w ._ : L • i II • La t eral co ntr ol p o wer in tran s i ti on a n d c ruis e ! , w as increas e d by u s ing fu l l s pa n aile rons a n d | sp o ilers compared to p ar t ia l sp a n ai l ero n s and I( s p o ilers use d in this simula t io n ' J I fli gh t mo de s w a s re d u ced .

• T hro t t l e s e ns itivity in th e hover a n d l ow s peed I i • The governor was mo di fie d for improv ed r esp on se m o dified w a s of • Repre s enta t ion engine dy n am ic s to mo re c l o se l_ m a _ch a c tual en g in e res pon se c haracteristi c s.

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• S tability A ug m entati o n S y st em (SAS ) re fin e d .

The h ov er an d l o w s p ee d SA S gains a n d sh api ng Crui se m o de r o l l and yaw S A S feedback loo p s I were mod ified t o obt ai n im pr ove d resp on s e.

-_ we r e added to e li min ate r o ll / s piral co u pli n g • and t o reduce t he h i g h dihedral eff ect . - _ - - - Q • u

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I it sh o u ld b e n o te d that the ab o ve list i s n o t all- i n c l us i ve, = but is provided t o i l lus t rate t he importance of early pilote d I si m ulation as an a i d in the air c raft design. A further dis- cus s ion of the piloted s i mula t ion results and pilot commen t s is in S ec tion 10.

I These pilote d studies were all conducted for a n aircraf t gro s s weight of 12,000 ib with the cen t er of gravity at 28% aerodynamic characteristics are as described in R o f a renc_ 1 I chord (most aft CG at this weight). Configuration details an d an d summarize d in Se c tion 3.0. It is to be emphasized that the aircraf t simula t e d d urin g this program is not the same as d e- I scribed in Boeing Document D222-I0050, Volumes 1 to ii (Stu d y of V / STOL Tilt R otor Resear c h Aircraft Program - Phase I).

£h e aircra f t g e ome t ry is essentially the same. Weight s , I inertiassaero d ynamic data, load alleviation, and SAS _ o nfigura- _ tions have been revis e d.

The pilo t c ommen t s are presented in the same o_der as shown in run plan. Generally inputs I the where the size of the control Z s not note d , the pilot was attemp t ing to put in l-second puls e s ( i inch of control at low spe e d an d i / 2 inch in the I cruise mode).

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g.1 FAMI L IARIZATION

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9.1.I General Comments on Cockpit Layout I The following items were noted concerning the cockpit layout.

a. Nacelle tilt indicator is a primary flight I instrument and should be located near the basic Tee. Rate of climb, airspeed, torque- meter and nacelle tilt indicator s_ould be I lo c ated on the same side of the basic Tee.

b. Nacelle tilt switch arrangement needs further or thrust lever), direction of travel, force I investigation. Switch position ( c y c lic sti c k gradient and breakout, proportional rate, fixed rate beep or two rate beep need I eva l ua t ion.

9.1.2 "Learn t o FI_" % I Mu c h time was spent in trying to determine the best way of _ ; _ a wide range of n a c elle tilt angles and body attitudes.

g ',_ flying through transition. The aircraft is very toleran_ _f i _ I The most c omfortable or convenient combination may be selected for the parti c ular task to be performed. The operation of the aircraf t in the c ruise mode is conventional.

I During this period the following items were n o ted.

I a. Operation with na c elle incidence above zero at spee d s above about 160 knots is undesirable.

Positive prevention of such operation shoul d be c onsidered. An automati c up stop at 90 to I 95 ° is desirable to facilitate reconversion.

Posi t ive pilot action should be required to go beyon d this se tt ing.

I b. The mag. _rake is too coarse to be used for trim in the cruise range an d shoul d be lo c ke d .

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9.2 HOVER MODE STUDIES 9.2.1 Height Control Ca pability in Hover I The height control capability of the Model 222 was evaluated by performing a series of vertical climbs and descents to specified altitudes. Adequacy and precision of control were I evaluated. The pilot comments for these maneuvers follow: - Initial Condition Maneuver Pilot Comments HOVER All SAS on(Roll, Perform rapid Difficult to control rate pitch, yaw). climbs to 50', of climb with existing System (LAS) on. altitude. Then The combination of high I Load Alleviation 150',and 250' power lever sensitivity.

descend rapidly throttle sensitivity and 150' to 50'. resulted in overshooting i from 250'to po o r external visual cues target altitudes by as much as 50'. A +10% torque I change in this man--euver typi c ally resulted in +i0 00 ft / min vertical rate. - As above but This was a more n atural at low verti- altitude change and was cal rates and much easier to control. • changes torque I 50' increments 2 to 3% in in altitude gave 200 to 300 ft / min verti c al rates. The - I maneuvers were fairly well "_ controlled although the i_." _ ' .

power lever was still quite ' _' - "_ " yaw were flown hands off and _ . _ I sensitive Pitch,roll, and _ . ,_ .- _ . :_ , these axes seemed well j " I Vertical control. S n tall altitude corrections _ : , _ Small height were 4 ifficult to a c hieve ........

stabilized. ! _i changes. Task and the difficulty was com .....

I was to hold al- pounded by poor visual cues.

titude as The slightest pressure on closely as the power lever was suffi- small c hanges by a few feet. The power I possible after c ient to change altitude in altitude lever sensitivity was not hover. Plus or minus 4 feet i were made. adequate for a pre c ision w as t h e bes t the pi lot could a c hieve•

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Initial Plus or minus 2 feet alti- -i I Condition Maneuver Pilot Comments _ tude could be held with a lot of work. Once trimmed, the aircraft held fairly I well.

9.2.2 Control Pulses, SAS Evaluation and C ontrol Sensitivity I in Hover r Control pulses (longitudinal, lateral stick and pedal), evalua- back loops off, and control sensitivity were evaluated, l_ote I tion of aircraft characteristics with various SAS and LAS feed- that the SAS evaluation was not a malfunction _nalysls Dut an evaluation of aircraft handling qualities with various compo- maneuvers I nents inoperative. The pilot comments for these follow: - I Initial i Condition Maneuver Pilot Comments i SAS and Load Pitch Pulse Aircraft behaves well, Alleviation returns to trim attitude System (LAS) On with no oscillation. C on- I trol sensitivity adequate.

Roll Pulse Well behaved response.

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I Control sensitivity _ adequate. _ _ Control sensitivity ade- !_!_ I Yaw Pulse Well behaved response. :_:_" quate. The yaw axis was / / _"i .

heavily damped and stopped _ input was rem o ved. _ > I immediately when the pedal _ HOV_R-LAS On Roll Pulses Fairly long period, ( 20 ,._. Ro--6_ Att itu d e sec), neutral, dampe d roll SAS Off oscillation All Roll SAS Roll Pulses Slow roll divergence Off

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-_ 33 Ini t ial Pi t ch A tt i t ude Pi t c h Pulses - Pit c h oscilla t ion developed.

