The effects of lateral aerodynamic uncertainties on the handling qualities of the space shuttle orbiter at Mach numbers of 1.5 and .6

NASA Technical Memorandum 83193

(NASA-TH-83 193! TJE EkE E C I S OF LATERAL N 8 2 - 1 lQ51 AEBODYNA8IC UYCEBTAIYTXES C N S H E H A N D L I N G Q U A L I T I E S OF THE S P A C E SHUTTLE ORBITER A T 4ACH YUlYBEHS OF 1.5 A N D - 6 ( N A S A ) 4 i p Unclas HC AJ3/ME A 9 1 CSCL 01C 63/05 08177

THE EFFECTS OF LATERAL AERODYNAM l C

UNCERTAINTIES ON THE HANDLING QUALITIES

OF Tc(E SPACE SHUTTLE ORBITER AT M A C H

NUMBERS OF 1.5 AND .6

Lawrence W. Brown

September 1981

Nat~onal Aeronautics and Space Aljmlnlstrat~on Lsngley R-rch Center Hampton Vlrgln~a 23665

THE EFFECTS OF LATERAL AERODYNAMIC UNCERTAINTIES ON THE HANDLING Q U A L I T I E S OF THE SPACE SHUTTLE ORBITER AT MACH NUFBERS OF 1.5 AND .6 Lawrence W . Brown The e f f e c t s o f aerodynamic u n c e r t a i n t i e s on t h e handling q u a l i t i e s o f t h e space s h u t t l e o r b i t e r were investigated w i t h t h e use o f s ix-degree-of- freedom, nonl inear equations o f motion on the h y b r i d computer system. F l i g h t condi t i o n c h a r a c t e r i s t i c s f o r Flach numbers o f 1.5 and .6 f o r the nominal and o f f nominal angle o f attack conditions were selected f o r t h i s i n v e s t i g a t i o n . Results revealed t h a t a t the low Mach number c o n d i t i o n ( M = .6) o n l y a few problems e x i s t e d f o r the angle of a t t a c k range and the many combinations o f l a r g e aero- Moreover, none o f these problems were consid- dynamic w r i a t i o n s considered.

For the angle o f a t t a c k ered t o be r e l a t e d t o poor hand1 i n g q u a l i t i e s .

conditions considered a t the high Nach number ( M = 1 . 5 ) , problems e x i s t e d w i t h reduction o f r o l l r a t e which can r e s u l t i n r o l l reversal conditions. I n many cases, s i d e s l i p hecarrie proverse and increased rudder d e f l e c t i o n s and yaw j e t s were required.

INTRODUCTION F l i g h t simulations are necessary i n the planning and d i r e c t i n g of f l i g h t t e s t programs f o r experimental and research-type a i r c r a f t . An accurate simula- t i o n o f the a i r c r a f t ' s motion response t o c o n t r o l i n p u t s necessitates a complete con~pi l a t i o n o f the c h a r a c t e r i s t i c s o f the aerodynamic d e r l v a t i ves which are i n d i c a t i v e of the actual a i r c r a f t . I n p r e d i c t i n g the aerodynamic d e r i v a t i v e s f r o m wind tunnel r e s u l t s , the accuracy of the p r e d i c t i o n i s dependent upon the Reynolds number difference between t e s t conditions and f l i g h t ; the ma nu factor^:^ d i fferencc between the model and actual c o n f i g u r a t i o n ; a f t nominal condi t i o n s such as v a r i a t i o n i n i n g l e o f attack and a l t i t u d e ; and ~ t h e r a n a m l i e s . It i s imporiant t h a t the p r e d i c t i o n o f the aerodynamic d e r i v a t i v e s o f the space s h u t t l e o r h i t e r be even more accurate than most vehicles tested since t h i s unpowered a i r c r a f t w i l l n o t have engines t o modulate the f l i g h t conditions.

I n the p i i o t e d simulat,ion o f the space s h u t t l e o r b i t e r f o r nominal f l i g h t conditions, s t t i s f a c t o r y f l y i n g quzl i t i e s have been predicted. However, d i f f e r - ences i n the aerodynamic d e r i v a t i v e s due t o the reasons above could cause s i g n i f i c a n t discrepancies i n the p r e d i c t i o n o f the o r b i t e r resTonses. These discrepancies cauld r e q u i r e a placard on t h e center-of-gravi ty p o s i t i o n opera- range. ledd t o c o n t r o l system r a t e 1 i m i t i n g ; cause excessive f u e l usage t i o n a l i n the r e a c t i o n c n n t r o l system (RCS); cause excessive p l l c t workload; and uossibly cause loss o f a i r c r a f t .

