Ground-based and in-flight simulator studies of low-speed handling characteristics of two supersonic cruise transport concepts

NASA TP c.1

NASA Technical Paper, 1240

Ground-Based and In-Flight

Simulator Studies of Low-Speed

Handling Characteristics of Two

Supersonic Cruise Transport Concepts

William D. Grantham, Luat T. Nguyen, Perry L. Deal, M. J. Neubauer, Jr., Paul M. Smith, and Frederick D. Gregory JULY 1978 TECH LIBRARY KAFB, NM

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013452b

NASA Technical Paper 1240

Ground-Based and In-Flight

Simulator Studies of Low-Speed

Handling Characteristics of Two

Supersonic Cruise Transport Concepts

William D. Grantham, Luat T. Nguyen, Perry L. Deal, M. J. Neubauer, Jr., and Frederick D. Gregory Langley Research Ceizter Hamptoiz, Virgi m z i a Paul M. Smith Vought Corpora tioiz H a vzp to 12 Techn ica I Ceiz ter Hamnpton, Virginia National Aeronautics and Space Administration Scientific and Technical Information Office SUMMARY Fixed ground-based and i n - f l i g h t s i m u l a t o r s t u d i e s have been conducted t o determine t h e low-speed f l i g h t characteristics of t w o advanced s u p e r s o n i c c r u i s e t r a n s p o r t concepts, each having an arrow wing, a h o r i z o n t a l t a i l , and f o u r d r y ( n o n a f t e r b u r n i n g ) t u r b o j e t s w i t h v a r i a b l e geometry t u r b i n e s . The major d i f f e r - ences between t h e t w o s i m u l a t e d t r a n s p o r t c o n c e p t s were t h a t t h e f i r s t , or base- l i n e , c o n c e p t i n c o r p o r a t e d f o u r under-the-wing e n g i n e s , whereas t h e second con- cept u t i l i z e d powered l i f t and w a s c o n f i g u r e d to i n c o r p o r a t e t w o under-the-wing e n g i n e s on t h e outboard p o r t i o n of t h e wing with t h e t w o inboard e n g i n e s l o c a t e d upper s u r f a c e to induce a d d i t i o n a l c i r c u l a t i o n l i f t . The p r i m a r y on t h e wing p i l o t i n g t a s k was t h e approach and landing.

The r e s u l t s o f t h e s t u d i e s i n d i c a t e d t h a t t h e s t a t i c a l l y ( l o n g i t u d i n a l ) u n s t a b l e t r a n s p o r t c o n c e p t s had u n a c c e p t a b l e l o n g i t u d i n a l low-speed h a n d l i n g q u a l i t i e s w i t h no augmentation. I n o r d e r to a c h i e v e " s a t i s f a c t o r y " h a n d l i n g q u a l i t i e s , c o n s i d e r a b l e augmentation w a s r e q u i r e d . Although t h e SCAS devel- oped i n t h i s s t u d y to achieve s a t i s f a c t o r y handling q u a l i t i e s w a s complex, it is w i t h i n c u r r e n t technology. A hardened s t a b i l i t y augmentation system (HSAS) w a s r e q u i r e d t o a c h i e v e "acceptable" h a n d l i n g q u a l i t i e s should t h e normal oper- a t i o n a l s t a b i l i t y and c o n t r o l augmentation system (SCAS) f a i l .

The a v a i l a b l e r o l l - c o n t r o l power w a s found to be i n a d e q u a t e t o meet e x i s t - for t h e b a s e l i n e concept: b u t r o l l c o n t r o l ing crosswind-land ing r e q u i r e m e n t s w a s a c c e p t a b l e for t h e p o w e r e d - l i f t concept. Other advantages o f t h e powered- l i f t concept over t h e b a s e l i n e c o n c e p t were t h e a b i l i t y to perform segmented- d e c e l e r a t i n g approaches for community n o i s e abatement, and t h e a b i l i t y t o per- form l a n d i n g approaches a t c o n s i d e r a b l y reduced a n g l e s of a t t a c k ; t h e r e b y , t h e p o s s i b i l i t y of e l i m i n a t i n g t h e "drooped-nose" requirement f o r acceptable p i l o t f i e l d of view was i n c r e a s e d .

The r e s u l t s of t h i s s t u d y i n d i c a t e t h a t t h e maximum allowable peak v a l u e s of l a t e r a l a c c e l e r a t i o n ( a ) a t passenger and p i l o t s t a t i o n s d u r i n g c o o r d i n a t e d Y t u r n s may be u n s a t i s f a c t o r y based on proposed requirements for large s u p e r s o n i c t r a n s p o r t a i r p l a n e s . It w a s f u r t h e r concluded t h a t a d d i t i o n a l r e s e a r c h is r e q u i r e d to o b t a i n s a t i s f a c t o r y r i d e q u a l i t i e s w h i l e m a i n t a i n i n g s a t i s f a c t o r y h a n d l i n g q u a l i t i e s for e i t h e r of t h e s u p e r s o n i c c r u i s e t r a n s p o r t c o n c e p t s a t l o w speeds.

INTRODUCTION During t h e N a t i o n a l S u p e r s o n i c T r a n s p o r t (SST) Program of t h e e a r l y 1 9 6 0 ' ~ ~ v a r i o u s aerodynamic r e s e a r c h s t u d i e s conducted a t t h e NASA Langley Research Center to develop a n e f f i c i e n t s u p e r s o n i c c r u i s e t r a n s p o r t a i r p l a n e r e s u l t e d i n a h i g h l y swept arrow-wing c o n f i g u r a t i o n d e s i g n a t e d t h e SCAT-15F. The arrow-wing concept o f f e r e d c o n s i d e r a b l e promise for s u p e r i o r s u p e r s o n i c c r u i s e performance; u n f o r t u n a t e l y , such c o n f i g u r a t i o n s u s u a l l y d o n o t possess good low-speed han- d l i n g c h a r a c t e r i s t i c s . E a r l y wind-tunnel and p i l o t e d s i m u l a t i o n s t u d i e s (for

I1 I1 I 1 IIIII I I

example, see refs. 1 and 2) identified some of the low-speed handling problems of the SCAT-15F. Later, in 1968, the Boeing Company made an in-depth study (ref. 3) of a supersonic cruise transport concept which was based on the NASA arrow-wing configuration, but which had a lifting canard and a small horizontal tail (fig. 1). That particular configuration promised good take-off and land- ing performance. Although the canard improved the trimmed lift-drag ratio, it reduced longitudinal stability.

Since the early 1970 ' s , the Langley Research Center has been conducting extensive wind-tunnel studies to improve the low-speed handling characteristics of the arrow-wing configuration without a canard. (See fig. 2.) Some improve- ments were achieved by careful attention to wing planform, wing leading-edge design, and high-lift devices. Performance calculations have shown that with such modifications and with 2- to 3-percent negative static margin, the result- ing configuration should produce lift-drag ratios as good as those of a stable concept with a forebody canard. However, the landing attitude was such that nose droop probably would be required for an acceptable pilot field of view. In addition, stability and control analyses indicated that the concept might have some deficiencies in the high-lift landing-approach configuration. Preliminary conceptual studies have indicated that these problems could be minimized by application of powered-lift principles. Whereas the baseline concept had four

under-the-wing turbojets with variable geometry turbines (fig. 2) , the simulated

powered-lift concept incorporated two under-the-wing engines on the outboard por- tion of the wing with the other two engines located inboard and on the wing upper surface to induce significant circulation lift. (See fig. 3.) The major advan- tage of this powered-lift concept is that it can provide the capability to approach at lower angles of attack; thereby, the possibility of eliminating the "drooped-nose" requirement for acceptable pilot field of view is increased, and more roll-control power at the approach lift coefficient is achieved.

Results obtained from the aforementioned configuration refinements were sufficiently promising to justify conducting piloted simulator investigations of the approach and landing characteristics of two of the most recent supersonic cruise transport concepts - conventional and powered lift.

The primary objectives of these studies were to evaluate the low-speed han- dling characteristics of the two SCAR concepts and to obtain sufficient informa- tion to provide guidance for future low-speed research requirements. Other major objectives of these studies were (1) Evaluate the general handling qualities of the unaugmented airplanes in the approach configuration.

(2) Develop the stability augmentation and flight control systems required to achieve satisfactory handling qualities.

( 3 ) Determine the control power required to meet established handling qual- ities criteria.

( 4 ) Evaluate the effects of various atmospheric conditions, including heavy turbulence, steady winds., and wind shear on the ability of the pilot to make a satisfactory approach and landing.

(5) Determine the advantages and/or disadvantages of the powered-lift, arrow-wing concept as compared with the conventional (baseline) arrow-wing concept.

SYMBOLS AND DEFINITIONS Values are given in both the International System of Units (SI) and U.S.

Customary Units. The measurements and calculations were made in U.S. Customary Units. Dots over symbols denote differentiation with respect to time. All calculations are based on the aircraft body axes.

an normal acceleration, g units lateral acceleration, g units aY CL lift coefficient lift-curve slope per unit angle of attack per rad

cLcl

CZ rolling-moment coefficient rolling-moment coefficient due to sideslip per deg

clB

c, pitching-moment coefficient

Cn yawing-moment coefficient CT thrust coefficient CX longitudinal-force coefficient CY side-force coefficient CZ vertical-force coefficient - C mean aerodynamic chord, m (ft) f n longitudinal short-period undamped natural frequency, Hz g acceleration due to gravity, m/sec2 (ft/sec2) h altitude, m (ft) Ix,IY,Iz moment of inertia about X, Y, and z body axes, respectively, kg-m2 (slug-f t2) product of inertia, kg-m2 (slug-ft2) Ixz K1 rudder aerodynamic effectiveness gain K 2 rudder flexibility gain

IllllllllI II I I I I I I

flexibility gain for 6 , flexibility gain for 6 , f flexibility gain for 6 , f i flexibility gain for 6 , f 0 flexibility gain for 6 , lift-drag ratio lift per unit angle of attack per unit momentum ( $ 3 / m V ) C , , per see airplane mass, kg (slugs) steady-state normal acceleration change per unit change in angle of attack for an incremental horizontal-tail deflection at constant airspeed, gravity units/rad period of Dutch roll oscillation, sec period of longitudinal phugoid oscillation, sec period of longitudinal short-period oscillation, sec rolling, pitching, and yawing angular velocities, respectively, deg/sec o r rad/sec roll rates at first and second peaks, respectively, deg/sec or rad/sec dynamic pressure, Pa (lbf/ft2) reference wing area, m2 (ft2) S Laplace operator T thrust, N (lbf) time to damp to one-half amplitude, sec t1/2 time to double amplitude, sec t2 time to achieve 30° bank angle, sec t@=300 V airspeed, knots (ft/sec) airplane weight, N (lbf) W

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longitudinal distance from aircraft center of gravity to pilot X station, m (ft) I l a t e r a l d i s p l a c e m e n t from l o c a l i z e r c e n t e r l i n e , m ( f t ) v e r t i c a l d i s t a n c e from a i r c r a f t c e n t e r of g r a v i t y to p i l o t s t a t i o n , p o s i t i v e when p i l o t l o c a t e d below c e n t e r o f g r a v i t y , m ( f t ) a n g l e o f a t t a c k , deg a n g l e o f s i d e s l i p , deg f l i g h t - p a t h a n g l e , deg Y A i nc r emen t a i l e r o n d e f l e c t i o n s , p o s i t i v e f o r r i g h t r o l l command, deg 'a commanded a i l e r o n d e f l e c t i o n , deg 'a,c f laper on de f l e c t i o n , deg 'a, inboard f l a p e r o n d e f l e c t i o n , deg b a f i outboard f l a p e r o n d e f l e c t i o n , deg 'a f o column d e f l e c t i o n , deg ' C

t r a i l i n g - e d g e f l a p d e f l e c t i o n , deg

'f l a t e r a l c o n t r o l s u r f a c e d e f l e c t i o n s (combination o f a l l r o l l - c o n t r o l % a t surfaces u s e d ) , deg pedal d e f l e c t i o n , c m ( i n . )

rudder d e f l e c t i o n , deg

d e f l e c t i o n o f spoiler-slot and i n v e r t e d spoiler-slot deflectors, deg h o r i z o n t a l t a i l w i t h geared e l e v a t o r d e f l e c t i o n , deg wheel d e f l e c t i o n , deg g l i d e - s l o p e error, m ( f t ) Dutch r o l l mode damping r a t i o l o n g i t u d i n a l phugoid mode damping ratio l o n g i t u d i n a l s h o r t - p e r i o d mode damping r a t i o damping ratio of numerator quadratic @/61,t t r a n s f e r f u n c t i o n p i t c h a t t i t u d e , deg TR roll mode t i m e c o n s t a n t , sec @ a n g l e o f r o l l , deg $ heading a n g l e , deg phase a n g l e expressed as a l a g for a c o s i n e r e p r e s e n t a t i o n of Dutch $B r o l l o s c i l l a t i o n i n s i d e s l i p , deg wd undamped n a t u r a l frequency o f Dutch r o l l mode, rad/sec undamped n a t u r a l frequency of phugoid mode, rad/sec WPh l o n g i t u d i n a l short-period undamped n a t u r a l frequency, rad/sec w SP undamped n a t u r a l frequency appearing i n numerator q u a d r a t i c o f w$ t r a n s f e r f u n c t i o n , rad/sec S u b s c r i p t s : a v average C command ed cg c e n t e r of g r a v i t y ge ground e f f e c t Ig l a n d i n g g e a r l a t l a t e r a l 0 a l l s u r f a c e s z e r o d e g r e e s osc osc i 1 l a t o r y r m s root mean s q u a r e PS p i l o t s t a t i o n SS s t e a d y s t a t e max maximum Abbreviations: AD1 a t t i t u d e director i n d i c a t o r ARI a i l e r o n - r u d d e r i n t e r c o n n e c t HSAS hardened s t a b i l i t y augmentation system I F R i n s t r u m e n t f l i g h t r u l e s I ILS instrument landing system PR pilot rating RAH roll attitude hold mode on

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RAH roll attitude hold mode off SAS stability augmentation system SCAR supersonic cruise aircraft research SCAS stability and control augmentation system SJT subsonic jet transport SST supersonic transport STOL short take-off and landing TIFS total in-flight simulator VFR visual flight rules WL wings leveler mode on - WL wings leveler mode off DESCRIPTION OF SIMULATED AIRPLANES* Both of the simulated airplanes (conventional and powered lift) were, in general, resized versions of the configuration described in reference 4. Three- view sketches of the two concepts are presented in figures 2 and 3; mass and dimensional characteristics, and the control-surface deflection and deflection rate limits for these concepts are presented in table I; and the aerodynamic data used in the study are presented in tables I1 and 111. The "conventional" supersonic cruise transport concept will hereafter be referred to as the base- line concept.

Baseline Concept The static aerodynamic data used for the baseline configuration were esti- mated on the basis of the various low-speed wind-tunnel test results (e.g., refs. 5 and 6) and corrected for configuration differences. The control sur- faces used for low-speed lateral control consisted of outboard ailerons, out- board spoiler-slot and inverted spoiler-slot deflectors, and inboard flaperons.

The lateral control system was designed in such a manner that all lateral con- *The work-up and analyses of the aerodynamic and geometric data packages utilized for this SCAR simulation program were performed under contract number NAS1-13500 by Paul M . Smith of Vought Corporation.

trol surfaces were driven by the commanded aileron deflection, and each surface was deflected so that each reached its limit simultaneously. The rigid aileron control data were estimated on the basis of unpublished wind-tunnel tests, and the flaperon, spoiler-slot, and inverted spoiler-slot deflector data were taken from reference 3 and modified to account for the size and location of the sub- ject airplane's control surfaces. A 40-percent-chord, full-span rudder was used for low-speed directional control. The rigid rudder effectiveness data were estimated by using the method presented in reference 4 . The reduction of lateral control effectiveness due to wing flexibility was estimated from the data presented in reference 3; and the reduction of directional control effec- tiveness due to fuselage side bending was based on unpublished data. (See fig. 4 for an indication of flexibility effects.) The methods presented in reference 7 were used to estimate the aerodynamic effects of ground proximity, and the data are shown in figure 5 .

The dynamic aerodynamic derivatives were estimated by using a combination of the forced oscillation test data of reference 1 and the estimation techniques of reference 8 .

An example of the engine response characteristics used for both the base- line and powered-lift concepts is shown in figure 6 .

Powered-Lift Concept The powered-lift airplane simulated had the same overall dimensions as the baseline airplane, except that the horizontal- and vertical-tail sizes were increased. The static longitudinal and lateral-directional aerodynamic charac- ter istics for the powered-lift airplane were developed by using the wind-tunnel test data of references 6 and 9, respectively, with the appropriate corrections applied to include the effects of the differences in horizontal- and vertical- tail volume coefficients. To obtain the same horizontal-tail trim download as for the baseline concept required that the center of gravity of the powered-lift concept be located at 0.72c. Configuration rebalance and main landing gear location analyses determined that the most aft center-of-gravity position was at 0.66c.

For the powered-lift airplane, the control surfaces used for low-speed lateral control consisted of outboard ailerons, outboard flaperons, and inboard flaperons. The lateral control system was designed in the same manner as the baseline concept; all surfaces were driven by the commanded aileron deflection.

Since no lateral control surface effectiveness data were measured with the power on for the powered-lift wind-tunnel model (ref. 61, the measured unpowered data of reference 9 were modified to approximate the effects of upper-surface blow- ing. This was achieved by conservatively assuming that the measured values of rolling- and yawing-moment coefficients as functions of thrust coefficient and outboard engine nozzle deflections (thrust vectoring) would represent the incre- ments in lateral control due to upper-surface blowing effects. For convenience and ease of implementation, these measured increments were added to the unpow- ered aerodynamic effectiveness of the outboard flaperon data. (Ekperience with powered-lift STOL wind-tunnel models has indicated that substantially more roll- control power can be achieved by utilizing deflection of trailing-edge surfaces of wings incorporating upper-surface blowing than that from use of thrust vec- toring alone.) An all-movable vertical tail was used €or low-speed directional control. (The aerodynamic effectiveness data were developed from the data of ref. 9 . ) The flexibility effects on lateral control were estimated from the data in reference 3; whereas, the flexibility effects on directional control were based on unpublished data. (See fig. 4.)

The aerodynamic effects of ground proximity (fig. 5) used for the baseline airplane were also used for the powered-lift airplane.

The dynamic stability derivatives used for the powered-lift airplane were estimated by using the data for the baseline airplane and corrections were made to include the effects of (1) the increase in horizontal- and vertical-tail size, and ( 2 ) the blowing of the two upper-surface engines.