SAS Off Nose Up Nose came up i0 ° , checked at I zero, then pitched down to . - 5 ° attitude. Oscilla t ion I_ _ C o n ditio n Man euver Pil ot C omments was neutrally damped.

_ I _ Pi tc h Pulses - Developed 3 degrees nose Nos e D o wn d own attitu d e on inpu t .

• A plus or minus 4 to 5 d e- gree neutrally damped | _ oscillation developed.

si gn i f icant c hang e I A l l Pi tc h SAS Pi tc h Pul s es No fr o m i Off S AS on. Pitch a xi s slowly : divergent and oscillatory.

I Yaw SAS Off Yaw Pulses Not much inherent damping in yaw alth o ugh s en s i t ivi t y !

I was good. Yaw axis ca m_ back to t rim and wen t di- ve r gen t in t he opp o s it e not repeatable and may be I direc t ion. This e ff e c t was a func t ion c f no t g e t t ing p edals ba ck t o t r im. F e e t I off pedals gave a left yaw rate that require d 1 / 8 inch righ t p ed al t o c he c k. Yaw I a xis very ligh t ly d amp ed .

HOV E R- L AS Off Pi t ch, Roll The pi t c h a n d roll axes A l l SA S On and Yaw Pulses looked ab o u t t he s a m e . _ I The yaw axis a gain w a s heavily damp ed and sto ppe d :_ i m m edia t e l y when p ed a l in - put was rem o v ed . There _ w e re no app a r e n t d iff e ren ces _ _ L AS on an d L AS o ff. _', H OV ER - L A S O ff Roll Pulses Ro l l osc i llations of R o ll SAS off shorte r p eriod than with LAS on, ne utral to con- v e rgent. Control r es p onse w a s sluggish.

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Ini t i a l I C ondition Maneuver Pilot Comments Pi t ch SAS Off Pitch Pulses P ower c hanges had qui t e an response to pulse inpu t s was i effect on pitch. The pitch axis slowly divergent and no t much different from that with LAS on.

I Response rather sluggish.

Yaw SAS Off Yaw Pulses Sensi t ivity an d response was I very similar to LAS on flight.

HOV E R-All SAS Pulses about The aircraft was manageable.

an d LAS Off all axes There was some yaw-pitch coup- right pedal I ling present; gave pitch down and lef t pedal gave pi t ch up.* The longitudinal I stick trim position was more . forward than with SAS on. In hands-off condition air c raft yaw left. SAS off control sen- I I was unstable with a tenden c y to sitivity was adequate. " HOVER-All SAS Pulses about Nc significant change from Off, LAS On all ax e s LA_ off.

9.2. 3 Response to Large Inputs in Hover comments for t he s e maneuvers are shown below:- U The response to large control inputs was evaluated. The pilot Initial l Condi t ion Maneuver Pilot Comments _- HOV E R-SAS an d Response to A 2 in c h longitudinal stick in- , : _' - Pit c h change. 2.5 inches of la t eral _,'_, l LAS On large inputs put produced a 5 degree attitude _,_ - Roll stick res u l t e d in 15 degrees of _ - ,_. . _.,- pedal; response was well d amped. _.._._l : l - Yaw bank. Applied 2.5 in c hes of _ , _l_,_ A d equacy of response dif fi cul t ......

t o evaluate b ec au s e o f motion I cu e s .

c o up li ng sin c e both engines t urn in th e sam e I * Note : This is the r esult o f e ngin e _n erti al d i recti o n .

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Q Initial _. Condition Maneuver Pilot Comments ' I _ HOVER-LAS On Large inputs in Except for la[ger e_<cur- 4. Pitch SAS Off pitch, roll and sions from trim pitch yaw attitude, response was |_ not very dif: u rent flom I! . SAS on. Roll inputs resulted in a small # , _ : amoun t of pitch up.

caused , ) Right pedal inputs % pitch down, left pedal, pitch up.

i Yaw SAS Off Large Inputs Yaw wa_ mo_e respor_s_ve than with SAS on. T_e (_ response was very sensi- tive and inherent damping was low. Yaw axis was un- stable in hover. Pitch inputs gave no coupling.

Roll inputs gave some proverse transient couplin g .

_i 1 Roll SAS Off Large Inputs Aside from roll instabi- _ ' lity, roll response was not on. With full l_tteral _ iillii very different from SAS ' " input , initial resp o nse was the same as SAS on. , Sensitivity was low for .. ! smaller inputs. Pit c h inputs gave no coupling. ' velocity and roll due to !

_ / i [ Yaw inputs gave lateral -, i dihedral effect. _ : _ t j 9.2.4 Response to Gusts in Hover • ){I The response of the Model 222 in random turbulence w en studied i during this piloted evaluation. Pilot comments are shown for 4 and 3ft / sec RMS turbulence. It should be noted t ] _t ti cs U i Ii are moderately severe random turbulence levels. Initial Condition Maneuver Pilot Conunents

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. | HOVER-All SAS Evaluation of Attempted to i_old alti- and LAS On response in gusts, rude at 95 feet. Used I _S gust velocity rat_ of cllmb instru- 3f% / sec ment to hold altitude.

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_i initial Condltion Maneuver Pilot Zomments | Torque varied between 33%

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and 70% as commanded by : .

the pilot. Started at 93 feet altitude and ended up

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at 195 feet. The task o f changing altitude and trying i to stabi l ize was diffic u lt with this level of gust.

Acceleration cues really successfully.

needed to modulate power : HOVER-LAS, Evaluation of Flown with hands off on , f - I! SAS On response in pitch, roll, an_ yaw. The , [_, gusts. RMS vertical axis was very g_st velocity difficult to hold. in -'_ |. 4 ft / sec severe gusts pi_ch holds ; [ +5 degrees, roll holds +3 ' - degrees, yaw was pretty- } _ much locked o n . It was _'%_ i _ impossible to trim out at _;_ : . i any particular altitude Power was 50% for hover :, . .. , plus or minus 20% to hold _ : al t itude. Rate of c limb : i _ wen u +300 feet per minute " _ t and occasionally to _i000 i_ feet per minute. R o tor RPM held well.

9.2.5 Ep_ine Out Operati o n in Hover =' " i During this maneuver, one engine was failed with the aircraft in a steady hover, to evaluate the engine out landing capability.

The pilot comments are n oted below. _'_ i Initial . •.

Condition Maneuver Pilot Comments ,,_ } HOV E R-All SAS On Single engine Dev e lops 1000ft / min rate of ] an d LAS On failure from descent in about 1 sec ioo ft. af t er failure. S t arting a t forward t hro tt le into over- i 8 0 ft of a l t i t ude, use of t ra v el region, checke d d e- - scent at 4 0 f_. i00 ft / min l rate of c l imb obtained at full forwar d thro t tl e . - - ,

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I Initial Condition Maneuver Pilot Comment i Detent position on throttl , - _ i ! checked rate of des c ent % c _ 500 feet per minute. '_'ra , ,- _!_ sient following failure i._id.