I n r e c o g n i t i o n o f these conditions , an i n v e s t i g a t i o n was undertaken on a h y b r i d computer t o determine the e f f e c t s these aerodynamic u n c e r t a i n t i e s would have on the handling q u a l i t i e s o f the space s h u t t l e o r b i t e r . A l a r g e number o f v a r i a t i o n s i n the aerodynamic d e r i v a t i v e s were a p p l i e d t o the nominal and off nominal f l i g h t conditions a t Mach numbers o f 1.5 and .6.

SYMBOLS The aerodynamic parameters a r e referenced t o a system o f body axes w i t h t h e o r i g i n a t the v e h i c l e c e n t e r - o f - g r a v i t y (Fig. 1). The p o s i t i v e sense of the angles, forces, moments, and angular v e l o c i t i e s are a1 so shown.

l a t e r a l acceleration, g ' s *Y b wing span, m ( f t ) wing mean aerodynamic chord, m ( f t ) r o l l i n g moment c o e f f i c i e n t , Mxlgsb

c I

yawing moment c o e f f i c i e n t , Mz/gsb s i d e l f o r c e c o e f f i c i e n t , Fylgs aerodynamic forces along the x, y, z body axes, n ( l b ) F ,F ,FZ X Y 2 2 acceleration due t o g r a v i t y , mlsec ( f t l s e c ) h a l t i t u d e , m ( f t ) moments of i n e r t i a about the a i r c r a f t body axes, 1 * I , I z X Y 2 2

kg - m ( s l u g - ft )

product o f i n e r t i a about the a i r c r a f t body axes, I xz 2 2 kg - m ( s l u g - f t ) K P 1 ,KP2,KP3,KRl l a t e r a l c o n t r o l system gains KR2,KR3,KNY ,KJ1 M Mach number aerodynamic moments about the x, y, z body axes,

m - n ( f t - l b )

m a i r c r a f t mass, kg ( s l u g s ) P r o l l r a t e about the body x axis, radlsec

r o l l r a t e coamand , r a d l s e c

a i 1 eron - rudder i n t e r c o n n e c t

p i t c h r a t e about t h e body y axis, r a d l s e c 2 2 dynamic pressure, N/m ( l b / f t ) 2 2 maximum dynamic pressure, N/m ( I b / f t ) yaw r a t e about t h e z body a x i s , r a d l s e c

s t a b i l i t y a x i s yaw r a t e , r - cos o s i n 4

2 '

wing area, m ( f t " ) time, sec 2 2 commanded yaw j e t s , mlsec ( f t l s e c ) v e l o c i t y , ni/sec ( f t l s e c ) a i r p l a n e body axes, o r i g i n a t c e n t e r - o f - g r a v i t y angle o f a t t a c k , deg t r i m angle o f a t t a c k , deg angle o f s i d e s l i p , deg angle o f r o l l , deg angle o f p i t c h , deg

- d

r o l l c o n t r o l i n p u t 6, = ( a )/2, p o s i t i v e i n d i r e c t i o n e l e , t o cause p o s i t i v e r o l l r a t e , deg commanded r o l l c o n t r o l i n p u t , deg rudder d e f l e c t i o n , p o s i t i v e d e f l e c t i o n cause l e f t yawi 59 moments, deg comnanded rudder d e f l e c t i o n , deg l e f t elevon d e f l e c t i o n , p o s i t i v e f o r t r a i l i n g edge down, deg r i g h t elevon d e f l e c t i o n , p o s i t i v e f o r t r a i l i n g edge down, deg

DESCRIPTION O F VEHICLE Physical and Control C h a r a c t e r i s t i c s The space s h u t t l e o r b i t e r ( F i g . 2) c o n s i s t s of a fuselage o f 33.77 m (107.53 f t ) i n l e n g t h w i t h a 45O swept wing and a v e r t i c a l t a i l . The mass and physical c h a r a c t e r i s t i c s a r e presented i n t a b l e 1. The o r b i t e r i s a reusable space v e h i c l e which f l i e s back from n e a r - e a r t h o r b i t a l o n g a p r e s c r i b e d t r a j e c t o r y ( F i g . 3 ) f o r an unpowered l a n d i n g a t a designated a i r f i e l d .

The o r b i t e r uses a combination o f spacecraft and a i r c r a f t c o n t r o l e f f e c t o r s .

A t low dynamic pressures i t i s c o n t r o l l e d u s i n g r e a c t i o n c o n t r o l t h r u s t e r s ( l i k e a s p a c e c r a f t ) . As dynamic pressure b u i l d s up, t h e r e i s a gradual t r a n s i t i o n from using t h r u s t e r s f o r c o n t r o l t o u s i n g t h e l a r g e aerodynamic surfaces.

