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velopment of an Advanced

Pitch Active Control System

and a Reduced Area

Horizontal Tail for a

Wide Body Jet Aircraft

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Executive Summary

- Wiley A . Guinn

LOCKHEED-CALI FO RN IA COM PANY BUR BANK, CALI FORNlA CONTRACT NO. NASl-IF;q76 . .

.__ _ _ - .-- -

(BASA-CB-772283) D E B E L O F d E E ? CP A N A D V A I C E D N87- 177 1 1 € I T C € i A C I I V E CCL'IfiCL S P S T E B A L L E itET)UCED A R E A H O H I Z G N T A Z TAIL E C R A W I L E - E O D Y J E T

AI.RCBAFT E x e c u t i v e Summary, CEC, 797d - Unclas

A p r . 1983 (Lockherd-California C c , , G3/08 43374 Space Administration .

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NASA Contractor Report 172283

Development of an Advanced

Pitch Active Control System

and a Reduced Area

Horizontal Tail for a

Wide Body Jet Aircraft

Executive Summary

Wiley A. Guinn

LOCKH EED-CALI FORN IA COMPANY BUR BAN K, CALI FO RN IA CONTRACT NO. NASl-153"" general release will be three (3) years from oare indicated o n thls aocumerit.

National Aeronautics and Space Ad ministration Langley Research Center Hampton Virginia 23665 Commercial t r a n s p o r t a i r c r a f t f u e l consumption can b e s i g n i f i c a n t l y reduced by r e l a x i n g t h e l o n g i t u d i n a l s t a t i c s t a b i l i t y and by d e c r e a s i n g t h e s i z e of t h e h o r i z o n t a l t a i l . However, both of t h e s e f u e l s a v i n g c o n c e p t s u s u a l l y r e s u l t i n degraded a i r c r a f t f l y i n g q u a l i t i e s , The f l y i n g q u a l i t i e s I can b e r e s t o r e d by u s i n g a P i t c h Active C o n t r o l System (PACS) t o p r o v i d e s t a b i l i t y augmentation.

T h i s r e p o r t summarizes work t h a t was accomplished f o r t h e NASA Aircraft Energy E f f i c i e n c y program by Lockheed toward development of a n advanced p i t c h a c t i v e c o n t r o l system (NASA CR 172277) and a reduced area h o r i z o n t a l t a i l (NASA CR 172278) f o r a commercial widebody t r a n s p o r t (L-1911).

I p e r c e n t mac provided f l y i n g q u a l i t i e s e q u i v a l e n t t o t h o s e of t h e b a s e l i n e I The reduced area h o r i z o n t a l t a i l design o b j e c t i v e w a s t o d e t e r m i n e t h e I maximum d r a g b e n e f i t t h a t can be achieved by r e d u c i n g t h e L-1011 t a i l a r e a .

The program i n c l u d e d e v a l u a t i o n of f o u r s m a l l h o r i z o n t a l t a i l c o n c e p t s w i t h t planform a r e a s on 30 and 33 p e r c e n t l e s s t h a n t h e s t a n d a r d L-1011 t a i l . Pro- f i l e changes which were e v a l u a t e d included l e a d i n g edge r a d i u s , camber, t h i c k - n e s s t o chord r a t i o , and h i g h - l i f t devices. Planform changes e v a l u a t e d were t i p c o n f i g u r a t i o n , l e a d i n g edge sweep, a s p e c t r a t i o , and t a p e r r a t i o . I n c l u d e d i n t h e r e p o r t are r e s u l t s of t h e high-speed and low-speed wind t u n n e l tests.

A n a i r p l a n e d r a g r e d u c t i o n o E approximately two p e r c e n t w a s o b t a i n e d w i t h t h e b e s t s m a l l t a i l d e s i g n . However, forward c.g. l i m i t a t i o n s would have t o b e imposed o n t h e a i r c r a f t because t h e maximum h o r i z o n t a l t a i l l i f t g o a l was n o t achieved and s u f f i c i e n t a i r c r a f t nose-up c o n t r o l a u t h o r i t y w a s n o t a v a i l a b l e .

T h i s l i m i t a t i o n would n o t be r e q u i r e d € o r a p r o p e r l y d e s i g n e d new a i r c r a f t .