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| i 9.3 TRANSITION MODE STUDIES 1 9.3.1 Slow and Rapid Acceleration and Deceleration Through Transition I Slow and rapid accelerations and decelerations were evaluated.

( These were con d ucted with the stability augmentation and the load alleviation system on. During the pilot familiarization portion of this program, the pilot flew a slow acceleration I through transition to 150 knots an d back to hover with the SAS and LAS inoperative. He stated that i t required considerable ! pilot effort and attention. It should be noted that this situation (all SAS an d L AS off) would require several malfunctions in the automatic stabili z ation system because of the dual and / or triple redundancy. The pilot comments of the slow and rapid , / I- _ accelera t i o n a n d d e ce ler at i o n cha racte ri s tics o f the Mo d el 222 , _ through transition with the SAS and LAS on, follow.

" _. _ r Initial ''' i I Condition Maneuver Pilot CoI,_.ents _ . _ HOVER-All SAS Slow acceleration Diffi c ult t o co z .trol alti- _% ( on LAS On from hover to tude. The motion-base _ L airplane flight postural t ilt gave decep- _" tive motion c ues. O_her- :_- ( wise docile.

Rapid acceleration S t rong pitch-down motion from hover to air- aggravated by postural .,_ plane flight tilt makes this maneuver ,. _ difficult. Full aft stick was required to hold alti- i tude near zero.

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"_ ' Rapid deceleration R e quires considerable tech- from cruise nique. With a fixed-rate ' i nac e lle t ilt con t rol t here i was a tendency to overshoot on required nacelle angle.

,_ AS minimum power spee d was was c on t roll e d wi t h pitch a tt i t u d e. Upon reaching use of throttle was required _ minimum power speed prompt j approached rate of climb --- to prevent a sharp increase i n d escent rate.

_. Slow d e c eleration No diffi c ulty encountered.

from cruise Close control of alti t ude re qu i re d .

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,Ib Q '7 i ! 9.3.2 Control Sensitivit_ and Aircraft Response to Lon@itu- dinal Sti ck, Lateral Stick and P edal Pulses in Transition Con t ro l sensi t ivi t y a n d ai rc r a f t response to lo n gitudinal stick, ! la t e ral s t i c k and pedal pulses in transit z on were evaluated f o r i 20 knot intervals through transition. 20, 80 and 120 knots have been selec t e d as the conditions to show the pilo t comments.

,i , } %he stability augmentation system (SAS) and load alleviati o n _'st e m were in oper a tion. The pilot c omments f o r these maneu- , _L_rs follow.

I Initial _ondition Maneuver Pllct Com_ , e n ts

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' V=20 Knots Pi t ch Pulses / I- %, -85 ° retur n to trim following I Na c elle Angle Pi t c h Up 3 ° attitud u c_.ange an d a ' , _, All SAS & LAS On a 5 ° overshoot

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_, Pitch Down Same as pitch up pulse f 8 - Roll Pulses Same as the response in _ ,. hover _" Pe d al P uls e s Same as the response in ;_ . _ hover J _' I _ V=80 Knots Pit c h Pulses Pitch - less apparent pi t ch = 6 0 ° stick forces. There was " / 1 Na c e lle Angle response, due to higher • _ I P_tch Attitu d e more damping than at 20 ": _ =4 ° knots. Sensitivity of the All SAS & LAS On response a bit sluggish ( with the h_gher damping.

_, Roll Pulses Aircraft rolls to a bank • angle and holds well.

Som e small sideslip a ngle de v e lops.

¢ Pe d al Pulses Aircraft is h : _ , : ; iy d amped.

V = 120 Knot s Pi t ch Pulses Response w e ll d amp ed an d I_ Na c elle Angl e sluggish, similar t o 8 0 k t s : = 2 0 _ Con t rol r_s p onse was adequa te .

All SAS & L AS On Roll Pulses Roll respons e seems be t t er j with good bank a ngle hold.

, " Pe d al Puls e s _esponse was a little weak, with high dih e dral e f fect.

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. 9.3.3 SAS Evaluation in Transit ion I_ %? . e response of the Model 222 to vari ou s SAS configurations in [ ! transition was evalu a ted at 80 knots. The pil o t c omments are p resented below.

li Initial C_ndition Maneuver Pilot Comment _i ' - u 0 Knots Pitch Pulses Trimmed at £0 _ na;elle inct- LAS On dence, 5 ° pitch attitude, and t o il & Yaw SAS 40% torque. The pitch axis was I _ On convergent and returns to trim Pi u ch rate and in about 2 seconds with a 1.5 longitudinal degree overshoot. Pitch damp- _tick pickoff on, ing was high.

bitch attitude off Same as above Pitch Pulses Pitch response onl_ slighuly : |_ with pit c h rate less stable but still convergent.

. [! _eedback off

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f - I| Same as above Pitch Pulses There was an apparent degrade- . % _ ; , . t : All pit c h SAS tion of sensitivity without the .D Off, LAS On pickoff• Pitch response had a • . , small overshoo_ and developed _ a convergent long perio_ oscil- " lation.

Same as above, Pitch Pulses Longitudinal st_:k trim moved LAS Off 1 / 2 to 3 / 4 of an inch more forward• On_ inch pitch pulse of atti- . . |i generated 3-4 degrees - tude. Long period os c illation developed with longer period I than LAS on case The pitch re- " - " ', sponse generally very similar ._ to LAS on case. _ .

V =80 Knots R oll Pulses Inpu * genera¢_ 7-_ degrees of ,[;_[_ Pit c h & Y a w r e turn to t rim• Roll c o ntr c _ " _ I SAS On, sensi t ivity seemed low . ,_ " Roll .tt_tu d e hol d Off _ [ J LAS On bank angle with a v_ry slow _ : [ Same as above R oll P ulses The basic aircraft had good w it h ra t e an d roll response. Poll rate ' att it u d e r oll da mpe d ou t re a s o n a bly well | SAS O ff , L AS O n wit h a very slo w re t u r n to -- - - trim.

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I Ini tial Condition Maneuver Pilot Comment i Roll SA S Of f and simila r t o S A S on i . e .,

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sl u gg i s h

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Same as a bo v e Roll Pu ls e s Ro ll response s imilar to L AS LAS O f f on case above. Slight roll re t urn to tr_m.

I a ng l e o v e rsh oo t and a slow Roll an d pi t c h d amped t o pedal inputs an d also i! 80 K nots LA S On Pe da l P u l se s Re sponse was stiff and h_gh l _ Yaw a t ti t u d e were unexplainably high [or th_s . / _ - ! hol d o f f case . (For c e feel sys t em I I SAS O n very slugg i sh. Pedal forces -_ _ m a lfunction) . _ Sa m e a s a b o v e Pedal P ul se s Peda l pulses generated 2 cycles , Yaw SAS Of f of fishtailing with high pe d al

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_ R o ll in to Ya w for c e s a gain. (Force feel SAS F u n c tion sys t em m a lfun ct ion)

II . O n

:_ " Sam e as abo v e Pe da l P ul se s Same response a s above '_ all Ya w SAS O ff <; ;. | ,_ Same as above Pedal P ulses Similar to LAS on case above.