The primary c o n t r o l surfaces a r e t h e elevons--deflected s y m n e t r i c a l l y f o r p i t c h c o n t r o l and d i f f e r e n t i a l l y f o r r o l l c o n t r o l --and conventional rudder f o r yaw c o n t r o l . The rudder i s s p l i t t o p r o v i d e a speed brake f o r improved d i r e c - t i o n a l s t a b i 1 i ty (C ) a t hypersonic/supersonic speeds and energy management (by "8 modulating 1 i f t / d r a g r a t i o ) i n t h e subsonic region. The body f l a p i s added t o supplement t h e elevons f o r p i t c h c o n t r o l . The c o n t r o l surface l i m i t s a r e presented i n t a b l e 1.

Lateral C o n t r o l System I n t h i s study o n l y Mach numbers o f 1.5 and lower were considered f o r which t h e l a t e r a l c o n t r o l system i s described i n f i g u r e 4.

I n t h e r o l l c o n t r o l

channel, as depicted i n the block diagram, the p i l o t ' s i n p u t from a center s t i c k i s converted t o a roll r a t e comnand and sumned w i t h t h e s t a b i l i t y a x i s roll r a t e t o create an a i l e r o n conand. The a i l e r o n surface d e f l e c t i o n i s 1 i m i t e d t o + lo0. The roll c o n t r o l s i g n a l i s a l s o fed t o t h e rudder channel by an a i l e r o n - rudder interconnect.

I n thc rudder c o n t r o l channel the l a t e r a l a c c e l e r a t i o n i s f i l t e r e d and combined w i t h the s t a b i l i t y a x i s yaw r a t e t o form the rudder command. The

commanded rudder signal i s fed t o the yaw j e t s through an on - o f f switching

l o g i c . The l o g i c turns two a f t mounted j e t s on when the s i g n a l equivalent t o 40 o f rudder i s comnanded. The commanded rudder s i g n a l i s a l s o combined w i t h t h e f i l t e r e d a i l e r o n - rudder interconnect, l i m i t e d t o + 22.80, and fed t o the rudder. The l a t e r a l c o n t r o l system gains a r e presented i n t a b l e 4.

TEST PROGRAM The o r b i t e r aerodynamics data o f December 1975 were used i n t h i s ir.vestiga- These data are based on wind tunnel t e s t s using models by Rockwell t i o n .

I n t e r n a t i o n a l Space D i v i s i o n w i t h c o r r e c t i o n s f o r c o n f i g u r a t i o n changes and operational

fl i ght conditions . The f o r c e and moment coefficients, as assembled

by the o r b i t e r program o f f i c e (Ref. 1) , are based on the wing reference l e n g t h and area.

Flying Quality Criteria The f l y i n g q u a l i t y c r i t e r i a used i n t h i s i n v e s t i g a t i o n are t h a t a recommended by Donald C. Cheatham, NASA Manned Spacecraft Center, Houston, TX.

The v e h i c l e r o l l accelerations and r o l l r a t e requirements were a r e s u l t of closed loop e n t r y guidance and c o n t r o l studies d e f i n i n g these requirements i n order t o maintain the v e h i c l e t r a j e c t o r y w i t h i n acceptable dynanic pressure 1 i m i t s .

The r o l l r a t e response c r i t e r i o n used has been defined i n terms o f a r o l l r a t e response envelope and i s presented i n f i g u r e 5 f o r t h e region o f i n t e r e s t f a r t h i s i n v e s t i g a t i o n . The roll r a t e response due t o a step r o l l r a t e comnand of 5 deg/sec s h a l l f a l l w i t h i n t h e response envelope presented. I n addition, the c r i t e r i o n o f l i m i t i n g the s i d e s l i p t o l e s s than 20 d u r i n g a change i n r o l l a t t i t u d e s o f up t o + 45O was a l s o used. Responses were judged unsatis- factory if r o l l r a t e was outside the envelope o f f i g u r e 5 and/or t h e sides? i p was greater than ZO.

Flight Conditions F l i g h t c o n d i t i a n c h a r a c t e r i s t i c s f o r Mach numbers o f 1.5 and .6 were obtained f o r t h e proposed nominal t r a j e c t o r y , f i g u r e 3. O f f nominal c o n d i t i o n s were computed f o r the proposed maximun t r i m at u n c e r t a i n t i e s o f i 4O.

The c o n d i t i o n at = -4' could n o t be obtained because o f t h e l i m i t s on the maximum

dynamic pressure (qm = 400 1 b / f t 2 ) i n t h e h y b r i d computer program. To provide as l a r g e a v a r i a t i o n o f angle o f a t t a c k as p o s s i b l e w i t h o u t exceedin the dynamic pressure l i m i t o f t h e program, angles o f a t t a c k o f 3.5 and 3 were chosen for the lower boundary off nominal f l i g h t c o n d i t i o n f o r Mach numbers o f 1.5 and .6, respectively. The f l i g h t c o n d i t i o n c h a r a c t e r i s t i c s f o r t h e nominal and o f f nominal t r i m angles of a t t a c k are presented i n t a b l e s 2 and 3.