DESCRIPTION OF SIMULATION EQUIPMENT Evaluations of the low-speed landing-approach handling characteristics were made at Langley Research Center by using a fixed-base ground simulator with a visual landing scene. After the ground-based study, a brief in-flight simulation program was conducted by using Calspan's Total In-Flight Simulator (TIFS) airplane in order to provide (1) points of reference for interpretation of the ground simulator results, ( 2 ) data for control system design trade-offs, and ( 3 ) data on the effects of motion cues not available in the fixed-base simulation.

Fixed-Base Simulator The fixed-base simulator had a transport-type cockpit which was equipped with conventional flight and engine-thrust controls and with a flight-instrument display representative of those found in current transport airplanes. (See fig. 7.) Instruments indicating angle of attack, sideslip, and flap angle were also provided. A conventional cross-pointer-type flight director instrument was used, and the command bars (cross pointers) were driven by the main com- puter program.

Real-time digital simulation techniques were used wherein a digital com- puter was programed with equations of motion for six degrees of freedom.

A visual display of an airport scene (fig. 8 ) was used in order to provide visual cues for the flare and landing. The display consisted of a closed- circuit television presentation, viewed through a collimating lens in the pilot's windshield, of the simulated approach to a 3505-m (11 500-ft) runway.

(See fig. 9 . ) Each flight was terminated at touchdown; the roll-out was not simulated.

In-Flight Simulator The TIFS is a fly-by-wire C-131 airplane with controllers for all six degrees of freedom and a separate evaluation cockpit forward and below the normal C-131 cockpit. (See fig.. 10.) When flown from the evaluation cockpit, the pilot control commands are the inputs to a model computer which determine the aircraft motion commands to be reproduced. These are combined with the TIFS motion sensor signals in another portion of the onboard computer to pro- vide TIFS controller commands. The simulated airplane motions are produced with maximum time lags of 50 to 150 msec in the frequency range of interest.

The evaluation cockpit instruments were mostly conventional and were posi- tioned as shown in figure 11. In addition to the conventional instruments dis- plays of sideslip angle and angle of attack were provided. Airspeed error was displayed as a tape motion on the left side of the ADI. Aircraft position rel- ative to the ILS glide slope was displayed (in ft) as a vertical bug motion on the left side of the ADI. A flight director computer producing the same func- tions as the computer used in the ground-based simulator was mechanized in the TIFS computer. This instrument was used in lieu of the conventional flight director on board the TIFS airplane in order to insure that the flight director was compatible with the simulated supersonic cruise transport dynamics.

Cardboard masking was used on the TIFS evaluation cockpit windshield to simulate the view expected from the cockpit of the supersonic cruise transport.

TESTS AND PROCEDURES Two research pilots participated in the simulation program and each used standard flight-test procedures in the evaluation of the handling qualities.

The primary piloting task was the approach and landing.

The tests consisted of IFR and simulated VFR landing approaches with cross- winds, turbulence, localizer offsets, glide slope offsets, and engine failure as added complicating factors. The ILS approach was initiated with the airplane in the power-approach condition (power for level flight), at an altitude below the glide slope, and on a 45O intercept course to the localizer. (See fig. 12.)

The pilot's task was to capture the localizer and glide slope and to maintain them as closely as possible while under simulated IFR conditions. At an alti- tude of approximately 91 m (300 ft), the pilot converted to VFR conditions and attempted to land the airplane visually (with limited reference to the flight instruments).

The results of these studies using the aforementioned evaluation procedures are in the form of time-history records of airplane motions and pilot comments regarding the low-speed handling qualities of the two supersonic cruise trans- port concepts and the effects of various stability and control augmentation sys- tems on these characteristics. The more significant results are reviewed in the following sections.

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RESULTS AND DISCUSSION The r e s u l t s o f t h e s e s t u d i e s are d i s c u s s e d i n terms o f t h e p r e v i o u s l y s t a t e d o b j e c t i v e s , and t h e p i l o t r a t i n g s l i s t e d f o r t h e v a r i o u s c o n d i t i o n s e v a l u a t e d are an average o f t h e r a t i n g s from a l l pilots who f l e w t h a t p a r t i c u - l a r c o n d i t i o n . (See t a b l e I V f o r t h e p i l o t r a t i n g system.) A l s o , t h e r e s u l t s d i s c u s s e d p e r t a i n to t h e d a t a o b t a i n e d on t h e b a s e l i n e s u p e r s o n i c cruise t r a n s - port concept u t i l i z i n g t h e fixed-base ground s i m u l a t o r u n l e s s s p e c i f i c a l l y noted.

N o S t a b i l i t y Augmentation a n e g a t i v e s t a t i c margin of approximately 4 per- The b a s e l i n e c o n c e p t had c e n t to improve t h e approach L/D, and t h e unaugmented h a n d l i n g q u a l i t i e s were r a t e d as u n a c c e p t a b l e (PR = 7 ) by t h e e v a l u a t i o n pilots. A s can be s e e n from t a b l e V, t h e t i m e to double a m p l i t u d e ( t 2 ) o f t h e l o n g i t u d i n a l aperiodic mode is 4 . 8 sec, which might be expected to be u n a c c e p t a b l e s i n c e t h e landing- approach minimum-safe (PR = 6.5) criterion o f r e f e r e n c e 1 0 s t a t e d t h a t a t 2 < 6 sec would be unacceptable. (See f i g . 13.) A comparison o f t h e p i t c h rate response of t h e unaugmented a i r p l a n e t o t h e desired response is p r e s e n t e d i n f i g u r e 1 4 and shows t h a t t h e response to a column step i n p u t a p p e a r s as an a c c e l e r a t i o n command i n s t e a d o f t h e d e s i r e d rate command. The p i t c h c o n t r o l power o f t h i s b a s e l i n e concept was r a t e d as acceptable i n s o f a r as t h e l o n g i t u - d i n a l c o n t r o l power r e q u i r e m e n t s for t h e approach and l a n d i n g t a s k s are con- c e r n e d , i n agreement w i t h t h e c o n t r o l power requirements c r i t e r i o n o f r e f e r - ence 11 as shown i n f i g u r e 15. Recent unpublished s t u d i e s have i n d i c a t e d t h a t an a c c e p t a b l e p i t c h a c c e l e r a t i o n criterion a t t h e minimum demonstrated a i r - speed is s a i d t o be a c c e p t a b l e i f @ 5 -0.05 r a d / s e c 2 and s a t i s f a c t o r y i f @ By u s i n g t h i s c r i t e r i o n t h e p i t c h c o n t r o l power w a s -0.08 rad/sec2.

determined to be a c c e p t a b l e , b u t n o t s a t i s f a c t o r y . (See t a b l e V I . ) A p i l o t r a t i n g o f 7 w a s a s s i g n e d to t h e unaugmented l a t e r a l - d i r e c t i o n a l h a n d l i n g q u a l i t i e s of t h e b a s e l i n e c o n f i g u r a t i o n . The major o b j e c t i o n s were (1) u n a c c e p t a b l e l a r g e a d v e r s e s i d e s l i p e x c u r s i o n s i n t u r n s ; ( 2 ) e a s i l y e x c i t e d , l i g h t l y damped Dutch r o l l mode; ( 3 ) poor r o l l and heading c o n t r o l ; and ( 4 ) s l u g - g i s h r o l l response w i t h l o w r o l l damping. The primary f a c t o r t h a t c o n t r i b u t e d to t h e poor p i l o t r a t i n g f o r t h e l a t e r a l - d i r e c t i o n a l c h a r a c t e r istics w a s t h e l a r g e a d v e r s e s i d e s l i p e x c u r s i o n s experienced d u r i n g r o l l i n g maneuvers. T h i s is i n d i c a t e d i n f i g u r e 16, and compared w i t h t h e d e s i r e d response c h a r a c t e r i s t i c f o r a l a t e r a l c o n t r o l step i n p u t . For a step i n p u t it is d e s i r a b l e to have (1) a r a p i d roll-rate r e s p o n s e t h a t r e a c h e s a r e a s o n a b l y s t e a d y - s t a t e v a l u e w i t h a minimum o f o s c i l l a t i o n ; ( 2 ) e s s e n t i a l l y z e r o s i d e s l i p produced by t h e roll- c o n t r o l i n p u t ; and ( 3 ) an immediate response i n heading. However, it is e v i d e n t from f i g u r e 16 t h a t f o r a l a t e r a l c o n t r o l step i n p u t for t h i s unaugmented con- f i g u r a t i o n , a l a r g e amount of a d v e r s e s i d e s l i p is e x p e r i e n c e d t h a t washes o u t t h e r o l l rate ( 4 ) i n a s h o r t p e r i o d o f t i m e and also c a u s e s an u n d e s i r a b l e l a g i n t h e i n i t i a t i o n o f t u r n rate (0). T h i s l a r g e a d v e r s e s i d e s l i p c h a r a c t e r istic, in combination with the low roll damping, required constant attention and con- poor siderable effort on the part of the pilot and still resulted in very lateral-directional control.

It must be noted that although the longitudinal and lateral-directional handling qualities of this nnaugmented supersonic cruise transport airplane were assigned a pilot rating of 7 when evaluated individually, the combination of poor characteristics resulted in an overall pilot rating of 10 for the air- plane. Therefore, it was apparent that considerable stability and control aug- mentation will be required to achieve satisfactory handling qualities for the landing-approach piloting task.

Normal Operational Stability and Control Augmentation System (SCAS) Based on the results obtained for the unaugmented configuration, the objective for the design of the SCAS was that the system should provide satis- factory handling qualities (PR 6 3.5) at all flight conditions evaluated during the study. A block diagram of the SCAS design obtained is shown in figure 17.

Longitudinally, a high-gain pitch rate command/attitude hold system was chosen because (1) stabilization of the unstable mode could be achieved with the pitch attitude feedback, (2) the system provided good short-period charac- teristics and rapid response to pilot inputs, and (3) the attitude-hold feature minimized disturbances due to turbulence or variations in thrust.

Laterally, a roll rate command/attitude hold system was employed to pro- vide a rapid roll mode and quick uniform response to pilot inputs; the attitude- hold feature resulted in a desirable neutrally stable spiral mode while counter- acting disturbances due to turbulence. In addition, a wings-leveler feature was provided to the pilot (to be used at his option) which automatically leveled the (4 = Oo) whenever the bank angle was less than 2O and the wheel was cen- wings tered. This feature relieved the pilot of the task of hunting for zero bank angle and was particularly useful when rolling out of a turn to a desired heading .

Directionally, roll-rate and roll-attitude feedbacks were used to provide turn coordination and improved Dutch roll characteristics. A roll control to rudder interconnect was also included to reduce adverse sideslip during turn entry and therefore minimize Dutch roll excitation during roll maneuvers.

An autothrottle that maintained the selected airspeed throughout the approach and landing was also used as part of the normal operational aug- mentation. Since the simulated engine dynamics (for example, see fig. 6) produced very rapid thrust response, the autothrottle generally maintained the desired airspeed within +3 knots and considerably reduced the pilot workload on the landing approach. Although this airplane is flown well up the "backside" of the thrust required curve at the approach speed of 153 knots p T f l / a V = -O.O023/knot) where normally the pilot would primarily use pitch attitude for airspeed control and thrust for glide-path control, the simulated quick engine-thrust response allowed the use of thrust (manually or automati- cally) for airspeed control and thus enabled the pilot to use pitch attitude for glide-path control - which is a very natural, simple technique.

The longitudinal SCAS (fig. 17) provided pitch rate proportional to column deflection, and produced the desired characteristics of rapid, well-damped responses to pilot inputs as well as inherent attitude stability. Figure 18 shows the improvement in pitch rate response provided by the SCAS, and it can be seen from table V that the time to double amplitude (t2) of the longitudinal aperiodic mode increased from 4.8 sec with no augmentation to infinity with the SCAS configuration. With this augmentation system operative, the average pilot rating for the longitudinal handling qualities on the ILS approach was improved from PR = 7 to PR = 2 .

Also shown in figure 17 is a block diagr.am of the lateral-directional SCAS.

Laterally, a rate command system provided roll rate proportional to wheel posi- tion, and the directional system consisted of several turn coordination features.

Table V shows that the Dutch roll characteristics were improved considerably; (W@/Wd) was increased from 0.565 to 1.004 (which indicates that the Dutch roll oscillation should be much less easily excited for roll-control inputs), and the damping parameter (?&od) was increased from 0.064 rad/sec to 0.197 rad/sec.

The improvement in the roll response and damping are indicated by the reduction of TR from 1.689 to 0.27 sec. (See table V.)

Figure 19 shows the improvement in the roll-rate response provided by the SCAS. By elimination of the large adverse sideslip, the roll-rate reversal was eliminated, and the heading response was immediate (no lag). The lateral SCAS also provided a desirable roll-attitude-hold feature which proved to be very beneficial, particularly during landing approaches made in simulated heavy tur- bulence. With this augmentation system operative, the average pilot rating for the lateral-directional handling qualities on the ILS approach was improved from PR = 7 to PR = 2 .

With the SCAS operative, the overall pilot rating of the simulated baseline SCAR concept for the landing-approach task was 2.

Hardened Stability Augmentation System (HSAS) As discussed previously, the baseline SCAR concept had unacceptable low- speed handling qualities with no augmentation. A hardened stability augmenta- tion system (HSAS) was therefore required to achieve acceptable handling quali- ties should the normal operational augmentation (SCAS) fail. (The term "hardened" SAS implies sufficient redundancy to negate loss of the system.)

The HSAS design objective was to provide improved handling qualities so that acceptable pilot ratings (PR 6 6.5) could be obtained for the approach and so that the system could be kept as simple as possible to max- landing task and imize reliability and ease of implementation. A block diagram of the HSAS design is shown in figure 20. Longitudinally, a filtered pitch rate feedback signal acting through a relatively high gain was used to reduce the instability of the unstable mode and to enhance the short-period characteristics. Laterally, a simple roll damper provided a smaller roll mode time constant and increased Dutch roll damping. Directionally, roll-rate feedback was used to provide: (1) improved turn-entry coordination; (2) reduced Dutch roll coupling during roll maneuvers (increased W+/Cod); and (3) further enhancement of the Dutch roll damping. Note that only two angular rate signals (pitch rate and roll rate) were required for the HSAS implementation so that sensor reliability problems and mechanization complexity would be minimized. The autothrottle was also considered to be part of the HSAS.

The average pilot rating assigned to the longitudinal handling qualities when the HSAS was operative was 4 . The primary objection was the less-than-

desired pitch damping. Table V shows that the short-period damping ratio (cSp)

for this configuration is 0.693, which would normally indicate adequate damping; however, the slowly divergent aperiodic mode (t2 = 44 sec) superimposed on the short-period response caused the motions to appear to the pilot as being inade- quately damped. It should be noted that reference 10 also indicated acceptable pilot ratings (PR 6 6.5) when t2 was greater than 6 sec. (See fig. 13.) Fig- ure 21 compares the pitch response to a column step for the unaugmented airplane with SCAS operative and with HSAS operative. The reason a higher gain was not implemented for the pitch rate damper, in order to satisfy the pilot's objection of low pitch damping, was that more damping would make the pitch axis unaccept- ably sluggish. It is evident from figure 21 that the HSAS configuration is already very sluggish in pitch, compared with the SCAS configuration.

The average pilot rating assigned to the lateral-directional handling qual- ities with the HSAS operative was 4 . The primary objections were sluggish roll response, Dutch roll excitation during turns, less than desired roll damping, and a lack of steady-state turn coordination. Figure 22 shows a comparison of the roll response to a lateral control step input for the HSAS, SCAS, and unaug- mented configurations.

Effects of Center-of-Gravity Location As previously stated, the airplane was configured to be slightly statically unstable at a center-of-gravity position of 0.56; t o minimize the required trim download of the tail. The resulting negative static margin was approximately 4 percent. Handling qualities evaluations for the landing-approach task at this llbasic" center-of-gravity position (0.56;) resulted in a pilot rating of 2 with the SCAS operative and a pilot rating of 4 with the HSAS operative. To evaluate the effects of center-of-gravity location on the low-speed handling qualities, the airplane was flown with increasing levels of negative static margin. The technique used to determine the most tolerable aft center-of-gravity location was to determine the center-of-gravity position at which the pilots evaluated the low-speed handling qualities as being "satisfactory" with the SCAS opera- tive, PR 6 3.5, and also as being "acceptable" with the HSAS operative (PR 5 6.5). It was determined that for a center-of-gravity location of 0 . 6 6 ; , the pilots rated the landing-approach task as being marginally satisfactory (PR = 3.5) with the SCAS operative and marginally acceptable (PR = 6.5) with the HSAS operative. Therefore, from handling qualities considerations, the aft center-of-gravity limit was said to be 0.66; (approximately 14-percent negative static margin).

To i l l u s t r a t e t h e e f f e c t o f c e n t e r - o f - g r a v i t y p o s i t i o n on t h e low-speed a i r p l a n e performance, t h e d a t a i n f i g u r e 23 are p r e s e n t e d . N o t e t h a t as t h e c e n t e r - o f - g r a v i t y l o c a t i o n is moved rearward, t h e approach l i f t - d r a g r a t i o i n c r e a s e s o n l y s l i g h t l y ; however, t h e t r i m a n g l e o f attack d e c r e a s e s s i g n i f i - c a n t l y . Although it is n o t p r e s e n t e d i n f i g u r e 23, it should be mentioned t h a t f o r f l a p s e t t i n g s less t h a n t h e landing-approach f l a p s e t t i n g (6f = 40°), t h e i n c r e a s e i n a i r p l a n e performance ( l i f t d r a g ) as t h e c e n t e r o f g r a v i t y is moved rearward is much more pronounced.

Crosswind Landings Both s t e a d y c r o s s w i n d s (up to 20 k n o t s ) and crosswinds w i t h h o r i z o n t a l s h e a r (8 k n o t s per 30 m ) were s i m u l a t e d . The p i l o t i n g t e c h n i q u e used f o r mak- i n g t h e approach and l a n d i n g c o n s i s t e d o f an i n i t i a l crabbed approach, and a t a nominal a l t i t u d e ( u s u a l l y a b o u t 15 m (50 f t ) ) , t r a n s i t i o n i n g to a wing-down si d e s l i p .