I, LA S Off NO appar e nt change in yaw re - i, _; s po nse o c c urred with L AS off.

in Transition . ( I i_ 9 . 3 . 4 Control Response to Large Inputs A i r c ra ft r es pons e c harac te risti c s to large inputs were eva- ', 11 f o r lU a te dand f o und t o b e si m ilar to s m a ll in pu t r e s po nse s ex c ePt a mp litude . !

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_i 9 . 3 . 5 Aircraft Response to Gusts in T ra n s i tion " The resp ons e of th e Mo d e l 22 2 t o r a n do m t u r bu l ence at 80 kn ot s m_ i n tra n siti o n wa s e v al ua ted . An R MS gu st v al ue of 4 fe z sec w a s i [ _ used . T h e pil o ts c omm e n ts a re n oted b el ow.

U Ini t ial Condition Maneuver Pil o t C om m e nts V-8 0 K n o_s Resp o nse t o H o lding v e rti c a l airs pee d w a s T A ll S A S On a RM S g ust the on ly pr oolem. He l d r ate LAS On o f climb with p it c h at t it ud e .

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Con dition Maneuver Pilot Comments _ Ini t ial r S am e a s a b o ve, P i tc h and ro ll at t l t ude hands off disturbances averaged about +i d_gree from tr_ . Air- speed varied about + 5 kno t s

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from t rim. Rate of-climb varie d +500 ft / min about tzlm.

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9•3.6 Flight Path Control Evaluation _n Transition Flight path control c apability of the Model ; 2 2 was ev_aated for s everal c ondi t ions. The pilo t s c omments for these maneuvers is no t e d below• • / .... 1 Initial Con dition Maneuver Pilot Comments j , _ " _ V = 230 Knots Decelerate to Decelerated tc 150 knots 3000 ft alti- 1 6 0 knots, 2500 ft with no difficulty and !_ I _ t u d e All SAS On of altitude af t er brought na c elles up to _! LAS On a 90 ° head i ng 25 degrees inci d ence. At !

_,, change , t hen 100 knots altitude or rate .... stabilize on a of climb required pil o t _ : _ 500 ft / min rate attention Stabilized _ |I of des c en t , and on a 500 ft / min rate of _ deceler a te through d escent at 120 knots and _ |_ transi t ion t o hover continued to beep nacelles _ _ at i000 ft alti- up reaching 80 knots witho_t ru de . d iff i culty. Faile d to apply power and lost con - |Y |i t r o l at approxima t ely 80 u d egrees nace l le incidence.

I_ _ am e as above S a me as above mecelerationthrough transi- |_ b u t w it h maxi- t ion a c cor:,plished with a mum n a ce l l e gain in alti t ude of i00 t o , - _ beep rate set 200 ft. Lower maximum a t 5° / se c f or n ace lle rates woul d b e de-

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t ransiti o n s irable a t th e hi g h spee d e nd of t r an sit ion.

I V- 6 0 kn o ts I nves ti g ate r a t es Set u p a r a t e o : descen t ___L SAS O n LAS O n o f d e s c e n t and o f 500 f t / m_ , at 35 % t orque 70 ° Nace lle m an e uv e rabi l ity ( 4 0 % re q uir e d f o _ l e vel to 1 5 _ a nd o bta i ned 700f t / I i n c i dence fli gh t ). Banke d air c raft m in r a t e of d es ce nt .

V Sa nke d a_rcr a f t to 2 0 ° i00 0 I and o b t a ine d ft / min ra te of d escent.

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I Initial Conditio., Maneuver Pilot Comments |i Aircraft was well behaved.

Lev e l e d of L an d reduced power to maintain i000 ft / min d u scent.

I! t 2 degre e l_ng period pltch os c illation developed.

Attempted a 1500 ft / min r ate I o f d es c ent u n satisfact o rily.

I L o s t c o ntr o l after pitch down, d ue t o wing stall and failure

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t to apply pow e r.

Same as Wings-level, Trimmed at zero rate of climb, above a t par t ial power 65 kncts, 5 deg. angle of S ! 65 Kno ts r ates of de- atta c k, 260 0 ft al t itu d e .

s ce n t Recorde d a_,_t u of atuack vs rate of descen t as follows: De c re a sing I 0 deg @ 500 ft / min Power 15 deg @ 1200 ft / min 18 deg @ ]50 0 ft / min d o wn du e to w i ng s t all.

L os t co ntr o l af t er pit c h - : 9 0 K n o ts , S AS A chi ev e r ates R ea ch ed 5 0 0 ft / min de s c ent i II On, LA S O n , 4 5 ° of desc en t at 9 0 kn ot s, 4 ° pi tc h a t t i- nacelle incA- in 500 f t / min r ude 7.5 deg_ - ees angle of d en ce 37% t o r qu e i ncre m e n t s by a t t a c k and 30% t orque with a nd 6 . 5 ° pitch r ed u ci ng p o _ , e r no d iffi c ulty.

att i t u d e at A d es ce nt o f i0 00 f t / m_n w a s _ , zero ra t e of obtained at 2_% torque , 7 1 degree pitch at t itu d e.

cl imb d e g r ees angl u of attack , an d I A 1 50 0 f t / min yi e l ded i 0 dQ - _- g rees & n z l e of attack and I! -2 d egr ee s pitch a tti t ude _ (no s e down).

F

A 2 0 00 ft / min rate of d_scent ._$_ li . o f 4 00 0 Pilotr _quir e d 1 2 ft / m i n de gr e es'e l e c t e dan rate angl e t o o f _° ° f d es c entt ° a tt a c k !_ir ' , '!!

an d no t ed i n cr eased I_ ' _._ do wn a tti t ude _,_[uire d n°SQ t o m aintain a i rs p eed.

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Initial Conditi on Maneuver Pilot Comments 80 Knots, SAS Same as above The following trim rates of On, LAS On, descent were flown: tude 6 ° angle rude and 6 degrees angle 5 ° Pitch atti- 500 ft / min @ 2 ° pit c h atti- of attack, of attack and 37% torque i000 f_ / min 91 ° pltch atti- 60 degree nacelle tude and l0 degrees angle incidence of attack 1500 ft / min @ 0 ° <level) 'i pitch attitude and 14 degrees angle of attack A 2000 ft / min point was flown.

_i 40 Knots S_S On, Same as The following trim rates of _ LAS On,80 degrees above descent were flown at the 5 degrees pitch 500 ft / min 9 4 ° angle of '_ H nacelle in c idence, listed condition_: : attitude, and a t tack zero rate of i000 ft / min at 40 angle of _' 'i climb attack _i _ [_ 1500 ft / min at 18 ° angle of !_ attack and 50 k n ots.