PRECEDING PAGE BLANK NOT FILMED’

iii TABLE OF CONTENTS

S U M M A R Y . . . . . . . . . . . . . . . . . . . . . . . . . . . iii

LISTOFFIGURES . . . . . . . . . . . . . . . . . . . . . . . v i i

LISTOFTABLES . . . . . . . . . . . . . . . . . . . . . . . i x

LIST OF SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . xi

1 . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . 1

1.1 Background . . . . . . . . . . . . . . . . . . . . . . . . 1

1 . 2 Program Objectives . . . . . . . . . . . . . . . . . . . . 2

1.3 Scope of Program . . . . . . . . . . . . . . . . . . . . . 3

2 . ADVANCED PACS DEVELOPMENT . . . . . . . . . . . . . . . . . . 5

2 . 1 C o n t r o l Law Design O b j e c t i v e s . . . . . . . . . . . . . .

C o n t r o l Law S y n t h e s i s . . . . . . . . . . . . . . . . . .

2.2

Feedback Loop Gain . . . . . . . . . . . . . . . . . . 5

2 . 2 . 1 2.2.2 Feed-Forward Loop . . . . . . . . . . . . . . . . . . .

i a

2.2.3 Primary Gain S c h e d u l i n g . . . . . . . . . . . . . . . .

2.2.4 Secondary Gain Scheduling . . . . . . . . . . . . . . .

Advanced PACS C o n t r o l Law . . . . . . . . . . . . . . .

2.2.5 2.3 F l y i n g Q u a l i t y A n a l y s i s . . . . . . . . . . . . . . . . .

2.3.1 Dynamic S t a b i l i t y . . . . . . . . . . . . . . . . . . .

2.3.2 Maneuver S t a b i l i t y . . . . . . . . . . . . . . . . . .

Speed S t a b i l i t y . . . . . . . . . . . . . . . . . . .

2.3.3 T r i m m a b i l i t y . . . . . . . . . . . . . . . . . . . . .

2.3.4 2.4 P i l o t e d F l i g h t S i m u l a t i o n Test . . . . . . . . . . . . . .

3 3 2 . 4 . 1 F l i g h t S i m u l a t o r . . . . . . . . . . . . . . . . . . .

2.4.2 S i m u l a t i o n Computer Program . . . . . . . . . . . . . .

2.4.3 S i m u l a t i o n Test C o n d i t i o n s . . . . . . . . . . . . . .

S i m u l a t i o n Test R e s u l t s . . . . . . . . . . . . . . . .

2.4.4 PACS System A r c h i t e c t u r e . . . . . . . . . . . . . . . . .

2.5

REDUCED AREA HORIZONTAL TAIL . . . . . . . . . . . . . . . . 41

3 .

V PRECEDING PAGE BLANX NOT FILMED- TABLE OF CONTENTS (Continued) . . . . . . . . . . .

Design Criteria . . . . . . . . . .

3.1 . . . . . . . . . . .

High-speed Design Criteria . . .

3 . 1 . 1 . . . . . . . . . . .

Low-Speed Design Criteria . . .

3.1.2 . . . . . . . . . . . .

3 . 1 . 3 S p e c i f i c Design Requirements . .

. . . . . . . . . . .

S m a l l T a i l C o n f i g u r a t i o n s E v a l u a t e d 3.2 S m a l l T a i l Design P r o c e d u r e s . . . . . . . . . . . . . .

3.3 H I 6 H o r i z o n t a l T a i l A n a l y s i s . . . . . . . . . . . . .

3 . 3 . 1 H 1 7 H o r i z o n t a l T a i l A n a l y s i s . . . . . . . . . . . . .

3.3.2 H18 H o r i z o n t a l T a i l A n a l y s i s . . . . . . . . . . . . .

3 . 3 . 3 H1g H o r i z o n t a l T a i l A n a l y s i s . . . . . . . . . . . . .

3 . 3 . 4 Wind Tunnel Tests . . . . . . . . . . . . . . . . . . . .

3.4 3.5 Wind Tunnel Test R e s u l t s . . . . . . . . . . . . . . . .

CONCLU S I O N S . . . . . . . . . . . . . . . . . . . . . . . .

4 .

5 3 R E F E R E N C E S . . . . . . . . . . . . . . . . . . . . . . . . .

v i L I S T O F FIGURES Page F i g u r e . . . . . . . . . . . . .