The requirements o f r e f e r e n c e 12 s t a t e t h a t t r a n s p o r t a i r p l a n e s w i t h o u t crosswind-landing gear should be c a p a b l e o f l a n d i n g i n 90° crosswinds u p t o 30 k n o t s , and t h a t t h e l a t e r a l c o n t r o l used s h a l l n o t exceed 75 p e r c e n t o f t h e c o n t r o l power a v a i l a b l e . F i g u r e 24 i n d i c a t e s t h e amount o f s t e a d y - s t a t e side- s l i p , bank a n g l e , rudder d e f l e c t i o n , and l a t e r a l c o n t r o l d e f l e c t i o n r e q u i r e d f o r s i d e s l i p p i n g crosswind approaches a t an a i r s p e e d o f 153 k n o t s ( t h e nominal approach s p e e d ) . It can be seen t h a t 75 p e r c e n t o f t h e a v a i l a b l e l a t e r a l con- t r o l was r e q u i r e d for a crosswind component of approximately 20 knots. It is, t h e r e f o r e , obvious t h a t t h i s b a s e l i n e s u p e r s o n i c c r u i s e t r a n s p o r t a i r p l a n e c o u l d n o t be landed w i t h an adequate l a t e r a l c o n t r o l margin i n 90° c r o s s w i n d s h i g h e r than approximately 20 k n o t s . A l s o , from a p i l o t i n g s t a n d p o i n t , t h e l a t e r a l - d i r e c t i o n a l c o n t r o l c o o r d i n a t i o n required for t h e t r a n s i t i o n from a crabbed-approach c o n d i t i o n to a s i d e s l i p p i n g , wing-down c o n d i t i o n becomes i n c r e a s i n g l y d i f f i c u l t as t h e 90° crosswind i n c r e a s e s above approximately 15 knots. It is, t h e r e f o r e , concluded from these ground-based, f i x e d - c o c k p i t s i m u l a t o r r e s u l t s t h a t t h e s u b j e c t s u p e r s o n i c cruise t r a n s p o r t a i r p l a n e concept should be equipped w i t h crosswind gear and/or provided w i t h a d d i t i o n a l r o l l - c o n t r o l p o w e r .

It should be mentioned t h a t a l t h o u g h t h e accuracy of t h e c o n t r o l coordina- t i o n w a s t h e prime f a c t o r t h a t affected t h e p i l o t ' s a b i l i t y to make "precise" l a n d i n g s i n high crosswinds, d e f i c i e n c i e s of t h e v i s u a l p r e s e n t a t i o n ( l a c k of p e r i p h e r a l v i s i o n and adequate h e i g h t c u e s ) and p o s s i b l y t h e l a c k of cockpit motion also a f f e c t e d t h e p i l o t ' s a b i l i t y to make s a t i s f a c t o r y l a n d i n g s i n l a r g e crosswinds.

E f f e c t s o f Turbulence on Landing Approach F l i g h t i n rough air w a s e v a l u a t e d by u s i n g a t u r b u l e n c e model based on t h e Dryden spectral form. The root-mean-square v a l u e o f t h e l o n g i t u d i n a l , l a t e r a l , and v e r t i c a l g u s t - v e l o c i t y components w a s v a r i e d from 0.61 m/sec (2 f t / s e c ) to 2.7 m/sec ( 9 f t / s e c ) . These v a l u e s were d e s c r i b e d by t h e p i l o t s as being repre- s e n t a t i v e o f l i g h t and heavy t u r b u l e n c e , r e s p e c t i v e l y . The p i l o t s commented t h a t t h e p i l o t r a t i n g for t h e approach t a s k on t h e b a s e l i n e s u p e r s o n i c cruise transport concept was degraded by one r a t i n g when t h e l a n d i n g approach w a s made i n t h e s i m u l a t e d heavy t u r b u l e n c e because of t h e i n c r e a s e d workload r e q u i r e d t o m a i n t a i n ILS t r a c k i n g .

F i g u r e 25 p r e s e n t s plots .of t h e root-mean-square v a l u e s o f t h e v e r t i c a l and l a t e r a l a c c e l e r a t i o n s a t t h e a i r p l a n e c e n t e r o f g r a v i t y experienced d u r i n g ILS approaches made i n v a r i o u s l e v e l s of s i m u l a t e d t u r b u l e n c e for both t h e b a s e l i n e s u p e r s o n i c c r u i s e t r a n s p o r t s i m u l a t e d and a typical s u b s o n i c j e t t r a n s p o r t . The root-mean-square values are compared w i t h t h e r i d e q u a l i t y c r i t e r i o n of r e f e r - ence 13. A s can be seen, t h e normal and l a t e r a l a c c e l e r a t i o n root-mean-square values are l o w e r f o r t h e s u p e r s o n i c c r u i s e t r a n s p o r t t h a n for t h e s u b s o n i c j e t t r a n s p o r t . T h e r e f o r e , t h e response of t h e s i m u l a t e d s u p e r s o n i c cruise a i r p l a n e to a t m o s p h e r i c t u r b u l e n c e would n o t be expected to be any worse than t h e r e s p o n s e of present-day s u b s o n i c transport a i r p l a n e s , e f f e c t s o f a i r f r a m e f l e x i b i l i t y being n e g l e c t e d .

Comparison of B a s e l i n e Concept With Powered-Lift Concept P o w e r e d - l i f t c o n c e p t s , such as t h o s e u t i l i z i n g upper-surface e n g i n e blow- i n g , o f f e r s e v e r a l aerodynamic improvements over t h e b a s e l i n e concept. P r a c t i - c a l l y a l l t h e s e improvements are achieved from t h e i n c r e a s e d c i r c u l a t i o n l i f t t h a t can be o b t a i n e d by blowing t h e j e t e f f l u x over l a r g e r t r a i l i n g - e d g e f l a p s .

(The t r a i l i n g - e d g e f l a p s i z e , and hence l i f t g e n e r a t i o n , is l i m i t e d on t h e base- l i n e concept because of t h e l o c a t i o n of t h e aft-wing-mounted t u r b o j e t e n g i n e s . ) Aerodynamics.- F i g u r e 26 i n d i c a t e s t h e i n c r e a s e i n l i f t c o e f f i c i e n t t h a t w a s ag6ieved w i t h t h e s i m u l a t e d p o w e r e d - l i f t concept. N o t e t h a t a CL of 0.66, which c o r r e s p o n d s to a trimmed approach speed o f 153 knots, is achieved a t ct = Oo on t h e p o w e r e d - l i f t concept compared w i t h ct = 8O on t h e b a s e l i n e con- cept. T h i s allows t h e p i l o t to f l y t h e l a n d i n g approach w i t h t h e p o w e r e d - l i f t concept a t a s i g n i f i c a n t l y reduced p i t c h a t t i t u d e which minimizes t h e l e n g t h o f t h e main landing-gear s t r u t s , and it also o f f e r s t h e p o t e n t i a l o f e l i m i n a t i n g t h e drooped-nose requirement f o r an a c c e p t a b l e p i l o t f i e l d o f view. F i g u r e 27 p r e s e n t s a view of t h e runway as seen by t h e p i l o t prior to touchdown f o r both t h e b a s e l i n e and p o w e r e d - l i f t c o n c e p t s w i t h t h e nose o f t h e a i r c r a f t drooped f o r maximum p i l o t v i s i b i l i t y , as w e l l as w i t h t h e nose i n t h e llup'l p o s i t i o n .

N o t e t h a t w i t h t h e nose drooped, t h e p i l o t f i e l d o f view i n d i c a t e d f o r both c o n c e p t s appears to be s u f f i c i e n t to m a k e a s i m u l a t e d approach and l a n d i n g , somewhat b e t t e r v i s i b i l i t y being i n d i c a t e d f o r t h e p o w e r e d - l i f t concept because o f t h e d i f f e r e n c e i n t h e approach a t t i t u d e . The nose-up s c e n e s are p r e s e n t e d to f u r t h e r i n d i c a t e t h e advantage of t h e lower approach a t t i t u d e . N o t e t h a t t h e p i l o t cannot see t h e runway w i t h t h e nose i n t h e up p o s i t i o n w h i l e f l y i n g t h e b a s e l i n e concept.

The r e d u c t i o n i n approach a n g l e of a t t a c k also r e d u c e s t h e d i h e d r a l e f f e c t ( f i g . 28) which, i n t u r n , improves t h e i n h e r e n t l a t e r a l - d i r e c t i o n a l h a n d l i n g q u a l i t i e s . The more e f f e c t i v e wing t r a i l i n g - e d g e f l a p s on t h e p o w e r e d - l i f t concept also o f f e r a means of i n c r e a s i n g t h e a v a i l a b l e r o l l - c o n t r o l power ( f i g . 2 9 ) . A s shown i n f i g u r e 30, t h e combination of reduced d i h e d r a l e f f e c t I -1 I and increased roll-control power result in acceptable crosswind-landing capa- bility, which was not the case for the baseline concept.

A segmented landing approach (for community noise abatement considera- tions) could also be readily flown on the powered-lift concept and still main- tain a relatively low pitch attitude. As indicated in figure 31, the powered- lift concept could be flown on an approach angle of -5O at an airspeed of 170 knots for one segment: then, at some designated altitude (nominally 152 m (500 ft)) transition to an approach angle of -2.7O and an airspeed of 153 knots could be made. In addition, if the drooped-nose consideration is ignored, the transition could be made to an approach angle of -2.70 and an airspeed of

136 knots - which is the nominal approach speed of present-day subsonic jet

transports.

It should be mentioned that although the aforementioned advantages of the powered-lift concept are considerable for terminal area operations, this concept does have some disadvantages during cruise. Potentially, the disadvantages of the powered-lift concept during cruise are (1) upper-surface wing-nacelle inter- ference drag, (2) increased wave drag due to the increase in slope of the for- ward cross-sectional area distribution curve, ( 3 ) airframe strength degradation due to thermal and acoustic effects, and (4) more complex engine inlet flow field.

Handling qualities. - The handling qualities of the unaugmented powered-lift

concept were, in general, the same (unacceptable) as those previously discussed

for the baseline concept. However , both satisfactory and acceptable handling

qualities were achieved by utilizing the same augmentation systems as discussed for the baseline concept. See tables V and VI for a comparison of the dynamic stability and control response characteristics of the two concepts.

Engine failure.- Lateral-directional control with a critical engine (out- board) failed has always been a prime consideration in the rudder design for multiengine airplanes. Control of asymmetries due to engine failure can be easily analyzed from static conditions by calculating the steady-state sideslip angle, bank angle, and control deflections for a straight flight path over the ground. The transient responses immediately following an engine failure, how- ever, present problems involving pilot reaction time, the manner in which con- trols are applied, and, of course, the altitude and configuration of the air- plane at the time of the failure. During the subject program, attempts were made to simulate the wave-off capabilities as well as continued approaches and landings after an outboard engine failure on both supersonic cruise transport concepts (baseline concept and powered-lift concept).

The manner in which an engine was failed during this simulation study was that which would be considered the most severe: that is, the engine failed instantaneously (a step form of thrust loss). Also, the configuration flown for both concepts incorporated what was considered to be the best stability and control augmentation system (SCAS) and autothrottle. The requirement used for evaluating the wave-off capability of the baseline concept after engine failure was determined based on the proposed airworthiness standards for supersonic

transports (ref. 14) - "With the approach flap setting, the aircraft shall be

capable of a 2.7 percent gradient (1.5O) climb, in rectilinear flight, with one ' engine inoperative at an airspeed no greater than the determined operational and performance speed." (The baseline concept has an operational approach speed of 153 knots which is equivalent to 1.22 times the minimum demonstrated speed.) The requirements used when documenting the wave-off capability of the powered-lift concept were based on NASA powered-lift flight experience (for

example, ref. 15) - "In the event of failure of one engine on approach, it

should be possible to arrest the descent and maintain level flight without change in flap setting or airspeed. It should also be possible after arrest- ing the descent to establish a sustained climb angle of 2O (3.5% gradient) by retraction of the flaps and without change in airspeed."

With the relatively high thrust-weight ratio available on both of these concepts (four-engine approach T/W = 0.5), the wave-off capability of both simulated concepts, from performance considerations, was no problem and met the aforementioned requirements with ease. However, typical of most multiengine aircraft, the increase in pilot workload caused by the necessity to retrim after an engine failure degraded the pilot ratings for the wave-off task to 3 for the baseline concept and 4 for the powered-lift concept. (A pilot rating of 2 was assigned to both concepts with no engine failure.) It should be mentioned that the amount of rudder required to trim the baseline and powered-lift concepts after an outboard engine failure was approximately 4O and 15O, respectively.

Attempts were made to simulate a continued approach and landing following the loss of an outboard engine on both the baseline and powered-lift concepts.

Typical approaches, for which the number four engine was failed during the approach, are presented in figure 32. The most interesting points indicated are the excursions from the localizer and glide slope immediately following the engine failure. As can be seen from figure 32 (a), the maximum lateral displace- ment from the localizer beam was aproximately 10 m (33 ft), and the maximum ver- tical displacement from the glide-slope beam was approximately 5 m (16 ft) for the baseline concept compared with 17 m (56 ft) and more than 12 m (39 ft) ,

respectively, for the powered-lift concept. (See figure 32 (b) . )

The pilots commented that the loss of a critical engine during an ILS approach on either concept posed no problems (insofar as tracking localizer and glide slope) but that the requirement of using rudder for trimming sideslip was bothersome, particularly for the powered-lift concept. For the continued approach task after an engine failure, the pilots assigned ratings of 2.5 and 4 to the baseline concept and powered-lift concept, respectively. In addition, the pilots commented that they would probably choose to perform a wave-off on the powered-lift concept if the engine failure occurred below an altitude of approximately 91 m (300 ft), whereas they would probably continue the approach and landing on the baseline concept regardless of the altitude at which the engine failed.

Comparison of Fixed-Base and In-Flight Results As stated previously, upon completion of the fixed-base ground simulator tests, a brief in-flight simulation program was conducted in order to provide (1) points of reference for interpretation of the ground simulator results; ( 2 ) data for control system design trade-offs; and (3) data of effects of motion cues not available in the fixed-base simulation. Only the baseline supersonic cruise transport concept was flown during the in-flight simulation program - the powered-lift concept was not simulated.

In general, the handling qualities assessments determined on the fixed-base simulator were substantiated during the in-flight simulator tests. Although the in-flight tests were more realistic (for example, the motions were realistic and the scene out of the window was the real world), these factors did not change the pilots' opinions of the handling characteristics of the simulated airplane.

However, it was determined during the in-flight tests that the SCAS produced unacceptable ride qualities (lateral accelerations) at the pilot station (cock- pit). As indicated in figure 33, the SCAS developed for the lateral-directional axes during the fixed-base tests provides a quick, uniform roll-rate (p) response to a lateral control input and at the same time provides good turn coordination (small produced). In addition, the lateral acceleration indicated for the center of gravity of the airplane ((ay)cg) is acceptable from ride qualities

considerations. However , the lateral acceleration indicated for the pilot's

station ((ay)ps) and particularly the rate of buildup of ((ay)ps) following a lateral control input was said to be unacceptable (uncomfortable) by the evalu- ation pilots during the in-flight tests. This unacceptable lateral acceleration at the pilot's station was produced primarily by the unusually long distance between the center-of-gravity location and the pilot's station on this super- sonic cruise transport airplane. Figure 34 indicates a comparison of pilot location, relative to the center of gravity, between the subject supersonic cruise transport airplane and the Boeing 747 subsonic jet transport airplane.

The relationship between the lateral acceleration at the pilot station and that at the center of gravity may be approximated as follows:

- (ayIcg +

32.17 As can be seen from figure 33, the term is the predominant factor. (The

distance - from the simulated center-of-gravity location to the simulated pilot

station x was 44.2 m (145.1 ft).) It may be erroneously concluded that any airplane that has a very long distance between the center of gravity and the cockpit will have unacceptable ride qualities, whereas compromises can be made between handling qualities and ride qualities and achieve satisfactory (or at least acceptable) characteristics for both. For example, the hardened stabil- ity augmentation system (HSAS) developed during the fixed-base tests produced acceptable (but not satisfactory) handling qualities during both fixed-base and in-flight tests and also had acceptable ride qualities ((ay)ps) during the in- flight tests. As shown in figure 35, the initial roll-rate response for a lat- eral control input is good for the HSAS configuration but the adverse sideslip continues to build up and "washes out" some of the roll rate. As stated pre- viously, the lateral-directional handling qualities of this airplane with the HSAS operative were assigned a pilot rating of 4 (acceptable). Since the HSAS did not produce good turn coordination (6 = O O ) , the yaw acceleration produced by a lateral control input was not appreciable (when compared with the SCAS response) and therefore the lateral acceleration at the cockpit was not as large as that produced with the SCAS operative.

Modified SCAS The results from the in-flight simulation tests implied that "acceptable" lateral acceleration characteristics could be achieved if the lateral- directional handling qualities were compromised. Therefore, an attempt was made to modify the lateral-directional part of the SCAS in such a manner as to maintain "satisfactory" handling qualities (PR 6 3.5) and at the same time attain "acceptable" (ay)ps characteristics. (The dynamic stability and response characteristics of the simulated airplanes with the modified SCAS operative are presented in tables V and VI.) These goals were accomplished by slowing the initial roll-rate response by applying a first-order lag to the roll-rate command signal, by reducing the wheel roll-rate command sensitivity, and by substantially reducing the ARI (aileron-to-rudder interconnect) gain.

The modifications to the initial lateral-directional SCAS are indicated in the block diagram presented in figure 36. Time histories of the motions obtained for a roll-control step input with the modified SCAS are presented in figure 37 and compared with the motions obtained with the initial SCAS. It can be seen that the roll-rate response for the modified SCAS is not as fast as that for the initial SCAS, but that good turn coordination is maintained (small .6) and an appreciable improvement in the lateral acceleration characteristics is achieved. The pilots assigned a rating of 3 to the lateral-directional handling qualities when the modified lateral-directional SCAS was used (compared with a PR of 2 for the initial SCAS) and said the lateral accelerations experienced for roll-control inputs were "acceptable," but not satisfactory.

Dynamic Stability Requirements and Criteria For several years the aircraft industry has been aware that many of the existing stability requirements of aircraft are outdated because of the expan- sion of flight envelopes and the increases in airplane size. Although research is presently being conducted in an effort to remedy this situation, to date essentially no clearly defined stability requirements and criteria have been established for aircraft similar to those for the supersonic cruise transport.

Therefore, in an effort to aid in the future establishment of new stability requirements, the low-speed handling qualities parameters of the supersonic cruise transport concepts are compared with some existing handling qualities criteria.