Aerodynamic Uncertainties V a r i a t i o n between wind tunnel and f l i g h t aerodynamic d e r i v a t i v e s has been noted i n e x i s t i n g a i r c r a f t ; and, i n many instances the d i f f e r e n c e s are q u i t e subs t a n t i a1 as i n d i c a t e d by Major General Thomas S t a f f o r d (AFFTC/Doy ) and J. Wiel (DFRC). These d i f f e r e n c e s coula cause s t a b i l i t y and c o n t r o l problems and are, therefore, a ccncern i n evaluating t h e handling q u a l i t i e s o f t h e space s h u t t l e o r b i t e r .

Wind tunnel and f l i g h t d e r i v a t i v e s were c o r r e l a t e d f o r a l a r g e number o f vehicles; and, a comparison o f maximum v a r i a t i o n s i n t h e d e r i v a t i v e s f o r conven- t i o n a l a i r c r a f t and 1 i f t i n g bodies was obtained. Based on s t a t i s t i c a l consider- a t i o n , the range o f u n c e r t a i n t i e s i n aerodynamic d e r i v a t i v e s was established.

The recommended increments o f the l a t e r a l d e r i v a t i v e are presented i n t a b l e 4.

A l a r g e number o f v a r i a t i o n s , i n s i n g l e and m u l t i p l e combinations, were made i n the aerodynamic d e r i v a t i v e f o r the augmented c o n f i g u r a t i o n .

The responses were viewed on a C R T screen and the ones of i n t e r e s t were recorded on a s t r i p c h a r t recorder. The v a r i a t i o n s included w i t h i n these recorded cases a r e of the same ma n i t u d e as t h e predicted u n c e r t a i n t i e s of t a b l e 5 i n many cases; i n some cases 9 i .e. , B d e r i v a t i v e s ) , the v a r i a t i o n s are as l a r g e as 200 percent 3f the predicted u n c e r t a i n t i e s . It was discovered t h a t l a r g e v a r i a t i o n s i n the r o t a r y and s i d e s l i p d e r i v a t i v e s alone have very l i t t l e e f f e c t on the responses.

The s j d e s l i p d e r i v a t i v e s showed some s i g n i f i c a n c e i n combinations w i t h the c o n t r o l d e r i v a t i v e s and are, therefore, included i n the cases o f i n t e r e s t .

A compi 1 a t i o n o f the selected cases o f aerodynamic d e r i v a t i v e incremental changes i s presented i n t a b l e 6. The basic and r e s u l t a n t (addinq t h e v a r i a t i o n i n tab1 e 6 ) values of the aerodynamic c h a r a c t e r i s t i c s f o r the c o n f i g u r a t i o n s i n v e s t i g a t e d are presented i n t a b l e 7. Table 7 a l s o summarizes the r e s u l t s o f each c o n f i g u r a t i o n tested.

RESULTS AND DISCUSSION Unaugmented Configuration I A d i g i t a l computer program was used t o compute t h e l a t e r a l response f o r a negative 20 a i l e r o n i n p u t for the unagumented o r b i t e r a t Mach numbers 1.5 and .6 f o r the nominal angles o f a t t a c k (a = 6.7O and u = 4.4O) and the o f f nominal angles o f a t t a c k (a = 3.S0 and a = 3 . 0 ~ f o r t h e lower boundary; o = 10.8O and a = 8.5O f o r the upper boundary), respectively, t o i l l u s t r a t e t h e need f o r s t a b i 1 i t y augmentation. The unagumented o r b i t e r responses a t Mach 1.5 show a

tendency f o r a r o l l r e v e r s a l c o n d i t i o n a t the higher angle o f a t t a c k (a = 10.80) There i s an appreciable w i t h a l a r g e adverse s i d e s l i p ( B > 2'). f i g u r e 6.

amount o f i n t e r a c t i o n between the Dutch r o l l and s p i r a l modes. The r o l l r a t e and the s i d e s l i p suggest d i f f i c u l t y i n c o n t r o l l i n g the bank angle a t the t h r e e angles o f attack.

A t Mach number .6, the response shows t h a t f o r t h e lower angles o f a t t a c k (CY = 3.00 and a = 4.4O) the r o l l r a t e reaches 35 deg/sec i n l e s s than f i v e seconds w i t h proverse s i d e s l i p o f about 2O, f i g u r e 7.

Augmented Configuration A h y b r i d computer system was programed w i t h six-degree-of-freedom, nonlinear equations o f motion t o i n v e s t i g a t e the e f f e c t s o f the aerodynamic u n c e r t a i n t i e s on the f l y i n g qua1 i t i e s o f t h e augmented o r b i t e r . The configura- t i o n was augmented w i t h the l a t e r a l c o n t r o l system o f 1975, f i g u r e 5. Time h i s t o r y responses are obtained f o r a r o l l r a t e command o f 5 deg/sec f o r t h e nominal and o f f nominal angle o f a t t a c k conditions. For t h e Mach number and angle of a t t a c k conditions considered, the responses show the r o l l r a t e i s t y p i c a l o f a f i r s t order system, f i g u r e 8. S i d e s l i p angles are small (6 < .5O).