_ |_ A slight pltch oscillation '* |_ developed at 1500 _t / min and _ 50 knots airspeed. - _ 9.3.7 Engine Out Operation in Transition _ Engine failures in transition were not evaluated per me. The . . : effect of reducing power is adequately covered in Section _ 9.3.6. Powers used were less than available with a single _ engine. The engine out transients and the ability to control _ _i altitude pre c isely at low speed (near hover)with single engine _._ power were not evaluated. These should be evaluated at a later _

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I 9.4 CRUIS______EE MOD___EE STUDIE________SSS • Cruise ............

This section presents the pilot comments obtained from control I 9.4.1 Longitudinal Stick, Lateral Stick and Pedal Pulses in pulses in the c ruise mode. Responses to longiuudinal stick, lateral stick and pedal pulses were examined over a range of does not have cruise SAS. These studies were all conducted with the load alleviation system ( L AS) on. The pulses in cruise were generally 1 / 2 inch for 1 sec c nd. Pilot co mments for these studies are presented below.

Initial i speeds from 140 knots to 260 knots. The simulated aircraft f " _! Condition Maneuver Pilot Comments i Control re- Trim Difficult to trim in roll - tends teristics at 140 Pitch Pulses sponse charac- to fall off _ - knots. Pitch Nose Up Reasonably well damped and attitude=6.5 ° returns to trlr_ Torque = 35% Nose Down Well damped and returns to trim Roll Pulses i Left Roll Rolled 5 or 6 degrees with a | very slow return to trim i Right Roll Rolled 5 degrees with a very slow return to trim Pedal Pulses Yaw rate reasonably well damped, i Control re- Pitch Pulses l sponse Nose Up Pitched up 5 ° , overshot to 2 ° teristics at below trim attitude. Developed I charac- D isymmetrical.hedral effect apparent and i_ I 160 knots, a very slow convergent pitch l _ , < " l Pitch atti- os c illation indicates poor trim- I _ rude =5 ° ability. Pitch oscillation not Torque =40% noticed at 140 knots.

reasonably symme t rical.

1 Nose Down S a m e cha r acteristics i . e . i_ ' Roll Pulses _i 1 Left Roll Reache d I0 ° bank angle an d _ 1 developed slight sideslip and _ ' " slowly returned to zero bank I angle with very long period. "_- _ - -

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I Ini t ial Condition Maneuver Pilot Comments Roll Pulses Right Roll Symmetrical roll response. In- no pitch inputs.

Pedal Pulses Input generated a 3 needle width ! _! puts essentially h ands off i.e.

_ yaw rate. Aircraft rolled 20 _ _. degrees after input. The yaw axis was well damped with a slight overshoot and slightly i os c illatory.

Control re- Trim Difficulty in attaining trim.

li istics at 180 of climb w i th pitch attitude.

knots. Pitch Pitch Pulses Very slight oscillation that _I i sponse character- Aircraft was sensitive to rate • was more noticeable for nose attitude = 3 5 ° . , Torque = 42% down inputs. There was a very LAS On slowly divergent long period oscillation.

Roll Pulses Left Roll Reached l0 ° bank angle an d gene- little bit of U rated a sideslip that returned to zero slip.

Riqht Roll Symmetrical response.

Pedal Pulses Inputs generated 2 needle width drove bank angle to 15 ° right I yaw rates. Dihedral effect -', for right inputs and I0 ° left for left inputs. '' ' Co ntrol re - Pi tc h Pulse " "_'_" sponse chara c ter- Nose Up I nput e xci t e d a pit c hing oscilla- I i s tic s a t _ 00 kno t s , ti cs o f l o ng period, plus o r minus _ ' _ Trim pitch a tt i- a couple of degrees at t itude, _, . . _( , ' _ ru d e = 1.5 ° plus or minus 5 knots airspee d , ......

° ' l Torque = 50% plus or minus 500 to 800 f t / min _ . _ • LAS On rate of climb. Long perio d seemed , neutrally damped.

I Nose Down Input gave .4g's, response similar to nose up input• I Rol l Pulses % L e ft Rol l A ch i e v es b a nk angl e a n d s lo wly r et urn s t o zero bank 47 _

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Q I Initial Condition Maneuver Pilot Comment Roll Pulses Right Roll Similar to ]eft input. Heavily damped oscillatory response I with first peak nose left• 4 Yaw Pulses Similar to response at 180 kts.

Ii single needle width rate gene- rated i0 degrees of bank angle• Beep Trim Vernier beep response was too for roll slow _ beep inputs. Be- cause of the difficuly of trim- • dency to overcontr o l. The ten- den c y to overcontr o l existed when the stick wasused for ver- il I li ming in roll there was a ten- nier control.

Control re- Pitch Pulses sponse charac- Nose Up Input generated .4g's and pitch Ii teristics at attitude of 4 degrees 255 knots• Trim Nose Down Input generated at 2 to 3 de- !

pitch attitude grees attitude change with a I ] =l ° 4 degree trin_ overshoot• Phugoid [ J Yorque = 58% slowly damped out.

LAS On tory and damped out and returned to trim slowly.

I! Roll Pulses Initial response was non-oscilla- I! Yaw Pulses A i needle width rate banked aircraft i0 degrees and was , i Control re- Pit c h Pulses i _* sponse charac- Nose Up Input generated 4g's 3 to 3.5 _eristics at degrees pitch attitude chan c ;_ 2 6 0 knots, and returned to trim, slight L_"_*_> t l J symme t rical. _._ rude = 0.0 ° to 300 ft / min. Airspeed c hanged ' ,, Torque = 75% about 5 knots• T o tal airspeed [ L LAS On variation was plus 2 kts to minus 4 kts.

I _ Trim pitch atti- undershoot in rate of descent --, _ _ Nose Down Input was .bg's, 3 degree attitude generated 1000 ft / min rate of descent. Returned to trim and _ overshot to 500 ft / min rate of | climb. Long term phugoid of +2 knots and _i degree pitch att[- Initial Condition Maneuver Pilot Commen ts " Roll Pulses Left Input generated i0 ° bank angle i that came back to 5 degrees ra_he_ I quickly, followed by slow return to 4 degree_, the air_raft re- _ mained at 4 degrees.

Pedal Pulses _ Right Pedal Generated a 1 needle width ya_ I : rate, I0 degree r l ght bank, _i then a i0 degr_ e left bank which was well co o r i lnated and drifte] to 13 or 14 degrees left bank . Le_ u Pedal Developed _" left roll angle.

Asymmetrl: response attributed Repeat Inputs Right Generated _ needle w_dtn rate H and 15 degrees u _ bank, then returned.

Left Generated 1 needle width rate i and 8 degrees of bank, then returned.