Two of the most widely used longitudinal handling qualities criteria are presented in figure 38. Figure 38(a) shows the short-period frequency require- ments of reference 12 and, as can be seen, the results predicted by the crite- rion agree reasonably well with the results obtained during the present simula- tion studies. Figure 38(b) shows the Shomber-Gertsen longitudinal handling qualities criterion of reference 16. This criterion relates the ability of the pilot to change flight path with normal acceleration to the factor &. By using this parameter and by recognizing that the pilot's mode of control is not constant for all flight regimes, a criterion for satisfactory short-period characteristics was developed that correlates well with current airplane experience and reasonably well with the results obtained during the present low-speed supersonic cruise transport simulation program. Figure 39 presents the longitudinal short-period criterion, for transport aircraft, of refer- ence 17. I n g e n e r a l , t h e r e s u l t s of t h e p r e s e n t s t u d y are s a i d t o be i n good agreement w i t h t h i s c r i t e r i o n , p a r t i c u l a r l y €or t h e unaugmented and HSAS con- f i g u r a t i o n s . As noted i n r e f e r e n c e 17, t h e l i m i t l i n e for t h e " a c c e p t a b l e unaugmented a r e a " of t h i s c r i t e r i o n is s u b j e c t to f u r t h e r research. It is b e l i e v e d , from t h e r e s u l t s o b t a i n e d f o r t h e SCAS c o n f i g u r a t i o n s d u r i n g t h e p r e s e n t study, t h a t t h e upper l i m i t l i n e €or t h e "acceptable augmented a r e a " could a l s o be extended to higher v a l u e s of t h e short-period damping r a t i o Tsp.

The low-speed p i t c h r a t e response c r i t e r i o n shown i n f i g u r e 4 0 , and r e p o r t e d i n r e f e r e n c e 18, w a s based on t h e Shomber-Gertsen c r i t e r i o n of r e f e r - ence 16. A s can be seen, t h e r e is e x c e l l e n t agreement between t h e r e s u l t s o b t a i n e d during t h e p r e s e n t s t u d y and t h i s low-speed p i t c h response c r i t e r i o n when t h e normal o p e r a t i o n a l augmentation (SCAS) w a s o p e r a t i v e . The c o n s t r a i n t s imposed upon t h e u s e of t h e r e f e r e n c e 18 c r i t e r i o n , however, negate its use f o r any of t h e o t h e r c o n f i g u r a t i o n s e v a l u a t e d du,ring t h e p r e s e n t study. For t h e m o s t part, t h e p i t c h divergence c r i t e r i o n of r e f e r e n c e 1 0 , with a time-to- double p i t c h a t t i t u d e of 6 sec or g r e a t e r for t h e most u n s t a b l e root, was con- s i d e r e d when t h e HSAS and unaugmented c o n f i g u r a t i o n s were e v a l u a t e d , and t h e s u b j e c t s i m u l a t i o n r e s u l t s agreed very w e l l w i t h t h e c r i t e r i o n . (For example, see f i g . 13.)

The r o l l - a c c e l e r a t i o n and r o l l - r a t e c a p a b i l i t y c r i t e r i a €or t r a n s p o r t a i r - c r a f t are p r e s e n t e d i n f i g u r e s 4 1 and 42, r e s p e c t i v e l y . (These c r i t e r i a were reported i n r e f s . 1 2 and 17, r e s p e c t i v e l y . ) The v a r i o u s c o n f i g u r a t i o n s evalu- a t e d d u r i n g t h e p r e s e n t s i m u l a t i o n s t u d y a r e i n d i c a t e d i n these p l o t s and, i n g e n e r a l , w o u l d n o t be considered t o be i n agreement with r e s u l t s p r e d i c t e d by these c r i t e r i a - p a r t i c u l a r l y f o r t h e r o l l - a c c e l e r a t i o n c a p a b i l i t y c r i t e r i o n p r e s e n t e d i n f i g u r e 41. For example, t h e r o l l - c o n t r o l power a v a i l a b l e f o r t h e powered-lift concept was determined t o be very s a t i s f a c t o r y for landing i n 90° crosswinds g r e a t e r than 30 knots, which was t h e most demanding p i l o t i n g t a s k e v a l u a t e d during t h e p r e s e n t low-speed s i m u l a t i o n s t u d i e s . (See f i g . 30.)

The bank a n g l e o s c i l l a t i o n l i m i t a t i o n s c r i t e r i o n of r e f e r e n c e 12 is pre- s e n t e d i n f i g u r e 43 and r e l a t e s t h e phase a n g l e of t h e D u t c h r o l l component of s i d e s l i p ($B) t o t h e measure of t h e r a t i o of t h e o s c i l l a t o r y component of bank a n g l e to t h e average component of bank a n g l e The v a r i o u s configura- t i o n s e v a l u a t e d during t h e p r e s e n t s i m u l a t i o n s t u d y a r e i n d i c a t e d i n t h i s p l o t , and it can be seen t h a t t h e simulated c h a r a c t e r i s t i c s a g r e e , reasonably w e l l , with t h e aforementioned c r i t e r i o n - p a r t i c u l a r l y , t h e f u l l y augmented (SCAS) and unaugmented c o n f i g u r a t i o n s .

I n g e n e r a l , it is concluded t h a t t h e r e s u l t s of t h e p r e s e n t s i m u l a t i o n s t u d y a g r e e with t h e e s t a b l i s h e d handling q u a l i t i e s c r i t e r i a used for compari- son i n t h i s paper, t h e major e x c e p t i o n being t h e r o l l - a c c e l e r a t i o n c a p a b i l i t y c r i t e r i o n of r e f e r e n c e 12.

R i d e Q u a l i t y C r i t e r i a The ride q u a l i t y c r i t e r i o n of r e f e r e n c e 13 r e l a t e s t h e root-mean-square v a l u e s of an and ay t o t h e root-mean-square v a l u e s of t h e g u s t i n t e n s i t y ( l e v e l of t u r b u l e n c e ) . As d i s c u s s e d p r e v i o u s l y and shown i n f i g u r e 25, t h e r e s p o n s e of t h e s i m u l a t e d s u p e r s o n i c cruise b a s e l i n e concept to atmospheric t u r b u l e n c e compared f a v o r a b l y w i t h t h e aforementioned c r i t e r i o n when t h e air- frame f l e x i b i l i t y e f f e c t s were n e g l e c t e d - p a r t i c u l a r l y , t h e c r i t e r i o n f o r t h e (ay)rms, which w a s e q u a l to or less than 0.055g f o r a c c e p t a b l e passenger r i d e comfort. It should be noted, however, t h a t t h e root-mean-square v a l u e s of an and p r e s e n t e d i n f i g u r e 25 f o r t h e SCAR and SJT a i r p l a n e s were t h e v a l u e s ay measured a t t h e c e n t e r of g r a v i t y o f t h e a i r c r a f t , whereas t h e c r i t e r i o n of r e f e r e n c e 1 3 p e r t a i n s to t h e v a l u e s a t any passenger l o c a t i o n . A l s o as d i s - cussed p r e v i o u s l y , t h e s p e c i f i c p e a k v a l u e s (as opposed to t h e root-mean-square v a l u e s ) o f experienced a t l o c a t i o n s f a r removed from t h e c e n t e r o f g r a v i t y ay of t h e a i r c r a f t were.found t o be u n s a t i s f a c t o r y d u r i n g t h e i n - f l i g h t s i m u l a t i o n part o f t h i s study.

A c r i t e r i o n f o r t h e maximum a l l o w a b l e a t any passenger s t a t i o n as ay w e l l as t h e p i l o t s t a t i o n (cockpit) h a s been proposed by t h e Boeing Company as a r e s u l t o f t h e i r a n a l y s e s d u r i n g t h e N a t i o n a l SST Program and is r e p o r t e d i n r e f e r e n c e 11. T h i s c r i t e r i o n states i n part - "Lateral a c c e l e r a t i o n a t t h e p i l o t s t a t i o n s h a l l n o t exceed a l e v e l of k0.075g p e a k , and t h e c r i t i c a l pas- senger s t a t i o n s h a l l n o t exceed +O.O5g peak. These l e v e l s s h a l l be m e t f o r a l l normal maneuvers i n c l u d i n g 30 d e g r e e bank and c a p t u r e using an average r o l l rate o f 5O/sec i n cruise and 10°/sec a t landing. I f u n p i l o t e d t i m e s t u d i e s are con- d u c t e d , t h e wheel i n p u t should be a 0.5-second ramp o f magnitude s u f f i c i e n t t o produce t h e s p e c i f i e d average r o l l rates. I' I n o r d e r to compare t h e p e a k v a l u e s o f t h a t would be experienced on ay t h e b a s e l i n e s u p e r s o n i c cruise t r a n s p o r t concept w i t h t h e aforementioned cri- t e r i o n , f i g u r e 44 w a s prepared. F i g u r e 4 4 p r e s e n t s t h e peak v a l u e s o f ay as a - f u n c t i o n o f t h e l o n g i t u d i n a l displacement from t h e a i r c r a f t c e n t e r of g r a v i t y ( x ) . (The v e r t i c a l displacement from t h e c e n t e r of g r a v i t y w a s maintained c o n s t a n t as t h a t r e p r e s e n t i n g t h e p i l o t s t a t i o n . ) As can be seen, t h e proposed c r i t e r i o n f o r cannot - be s a t i s f i e d even when no t u r n c o o r d i n a t i o n is pro- ay vided for any v a l u e of x c o n s i d e r e d . That is, c o n s i d e r i n g t h a t an approxi- mation o f (ay)ps is w i t h no yaw a c c e l e r a t i o n ( i ) f o r t u r n c o o r d i n a t i o n , and even n e g l e c t i n g t h e c o n t r i b u t i o n , t h e 6 ( 2 ) / 3 2 . 1 7 term ( r o l l a c c e l e r a t i o n t i m e s t h e (a,) : g v e r t i c a 1 displacement of t h e p i l o t from t h e a i r c r a f t c e n t e r o f g r a v i t y ) is O.O8g, which is l a r g e r than t h e a c c e p t a b l e l e v e l of t h e aforementioned crite- r i o n . I n a d d i t i o n to no t u r n c o o r d i n a t i o n (which is u n r e a l i s t i c ) , t h e s e v a l u e s o f (ayIps were o b t a i n e d f o r a r i g i d a i r f r a m e . It is b e l i e v e d t h a t i f a i r f r a m e were included, t h e peak v a l u e s of ay would be even l a r g e r .

f l e x i b i l i t y e f f e c t s It is also b e l i e v e d t h a t some o f t h e l a r g e r s u b s o n i c t r a n s p o r t s o f today c o u l d n o t meet t h i s proposed l a t e r a l a c c e l e r a t i o n c r i t e r i o n , simply because o f t h e geometry o f t h e problem. Therefore, it is concluded t h a t t h e requirements o f t h i s proposed c r i t e r i o n m u s t be r e l a x e d or t h e r o l l maneuvers o f a l l v e r y l a r g e a i r p l a n e s m u s t be c o n s t r a i n e d i n o r d e r to have a c c e p t a b l e low-speed r i d e q u a l i t i e s .

CONCLUDING REMARKS Fixed-base s i m u l a t o r and i n - f l i g h t s i m u l a t o r s t u d i e s have been conducted to determine t h e low-speed f l i g h t c h a r a c t e r istics of t w o advanced s u p e r s o n i c cruise t r a n s p o r t c o n c e p t s (a c o n v e n t i o n a l concept and a p o w e r e d - l i f t c o n c e p t ) , each having an arrow wing, a h o r i z o n t a l t a i l , and f o u r d r y t u r b o j e t s w i t h v a r i - a b l e geometry t u r b i n e s . The primary p i l o t i n g t a s k was t h e approach and landing.

T h i s paper has attempted to summarize t h e r e s u l t s o f t h e s e s t u d i e s which support t h e f o l l a w i n g major c o n c l u s i o n s .

The s t a t i c a l l y u n s t a b l e ( l o n g i t u d i n a l l y ) s u p e r s o n i c c r u i s e t r a n s p o r t con- cepts s i m u l a t e d had u n a c c e p t a b l e (pilot r a t i n g o f 1 0 ) low-speed h a n d l i n g q u a l i - ties w i t h no augmentation.

The l o n g i t u d i n a l normal o p e r a t i o n a l s t a b i l i t y and c o n t r o l augmentation sys- t e m , c o n s i s t i n g o f a high-gain p i t c h rate command/attitude h o l d system and an a u t o t h r o t t l e , e s s e n t i a l l y e l i m i n a t e d t h e l o n g i t u d i n a l c o n t r o l problems. The l a t e r a l - d i r e c t i o n a l SCAS, c o n s i s t i n g o f a r o l l rate command/attitude h o l d system and o f roll-rate, r o l l - a n g l e , and r o l l - c o n t r o l s u r f a c e d e f l e c t i o n feedback s i g - n a l s to t h e rudder, made t h e l a t e r a l - d i r e c t i o n a l h a n d l i n g c h a r a c t e r i s t i c s s a t i s - f a c t o r y . With t h e s e augmentation systems o p e r a t i v e , t h e average p i l o t r a t i n g f o r t h e i n s t r u m e n t approach t a s k was 2 for both t h e b a s e l i n e and p o w e r e d - l i f t concepts.

The hardened s t a b i l i t y augmentation system (HSAS), designed to p r o v i d e a c c e p t a b l e h a n d l i n g q u a l i t i e s w i t h maximum s i m p l i c i t y ( f o r r e l i a b i l i t y and ease o f implementation) c o n s i s t e d o f a f i l t e r e d p i t c h rate feedback s i g n a l to t h e l o n g i t u d i n a l c o n t r o l s u r f a c e f o r a d d i t i o n a l p i t c h damping, and a roll-rate feed- back s i g n a l to t h e r o l l - c o n t r o l s u r f a c e s , as w e l l as to t h e r u d d e r , for addi- t i o n a l r o l l damping and improved t u r n - e n t r y c o o r d i n a t i o n . With t h i s HSAS oper- a t i v e , t h e average p i l o t r a t i n g f o r t h e i n s t r u m e n t approach t a s k w a s 4 for both t h e b a s e l i n e and powered-lif t concepts.

I n an e f f o r t to e v a l u a t e t h e e f f e c t s of c e n t e r - o f - g r a v i t y l o c a t i o n on t h e low-speed h a n d l i n g q u a l i t i e s , t h e b a s e l i n e s u p e r s o n i c cruise t r a n s p o r t concept w a s flown w i t h i n c r e a s i n g l e v e l s of n e g a t i v e s t a t i c margin. I t w a s determined t h a t w i t h t h e SCAS or HSAS o p e r a t i v e , t h e landing-approach t a s k c o u l d be per- formed w i t h a n e g a t i v e s t a t i c margin as high as 14 p e r c e n t . (The p i l o t r a t i n g s to t h e SCAS and HSAS c o n f i g u r a t i o n s f o r t h e landing-approach t a s k were a s s i g n e d 3.5 and 6.5, r e s p e c t i v e l y . ) The a v a i l a b l e r o l l - c o n t r o l p o w e r r e q u i r e d t o meet t h e e x i s t i n g crosswind- l a n d i n g r e q u i r e m e n t s w a s found to be i n a d e q u a t e f o r t h e b a s e l i n e c o n c e p t b u t adequate f o r t h e p o w e r e d - l i f t concept.

The response o f t h e s u p e r s o n i c cruise t r a n s p o r t c o n c e p t s to a t m o s p h e r i c t u r b u l e n c e would n o t be expected to be any worse than t h e r e s p o n s e o f p r e s e n t - day s u b s o n i c t r a n s p o r t a i r p l a n e s , f l e x i b i l i t y d i f f e r e n c e s being n e g l e c t e d .

However, t h e p i l o t s commented t h a t t h e r a t i n g f o r t h e landing-approach t a s k on t h e t r a n s p o r t c o n c e p t s w a s degraded by one r a t i n g when t h e l a n d i n g approach w a s made i n t h e s i m u l a t e d heavy t u r b u l e n c e s i n c e t h e g l i d e - s l o p e t r a c k i n g t a s k r e q u i r e d h i g h e r p i l o t workload.

The m o s t a p p a r e n t advantages o f t h e p o w e r e d - l i f t c o n c e p t (over t h e base- l i n e c o n c e p t ) were t h e a b i l i t y to perform segmented-decelerating approaches (for community n o i s e a b a t e m e n t ) , and t h e a b i l i t y to perform l a n d i n g approaches a t reduced a n g l e s of attack; t h e r e b y , t h e p o s s i b i l i t y o f e l i m i n a t i n g t h e "drooped- nose" requirement for a c c e p t a b l e p i l o t f i e l d o f view w a s i n c r e a s e d .

The wave-off c a p a b i l i t i e s , as w e l l as c o n t i n u e d approaches and l a n d i n g s , were s i m u l a t e d a f t e r t h e f a i l u r e o f an o u t b o a r d e n g i n e on b o t h t h e b a s e l i n e and p o w e r e d - l i f t concepts. With t h e r e l a t i v e l y h i g h t h r u s t - w e i g h t ratio a v a i l a b l e on both o f t h e s e c o n c e p t s (four-engine approach T/W = 0.5), t h e wave-off cap- a b i l i t y was no problem, from performance c o n s i d e r a t i o n s , and m e t t h e e s t a b l i s h e d r e q u i r e m e n t s w i t h ease. The p i l o t s commented t h a t t h e loss o f a c r i t i c a l e n g i n e d u r i n g an i n s t r u m e n t approach on e i t h e r c o n c e p t posed no problems i n s o f a r as t r a c k i n g l o c a l i z e r and g l i d e slope. The p i l o t s f u r t h e r commented t h a t t h e y would probably choose to perform a wave-off on t h e p o w e r e d - l i f t concept i f t h e e n g i n e f a i l u r e o c c u r r e d belaw an a l t i t u d e o f a p p r o x i m a t e l y 9 1 m (300 f t ) , whereas t h e y would p r o b a b l y c o n t i n u e t h e approach and l a n d i n g on t h e b a s e l i n e c o n c e p t r e g a r d l e s s o f t h e a l t i t u d e a t which t h e e n g i n e f a i l e d .

I n g e n e r a l , it w a s concluded t h a t t h e r e s u l t s of t h e s i m u l a t i o n s t h d y a g r e e w i t h t h e e s t a b l i s h e d h a n d l i n g qualities c r i t e r i a used f o r comparison i n t h i s paper. However, it is b e l i e v e d from t h e r e s u l t s of t h i s s t u d y t h a t t h e proposed requirements f o r t h e maximum a l l o w a b l e p e a k v a l u e s of l a t e r a l a c c e l e r a - t i o n (a,) a t any passenger s t a t i o n , as w e l l as t h e p i l o t s t a t i o n (cockpit) dur- ing c o o r d i n a t e d t u r n s must be r e l a x e d or t h e r o l l maneuvers of a l l v e r y l a r g e t r a n s p o r t a i r p l a n e s m u s t be c o n s t r a i n e d i n o r d e r to have s a t i s f a c t o r y r i d e qualities.

It is f u r t h e r concluded t h a t a d d i t i o n a l low-speed r e s e a r c h is r e q u i r e d to a c h i e v e s a t i s f a c t o r y r i d e q u a l i t i e s and a t t h e same t i m e m a i n t a i n s a t i s f a c t o r y h a n d l i n g q u a l i t i e s on e i t h e r o f t h e s u b j e c t s u p e r s o n i c cruise t r a n s p o r t concepts.