A t the higher angle o f a t t a c k conditions (a = 1 0 . 8 ~ and a = 8.5O f o r M = 1.5 and M = .6, r e s p e c t i v e l y ) there i s an increase i n t h e a i l e r o n and rudder d e f l e c t i o n s .

A t Mach 1.5, the r o l l r a t e response suggests sluggishness f o r t h e higher angle of attack condition.

Variations i n A i 1 eron and Rudder D e r i v a t i v e s The responses f o r the configurations i n v e s t i g a t e d i n t h e study (Table 7) were compared t o the responses of the augmented vehicle, f i g u r e 8.

The e f f e c t s o f the a i l e r o n and rudder c o n t r o l d e r i v a t i v e u n c e r t a i n t i e s are presented i n f i g u r e 9. Configuration 1, f i g u r e 9a, shows a reduced r o l l r a t e w i t h some acceleration. Further reduction i s seen i n the r o l l r a t e w i t h an increase i n proverse s i d e s l i p , a i l e r o n , and rudder d e f l e c t i o n w i t h an increase i n angle of attack (compare Fig. 9a, a = 6.7O, a = 10.8O). Reversal o f the r o l l r a t e command increases the demand f o r rudder d e f l e c t i o n which requires a c t i v a - t i o n o f the yaw j e t s , as i n d i c a t e d by the data i n t h e rudder channel a t t h e higher angles o f attack.

Configuration 2, f i g u r e 9b, shows o n l y small o r no effects on the response due t o aerodynamic u n c e r t a i n t i e s a t the low angle of attack.

For c o n f i g u r a t i o n 2c (compare Fig. 9b, CI = 8.50), s i d e s l i p becomes proverse and the demand on the rudder d e f l e c t i o n requires yaw j e t a c t i v i t a t i o n w i t h r o l l r a t e comnand i n p u t .

There i s also an increase i n yaw j e t a c t i v i t a t i o n w i t h reversal o f the command.

For c o n f i g u r a t i o n 3, f i g u r e 9c, an unsatisfactory c o n d i t i o n e x i s t s a t t h e lower angle o f a t t a c k (3a, n = 3.5O).

Even before the r o l l r a t e command i n p u t was applied, a r o l l r a t e developed, i .e., the v e h i c l e began t o r o l l v o l u n t a r i l y .

This i s why there i s an i n i t i a l r o l l r a t e when the c o n t r o l i n p u t was applied.

Upon a p p l i c a t i o n o f t h e c o n t r o l input, the r o l l r a t e s t a r t s i n t h e r i g h t d i r e c t i o n b u t immediately turns around i n d i c a t i n g a r o l l reversal c o n d i t i o n .

Configuration 3b (a = 6.7O) shows a reduction i n the r o l l rate, compared w i t h f i g u r e 8, w i t h some o s c i l l a t i o n and proverse s i d e s l i p . With reversal o f the r o l l r a t e comnand, rudder d e f l e c t i o n s increase and the yaw j e t a c t i v a t i o n i s required. For c o n f i g u r a t i o n 3c, (a = 1 0 . 8 ~ ) . rudder demand i s high r e q u i r i n g yaw j e t s w i t h i n i t i a l r o l l rcite command i n p u t . With the reversal o f t h e conmand, s i d e s l i p i s l a r g e ( 6 2O) and the rudder d e f l e c t i o n s become excessive ( 6 r > l o 0 ) w i t h increased yaw j e t a c t i v a t i o n required.

For configuration 4, f i g u r e 9d, o n l y a l i t t l e change from f i g u r e 9b i s rtT:tmd . ~ t the high angle o f a t t a c k ( a = 8.50) where the rudder d e f l e c t i o n i ~~c::*tiui-.c~ r e q u i r i n g the yaw j e t s .

A i l e r m and rudder s i d e f o r c e u n c e r t a i n t i e s were considered along w i t h t h e r o l l i n g and yawing moments f o r the high Mach number and are presented i n figure IG. Configuration 5a, (a = 3.50) f i g u r e 10a, shows a s l i g h t increase i n tfie ro'!f r a t e , compared t o f i g u r e 8, w i t h some o s c i l l a t i o n and a small proverse sidesl i p . For c o n f i g u r a t i o n 5b ( a = 6.70) the yaw j e t s were n o t allowed t o f i r e . For t h i s c o n f i g u r a t i o n there i s an increase i n r o l l r a t e o s c i l l a t i o n s , proverse s i d e s l i p and rudder d e f l e c t i o n s . Upon reversal o f the comnand, there i s an increase i n s i d e s l i p ( 6 = l o ) and rudder d e f l e c t i o n s ( 6 r 2 6O). For c o n f i g u r a t i o n 5c ( a = 1 0 . 8 ~ ) w i t h r o l l r a t e comnand i n p u t , s i d e s l i p increases and the demand f o r rudder d e f l e c t i o n s greater than 40 requires yaw j e t s . Upon reversal o f the command, l a r g e rudder d e f l e c t i o n s are required ( 6 r > SO) along w i t h the yaw j e t a c t i v a t i o n and s i d e s l i p i s l a r g e ( B = 1.80).