9.4.2 SAS Evaluation in Cruise augmentation sys t em , aircraft response to longitudinal stick, Since the simulated aircraft does not have a cruise stability lateral stick an d pedal pulses were evaluated with the load alleviation sys t em (LAS) off. These runs were conducted at 140, 180 , an d 2 6 0 knots. The load alleviation system simulated _. _< z eros ou t the rotor hub moments. The pulses in cruise were .....

generally 1 / 2 inch for 1 se c ond , except where noted. Pilot |_ _.• " _ I commen ts ar e s h o wn be lo w for thes e man e uvers. _ , _[ ; _ Ini t ial -' ; Con dition Maneuver Pilot Comments ;_ _ Control re- Pitch Pulses Pit c hed up 2 to 3 degrees, re- _-_ / _ sponse charac- Nose Up turned to trim and un d ersho t .

l teristics at No tendency to diverge an d te- l40 knots. LAS t_rns to trim airspeed and pi t ch off - Trimme d attitu d e well.

at 5 . 5 ° pit c h I att i tude , Tor q ue = 35 %

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I Initial Condition Maneuver Pilot Comments I Pitch Pulses Nose Down Pitche d down 2.5 degrees, nos e up overshoot with a very slight I airspeed c hange and returns to trim well.

inputs with a slight tendency I Roll Pulses Very slugglsh in r o ll for 1 inch for long period roll oscillation.

t Pedal Pulses Sluggish resp , nse with str o ng d ihedral effect. A 1 nee d le width ra t e generates 8 degrees _[ of bank angle. • / !

Control re- Pitch Pulses ' -,_I '_ s P Onse charac- Nose Up Generated 3 degree attitude change teris t i c s at with slow return. Airspee d fell , _ 180 knots LAS to 160 knots Neutrally damped _ of f long period between 180 and 160 _+ Trimmed at 3 ° kts. A t titude excursion reached _. _ pi t ch at t itude max. nose up 7 deg. possibly due • _ ; _ - : Torque 43% to mistrim. _ - .

: _"i _ Nose Down Pitched down 3 deg. from trim !

+_ and gained I0 kts airspeed, pitch _+ and reached 5° nose up, airspeed + - d ropped to 170 knots +i0 knot ; oscillation abou t trim.

_! attitude returned through trim i

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- Roll Pulses , Left Roll Aircraft continued to roll after | I r_mcval of input. Tendency t o , _ _iral instability.

response was sluggish and diffi - i_ Right Roll Response was symmetri c al. Roll " _ ) cult to trim. Required large _ +_+ and long inpu t s to maneuver.

II C on t rol re- Pe d al Pulses Response similar to l o wer ..- sponse charac- speeds, with high dihedral teristic s a t 2 6 0 e ffe c t.

I knots. LAS Off Trimme d a t 0 ° , pi t ch a t t itud e l an d 75% t ozque 5O Q

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I Initial Condition Maneuver Pilot Comments Pi t ch Pul s es ini t ially. Aircraft came back to trim an d overshot airspeed by i0 kts. Long te ; m os c illation was about the same as l o wer _ Nose Up Response was fairly well damped airspeeds, i.e., _i0 knots.

Nose Down Response reached .5g's initial- ly and was well damped. LOw amplitude long period oscilla- tion was apparent.

_ Roll Pulses . / _ Left Roll Aircraft was difficul t to trim in roll with LAS off. Initially , rolled to i0 ° left with a slow, F very slow, return to wings level.

Response was s l ightly uns_mme- _ trical, probably due t o mistrim.

_ _ Roll axis tended to fall off one _ way or the other at random.

_ teristics as lower speeds. _' .... _ Pedal Pulses Similar dihedral effect charac- _ ' _ . 9.4.3 Maximum Accel eration and Deceleration in Cruise : Maximum ac c eleration and deceleration characteris£ics of the i from 140 knots t o 250 knots an d then d e c elerated to 140 knots.

Model 222 were investigate d . The aircraft was accelerate d The pilo t com m en t s for this maneuver is shown below.

_ Initial Condi tion Maneuver Pilot Comments I Level flight A c c eleration at The aircraft accelerated at 140 kno t s maximum power t o slower than an t icipa t ed LAS On 25 0 kno ts , t hen d eveloping 500 foot per t o 140 kno ts of climb was sensi t ive to I decelerate back minute rate of climb. Ra t e _ pi t c h at t itu d e changes.

The roll axis was uns t ea d y but controllable. The con- t rol harmony between roll an d " pi t ch was no t very good. The I force feel system di d no t have

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a po s i t ive fe e l around zero force. Transien t s easily _ r controllable.

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Aircraft response to large control inputs in the cruise mode were investigated at 140 knots. The load alleviation system was on. Pilot comn%ents are shown below.

1 Initial Condi tion Maneuver Pilot Comments ! Level flight 2" pitch pulse Control response was adequate LAS On 2.5" pedal pulse undesirable coupling except i t _t 140 kts. 2" roll pulse in all axes. There was no for the high dihedral effect.

Trimmability was quite poor, particularly in roll.

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9.4.5 Response to Gusts in the Cruise Mode ' _{ The response of the Model 222 to random turbulence in the cruise mode was investigated at 140, 180, 225 and 250 knots.

_i The load allev i ation sys t em was on. Pilot comments for these I maneuvers are as follows.

- Initial Condition Maneuve__r Pilot Comments L evel Flight 4 ft / sec Most active in pitch axis with _ .

at V=I40 K t s RMS gust very little yaw disturbance.

I LAS On Trimmabili t y was poor, however, _.

the aircraft returned to trim after upsets. 1 Level Flight a t Gust response similar to 140 _ V=I80 Kts. knots. _ i LAS O n i : Level Flight Response similar to 140 knots, .....

at V = 225 K t s but t he ver t ical upsets became _ LAS On mere abrupt. Aircraft r et urn e d _ I t o t r i m a fte r gus_ upse t s. _ , _ a t V = 250 K t s Aircraf t did not diverge due to _*_ ' " II L evel Flight R es ponse similar to 22 5 knots. ,_ :._ LAS On gusts, and airspee d drifte d +1 5 kno ts .

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I 9.4.6 Climbs and Descen ts in the Cruise Mode the cruise mode were evaluated. The pilot started from sea I Aircraft response during steady state climb and descents in level and climbed to 10,000 feet holding airspeed at 150 knots.

I n itial Condition Maneuver Pilot C 9mments _! . evel Flight Climb to Aircraft cllmbed at 1700 ft / @ V=IS0 Knots 10,000 ft min, no pr<blem in holding LAS On at 150 K t s. airspeed• ! Level Flight Descend to Pulled off power and descended V = 150 Knots Sea Level at 1000 ft / mln. Pitch attitude L ; S On between 2 and 3 degrees. Air- . f- ' craft well behaved.

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_ ' _ l i 9.4.7 Engine Out Operation in Cruise Mode |' : Aircraft response to engine failure in the cruise mode was !_ evaluate d . A t 250 knots one engine was failed, wi t h the re- _: sponse no t ed. This was repeated a t 250 kts wi t h two engines _ , i _ f a iled. P ilot c o mm e nts ar e no t e d . !

',_., I n itial _ , Condition Maneuver Pilot Comments _ |_ Level Fligh t Fail one With one engine failed and ' * - _ @ 250 Kno t _ engine hands off, the aircraft is LAS On well behaved. The roll axis is unstable but aircraf t is [ , easily controllable. Tran- sients mild. The aircra f t " stabili z ed at 150 knots.