Langley Research Center N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n Hampton, VA 23665 A p r i l 28, 1978 REFEl?ENCES 1 . Freeman, Delma C., Jr.: Law Subsonic Flight and Force Investigation of a Supersonic Transport Model With a Highly Swept Arrow Wing. NASA TN D-3887, 1967.

2 . Grantham, William D.; and Deal, Perry L . : A Piloted Fixed-Base Simulator Study of Law-Speed Flight Characteristics of an Arrow-Wing Supersonic Transport Design. NASA TN D-4277, 1967.

3. Mach 2 . 7 Fixed Wing SST Model 969-336C (SCAT-15F) . Doc. No. D6A11666-1

(Contract FA-SS-67-3), Boeing Co., July 1969.

4. Baber, Hal T., Jr.; and Swanson, E . E . : Advanced Supersonic Technology Concept AST-100 Characteristics Developed in a Baseline-Update Study.

NASA "M X-72815, 1976.

5 . Lockwood, Vernard E . : Effect of Trailing-Edge Flap Deflection on the Lat- eral and Longitudinal-Stability Characteristics of a Supersonic Transport Model Having a Highly-Swept Arrow Wing. NASA TM X-71936, 1974.

6. Coe, Paul L., Jr.; McLemore, H. Clyde; and Shivers, James P.: Effects of Upper-Surface Blowing and Thrust Vectoring on Low-Speed Aerodynamic Characteristics of a Large-Scale Supersonic Transport Model. NASA TM X-72792, 1975.

7. LTV Hampton Technical Center: Advanced Supersonic Technology Concept Study - Reference Characteristics. NASA (33-132374, 1973.

8. USAF Stability and Control Datcom. Contracts AF 33(616)-6460 and F33615-75-(2-3067, McDonnell Douglas Corp., Oct. 1960. (Revised Apr. 1976.)

9. Coe, Paul L., Jr.; Smith, Paul M.; and Parlett, Lysle P.: Low-Speed Wind Tunnel Investigation of an Advanced Supersonic Cruise Arrow-Wing Config- uration. NASA TM 74043, 1977.

10. Rising, J. J . : A Study of the Effects of Relaxed Static Stability on Sta- bility Augmentation System Reliability Requirements. LR 26833, Lockheed California Co., Dec. 20, 1974.

11. Stability and Control, Flight Control, Hydraulic Systems and Related Structures Criteria. Doc. No. D6-6800-5, Boeing Co., Jan. 1970.

12. Chalk, C . R.; Neal, T . P.; Harris, T . M.; Pritchard, F. E.; and Woodcock, R. J. : Background Information and User Guide for MIL-F-8785B(ASG),

"Military Specification - Flying Qualities of Piloted Airplanes."

AFFDL-TR-69-72, U.S. Air Force, Aug. 1969. (Available from DDC as AD 860 856. ) 13. Low-Wing-Loading STOL Transport Ride Smoothing Feasibility Study.

D3-8514-2 (Contract NAS1-10410) , Boeing Company, Feb. 8, 1971.

(Available as NASA CR-111819.)

14. Tentative Airworthiness Standards for Supersonic Transports. Flight Standards Service, FAA, Nov. 1, 1965, Revision 7, Jan. 1, 1971.

15. Parlett, Lysle P.; Fink, Marvin P.; and Freeman, Delma C., Jr. (With appendix B by Marion 0 . McKinney and Joseph L. Johnson, Jr.): Wind- Tunnel Investigation of a Large Jet Transport Model Equipped With an External-Flow Jet Flap. NASA TN D-4928, 1968.

16. Shomber, H. A.; and Gertsen, W. M . : Longitudinal Handling Qualities Criteria: An Evaluation. AIAA Paper No. 65-780, Nov. 1965. .

17. Aerospace Recommended Practice: Design Objectives for Flying Qualities of Civil Transport Aircraft. ARP 842B, SOC. Automot. Eng., Aug. 1, 1964.

Revised Nov. 30, 1970.

18. Sudderth, Robert W.; Bohn, Jeff G.; Caniff, Martin A.; and Bennett, Gregory R . : Development of Longitudinal Handling Qualities Criteria for Large Advanced Supersonic Aircraft. NASA CR-137635, 1975.

V TABLE I.- M A S S AND DIMENSIONAL CHARACTERISTICS OF SIMULATED SUPERSONIC CRUISE TRANSPORT AIRPLANES (a) B a s e l i n e concept

Weight, N ( l b f ) . . . . . . . . . . . . . . . . . . . . . . 1 924 479 ( 4 3 2 6 4 0 )

Reference wing area, m2 ( f t 2 ) . . . . . . . . . . . . . . . . 784.75 ( 8 4 4 7 )

Wing span, m ( f t ) 38.66 ( 1 2 6 . 8 3 ) . . . . . . . . . . . . . . . . . . . . . .

Wing leading-edge sweep, deg (see f i g . 2) . . . . . . . . . . 74.00/70.84/60.00

Reference mean aerodynamic chord, m (ft) . . . . . . . . . . 27.00 ( 8 8 . 5 9 )

Center-of-gravity l o c a t i o n , p e r c e n t c . . . . . . . . . . . . . . . . 56

Static margin, p e r c e n t . . . . . . . . . . . . . . . . . . . . . . . . - 3 . 9

IX, kg-mi ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 6 887 550 ( 5 080 000)

Iy, kg-m ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 67 994 260 (50 150 000)

12, kg-m2 ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 72 902 230 ( 5 3 770 000)

IxZ, kg-m2 ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . -2 833 660 (-2 090 000)

Maximum c o n t r o l s u r f a c e d e f l e c t i o n s :

B t f d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +20

6 f , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . O t o 4 0

6 a , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +30

6,f, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +22.5

6 s , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +SO

6 r , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +35

Maqimum c o n t r o l surface d e f l e c t i o n rates:

it, deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +SO

$ f , deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . 510

G a l deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +70

G a f f deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +40

is, deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . 550

6, . deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +50

H o r i z o n t a l t a i l : Gross h o r i z o n t a l - t a i l area, m2 ( f t 2 ) . . . . . . . . . . . . . 4 9 . 8 0 ( 5 3 6 )

Mean aerodynamic chord, m ( f t ) . . . . . . . . . . . . . . . . 6 . 0 4 ( 1 9 . 8 0 )

Distance from c e n t e r of g r a v i t y to h o r i z o n t a l - t a i l 0 . 2 5 c r m ( f t ) . . . . . . . . . . . . . . . . . . . . . . 3 2 . 9 0 ( 1 0 7 . 9 3 ) V e r t i c a l t a i l : Exposed v e r t i c a l - t a i l area, m2 ( f t 2 ) . . . . . . . . . . . . . 1 6 . 7 2 ( 1 8 0 )

Mean aerodynamic chord, m ( f t ) . . . . . . . . . . . . . . . . 6 . 3 5 ( 2 0 . 8 3 )

Distange from c e n t e r of g r a v i t y to v e r t i c a l - t a i l

0 . 2 5 ~ ~ m (ft) . . . . . . . . . . . . . . . . . . . . . . 3 6 . 4 1 ( 1 1 9 . 4 6 )

TABLE 1.- Concluded (b) Powered-lift concept

Weight, N ( I b f ) . . . . . . . . . . . . . . . . . . . . . . 1 924 479 (432 640)

. . . . . . . . . . . . . . . .

Reference wing area, m2 ( f t 2 ) 784.75 (8447)

Wing span, m ( f t ) . . . . . . . . . . . . . . . . . . . . . . 38.66 (126.83)

Wing leading-edge sweep, deg (see f i g . 3) . . . . . . . . . . 74.00/70.84/60.00

. . . . . . . . . .

Reference mean aerodynamic chord, m ( E t ) 27.00 (88.59)

Center-of-gravity l o c a t i o n , p e r c e n t c . . . . . . . . . . . . . . . . 66

S t a t i c margin, p e r c e n t . . . . . . . . . . . . . . . . . . . . . . . . -9.5

Ix, kg-mi ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 6 937 629 ( 5 117 000)

Iy, kg-m ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 65 860 697 ( 4 8 577 000)

IzI kg-m2 ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . 70 713 105 (52 156'000)

1x2, kg-m2 ( s l u g - f t 2 ) . . . . . . . . . . . . . . . . . -2 779 390 (-2 050 000)

Maximum c o n t r o l s u r f a c e d e f l e c t i o n s :

6 t , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +20

6 f , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . O t o 4 0

d a I d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 230

dafo, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +20

6 a f i , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +30

6 = , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . + 2 5

Marimum c o n t r o l s u r f a c e d e f l e c t i o n rates:

itI deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . &SO

o f , deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . 510

$a, deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +70

Qafo, deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +40

f a f i r deg/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . +40

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6, deg/sec +50 H o r i z o n t a l t a i l : Gross h o r i z o n t a l - t a i l area, m2 ( f t 2 ) . . . . . . . . . . . . . 88.83 (956)

Mean aerodynamic chord, m ( f t ) . . . . . . . . . . . . . . . . 7.13 (23.39)

Distance from c e n t e r o f g r a v i t y to h o r i z o n t a l - t a i l

0.25:, m ( f t ) . . . . . . . . . . . . . . . . . . . . . . . 29.76 (97.62)

V e r t i ca 1 t a il :

Exposed v e r t i c a l - t a i l area, m2 ( f t 2 ) . . . . . . . . . . . . . 34.84 (375)

Mean aerodynamic chord, m ( f t ) . . . . . . . . . . . . . . . . 9.16 (30.07)

Distance from c e n t e r of g r a v i t y to v e r t i c a l - t a i l

0.25C, m ( f t ) . . . . . . . . . . . . . . . . . . . . . . 33.27 (109.16)

. ...............

TABLE 11.- AERODYNAMIC INPUTS USED IN SIMULATION OF BASELINE CONCEPT

Aerodynamic inputs for - 1

-4 ' -0.0484 -0.0038 0.0734 0.00059 -0.00039 -0.00769 I -0.0019 -0.0085 0.00059

-. 0038

0 -.0340 -. 0920 -. 00085 -. 00045 -. 00720 -. 0085 -. 0018 0

-. 0203 -. 2551, -. 00031 -. 0038 -. 00053 -. 00675 -. 0085 -. 0017 -. 00059

ii 1 -.0071 -. 4159 ,00018 -. 0038

-. 00061 -. 00645 -. 0084 -. 0016 -. 00119

.0031 -. 5820 .00058 -. 0037 -. 00619 -. 0016

-. 00073 -. 0083 -. 00177

.0180 -. 7414 .00095 -. 0036

-. 00096 -. 00632 -. 0082 -. 0014 -. 00235

20 .0235 -. 9301 .00128 -. 0035 -. 00095 -. 00588 -. 0080 -. 0008 -. 00291

I

! I

for tail deflections of -

C,.,.,(df=o) a, deg?

I I I i

I ' -200 -150 -100 , -50 00 i 50 i l o o I 15O ! 200 1

I --I_i

i -4 ~0.1080~0.1060~0.1010~0.0843 0.0447 0.0073 -0.0237!-0.0478'-0.0645

0 .1318 ", .1255 .1113 : .0858 .0460 : ,0083' -. 0225 -. 0465 -. 0630

.0870 -0475 .0100, -.02131 -.0453 -.0615 h , W w TABLE 11.- Continued ~~ Aerodynamic inputs for - _I I deg-l Cy& , deg-l C z 6 deg- C deg-l C deg-l C , deg-l Iy6af a S "6s "6af " 6 , (a) -0.00029 -4 -0.00029 0.00019 -0.00009 0 0.00012 0.00017 0.00044 0.00039

0 .00019 -. 00009

-. 00029 -. 00028 0 .00017 .OOOll - 0 0 0 4 4 .00038

-. 00028 4 -. 00028 .00017 -. 00008 0 .00016 .OOOll .00042 ,00038

-. 00026 8 -. 00026 .00012 -. 00008

0 .00009 .00014 .00040 .00036

-. 00023 1 2 -. 00023 .00004 -. 00007 0 .00007 .00010 .00036 .00033

I

-. 00008 -. 00008 .00002 1 6 -. 00003 0 .00028 .00029

.00003 ,00004

-. 00007 0 . 0 0 0 0 1 20 0

0 .00002 -. 0 0 0 0 1 .00016 .00022

I - -u Aerodynamic inputs for -

1 % deg

-4 0.00100 -0.00008 -0.00120 -0.00646 -0.00128 0.00183

0 .00100 0 -. 00120

-. 00654 -. 00150 .00176

4 .00100 .00008 -. 00120 -.00681 -. 00179 .00169 1

8 .00100 .00017 -. 00119

-. 00723 -. 00219 .00160

1 2 .00099 .00026 -. 00116 -. 00789 -. 0 0 1 9 1 . 0 0 1 3 1

1 6 . 0 0 0 9 1 .00044 -. 0 0 1 0 1

-. 00649 -. 00199 .00125

20 .00059 .00043 -. 00059 -. 00770 -. 00200 .00119

L L -L aRigid derivatives.

TABLE 11.- Concluded _ _ _ _ _ ~ ~ ~ ~~

Aerodynamic inputs for -

a, deg rad-l C % , rad-l CypI rad-l I rad-l;CnpI rad-l CyrI rad-l;ClrI rad-l CnrI r a d ' l

' c%r 1 c2P ~

-4 i -1.2716 -0.1279 -0.1113 , 0.2291 ~ 0.0257 -0.3225 -0.1621 0.1233

-. 3120

-1.2680 -. 1528 .5094 -. 1209 -. 1074 .2402 .0879

-. 1404 .E368 -. 1237 -. 0985 .3016 .1432 -. 3009

-1.2635

-. 1184 1.1793 -. 1389 -. 0747 .4154 -. 2941

-1.2590 .1946

-. 0844 1.5157 -. 1796 .5815 .2439 -. 2712

-1.2555 .0057

-. 2256 -. 2389

16 -1.2475 .0165 1.8106 ,0287 .7817 .2848

.0672 2.0982 -. 2880 .0567 .9837 .3293 -. 2003

-1.2550 w r W h) TABLE 111.- AERODYNAMIC INPUTS USED I N SIMULATION OF POWERED-LIFT CONCEPT [Trailing-edge f l a p s d e f l e c t e d 40°3 (a) All engines o p e r a t i n g

f o r t a i l d e f l e c t i o n s of -

CX -100 50 15O 200 -15O -50 00 100 -0.0722 -0.0802 -0.0900 -0.1106 -0.0949 -0.0837 -0.0750 -0.0692 0 -4 -0.1307

-. 0687 -. 0772 -. 0877

-. 1072 -. 0910 -. 0792 -. 0702 -. 0651

10 I 0 ~ -.1275

-. 0624 -. 0715 -. 0822

-. 1005 -. 0840 -. 0721 -. 0630 -. 0585

0 4 -.1216

-. 0662 -. 0774

-. 0937 -. 0772 -. 0653 -. 0562 -. 0522 -. 0565

0 8 ~ -.1148

-. 0628 -. 0749

-. 0696 -. 0590 -. 0517 -. 0479 -. 0524

12 -.lo70 -. 0858

-.0745

-. 0654 -. 0559 -. 0503 -. 0468 -. 0515 -. 0623

0 16 -.lo15 -. 0808

-. 0772

-. 0633 -. 0556 -. 0521 -. 0495 -. 0536 -. 0646

20 -.0983 -. 0776

-.0197 -.0246 -.0134 -.0047 .0011 -.0019 -e0099

-4 -.0604 -. 0403

.15

-. 0167

-. 0200 -. 0082 .0008 -0059 -0023 -a0062

0 -.0565 -a0362 .15

.0133 .0094 .0003 -. 0104

4 -.0498 -.0287 -.0122 -.0003 .0088 .15

,0188 ,0091 -. 0021

8 -.0395 -e0184 -.0019 .0100 .0191 ,0231 .15 .0201 .0080 12 -.0241 -.0029 .0133 .0239 .0312 ,0350 ,0305 .15 .0451 -0343 .0221 16 -a0049 -0158 .0312 .0407 ,0443 ,0498, .15 .0584 .0474 .0348 20 .0137 .0344 .0487 .0564 .OS99 .0625 .15 .0806 .0399 .0600 .0757 .0869 .0956 .lo14 ,0984 ,0904 .30 -4 ,0940 .0835 .0640 .0802 ,0920 .lo10 .lo61 .lo25 .3Oj 0 .0437 .0892 4 .0498 .0709 .0874 .0993, .lo84 .1129 .1090, .0999 .30 .30 8 ,0597 .0808 .0973 .lo92 .1183 ,1223 .1180 -1083 -0971 .30 12 .07621 e0974 -1136 -1242 -1315 -1353 .1308 -1204 -1083 .30 16 .0984 I .1191 .1345 .1484 .1376 1254 -1210 I -1417 .1560 -1657 -1547 ,1421 .30 20 1- --ui-.