For c o n f i g u r a t i o n 6, f i g u r e lob, compare w i t h f i g u r e 8a, there i s a reduction i n the r o l l r a t e w i t h some o s c i l l a t i o n a t the lower angle o f a t t a c k ( 3 = 3.5O). Configuration 6b shows an u n s a t i s f a c t o r y c o n d i t i o n i n which r o l l r a t e has been reduced t o zero. The v e h i c l e would n o t r o l l w i t h almost constant a i l e r o n and rudder d e f l e c t i o n f o r t h i s r o l l r a t e comnand i n p u t . Configura- t i o n 6c ( a = 10.80) would be u n s a t i s f a c t o r y i n r o l l because o f t h e reduced r o l l r a t e . There i s an increase i n a i l e r o n and rudder d e f l e c t i o n w i t h reversal o f the r o l l r a t e command.

V a r i a t i o n i n A i 1 eron, Rudder, and Sides1 i p Derivatives The s i d e s l i p d e r i v a t i v e u n c e r t a i n t i e s had l i t t l e o r no e f f e c t on the l a t e r a l responses alone; therefore, they were included w i t h the a i l e r o n and rudder d e r i v a t i v e u n c e r t a i n t i e s and are presented i n f i g u r e s 11 and 12.

Configuration 7, f i g u r e I l a , shows an increase i n s i d e s l i p o s c i l l a t i o n s and rudder d e f l e c t i o n s w i t h yaw j e t s required as angle o f a t t a c k increases. For c o n f i g u r a t i o n 7c ( a = 10.8O), the demand on the rudder d e f l e c t i o n i s excessive ( 6 r . 11') w i t h added requirement on the yaw j e t s and an increase i n s i d e s l i p w i t h o s c i l l a t i o n s .

The r o l l r a t e shows o n l y a small change from the augmented conditions (Fig. 8a).

Configuration 8, f i g u r e l l h , shows o n l y a s l i g h t increase i n the r o l l r a t e , s i d e s l i p, and rudder d e f l e c t i o n s , compared t o f i g u r e 8b.

The aerodynamic u n c e r t a i n t i e s have very 1 i ttl e e f f e c t on the response f o r t h i s c o n d i t i o n .

The sideforce dwivstives were considered along with the rolling and yawing moment derivative: and the results are shown in figure 12. For configuration 9a, ( a = 3.5O) figure 12a, the roll rate i s reduced with roll reversal tecdencies and proverse sideslip for an unsatisfactory condition.

The aileron and rudder deflections are almost constant. Configuration 9b (a = 6.7O) shows sidesl ip and rudder deflection increases requiri ng yaw j e t s with roll rate command inputs. Upon reversal of the comnand, sidesl i p and rudder deflection became large ( 0 = ZO, Sr > 8O) increasing the requirements Configuration 9c ( a = 10.8O) shows an unsatisfactory condition for yaw jets.

where the sideslip indicates a n aperiodic mode. Roll rate i s reduced with roll

reversal tendencies and rudder deflection increases . Upon reversal of the

command, roll rate i s nulled, the rudder deflection i s divergent, and the sidesl ip is 1 imi ted.

Configuratior! 10, figure 12b, stlows l i t t l e or no effect due to aerodynamic uncertainties. Configuration 11, figure 1 Zc, shows unsatisfactory responses for a l l three angles of attack. The roll rate i s restrained with constant aileron and rudder deflections a t the lower angle of attack (a = 3.5O). The sideslip, the reduced roll rate with roll reversal tendencies, and the divergent rudder deflections indicate an aperiodjc mode for the higher angles of attack (a = 6.70 and a = 10.8O).