; L evel Fligh t Fail two Same as with single engine " * @ 250 Knots engines failure except pilot applied , ' " , LAS On power t o stabilize at 150 _ ' _[

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9.5 EVALUATION OF MAXIMUM NACELLE RATES IN TRANSITION Th is wa s evaluated in Section 9.3 . 6 . The maximum nacelle ?

{ t7 ra t e capability was z e d uced to 5 d eg / sec (from a nominal l! i0 d eg / sec). The pilot indicated that lower maximum rates might be desirable in the high speed en d of transition (100 to 140 kno t s) to minimi z e pit c h attitude changes at these c on- I _ d i t ions, wh i le higher na c elle rates are acceptable at low e r

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area to sp e eds.

required i Addi t ional work is in this evaluate tilt r ates .

i the d esirab i lity of establishing a schedule of maximum nacelle Evalua t i o n of t he hel ic op t er fligh t m o de was c o n d u c t ed for I several in c i d en c e angles . The pro c edure was to establish a nacelle inci d en c e angle an d then accelerate a t that incidence angle. The pilots comments are shown below. It shoul d be down pi t ch a tt i t udes at t he higher spee d s an d some nega t iv e I note d that t his type of operation results in extreme no s e spee d s t ability.

I I ni t i al Condition Maneuver Pilot Comments i > All S A S and Acc e ler at e to i0 d eg nose down @ 80 kn ots LAS On 160 knots at 15 deg nose down @ 100 knot s N a c el l e inci- c ons t an t a l t i - !

I I de nc e = 9 0 _ t u d e _ Na c elle Same Longitudinal stick trim aft _ - inci de n c e = 80 ° wi t h in c r e asing a ir s p eed b e- _ I t w ee n 50 k no ts a n d 8 0 k no ts _ ( re v e rs a l) . T h e st i c k m o v es _ f o rw a rd be t w ee n 80 a n d i00 - _ kn ots . _. .

_ N ace l le S a m e Longi t u d inal s tick re v ersa l i in c i d en c e 70 ° a t about 8 0 kno t s start s for- ,.

w a r d a t 9 0 kn ots , i_ _ i00 knots 5 o n o se down , s t i c k is m o ving fwd. 160 I kn o ts i0 ° n ose do w n n a c elles b ee p down au t o m a t ically.

in c iden c e = 45 ° Tri mm ed at a 4 ° n o se d o wn I N acelle Same Accelerated to 1 60 kno ts .

a l t i t u de at that speed .

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i0.0 CONCLUSIONS AND R ECOMMENDATIONS i. The ma t h model was success f ully programmed for the hybrid Simula t or.

l compu t er and c hecked out on the Small Motion Base Flight 2. Na c elle tilt indicator is a primary flight instrumen t I an d should be loc a ted near the basic Tee. Rate of climb, airspeed, torqueme_er and nacelle tilt indicator should be located on the same side of the basic Tee.

I _. Nacelle tilt switch arrangement needs further investiga- tion. Swit c h position (cyclic stick or thrust lever), . portional rate, fixed rate beep or two rate beep need I d ire c t i on of travel, force gradient and breakout, pro- evaluation.

I above about 160 knots is undesirable. P o sitive prevention of such operation should be considered. An automatic IR 4. Operation with nacelle in c idence above zero at speeds I up s t o p at 9 0 t o 9 5 ° i s d e sirable t o facili t a t e r ec onver- } sio n. Pos i t iv e pilot ac t i on sh o ul d be re quire d t o g o • bey o n d th i s s e tt ing.

t o o coarse to mag.

I 5 . The bra ke is be used for t rim in th e cruise range and should be locked.

I 6 . G o v ernor c ha ra cteris t i c s an d th rus t / c olle c tive pi tc h lever se nsi t ivi t y nee d t o be carefully tailore d t o f a c i- li t ate a c c ura t e alti t ude c on t rol in hover.

I 7. H o ver a n d tran si t i o n c h a ra ct eri st i c s are ac c e p ta bl e S AS off ing c o ntrol sensitivi t y in pitch and roll. The aircra f t I was successfully flown through transition SAS off.

and go od SAS on . The y can be f ur t her impr o ve d by i n creas - ! i___i 8 . Ove rt rav e l o f th e t h r us t / co l l e c t ive lever t o pr o vi d _ p osi- ,_ , _' I t i o n c on trol o f pi tc h fo r a c o lle ct ive f la re is highly d e sira b l e for u se in even t o f e ngine failure. _ , _i_ "_ H i gher descent r a t e s at s p eeds a r o u nd 60 k ts m ay re sult i n _ i 9 . De scent r at e s o f 1 500 f t / m i n a c h ie v a ble wi t h o ut pr o b le m. _ wing s tal l and rapid further increas e in descent rate if po wer is not a pp li e d rapidly. F urthe r inv es tigation is I n eed e d i n t h is ar e a .

i 0. L ong it u di n al c h ar a ct e ri s tics in c r ui s e s at i s f ac tor y . ._-_._ making it hard to tr i m in r o ll. It a l so s h owed very l arg e I The aircraf t mo d el led ha d a r o ll / s pi ral c o up l ing in cruis e dihedral e ff e c t. A cr u is e S A S sh o u ld be p rov ided i n the l at e ra l / dire c ti o na l ax e s.

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i i . T i lt rates a r o und 1 0 ° / sec a ppeare d qui t e a c cep t ab l e at i high nacel le a ngles, bu t l ow e r r at e s a re prefe r re d at the high speed end of t ra n si t ion.

I 12. The aircraf t can b _ flown in the helicopter mo_e up to abou t i00 k t s bu t this results in larje nose down a t ti t udes.

Early in i tiation of nacelle tilt provi d es _ more comfor t -

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I able and easily controlled transition.

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. ii.0 R EFERENCES Design o f R e sear c h Aircraf t ", NASA CR-I14438, Boeing Company, Ver t ol D_vision, Philadelphia, Pa., March 1972.

_ _ 1. "V / STOL Tilt Rotor Aircraf t Study, Vol• If, Preliminary 2 . US A F Stability and Control DAT C OM, Air ?o _ce Flight _ _ l Dyn a mics 1970)• Labora t ory, October 19 6 0, (Revis u d September | 3 . Re e d, T. J•, "User Report" Prop / R o tor Dynamxc Derivative | Program C41 J.N." , Boeing Documen t D210-I0116-I, Vertol D i vision, The Boeing Company, Philade l phia, Pa., June 1970.

4.

I A m os, A. K. ; Miao, W., "Program C-49: k_ u or Stability Derivatives" , Boeing Interoffice Memorandum, 8-745 3 -1-2483, Vertol Division, The Boeing Company, Philadelphia, Pa., I July 1971.

5 Davenpor t , F.J , "Analys : of Propeller and Rotor P Qr- I nfluen c e T e chnique" Boeing Document R-372 Boeing Company I formance in Static and Ax z al V light by aN Explicit Vor t ex .

V e rtol Division , P hiladelphia, Pa., February 19 6 5.