TABLE 111.- Continued [Trailing-edge flaps deflected 40°] h (a) Continued

for tail deflections of -

Cz 2 00

' 7

4 -.3286 -.3450 -.3728

-. 41331 -. 4694 -. 5235' -. 5666 -. 5987 -. 6211

0 1 I

1 0 ' 8 I -.4610 -.47911 -.5092 -.5517 -.6074 -.6615 -.7044 -.7360 -.7577

0 12 i -.6093 -.6271 -.6555 -.6970 -.7530 -.E079 -.E502 -.8806 -.go13

0 1 16 1 -.7691/ -.7839, -.E104 -.E514

-.9078l -.9622 -1.0030 -1.0311 -1.0494 0 20 ' -.9410 -.9558 -.9814 -1.0207 -1.0751 -1.1276 -1.1666 -1.1944 -1.2123 I I I .15/ -4 -.4098 -.4119 -.4192 -.4430 -.4994 -.55361 -.5965 -.6293 -.6526 .15\ 0 -.5375 -.5475 -.5671

-. 6034 -. 6596 I -. 7136 I -. 7572 -. 7895 -. 8125

, .15, 4 -.6776 -.6940 -.7218

-. 7623 -.E184 -.E725 -.9156 -.9477 -.9701

.15 8 -.E287 -. 8468 -. 8769

-. 9194 -.9751 -1.0292 -1.0721 -1.1037 -1.1254

.15 12 -.9902 -1.0080 -1.0364 -1.0779 -1.1339 -1.1888, -1.2311 -1.2615 -1.2822 .15 16 -1.1453 -1.1601 -1.1866 -1.2276'

-1.2840 -1.3384 ' -1.3792 -1.4073 -1.4256

-15 20 -1.3031 -1.3179 -1.3435 -1.3828 -1.4372 -1.4897 -1.5287 -1.5565 -1.5744 I

-30 -4 , -.5305 -. 5326 -. 5399 -.6201 -.6743 -.7172 -.7500 -.7733 -. 5637'

0 -.6675

.30 -. 7334 -. 7896 -. 8436

-. 8872 -.91951 -. 9425

-. 9056 -. 9617 -1.0158

-1.0589 -1.0910 -1.1134

-. 9943 -1.0244 '-1.0669 -1.1226 -1.1767 -1.2196 -1.2512 -1.2729

-1.1388 -1.1566 -1.1850 -1.2265 -1.2825 -1.3374 -1.3797 -1.4101 -1.4308 -1.3227 -1.3492 -1.3079 -1.3902 -1.4466 -1.5010 -1.5418 -1.5882 -1.5699 -1.4885 -1.5141 -1.5534 -1.6078 -1.6603 -1.6993 -1.7450 -1.7271 W W w l b TABLE 111.- Continued [Trailing-edge flaps deflected 4001 (a) Concluded I

for tail deflections of -

c, 1

-200 -15O -100 -50 00 50 100 15O 0.1173 0.1148 0.1063 0.0785 0.0126 -0.0508 -0.1009 -4 -0.1393 -0.1665

-. 1503

.2087 .1896 4

.1571 1 .lo97 1 .04411 -.0191 -. 0694 -. 1070 -. 1332

~ ,2326 .2114 8 .1762 1 .1266 1 .0615 -.0018 -. 0519 -. 1142

-. 0889

0 .2476 12

.2268 ,1936 .1451 I .0797 .0155 -. 0339

-. 0695, -. 0937

0 16

.2119 .1640 ' .0980 .0344 -. 0132 -. 0461

-. 0675

0 20

.2245 .1785 ' .1150 .0536 .0080 -. 0245 -. 0454

I .15 -4 .0623

.0598 .0513 .0235 -.0424, -.lo58 -. 1559 -. 1943 -. 2215

.15

0 j ,1159 .lo42 .0813 .0388 ! -.0269 j -.0899

-. 1409 -. 1787 -. 2056

.15 4 .1528 .1337 ,1012 .0538 -.0118 -.0750

-. 1253 -. 1629 -. 1891

.15 8 .1770 ,1558 .1206 .0710 .0059 -.0574

-. 1075 -. 1445 -. 1698

.15 .1947 .1739 .1407 .0922 I .0268 -.0374

-. 0868 -. 1224 -. 1466

.15 16

.2076 .1903 ,1594 ' .1115 .0455 -.0181 -. 0657

-. 0986 -. 1280

.15 20 ,2244 .2071 .1772 ' .1312 .0677 .0063

-. 0393 -. 0718 -. 0927

I .30 -4 .0462 .0437 .0352 ' .0074 -.0585 -.1219 -.1720 -.2104 -.2376 .30 0 1 -0966 .0849 .0620 .0195 -.0462 -.lo92 -.1602 -.1980 -.2249 .30 4 .1323 I .1132

e0807 -0333 I -.0323 -.0955 . -.1834

-.1458 -.2096 .30 8 .1545 .1333 -0981 -0485 -.0166 -.0799 -.1300 -.1670 -.1923 .30 12 .1709 .1501 .1169 .0684 ' .0030 -.0612 -.1106 -.1462 -.1704 .30 16 .1839 ' .1666 I -1357 .0878 I ,0218 -.0418 -.0894 -.1223 -.1437 .30

20 , .2012 .1839 ' .1540 .lo80

.0445 I -.0169 -.0625 -.0950 -.1159 TABLE 111.- Continued [Trailing-edge flaps deflected 4001 (b) Number 3 engine failed I 10 ~ -4 -0.1307 ~-0.1106~-0.0949~-0.0837~ -0.0750'-0.06921-0.0722 -0.0802 -0.0900 0 ' 0 -.1275 -.lo72 -.0910' -.0792 -.0702 -.0651 -.0687 -.0772 -.0877

0 4 -. 1216 -. 1005 -. 0840 -. 0721 -. 0630, -. 0585 -. 0624 -. 0715 -. 0822

0 8 -.1148 -.0937 -.0772 -.0653 -.0562 -.0522 -.0565 -.0662 -.0774

0 12 -. 1070 -. 0858 -. 0696 -. 0590, -. 0517 -. 0479 -. 0524 -. 0628 -. 0749

, o -.io15 -.oaoai - . o m -.05591 -.0503' -.0468 -.0515 -.0623 -.0745

' 0 20 -. 0983 -. 0776 -. 0633 -. 0556 -. 0521 -. 0495 -. 0536 -. 0646 -. 0772

I

I .15 -4 -. 0425 -.0224 -.0067 .0045 .0132 .0190 .0160 .0080 -.0018

I .15 0 -. 0394 -. 0191 -. 0029 ,0089 ,0179 .0230 .0194 .0109 ,0004

.0294 .0255 .0164 .0057

' .15 4 -. 0337 -. 0126, .0039 .0158; ,0249

.15 -. 0250 -. 0039 ' .0126 .0245' .0336 .0376 ,0333 .0236 .0124

.15 12 -. 0115 .0097 .0259 .0365; .0438 .0476 .0431 .0327 .0206

.15 16 .0055 .02621 .0416 ,0511 .0567 .0602 .0555 .0447 .0325 .15 20 ,0222 .0429 ! .0572 .0649 .0684 .0710 ,0669 .0559 .0433 .30 -4 .0988 .1189 .1346 .1458 .1545 .1603 .1580 .1495, .1390 0 .0992 .1195 .1357 .1475 .1565 .1616 .1616 .1525 .1418 4 ' .lo24 .1235 .1400 .1519 .1610 .1655 -1652 ,1555 .1443 8 .lo69 .1280 .1445 .1564 .1655 .1695 .1683 .1579 .1458 12 .1137 .1349 .1511 .1617 .1690 .1728 .1729 .1621 .1499 .30 .1229 .1436 .1590 .1741 .1776 16 .1685 .1789 .1679 .1553 .30 .1342 .1549 .1692 .1769 .la04 .1830 W m TABLE 111.- Continued [Trailing-edge flaps deflected 4001 (b) Continued

for tail deflections of -

CZ ~~ ar deg CT

-15O 1 -100

-50 00 50 100 15O 200.

-- -0.0912 -0.1223 -0.1787 -0.2329 -0.2758 -0.3086 -0.0985 -0..3319

-.2654 -.3216 -.3756 -.4192 -.4515 -. 2095 -. 2291 -. 4745

-. 4133 -. 4694 -. 5235 -. 5666 ' -. 5987 -. 3450 -. 3728

-. 6211

0 8 ' -.4610 -. 4791

-.5517 -.6074 -.6615 -.7044 -.7360 -. 5092 -. 7577

0 12 -.6093 -. 6271

-. 6555 -. 9013

0 16 -.7691 -. 7839

-. 8104 -. 8514 -1.0494

0 I 20 , -.9410 -. 9558

-. 9814 -1.0207 -1.2123

1 .151 -4 -.3390 -. 3411

-. 3484 -. 3722 -.4286 -.4828 -.5257 -.5585 -. 5818

I -151 0 ~ e.4725 -. 4825

-. 5021 -. 5384 -. 5946 ' -. 6486 -. 6922 -. 7245, -. 7475

4 1 -.6164 -. 6328

-. 6606 -. 7011 -.7572' -e8113 -.8544 -.8865 -.go89

'i : : : I 8 , -.7692 -. 7873

-. 8174 -. 8599 -. 9156 ' .9697 -1.0126 -1.0442 -1.0659

I 1 .15/ 12 ' -.9324 -.9502: -.9786 -1.0201 -1.0761 -1.1310 -1.1733 -1.2037 -1.2244

j .15 ~ 16 , -1.0901 -1.1049 -1.1314, -1.1724 -1.2288 -1.2832 -1.3240 -1.3521 -1.3704

I .15' 20 '-1.2496 -1.26441-1.2900:-1.3293 -1.3837 -1.4362 -1.4752 -1.5030 -1.5209 .j

j .30 -4 I -.3120, -.3141: -.3214 -.3452 -.4016 -.4558 -.4987 -.5315 -.5548

I -30 0 ~ -.4565 -. 4665 -.4861: -.5224. -.5786 -.6326 -.6702 -.7085 -.7315

-30 4 , -.6154 -. 6318

-. 6596 -.7001 -.7562 -.8103 -.8534 -.8855 -.go79

- 3 0 , 8 -.7742 -. 7923

-. 8224 -.8649 -.92061 -.9747 -1.0176 -1.0492 -1.0709

-30 12 -.9372 -. 9550

-. 9834 -1.0249 1-1.0809 -1.1358 -1.1781 -1.2035 -1.2292

-30 16 -1.1157 -1.1305 -1.1570 -1.1980,-1.2544 j -1,3088 -1.3496 -1.3777 -1.3969 - 3 0 , 20 -1.2898 -1.3046 -1.3302 -1.3695,-1.4239j-1.47644.51541-1.5432 ,-1.5600 1 - TABLE 111.- Continued [Trailing-edge flaps deflected 4 0 ° 1 (b) Concluded I I I

for tail deflections of -

,

-0.0508 -0.1009'-0.1393 -0.1665 0 ' -4 i 0.1173 I 0.1148 1 0.1063 ' 0.0785 0.0126

0 1 0 .1712 .1595 I .1366 .0941 .0284 0 4

.2087 .1896 ' .1571 .lo97 .0441

.2326 .2114 I .1762 .1266 .0615 ; o 8 .2476 .2268 ' .1936 .1451 .0797 i o I 12 ' 0 ' 16 .2601 .2428 .2119 .1640 .0980 I 0 I 20 .2717 .2544 .2245 .1785 .1150 I -4 .0443 ~ -.0216' -.0850 ! -.1351 -.1735 -.2007

, .15 I .0831 .0806 .0721

.15 0 1 .1357 .1240 .loll -0586 -. 0071 -. 0701 -. 1211. -. 1589 -. 1858

I .15 4 .1725 .1534 .1209 .0735 ~ .0079 -.0553 -.lo56 -.1432 -.1694 .15 8 .1948 .1736 .1384 ,0888 ! .0237 -.0396 ' -.0897 -.1267 -.1520 I -.0684 -.lo40 -.1282 .15 12 .2131 .1923 .1591 ,1106 .0452 -.0190

-. 0461 -. 0790 -. 1004

.15 16 .2272 .2099 .1790 I ,1311 .0651 .0015

.15 20 , .2443 .2270 .1971 .1511 .0876 .0262 -. 0194 -. 0519 -. 0728

I

' .30 -4 .OB51 .0573 -.0086: -.0720 -. 1221 -. 1605 -. 1877

0 .1106 .0681 .0024 -.0606 -. 1116 -. 1494 -. 1763

.30

4 .1282 , .0808 .0152 -.0480 -. 0983 -. 1359 -. 1621

.30

8 .1427 .0931 .0280 -.0353 -. 0854 -. 1224 -. 1477

.2155 12 -. 0660 -. 1016 -. 1258

.1947

.2131 16 -. 0429 -. 0758 -. 0972

.2304

.2490 -. 0147 -. 0472 -. 0631

.2317 .30 20 w W 0) TABLE 111.- Continued [Trailing-edge flaps deflected 4 0 ° 3 (c) Number 4 engine failed -

for tail deflections of -

CX CT -50 00 -100 - -0.0837 -0.0750 -0.0692 -4 -0.0949 -0.0722 -0.0802 -0.0900

-.0792 -.0702 -.0651 -.(I910 -.0687 -.0772 -. 0877

0 0

-. 0721 -. 0630 -. 0585 -. 0624 -. 0715 -. 0822

0 4 -. 0840

-.0653 -.0562 -.0522 ! -.0565 -. 0524, -.0662 -. 0628 -. 0774

0 8 -. 0772

-. 0749

0 12 -. 0696

-.0515 -.0623 -. 0745

0 16 -. 1015 -. 0808 -. 0654

-. 0983 -. 0776 -. 0633 , -. 0556 -. 0521 -. 0495 -. 0536 -. 0646 -. 0772 I

0 20 -4 -.0798 -.0597 -.0440 -.0328 -.0241 -.0183 -.0213 -.02931 -.0391 .15

.15 0 -. 0762, -. 0559 -. 0397 -. 0279 -. 0189 -. 0138 -. 0174 -. 0259 -. 0304

4 -.0694 -.0483 -.0318 -.0199 -. 0x08 ! -. 0063 -. 0102 -.0193 -.0300

.15

.15 8 -. 0586 -. 0375 -. 0210 -. 0091 0 .0040 -. 0003 -. 0100 -. 0212

.15 -.0414 -.0202 -.0040 .0060 .0139 .0177 .0132 .0028 -.0093 -.0192 .0015 .0169 .0264 .0320 I .0355 .0308 .0200, .0078 .15 16 .0520 .0243

.15 20 .0032 .0239 .0382 .0459, .0494 ' .0479 .0369

-4 .0240 .0441 .0598 .0710 ,0797 .0855 .0825 .0745 .0647 .30 .30 0 .0257 .0460 .0622 .0740 -0830, .0881 .0845 .0760 .0655 4 .0307 .0518 .0683 .0802 .0893 .0938 .0899 .0808 .0701 ' .30 .30 8 .0397, ,0608 .0773 .0892 .lo23 .0883 .0771 .0983 .0980 .0538 .lo18 .1129 .lo84 .0980 .0859 .30 12 .0750 .0912 .lo91 .30 .07331 .0940 .lo94 .1189 .1245, .1280 .1233 .1125 .lo03 ' .30 20 -0962 -1169 -1312 ,1389, .1424 .1450 -1409' .1299! .1173 TABLE 111.- Continued [Trailing-edge flaps deflected 4001 (c) Continued

I

1 Cz for tail deflections of -

' 0 -4 , -0.0891 -0.0912 -0.0985'-0.1223 -0.17871 -0.2329 -0.2758 -0.3086 -0.3319 0 0 1 -.1995 -.2095 -.2291 -.2654 -.32161 -.3756 -.4192 -.4515 -.4745

0 4 , -.3286 -.3450 -. 3728 -. 4133' -. 46941 -. 5235 -. 5666

-. 5987 -. 6211

' 0 8 i -.4610 -.4791 -.5092 -.5517 -.6074 -.6615 -.7044 -.7360 -.7577

0 12 -.6093 -.6271 -. 6555: -. 6970 -. 7530 -. 8079 -. 8502

-.8806 -.go13

0 16 -.7691 -.7839 -. 8104, -.8514 -.go78 -.9622 -1.0030

-1.0311 -1.0494

0 20 -.9410 -.9558 -. 9814 -1.0207 -1.0751 -1.1276 -1.1666

-1.1944 -1.2123

.15 -4 -.4645 -.4666 -. 4739

-.6512 -.6840 -.7073

.15 0 -.5945 -.6045 -. 6241

-. 8142 -.8465 -.8695

.15 4 -.7349 -.7513 -. 7791 -. 9729 -1.0050 -1.0274

8 -.8843 -.go24

.15 -. 9325 -1.1277 -1.1593 -1.1810

.15 12 -1.0440 -1.0618 -1.0902 -1.2849 -1.3153 -1.3360 .15 16 -1.1984 -1.2132 -1.2397 -1.4323 -1.4604 -1.4787 -1.4342 -1.4886 -1.5411 .15 20 -1.3545 -1.3693 -1.3949 -1.5801 -1.6079 -1.6258

-. 5962 -. 6526 -. 7068 .30 -4 ~ -.5630i -.5651 -. 5724 -. 7497 -. 7825 -. 8058

-.7664 -.8226 -.8766 -. 7105 -. 7301 -. 9202 -. 9525 -. 9755

-. 9371 -. 9932 -1.0473 -. 8688 -. 8966 -1.0904 -1.1225 -1.1449

-1.0949 -1.1506 -1.2047 -1.0042 -1.0223 -1.0524 -1.2476 -1.2792 -1.3009 -1.2481 -1.3041 -1.3590 -1.1604 -1.1782 -1.2066 -1.4013 -1.4317 -1.4524 -1.4144 -1.4708 -1.5252 -1.3321 -1.3469 -1.3734 -1.5660 -1.5941 -1.6124 -1.5793 -1.6337 -1.6862 -1.4996 -1.5144 -1.5400 -1.7252 -1.7530 -1.7709

I

IP

I

TABLE 111.- Continued [Trailing-edge flaps deflected 4001 (c) Concluded

C, for tail deflections of -

CT % deg -200 -15O -100 -50 00 50 100 150 200 t o -4 0.1173 0.1148 0.1063 0.0785 0.0126 -0.0508 -0.1009 -0.1393 -0.1665 ' 0 0

.1712 .1595 .1366 .0941 .0284 -.0346 -.0856 -.1234 -. 1503

-1097 .0441

0 4 ' -2087 i -1896 i .1571 -. 1332

-. 1142

-. 0937

' 0 16 .2601 .2428 ,2119 -1640 .0980, -0344 -.0132 -.0461 -.0675 10 20 .2717 .2544 ' .2245 -1785 .1150 .0536 .0080 -.0245 -.0454 .15 -4 .0535 .0510 .0425 ,0147 -.0512 -.1146 -.1647 -.2031 -.2303 I .15 0 .lo61 .0944 ' .0715 -0290 1 -.0367 -.0997 -.1507 -.1885 -.2154 , .15 4 .1431 .1240 .0915 -0441 ~ -.0215 -.0847 -.1350 -.1726 -.1988

, .15 8 .1664 .1452 .1100 -0604 ' -. 0047 ' -. 0680 -. 1181 -. 1551 -. 1804

.15 12 .1849 .1641 .1309 .0824 .0170 -.0472 -.0966 -.1322 -.1564 .15 16 .1994 .1821 .1512 -1033 .0373 -.0263 -.0739 -.lo68 -.1282 ~ .15 20 .2167 .1994 ,1695 ,1235 .0600 e.0014 -.0470 -.0795 -.lo04 ~ .30 -4 .0369 .0344 .0259 -.0019 I -.0678 -.1312 -.1813 -.2197 -.2469 ~ .30, 0 .0859 .0742 ,0513 ,0088 -.0569 -.1199 -.1709 -.2087 -.2356 .30 4 .1211 .lo20 .0695 ,0221 -.0435 -.lo67 m.1570 -.1946 -.2208 .30 8 .1422 .1210 .0858 .0362 -.0289 -.0922 -.1423 -.1793 -.2046 .30 ' 12 .1591 .1383 .lo51 ,0566 -.0088 -.0730 -.1224 -.1580 -.1822 .30 16 .1747 .1574 .1265 .0786 .0126 -.0510 -.0986 -.1315 -.1529

-301 20 , .1939 .1467 , ,1007 I .0372 -.0242 -.0698 -.lo23 -.1232

I - b LA TABLE 111.- Continued [Trailing-edge flaps deflected 4 0 ° 1 (d) Aerodynamic inputs

I Aerodynamic inputs for -

- j c1 % , I 6, -4 -0.00616 -0.00038 0.00176 0.00161 0.00019 -0.00173

0 -. 00510 -. 00140 .00206 .00160 .00013 -.00179

0 4 -. 00420 -. 00212 .00217 .00155 .00009 -.00179

0 8 -. 00359 -. 00278 .00206 .00144 .00007 -. 00170

0 12 -. 00324 -. 00329 .00167 .00106 .00007 -.00150

0 16 -. ooigi -. 00382 .00223' .00061 .00012 -.00119

0 20 .00189 -. 00408 .00376 .00045 .00020 -.00105

.00016 -,. 00173 .15 -4 -. 00649 -. 00035 .00176 .00194

-. 00180 .00207 .00194

.00013 .15 0 -. 00544 -. 00140

.00011 -.00181 .15 4 -. 00453 -. 00214 ,00219 .00188

.00009 -.00175 .15 8 -. 00389 -. 00280 .00211 .00174

.00008 -.00158 .15 12 -. 00346 -. 00330 .00175 .00128

.00010 -.00127 .15 16 -. 00203 -. 00380 .00231 .00073

.00016 -.00112 .15 20 .00188 -. 00404 .00383 .00047

.00014 -.00173 -4

.30 -. 00681 -. 00033

-00013 -. 00181 .30 0 -. 00578 -. 00140

.00013 -.00182 4 -. 00487 -. 00216

.OOOll -. 00180 .00216 .00205 8 -. 00420 -. 00282

.00009 -.00166 .00183 .00150 12 -. 00368 -. 00331

.00008 -.00134 .00238 .00085 -. 00215 -. 00378

.00013 -. 00119 .00390 .00048

20 .00186 -. 00401

aRigid derivatives.