COfiCLUDI N G REMARKS The effects of aerodynamic tincertainties on the handling qualities of the space shuttle orbiter were invcsti~ated with the use of six-degree-of-freedom, nonl i near equations of aotion on a hybrid computer system. Flight conditions characteristic of Mach numbers of 1.5 and .6 for the nominal and off nominal angle of attack conditions ( a ' s = 3.5', 6.7', 10.8' and a ' s = 3.0°, 4.4', 8.5', respectively), were selected for this investigation. Results revealed that a t the low Wch number condition ( M = .6) only a few problems exist (i.e.!

existenie 01. proverse sideslip and an increase in rudder deflection which required the yau jets) for the angles of attacr, and the combinations of large aerodynamic variations considered , b u t n o t any that would be considered unsatisfactory. A t the higher Mach number ( M = 1.5) and angle of attack condi- tions considered. problems resulted from various cases of reduced roll rate, 1 arge value of proverse sidesl ip, and increased rudder deflections and yaw j e t activitation with initial roll rate command inputs. Unsatisfactory conditions exist consisting of roll reversal problems and increased proverse sidesl ip in addition t o long periods of large rudder deflections requiring extended use of the yaw jets. There seemed t o be an aperiodic mode developing in some instances.

R E F E R E N C E S 1. Rockwell I!iternational Space Division: Space Shutt:e Program, Aerodynamic Design Data Book, Vol. 1, Orbiter Vehicle, December 1975.

T a b l e 1. Space S h u t t l e O r b i t e r Mass and P h y s i c a l C h a r a c t e r i s t i c s !/eight, N ( l b )

I

c . g . , p e r c e n t body l e n g t h

I Moments o f i n e r t i a

I

Span, m ( f t )

2 '

Area, m ( f t " ) Cord, m ( f t )

I

Surface D e f l e c t i o n Lirni t s Elevons , deg - 3 5 , +20 Rudder, d e ~ -228, +22.8 Speed Grakes, deg - 8 7 . 2 , 0 Body F l a p s , deg - 1 1 . 7 , +16.3 --

T a b l e 2. F l i g h t Conditions C h a r a c t e r i s t i c f o r Mach Number 1.5

3 -

50,000 h, ft 70,000 62,000

-

1,452.1 1,456.35 V , f t / s e c 1,529.56 383.68 239.86 147.62

6 , #/ft2

- 3.5 6.7 10.8 a , deg -6.68 -8.24 -19.5 8, deg - --

- .0875 - .087 - .08&3

C i g , per r a d .-

- .307

-. 293 - .284

C l p , per r a d .I15 .I38 .I03 C l r . p e r r a d

-

.00619 ,019 .0375 CnB , per r a d ~. -- ~- - ,113 .I58 ,133 Cnp , p e r r a d

- .384

- ,453 - .433

C n r , per r a d --

- ,955 - .946 - .96E

Cys , p e r r a d - .0841 .0897 .081 C, 6a. per r a d - - .034 ,031 .33[17 C, 6 i p e r r a d ---

.008 . 008

- ,0036

Cndd Per r a d . - -

. ------ -- --

- .059 - .056 - .055

c , , ; p e r r a d -

- .018 -.01445

- ,0183

.- - - - .- --- - -- .I03 .lo14 #

T a b l e 3. F l i g h t Conditions C h a r a c t e r i s t i c f o r Mach Number .6 h , ft 9,000 18,000 33,000 V , f t l s e c 627.08 589.13 648.82 266.484 381 .257 j 138.234

9, # l f t

3.0 4.4 8.5 a , deg - .-- - - -18.7 -19.3 -13.8 0 , deg

- .077 - .093 -.I23

c , ~ , p e r r a d

- .282

-. 277 - .309

C , p , p e r r a d -- ,144 ,153 .I79 c p e r r a d ----- -- .095 c , ~ . p e r r a d .I09 .09 t , p e r r a d .I59 .I88 P --

-.305 - .263 - .262

, , per r a d

-1.081 - 1 .I35 -1.125 Cy,. Per r a d - .

.215 .225 .210 C , 6 a , erra ad ---- -- , p e r r a d .044 .047 .046 -- .034 .05; .031 C n s i Per ~ a d . - ~- ---

- .082

cn6; Per *ad cyd; Per *ad -.I80 -- C Y 6 + p e r r a d

- - ----

Table 4. Space S h u t t l e O r b i t e r L a t e r a l C o n t r o l System Gains Mach No. 1 . 5 1 . 5

. C . G

I 70,000 50,000 62,000 h , ft. 18,000 9,000 33,000 - 10.8 3.5 4 . 4 6.7 3.0 C . 5 , deg.

- KP1 1 1 1 1 1 1 .543 .543 KP2 ,543 .543 .658 .543 KP3 .970 .391 .563 .393 .625 ,970 KR1 5.9 5.9 5 . 9 5.9 5.9

5 * 9 1

.8SO .469 ,675 ,472 .850 KR2 .75

- .-

.O KR3 .7223 .I455 .O .3731 .O I KNY 1.11 1 . 1 1 .419 .419 1.233 .419 I .25 .25 .25 KJI .25 .25 .25 i m

I . , 4 LA 10 -f I .