I 6 . T a rz anin , F. an d Th o mas, E . , "Aeroelastic R oto r An a ly si s", Bo_ing Documen t D8-0 6 14, Boeing Company, Vertol Division, P hiladelphia, Pa. , May 1967.

| I 7. Mag_ e , J.P. e t al , "T es t Progr a m II , Wind Tunnel T e s t of a Powered Tilt Rotor Performance Model , Volume VI, R esults Company, Vertol Division, Phila d elphia , Pa., Augus t 1970.

i and Analysis", Boeing Docume n t D213-100 0 0- 6 , Boeing { i 8 . Smi t h M.C., "University of Maryland Win d Tunnel Tes t 489 , i I Force, Moment and Downwash Measurements on a Rigid Ro t or

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an d S e misDan Wing" , (4 volum es ) , Boeing D oc umen t D 8- I 0 62- I , Th e Bo e ing Company, V e r t ol Division , Phila d elphia, Pa., I M arc h 1968 .

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Q Aircraft time histories of the Boeing Verto L Mode] 222 Tilt appendix. They were selected to illustrate seme of the per- " tinent pilot comments descrlbed in Section 9.0. These piloted I Rotor for selected piloted maneuvers are presented in this studies were all conuucted for an aircraft gr_s weight of 28% mean aerodynamic chord (most aft C .G. at this weight).

i 12,000 ib with the nacelles down (iN=0) c ent_'r of gravity at C onfiguration details and aerodynamic characteristics are aP described in Refer_ncL 1 and summarized in Section 3.0.

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Figure A.l(a) throug;l A.l(c) show a slow transition from hover to cruise (approximat_ly 80 seconds) and a reconversion to the ' ation concerning the l o ngitu d inal and lateral directional axes I hover mode (approximately 140 seconds). All pertinent inform- are shown. In addition, information on the SAS motions , angle , nacelle angle, and governor behavior a£e shown. Figures i power / collective lever travel, r o tor c oilec t oive pitch, flap A. (a) through A.2(c) show the same information for a rapid I mately 13 seconds) and a reconversion to the hover mode (approxi- mately 15 seconds). For slow transitions, t he aircraft is docile. As can be noted for the rapid transitions, holding is anticipat - d it wo_id be c ome easier with additional pilot ex- I altitude would require considerable pilot effort, although it i!_ transition from the hover mode to the cruise mode (app r oxi- perience with this vehicle. All stability a u %mentation (SAS) and load alleviation systems (LAS) were operating for these I runs.

Figures A.3 and A.4 show the Model 222's r sponse to ]ongitu- i u inal and lateral directional control pulses in hover with all SAS and LAS sys t ems functioning. Figures A.5 and A.6 show this informati o n with the SAS and LAS off. The pitch and yaw axes exhibit inherent d a mping provided by the h[ngeless r o tors. £he I _ yaw axis is u nstable. This is attributed to the lack cf inherent damping in this axis. All tilt rotor aircraft, however, regard- less of the type of zotor system would have this characteristi c . . _ , - It Should be noted t_at fuselage angle of a ttack and sideslip _ ¢ _ U are undefined for the hover mode in the mathematical model, i and therefore t hose t races sho,_id be disregarded for these --_ Figures A. I and A.8 show aircraft resp o nse to l o ngitudinal and lateral direction_l control pulses in the transitio,_ mode i at 80 knots. The nacelle angle is 70 degrees. Figures A.5 _a) and A.9(b) show helicopter mode maneuvers. With tLe nacelles a t 90 degrees, t he Model 222 is at a 15 dt_Trqe nose down a tt i- _"- I rude a t I00 knots, and t he longi tu dinal stic_ is approximately 3 inches forward. Figure A.10 shows a series of p a rtial power de£cent8 at v0 knots with the nacelle angle at 70 degrees.

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Descent rates in excess of approxi m ately 1500 ft / min could be I achieved before wing stall o c curred Figures A.7 through A.10 were obtained with the SAS and LAS systems o perating.

I Fi g ures A.II through A.14 show longitudinal and lateral direc- tional pulses at 140 knots in the cruise mode (iN=0) with the load alieviatLon system on and off. It should be noted that at this time, the Mode] 222 did not have a cruise SAS. There I is virtually n_ difference in response LA_ on o r off. The air c raft is heavily damped in the longitudinal axis. In the lateral d irectiona± _' , es there is roll / spiral c o upling which I makes it difficult to t_in and a large dihedr_i effect. Both of these characteristics were eliminated in B,,eing Vertol's January 1973 propos_l with a cruise mode stability augmentation for the cruise mode at 260 knots.

I system. Figures A.'5 through A.18 show the same information

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llD.Ia x )ill _tlt_l_ : Figur e A . l(a ). Slow Piloted Transition and Reconversion SAS and LAS On 6O m i

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F i g d reA . l(b). Slow Pi l otedTrans i tionand Reconve l sion SASand LAS On ' ] ,, / / -, Figure A . l(c) . S l ow Pi l oted Transition and Reconversion SAS and LAS On 4i

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l; - "+"° ' " ' c " °' +"" ...... , . :: lil _"+iii+ +++!+I_{_ _.,_'+ _ + + _'+'" .... + i++ I+,+_ - + I +' + +'_ _+_ + _+ + ++ ++ +' +_:+ +; + + + _ : :'+' + _, +., + + +++, +_ ' ++++++ ' "+' +"+"++ + , + . ,. , ,., ,, ,. , o ....... + + _+ : _ +++ v, , .. , :x:, , - + , ,,, o . . , .,_ _ , J _ + LI] _ L tILL + + '

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i + t ' + + i ".'"'- ' ° ' ' ttt.t.,+.i+ i',+u .., t _._t+ +++,, • . + , .-, ' + I Ii t .,l.,] t_r t -.o ,o,, - , - , _-- : i - _ : I : _. , +!_ : + i_ . + . +.++++I_ , ,+,_ . +: :

-m _t - -l - , , :+++:+, ; '+ ; :+:,'+;+.,+ , .+ t +. + ,.. + '+ t + t++ ' _ ; j+]tj.l_

' " , T T ' ....... T L " " " ' + + . + + ....... ' + . p ' ..... + + +.._1_1 , ' + ' +1 " +" _1 ' ' ' + . + : ''t 1+' ' + . + . +. + +. +.. _ . _ + 1+ " + 1 : I F i gureA . 2( l ), R =p id Pdo t_ dTrln$itmn = ndRlco n versmn SIS andLAS On

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Figure A.8. Piloted Time History - Responseto L ateral Stick a nd Rudder Pedal Pulsesin Transition , SAS end LAS On , iN = 7 0 ° V = 80 Knots ?1 'W es po ns eto Late r al Stick a n d Rudder PedalPul s e s i n Tran s ition, _L , " V = 80 K n ot s 7 1 db • ".. 4 w,,_ % .

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F_ 9 _ re A.1 3 . Pdo t e d T i m* . H_$1 o _- R etpo n l_t o Li tera l S lic k I_ I R _l d _ r Pedal P ul se lIt 1 40 Knots .

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