TABLE 111.- Continued [Trailing-edge f l a p s d e f l e c t e d 4Qo] ( d ) Continued

I n p u t due t o NO. 3 1 I I n p u t due t o N o . 4

engine f a i l u r e engine f a i l u r e

1 cT . I a ' deg l--l---

0 0 0 0 I -4 0

0 0 0 0 0 O I : 0 0 0 4 0 0 0 4 0 1 0 0 8 0 0 0 8 0 1 0 0 0 c 12 0 0 0 12 0 1 0 0 0 16 0 0 0 0 16 O i O 0 0 20 0 0 0 ' 0 20 0 , 0 0 I -4

.15 -. 0030 .0153 .0034 .15 -4 .0267

0 , ,0018

.15 0 -. 0042 .0154 .0034 .15 ~ 0 0 ' .OOOl .0269

.15 4 -. 0048 .0160 .0033 .15 4

0 -. 0015 .0267

.15 8 -. 0052 .0170 .0033

.15 8 0 -. 0033 .0264

.15 12 -. 0071 .0182 .0034 12

.15 0 -. 0048 ' .0259

.15 16 -. 0122 .0199

.0034 .15 16 0 -. 0063 .0250

.15 20 -. 0204 .0217

.0046 .15 20 0 -. 0077 .0241

I

.30 -4 -. 0060 .0307 .0068

.30 , -4 0 .0032 ~ .0535

.30 0 -. 0084 ,0309 .0067

.30 0 0 .0002 ,0537

.30 4 -. 0097 .0321 .0066 .30 , 4

0 -. 0028 I .OS35

.30 8 -. 0103 .0340 .0066 .30 8

0 -. 0058 .0529

.30 -. 0142 .0365 .0067 .30 12

0 -. 0087 .0519

.30 16 -. 0244 .0399 .0068 .30 16 0 -. 0114

.0505 .30

-. 0409 .0434 ' .0091 .30 20 0 -. 0139

.0487 I 20 I TABLE 111.- Continued [Trailing-edge flaps deflected do0] (d) Continued -i

Aerodynamic inputs for -

-0.00038 0.00009 -0.00015 0.00050 0.00007 -0.00008 0 -4 0.00042

-. 00040 .00046 ' .00009 -. 00027 .00049 .00008 -. 00007

0 0

0 4 -. 00041 .00046 .00009 -.00033 .00048 .00008 -. 00006

-. 00035 -. 00006

0 8 -. 00041 .00043 .00009 .00047 .00007

-. 00033 .00048

0 12 -. 00039 .00036 .00008 .00006 -. 00005

-. 00035 .00028 .00007 -. 00029 .00048 .00006 -. 00004

0 16

-. 00013 .00010 -. 00025 .00041 .00008 -. 00004

0 20 .00024

.00007 -. 00038

.15 -4 -. 00038 .00042 .00009 -. 00015 .00050

-.00027 .00049 .00008 -. 00036

.15 0 -. 00040 .00046 .00009

.15 4 -. 00041 .00046 .00009 -. 00033 .00048 .00008 -. 00035

.15 8 -. 00041 .00043 .00009 -. 00035 .00047 .00007 -.00042

.00006 -. 00045

.15 12 -. 00039 .00036 .00008 -. 00033 .00048

-.00035 , .00028 .00007 -.00029 .00048 .00006 -.00042 .15 16 ,

-. 00013 .00024 .00010 -. 00025 .00041 .00008 -. 00040

.15 20 ,

.00050 .00007 -. 00068

.30 -4 -. 00038 .00042

.00137 .00039 .30' 0 -. 00040 .00046 .00049 .00008 -. 00065

.00134 .00041 .00008

.30 4 -. 00041 .00046 .00048 -.00065

.00131 .00043 .00047 .00007

.30 8 -. 00041 .00043

.00125 .00039 .00008 -.00033 -. 00039 .00036 .00048 -. 00085

-301 12

.00116 .00029 .00007 -. 00029 .30 16 -. 00035 .00028 .00048 .00006 -. 00080

.OOllO .00027 1 .00008 -oooo6

.30 20 -.00013 .00010 -. 00025 .00024 .00041 -. -.00078 00077 I

I I 1 I 1 TABLE 111.- Concluded [Trailing-edge flaps deflected 4001 (d) Concluded I I

Aerodynamic inputs for -

hf r a d ' l :ypf rad-l ! I p f rad-l

I -0.4239 0.0897 -0.1206

-0.1150 0.2328 0.0117 1 -0.3848 1

-. 1074 -. 4146 .5094

0 -1.5604 -. 1209 .2339 .0917 1 -.4057 I

-. 0909 -. 4022 .E682 -. 1204 .1373 1

0 4 -1.5560 .2823 -. 3899

1.2427

-. 0661 0 1 8 ' -1.5515 -. 3749 -. 1171 .3959 .1822 , -.3227 1

-. 3257 1.6091 -. 0922

-.0065 ' 0 12 -1.5480 .5518 .2048 -. 1692

-. 2049 1.9096 .2740

0 ' 16 -1.5400 -. 0970 .7052 -.3599 '

i 0 20 I -1.5475 -. 1279 2.0000 .0112 .7511 .2076 .2884

I I -1.5641

' .I51 -4 -. 4239 ,4429 -. 1876 .2328 .1292 -. 3467

-1.5604 -. 4146

.15 0 ,8637 -. 2243 .2339 ,2080 -. 3510

-15 4 -1.5560 -. 4022 1.2102 -. 2595 .2823 .2424 -. 3198

-15 8 -1.5515 -. 3749 1.5630 -. 2930 .3959 .2704 -. 2384

-15 12 -1.5480 -. 3257 1.8732 -. 3004 .5518 .2664 -.0754

.15! 16 -1.5400 -. 2049 2.0663 -. 3212 .7052 [ .3002 -. 2668

-1.5475 -. 1279 2.1070 .3722

.15 20 -. 1863 .7511 ' .1947

-30 -4 -1.5641 -. 4239 .5691 -. 2216 -. 3099

.2328 .1747 -. 3237

.30 0 -1.5604 -. 4146 1.0000 -. 2755 -.3259 1 .2527 -. 3199

.2339

-30 4 -1.5560 -. 4022 1.3601 -. 3315 -.3437 ; -. 2795

-2823 1 .2837

- 3 0 ' 8 ' -1.5515 -. 3749 1.7161 -. 3801 -.3642

.3959 ! .3020 -. 1927

I

- 3 0 ' 12 I -1.5480 -. 3257 2.0255 -. 3984 -.3566 .2841 -. 0271

.5518 1

.30 16 -1.5400 -. 2049 2.2301 -. 4373 .3016 -. 2148

.7052 ,

.30 20 -1.5475 -. 1279 2.2750 -. 3433 .4222

.7511 .1761 TABLE 1V.- PILOT RATING SYSTEM SATISFACTORY Excellent, highly desirable. 1 feasible pilot compensation.

Capable of being controlled Very objectionable deficiencies. Major 6 or managed in context of improvements are needed. Requires best mission, with available available pilot compensation to achieve pilot attention. acceptable performance.

Major deficiencies which require improvement 7 for acceptance. Controllable. Performance inadequate for mission, or pilot compensa- tion required for minimum acceptable per- UNACCEPTABLE formance in mission is too high.

Deficiencies which require improvement. Inadequate Controllable with difficulty. Requires sub- 8 performance for mission even with maximum stantial pilot skill and attention to retain feasible pilot compensation. control and continue mission.

IMarginally controllable in mission. 9 Requires maximum available pilot skill and 1 attention to retain control.

UNCONTROLLABLE luncontrollable in mission.

I

/Po I

Control will be lost during some portion of mission.

I TABLE V . - DYNAMIC STABILITY CHARACTERISTICS OF SIMULATED SUPERSONIC CRUISE TRANSPORT AIRPLANES [Approach speed was 1 5 3 knots] (a) Baseline concept Augmentation Satisfactory Acceptable = A S Iblodified Parameters criterion criterion Shor t-per iod mode -~

wSpr rad/sec . . . . . 0 . 1 7 1

1 . 5 3 4 1 . 5 3 4 0 . 7 5 1

I See figure 38 lsee. figure 3t

pSpr sec . . . . . 42.72

1 5 . 1 2 15.12 8 . 7 1

1.036 csp . . . . . . . . . 0.507 0.693

1 . 0 3 6 0.35 to 1.30 0.25 to 2.0(

LcL/wsp . . . . . . . . 2.32

0.259 0.259 0.529 See figure 3 8 See figure 3E

n/a, g units/rad . . . 3.19 3.19

3.19 3.19 See figure 38 See figure 3E Long-period (aperiodic) t 2 r sec . . . . . .

Long-per iod (periodic) mode 0.067 0.080 0.080 1 2 5 . 2 98.9 98.9 0.649 0.609 0.609 odr rad/sec . . . . .

0.522 0 . 7 4 1 0.562 0.805 0.079 : d . . . . . . . . . .

0.450 0.266 0.259 0.064 ; , j w d r rad/sec . . . .

0.235 0.197 0.146 sec . . . . . . .

1 3 . 4 7 8.79 1 1 . 5 8 7.83 2.5 )/f3 . . . . . . . . .

2.10 0.80 0 . 7 1 Roll-control parameters aAutothrottle on.

TABLE V.- Concluded [Approach speed was 153 knots] (b) Powered-lift concept Satisfactory Acceptable HSAS SCAS Modified None criterion criterion SCAS (a) (a) (a 1

I

I Short-per i d mode

I

I 0.756 See figure 38 See figure 38 0.185 0.962 0.756 wSp, rad/sec . . . . .

_ _ _ _ _ _ _ _ _ _ _ _ _ -_-----------

34-70 34.70

Pspr sec . . . . . 45.14 9.09

0.971 0.35 to 1.30 0.25 to 2.00 0.971

TSP . . . . . * . . . 0.259 0.697

0.550 See figure 38 See figure 38 2.25 0.550

LQ/wsp . * . . . . . . 0.432

3.34 See figure 38 See figure 38 n/a, g units/rad . . . 3.34 3.34 3.34

Long-per iod (aperiodic) mode

I ------------- --El

t2, sec . . . . . . .I 2.98 I 90.94 1

m l Long-per iod (periodic) mode

I

rad/sec . . . . . ----- 0.066 0.066 0.066

----- 110.32 106.12 106.12

Roll mode

S3.0 sec . . . . . . .I 1.017 1 0.351 I 0.238 -- 1 0.218 [

Spiral mode

sec . . . . . .I128.6 I 178.21

ml Dutch roll mode 0.685 0.524 0.617 0.528 wd, rad/sec . . . . .

. . . - 0.051 0.205 0.269 0.254

0.035 0.107 0.166 0.134 <dud, rad/sec . . . .

Pdr sec . . . . . . . 9.18 12.24 10.57 12.01

+/B . . . . . . . . . 2.0 1.1 = O = O

. .

Roll-control parameters

I

w$/wd . . . . . . . 0.695 0.908 0.994 0.996

I <+/<a . . . . . . I 4.651 1.161 0.9701 0.961

aAutothrottle on.

P a , TABLE V I . - CONTROL RESPONSE CHARACTERISTICS OF SIMULATED SUPERSONIC CRUISE TRANSPORT AIRPLANES [Approach speed was 153 knots] (a) B a s e l i n e concept r Augmentation S a t i s f a c t o r y Acceptable Parameters None HSAS SCAS Modified c r i t e r i o n c r i t e r i o n SCAS (a) ( a ) (a) L o n g i t u d i n a l

OmaxI rad/sec2 . . . . b-O. 06 b-O. 05

b-O. 08 b-0.05 ' 0 .

o/oss . . . . . . . . .

See f i g u r e 40 See f i g u r e 40 I Aan/6, g/deg/sec2 . . .

See f i g u r e 1 5 Later a1 0.190 See f i g u r e 41 0.190 See f i g u r e 4 1 : .

---_------_--

Qmaxl deg/sec . . . . . 14.94 9.3

15.7 19.9 See f i g u r e 42 0.992 20.60 p2/p1 . . . . . . . . . -0.155 0 . 8 0 3 0.940 20.25, 0.015 See f i g u r e 43 0.801 0.012 0.011 See f i g u r e 43 $osc/@av 2.9