! , : " ' , ' ' sf 0 0 0 \L3 03

q) C Q . q ' m < ' i

! -

, m a 0 C 0 .o10 0 i I I ' I , i I i ! ' I

" I / < . 8 8 I , '

; . ' , ;; 6 ,

4 1 : 1 . I i I . I

I - I / ' I / PAGE IS

i - --c -2 ..-..-- -.-.. - ..-,. , ..--.?, i.-L.

QUAUlY

> f , ' 4 , .-. , 4 9' I

Q I ~ C ~ I ~ I ,COI- N I ig;~lsl ' W ' 1-!01.- . .I .j I l l . 1.. i.+ 4 im~-~rn~ I l l , ;'&ld:

i 1 - 1

r - t - ! , I N i - t

R udder-speed -brake panels

I \- ; icjure 1.- System of body axi?s used. Positive directions are indicated.

PROPOSED ALTITUDE SCHEDULE WITH MACH NO.

Mach Na Figure 3.- proposed a l t i t u d e s c h e d u l e w i t h Mach number.

ROLL CONTROL I - - 6a LIMIT -

+a0 A

p cos a

-----

I ' -0 r'sin o + I !

' Pi!

i RUDDER COi~!TkGl.

I I

--,' K Jl i. -- -$ LOGIC - . - UZC

DRRC I 1 , - - - (JETS) .

I- . - - 7 A -.

1 2 1 {=,'.++?I AKR2' ~>~---.-l $2 1

= br

I r-- i,,'~

s + 2 C

I LI:.ilT i

- t

i c o s a p sin a Figure 4 . - Space shuttle orbiter lateral control aystm.

(late system)

p, deg sec Time. t, sec ( a ) Mach number 1 . 5 Figure 6 . - Response f o r t h e unagumented space s h u t t i e o r b i t e r for a two degree a i l e r o n input.

r , deg sec 2 3 4 Time, t, sec (b) Mach number 1.5 Figure 6 . Concluded.

ORiG'rNAL PAGE IS pf n.--.- n'1ALlly

Tine, t, sec

(a) Mach number .6 Figure 7.- Response for the u n a w e n t e d s p e e s h u t t l e o r b i t e r for a two degree aileron input.

C J I 1 1 0 1 2 3 4 5 6 Ti me, t, sec (b) Mach number . 6 Figcre 7. - Concluded.

k

DEG

i8

d q l sec

-

-20 m N A L PAGE I S OF POOR QUALITY lime, sec Ti.me,sec lime sec - - -L (a) Mach number 1 . 5 Figure 8.- Response for the augmented space shuttle orbiter for a five deg/sec r o l l rate command.

&r, 0 DEG PC* deg J sec

u

0 4 8 1 2 1 6 20 0 4 8 12l6N Time, sec Time, sec Time, sec ([b) Mach number .6 Figure 8.- Concluded.

P n a m 148~. h=?O,O[Ilfi

I F -

l r -

I , I I I I L 1 1 1 ~ 1 l l l ' i J

1 4 . 8 1 2 1620

0 4 8 1121620

TIME, SEC TIME, SEC ac

(.I r n b n m w I.'

W a r n J.- mr effwctm o f wulmtimm tn thr allar04 md W W mkliag en8 JnYlnr: l r a s n f s tor n rlrr drltlnw 6 *kt* c o u n d .

a4 OEG . . " I ,-r, : : :I;, p:: .L .,,, .. .I ,;- - :: ..I. ..li I.. .. ,,

"[r'V "r *,. "...,.--- ,kq,sec O r , . 1 . , J ! : ; , . I ' Tine. s~

DEGl SEC

LI

- z 5 , ; , , , , u

0 4 8 12116 2O 0 4 8 12 1620 0 4 8 12 16, 20 TIME, SIC TIME, SEC TIM, a c

v~ deql sec

AND br. DEG

o...fc-

M G I SEC

L 1

-2 5

- w

0 4 8 1 2 1 6 2 0 0 4 8 1 2 1 6 2 0 0 - 2 0 rime, SEC TIME. SEC TIME, SEC (81 m h mmbn 1 . 5 r1Fm 11.- n+ .rhr~m a? rvi.tlsn, i n 'he .ldrsltp. 81lrmn.

uld rudder mlllna end )rrin# l u n ! . .

I ORIGINAL PAGE IS

OF POOR QUALlW

DEG DEG Pc' 0 deg lsec -20 'f lme, 4ec

h a d 4 DEG I& DEG 4 DEG 4 DFC q M C AND

6% okL------ A 1 -

DEG ;

C "8 0 f a d

-. 64

0 8 16 ....., ..... ... 8 - .., . . ;_,_.../. i ..I ?.I I . I.

OFC/ S E C I , i.. / 'I '; , . . . , I... . . . a -20 L-L- 0 4 8 1216 20 TIME. S t C TIME, S E C TIME, Sf C h t . zub*r 1.3 . .,,.... f j rp-.' ,w.