52.5 t$=30 o sec . . . . . . 2.9 4.0 2.7 53.2

~~~~ aAuto t h r o t t l e on.

h i n i m u m demonstrated speed o f 125 knots.

TABLE V1.- Concluded [Approach speed was 153 knots] (b) Powered-lif t concept Augmentation Satisfactory Acceptable

Parameters i-

criterion Longitudinal

. . . . b-0.13

Gmaxr rad/sec2 b-0.10 Same as b-O. 08

b-O. 05 S C A P 0 .

-__--_

o/oss . . . . . . . . . ------ See figure 40

See figure ,40

------

ban/6, g/deg/sec2 . . . ------ See figure 15

See figure 15 Lateral 0.433 0.470 0.400

imaxr rad/sec2 . . . . 0.299 See figure 41 See figure 41

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

>30 25.9 25.0 17.0

$maxr deg/sec . . . . . See figure 42

0.792 0.196 0.988

p2/p1 . . . . . . . . . 0.995 20.60 LO. 25

0.531 0.037 0.004 -0 See figure 43 See figure 43 $osc/@av 9 9 + 1.78 2.07 2.54

t$=300r sec . . . . . . 1.88 42.5 63.2

aAutothrottle on.

hinimum demonstrated speed of 125 knots.

cn

-- - 40.39

(132.50)

89.92 (295.00) - , -

Figure 1 . - Boeing model 969-336C based on NASA SCAT-15F configuration.

All linear dimensions are in meters (feet).

1 , I II I /7

k 7 0 0 (88.59)

I *

-

46.37 (152.13)

38.66

P (126.83) - - - I

i W I NG REFERENCE PLANE

39.32 (129.00)

152.50) 96.01 '(315.00)

TLT

(34.17 1

Figure 2 . - Baseline supersonic cruise transport simulated.

All linear dimensions are in meters (feet).

IU

38.66 (126.83) --l

r CENTER OF -

-

G RAV I TY AT .66C

W I N G REFERENCE PLANE 4

.- nn

- 96.01 (315.00)

Figure 3.- Powered-lift supersonic cruise transport simulated.

All linear dimensions are in meters (feet).

1.0

-

K 1 . 9 I I / I I I A- I . - . 8 f Rudder d e f l e c t i o n , deg 1.00

\q = o o

. 9 8 K2 . 9 6 . 9 4 . 9 2 1 5 0 2 0 0 5 0 1 0 0

E q u i v a l e n t a i r s p e e d , k n o t s

(a) Flexibility effects and rudder effectiveness for baseline concept.

4 . - Flexibility effects on lateral and directional control Figure effectiveness.

I I I L J . 8 0 1 0 2 0 3 0

t Rudder d e f l e c t i o n , deg

1.00 O 0

16, I

. 9 8 K 2 . 9 6 . 9 4 2.5' I. , I . . I . I I . I . 9 2 0 5 0 1 0 0 1 5 0 2 0 0 0 5 0 1 0 0 1 5 0 2 0 0

E q u i v a l e n t a i r s D e e d , k n o t s

(b) Flexibility effects and rudder effectiveness for powered-lift concept.

Figure 4.- Continued.

1 . o -

-

. 9 - K g K g S S

.8 -

L z , e l a s t i c ' 1 , r i g i d . 7 -

. 6 -

. 5 -

K g K g a a 1 I I . 4 ~ -1 = . - I . . I 1 I I I . . I

. 4 I ~ -1 = . -

1 5 0 1 6 0 1 7 0 1 8 0 1 9 0 2 0 0 1 4 0 1 5 0 1 6 0 1 7 0 1 8 0 1 9 0 2 0 0 E q u i v a l e n t a i r s p e e d , k n o t s (c) Flexibility effects on roll-control effectiveness for baseline concept.

Figure 4 . - Continued.

1 . 0 . 9 . 8 C 2 , e l a s t i c 1 , r i g i d . 7 . 6 . 4 2 0 0 1 4 0 1 5 0 1 6 0 1 7 0 1 8 0 1 9 0 E q u i v a l e n t a i r s p e e d , k n o t s (d) Flexibility effects on roll-control effectiveness for powered-lift concept.

Figure 4.- Concluded.

(a) Incremental changes in pitching-moment coefficient due to ground effects.

Figure 5 . - Incremental changes in pitching-moment, longitudinal-force, and vertical-force v) 4 coefficients due to ground effects.

. 0 2 A l t i t u d e , h l g , m ( f t ) .01 C x , g e

-. 0 1

w - 4

- . 0 2 I) (b) Incremental changes in longitudinal-force coefficient due to ground effects.

Figure 5.- Continued.

A l t i t u d e , h l g , m ( f t ) ' g e , I -

-. 1 5

I

- . 2 0 , -

- . 2 5 -

(c) Incremental changes in vertical-force coefficient due to ground effects.

Figure 5.- Concluded.

I

loo[ 80

THRUST,

percent

40 I

FLIGHT IDLE

0 1 2 3 4 5 6

TIME, sec

Figure 6.- Example of engine response characteristics.

L-78-79 Figure 8 . - Photograph of landing scene equipment and airport model.

. ...

L-7 8-8 0 Figure 9.- View of runway as seen by pilot prior to touchdown.

Q, W (a) TIFS airplane.

E L E C T R O N I C COMPONENTS O F SENSOR. F E E L . AND SERVO SYSTEMS E L E C T R O N I C COMPONENTS OF D E L - F O L L O W I N G A N 0 SPOWSE-FEEDBACK SYSTEMS D I C I T A L TAPE SAFETY P I L O T S T E S T E N G I N E E R S .UAT I ON P I L O T S , CANOPY 4CCESS TUNN (b) Layout of TIFS.

Figure 10.- Photograph and layout diagram of Total In-Flight Simulator (TIFS).

(a) Overall view of TIFS cockpit.

(b) Closeup view of TIFS instrumentation.

L-78 -82 Figure 11.- TIFS cockpit and instrument display.

LOCALIZER7

CENTER LINE'

f

GLIDE SLOPE

\ b 1 4 9 . 4 m

\\ ( 4 9 y f t )

CENTER LINE

\

\ 6 0 9 . 6 m

( 2 0 0 0 f t ) 3 . 2 k m .

( 2 s t . m i . ) Figure 12.- Sketch i n d i c a t i n g a i r c r a f t p o s i t i o n r e l a t i v e to l o c a l i z e r and g l i d e slope a t time zero.

-

CRITERION OF REFERENCE 10

BASELINE CONCEPT UNACCEPTABLE

-

f S I M U l A T E D

P I LOT

ACCEPTABLE

RATING

-

1- -

- SAT1 SFACTORY

Figure 13.- Comparison of unaugmented baseline concept w i t h criterion of reference 10.

2.4

NAU GMENTED

2.0

- . -. -. ,-. - .

1.6

DESIRABLE

P I T C H RATE,

1.2

deglsec

. 8

.4

COLUMN

I

Figure 14.- Comparison of desirable pitch rate response characteristics w i t h those of unaugmented airplane.

0 Basel i n e concept s i m u l a t e d

0 Powered - l i f t concept s i m u l a t e d

0 C o n c o r d e

0 Boeing' s 969-336C

A Boeing' s 707 prototype

Unacceptable

I I I I 1 I I I L I I I I 1 . -

0 2 4 6 8 10 12 14

..

M a x i m u m p i t c h a c c e l e r a t i o n , 8 , deg/sec

Figure 15.- Comparison of longitudinal control characteristics of simulated SCAR concepts with control requirements of reference 11.

- . - . - DESI RED RESPONSE

RESPONSE OF UNAUGMENTED CONFIGURATION

W ’ 9,

PERCENT de9

.

.

@, 0,

d e g l sec

degl sec

./-

6 t

“IT I/ .,2-1

2 * /

L O 2 4 6 8 10 0 2 4 6 8 10

TIME, sec

T I ME, sec

Figure 16.- Comparison of desired lateral-directional response characteristics and those obtained for unaugmented airplane.

6 C - + p E

D I N A L

,6

a,c

LATERAL

+ - 6

r, c +

D I RECT I ONAL

.f - 0 . 8 ’ : -

P

+

0 e -0.35 -

+

-1.3 -

a

Figure 17.- Normal operational stability and control augmentation system (SCAS).

(All control-surface deflections had 0.1-second lag due to actuator servo.)

I

-

2.4

UNAUGMENTED (PR = 7 )

I

2.0

SCAS (PR = 2 )

1.6

8,

1.2

' deglsec

.8

.4

C

C '

I

0 1

2 3 4 5 6 7 8

TIME, sec

Figure 18.- Comparison of pitch rate response for unaugmented and SCAS configurations.

PERCENT

-

L d e g 2

4 : 1 -

0 , --

I------

-

-2

n .- in-

TI ME, sec TIME, sec

Figure 19.- Comparison of lateral-directional response characteristics for unaugmented and SCAS configurations.

- +

‘ LONGITUDINAL

r 1 J I

,6

0.75

a, c

6 W

, ,6

-10

r, c

P

-

1 - 1 1 DIRECTIONAL

Figure 20.- Hardened stability augmentation system (HSAS). (All control-surface deflections had 0.1-second lag due to actuator servo.)

2.4 I?

UNAUGMENTED (PR = 7)

2.0 -

e - -

\ SCAS (PR = 2)

\

/ i ------- - - - - -

1.6 -

1.2

i i i Y

6 , deg/ sec

.8

.4

.2

1c'

6 7 8

1 2 3 4 5

TIME, sec

Figure 21.- Comparison of pitch rate response characteristics for various control systems.

W’

PERCENT

0, deglsec SCAS 9, deg/ sec

/,,, = 2)

UNAUGMENTED7 lor

*r

. I I r H S A S (tJK = ‘y /

0 /’

c - -

0 0

t 2

~-

‘0 2 4 6 8 l o

O O 2 4 6 8 10

T I ME, sec

T I ME, sec

Figure 22.- Comparison of lateral-directional response to a lateral control step input for various control systems.

9 - 8 - 7 - 6 - c, 3 Normal c g

I

2 P R = Z . O ( S C A S ) P R = 3 . 5 ( S C A S )

V P R = 4 . O ( H S A S ) P R = 6 . 5 ( H S A S ) L a a. 5 . 1 n a 5 L , + , , t , 2 5 . 0 L 5 5 6 0 6 5 70 5 0

C e n t e r - o f - g r a v i t y p o s i t i o n , p e r c e n t C

Figure 23.- Indication of effects of center-of-gravity variation on low-speed airplane performance.

W Y

PERCENT LATERAL CONTROL

REQU I RED

/ / 1 ou

LATERAL

60 CONTROL

AVAILABLE,

p e r c e n t

I I -1, n

U

STEADY 9 0 ' CROSSWIND VELOCITY, k n o t s

Figure 24.- Indication of crosswind trim capability of simulated baseline SCAR concept.

I

criterion

(Aa 1

n r m s '

13 1

g units

I I I 1 I

0 SCAR (baseline)

SJT

Ride quality

criterion

.04-

( Ref.

13 1

-

( a y ) r m s ,

g units

I

0 1 2 3

Root - mean -square gust intensity , m / sec

0 2 4 6 8 10

Root - mean -square gust intensity , ft/ sec

Figure 25.- Acceleration responses during landing approaches in various levels of gravity.)

of turbulence. (Accelerations measured at aircraft center CD

-

1.4

D

- 0 0 '

1.2

-

1.0

-

.8

cL

-

.6

-

.4

-0- BASELINE CONCEPT ( bf =40° 1

--ow- POWERED -LIFT CONCEPT ( bf=40 ; CT'O. 15)

-

.2

SOLID SYMBOL DENOTES T R I M a

I I I I I I

4 0 4 8 12 16 20

ANGLE OF ATTACK, a, deg

Figure 26.- Comparison of lift coefficient characteristics for baseline and powered-lift concepts simulated.

(a) Baseline SCAR concept.

(b) Powered-lift SCAR concept.

L-7 8-03 Figure 27.- View of runway as seen by pilot prior to touchdown.

(hlg = 30.5 meters (100 feet); y = -2.7O.I

BASELINE CONCEPT

-0-

\ \

\ --o-- POWERED -LIFT CONCEPT

\

\ S O L I D SYMBOL DENOTES T R I M a

-.001 t -

Cl PER deg

B

-.002

-.003 -

!

I

-.004

-4 0

4 8 12 16 20

a, deg

Figure 28.- Comparison of effective dihedral coefficient characteristics for baseline and powered-lift concepts simulated.

+ BASELINE CONCEPT

--O-- POWERED-LIFT CONCEPT

-04r

SOLID SYMBOL DENOTES T R I M a

A C

2, lat

.01

Figure 29.- Comparison of maximum lateral control effectiveness for baseline and powered-lift concepts.

BASELINE CONCEPT

---

POWERED - LIFT CONCEPT

7- \\\ \\\ \\\ \\\ \\\ \\\ \\\

-

LATERAL CONTROL

4- REQUIRED, /-

PERCENT

/- I / / / I #

MI N I M U M REQU I REMENT

I J, I

0 5 10 15 20 25 30 35

STEADY 9 0 ' CROSSWIND VELOCITY, knots

Figure 30.- Comparison of crosswind trim capability for baseline and powered-lift concepts.

15 -

10 -

I a = O

5 -

FL I GHT- PATH

0 -

ANGLE, Y I deg

-r; c /

= 16'

-

- 10

-

- 15

-

- 20

I I I I I I I

0 .2 4 6 8 1 .o 1.2 1.4

TRIMMED LII? COEFFICIENT, cL,TRIM

Figure 31.- Indication of e f f e c t s of engine t h r u s t on trim l i f t c o e f f i c i e n t and f l i g h t - p a t h angle for powered-lift concept.

'r 'r

y'deg -5 O I r ' d q -5 -10 -

-12 I

- ' O r - loo

600 -

T, kN 500- -0 Y . fi 400 - T , Ib

300 -

-30 - - -100 200 - I I I I I -60 I 1-200 100- I I I I I I 0 20 40 60 80 loo 120 1 4 0 0 20 40 60 80 loo 120 1 4 0 Time, sec Time, sec (a) Baseline concept.

Figure 32.- Indication of lateral and vertical excursions experienced following failure of number four engine.

(Engine failed at t = 55 seconds.)

-5 L 10 - 5 - 6- -6- -12 30 ,- \

I

-30 t I-'"

+ I I I I , J-2m

0 20 40 60 80 1 w ,120 140 80 1 w 120 140 0 20 40 60 Time, sec Time, sec (b) Powered-lift concept.

Figure 32. - Concluded.

6 9 8

w 4 cg '

de9

g unitso

vy P ,

degl sec

I / P 9

de9

1,

-2

0'

" 0 2 4 6 8 10 2 4 6 8 10

TIME, sec

TIME, sec

Figure 33.- Calculated l a t e r a l response to a wheel step input w i t h SCAS operative.

a@-

% _

ITY

AIRCRAFT CENTER OF GRAV

L4.56 (14.88)

/

/

BOEING 747

P I LOT'S EYE P O S I T I O N

AIRCRAFT CENTER OF GRAVITY

L 4.78 ( 15.70 1

NASA SCAR CONCEPT

Figure 34.- Comparison of pilot location relative to airplane center of gravity for simulated SCAR and Boeing 747 airplanes. All linear dimensions are in meters (feet).

Q) W

g u n i t s L

-. 1

degl sec

p ' 'n 0

(aY)Ps3*1 r r r '

2 - g u n i t s 0

p ' 0 ,

de9

i- 6) ,

-2 - g u n i t s O

p (3 , ~

TIME, sec TIME, sec

Figure 35.- Calculated l a t e r a l response to a wheel step input w i t h HSAS operative.

.lr

r

q units

"

B A S I C SCAS -.1L

P ,

degl sec

...... ........................

g units

-2 g units0

................

units I

- 1 I I I I I i I 1 I J

10 2 4 6 8

-0 2 4 6 8 10

TIME, sec

TIME, sec

Figure 37.- Comparison of lateral response to a wheel step input for SCAS and modified SCAS configurations.

0 SCAS 0 HSAS Ooen symbols denote baseline conceot . . . .

Solid symbols denote powered -lift concept Solid symbols denote powered -lift concept Unacceptable Unacceptable Acceptable uSD, rad/sec / / / / Unacceptable // / / .3L I I I I I I l l ! I I I I 1 1 1 1 1 10 100 n l a , g u n i t s l r a d Longitudinal short-period frequency requirements of reference 12.

(Unaugmented configurations fell outside of plotted range.)

1. 2

r

0 SCAS Unacceptable 0 HSAS

1. G I

Open symbols denote baseline concept Solid symbols denote powered - l i f t concept IIIIIIIIII] Subsonic jet transports Acceptable La I w per rad . 2 I . 1 1 - I 1 1 I 1 I I I 1 I I I 1 0 . 2 . 4 . 6 . 8 1.0 1.2 1.4 1 . 6 ' Short - period damping r a t i o , < S P (b) Shomber-Gertsen longitudinal handling qualities criteria of reference 16.

(Unaugmented configurations fell outside of plotted range.)

Figure 38.- Longitudinal handling qualities criteria.

0 SCAS ( P R = 2 )

0 HSAS ( P R = 4 )

NO SAS ( P R = 7 )

. . . 0 ....

Subsonic jet t r a n s p o r t s

Is denote baseline concept

0 Pe

Solid symbols denote powered - l i f t concept

The location of t h i s l i n e

/-

- subject to f u r t h e r research

/

Acceptable a u g mented

e

a rea

I I 1

.1 . 2 . 3 . 4 . 5 . 6 . 7 . 8 . 9 1.0 1. 1

Longitudinal s h o r t -period damping ratio , SP Figure 39.- Longitudinal short-period criterion for transport aircraft.

Boundaries from reference 17.

3.5 -

I

3 . 0

- - -

Baseline concept

2.5

a a a a a Powered- lift concept

2.0

1.5

1.0

. 5

0 1 2 3 4 5 6 7 8 9 10

Time , sec

Figure 40.- Low-speed pitch rate response criterion of reference 18.

Boundaries for normal operation (PR 5 3.5).

0 SCAS

0 HSAS

0 N o S A S

A Modified SCAS

Open symbols denote baseline concept Solid symbols denote powered - l i f t concept

\ Satisfactory \ \

Roll power # ..

0 * 2

rad 1 sec

. 01

10.0 .10 1.0 Roll mode time constant, T~ , sec Figure 41.- Roll acceleration response boundaries for large aircraft.

Boundaries from reference 12.

0 SCAS

I

-

c3 HSAS

0 No SAS

~-

A Modified SCAS

- 0 Subsonic jet transports

Open symbols denote baseline concept

50 -

Solid symbols denote powered - l i f t concept

- .cc'r

- DC - 8

0 Acceptable area

. .

-

-

- -

L

- @

E

2 0 -

0 Unacceptable area

-

\ \ \ \ \ \

10 -

0 B -52

-

< I I I i I I I I I I I I I 1 I I I I 1

0 S C A S .

0 H S A S

0 N o S A S

A Modified SCAS

Open symbols denote baseline concept

Solid symbols denote powered - l i f t concept

Unacceptable

: -

--

-120 ~~ -160 -200 -240 -2 80 -320 -360

-40 -80

Figure 43.- Bank angle oscillation limitations of reference 12.

.28 c

-

- -

- Modified SCAS

.24 I C

-

.20

1- c

.,o s

-

.16

-

I

.12

Unacceptable

k’

Acceptable

.04

I

1 I I I I I I I I I I I 1 I I I I

0 20 40 60 80 100 120 140 160

Pilot location relative to a i r c r a f t center of gravity, i , ft

W W Figure 44.- Peak values of compared with criterion of reference 11.

(ay)ps . .- ___ 2. Government Accession No. 1. Report No. 3. Recipient's Catalog No NASA TP-1240

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4. Title and Subtitle 5. Report Date GROUND-BASED AND IN-FLIGHT SIMULATOR STUDIES OF LOW- J u l y 1978 SPEED IIANDLING CHARACTERISTICS OF 7 3 7 0 SUPERSONIC 6. Performing Organization Code CRUISE TRANSPORT CONCEPTS _ .

8. Performing Organization Report No, 7. Author(s1 W i l l i a m D. Grantham, Luat T. Nguyen, P e r r y L. Deal, M. J. Neubauer, Jr., P a u l M. Smith, L-12165 and Major F r e d e r i c k D. Greggry 10. Work Unit No.

9. Performing Organization Name and Address 74 3-0 4-13-0 1 NASA Langley Research C e n t e r 11. Contract or Grant No, Hampton, VA 23665 13. Type of Report and Period Covered T e c h n i c a l Paper 12. Sponsoring Agency Name and Address N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n 14. Sponsoring Agency Code Washington, DC 20546 . ..

15. Supplementary Notes W i l l i a m D. Grantham, Luat T. Nguyen, P e r r y L. Deal, and M. J. Neubauer, Jr.: Langley Research C e n t e r , Hampton, V i r g i n i a .

P a u l M. Smith: Vought C o r p o r a t i o n Hampton T e c h n i c a l C e n t e r , Hampton, V i r g i n i a .

Major F r e d e r i c k D. Gregory: H e a d q u a r t e r s A i r F o r c e Systems Command, a s s i g n e d to Langley Research C e n t e r .

- .. _ _ _ - - ____ 16 Abstract Fixed ground-based and i n - f l i g h t s i m u l a t o r s t u d i e s have been conducted to determine t h e low-speed f l i g h t c h a r a c t e r i s t i c s of t w o advanced s u p e r s o n i c c r u i s e t r a n s p o r t c o n c e p t s - c o n v e n t i o n a l and powered l i f t . The p r i m a r y p i l o t i n g t a s k s were approach and l a n d i n g .

The r e s u l t s o f t h e s t u d i e s i n d i c a t e d t h a t t h e t r a n s p o r t c o n c e p t s had unaccept- a b l e low-speed h a n d l i n g q u a l i t i e s w i t h no augmentation, and that i n o r d e r to a c h i e v s a t i s f a c t o r y h a n d l i n g q u a l i t i e s , c o n s i d e r a b l e augmentation w a s r e q u i r e d . The a v a i l a b l e r o l l - c o n t r o l power was found to be i n a d e q u a t e to meet e x i s t i n g crosswind- l a n d i n g r e q u i r e m e n t s for t h e c o n v e n t i o n a l c o n c e p t ; b u t r o l l c o n t r o l w a s a c c e p t a b l e f o r t h e p o w e r e d - l i f t c o n c e p t . The r e s u l t s also i n d i c a t e d t h a t a d d i t i o n a l r e s e a r c h is r e q u i r e d to o b t a i n s a t i s f a c t o r y r i d e q u a l i t i e s w h i l e m a i n t a i n i n g s a t i s f a c t o r y h a n d l i n g q u a l i t i e s f o r e i t h e r o f the s i m u l a t e d s u p e r s o n i c c r u i s e t r a n s p o r t con- cepts a t l o w speeds.

18. Distribution Statement 7. Key Words (Suggested by Author(s)l

Powered l i f t U n c l a s s i f i e d - Unlimited S u p e r s o n i c t r a n s p o r t

Ride q u a l i t i e s Handling q u a l i t i e s S i m u l a t i o n Approach and l a n d i n g S u b j e c t Category 0 - 22. R i c e ' 20. Security Classif. (of this page) 21. NO. of Pages 9. Security Classif. (of this report1

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