C . E . management system f u e l s a v i n g s Reduced area h o r i z o n t a l t a i l f u e l s a v i n g s . . . . . . . . . .

. . . . . . . . . . .

F l i g h t test a i r p l a n e (L-1011 S / N 1001) . . . . .

L o n g i t u d i n a l c o n t r o l system with t h e advanced PACS Schematic of t h e L-1011 c o n t r o l system w i t h t h e advanced PACS series s e r v o . . . . . . . . . . . . . . . . .

. . . . . . . . . .

PACS dynamic s t a b i l i t y d e s i g n o b j e c t i v e s PACS column f o r c e g r a d i e n t d e s i g n o b j e c t i v e . . . . . . . . .

Blended normal-acceleration/pitch-rate (C*) r e s p o n s e o b j e c t i v e . . . . . . . . . . . . . . . . . . . . .

Advanced PACS c o n t r o l l a w s y n t h e s i s . . . . . . . . . . . . . 11

Advanced PACS c o n t r o l math model i n s t a t e - s p a c e form . . . . 14

11 L-1011 s t a b i l i z e r r o t a t i o n (6~) r e l a t i v e t o c o n t r o l column displacement (&) 1 7 . . . . . . . . . . . . . . . . . .

1 2 Scheduled p i t c h rate feedback g a i n c u r v e s ,

f l a p s - u p c o n d i t i , o n s . . . . . . . . . . . . . . . . . . . . . 20

Scheduled feed-forward g a i n curves f l a p s - u p c o n d i t i o n s . . . 2 1

1 3

1 4 Advanced PACS secondary g a i n c o n t r o l l e r . . . . . . . . . . . 23

15 Advanced PACS c o n t r o l l a w block diagram . . . . . . . . . . .

1 6 Comparison of a i r c r a f t response w i t h and w i t h o u t PACS engaged f o r v a r i o u s l e v e l s of c o n t r o l column s t e p i n p u t s . . . . . . . . . . . . . . . . . . . . .

Comparison of a i r c r a f t response w i t h and w i t h o u t PACS engaged f o r a severe v e r t i c a l g u s t . . . . . . . . . . .

18 Blended normal-acceleration/pitch-rate r e s p o n s e 29 . . . . . . .

1 9 B a s e l i n e a i r c r a f t maneuver s t a b i l i t y column f o r c e

g r a d i e n t s , c r u i s e . . . . . . . . . . . . . . . . . . . . . . 30

PACS c o n f i g u r e d a i r c r a f t maneuver s t a b i l i t y column

f o r c e g r a d i e n t s , c r u i s e . . . . . . . . . . . . . . . . . . . 30

Speed s t a b i l i t y column f o r c e s , c r u i s e 32 . . . . . . . . . . . .

2 2 P i l o t e d f l i g h t s i m u l a t i o n test c o n d i t i o n s . . . . . . . . . . 34

Handling q u a l i t i e s r a t i n g s c a l e . . . . . . . . . . . . . . . 36

v i i L I S T OF FIGURES (Continued) Page F i g u r e 24 Cooper-Harper r a t i n g s f o r f l i g h t c o n d i t i o n 7 , moderate t u r b u l e n c e . . . . . . . . . . . . . . . . . . - .

Summary of advanced PACS p i l o t e d f l i g h t s i m u l a t i o n t e s t r e s u l t s . . . . . . . . . . . . . . . . . . . . . . .

3 9 26 Advanced PACS i n t e r f a c e b l o c k diagram . . . . . . . . . . .

27 Advanced PACS a r c h i t e c t u r e . . . . . . . . . . . . . . . .

28 4 3 Small h o r i z o n t a l t a i l c o n f i g u r a t i o n s e v a l u a t e d . . . . . .

2 9 46 Curvature a i r f o i l shaping d e s i g n system . . . . . . . . . .

30 H o r i z o n t a l t a i l d r a g c h a r a c t e r i s t i c s . . . . . . . . . . . 4 8 31 Comparison of small h o r i z o n t a l t a i l h i g h - l i f t c h a r a c t e r i s t i c s . . . . . . . . . . . . . . . . . . , . . .

v i i i LIST OF TABLES I T a b l e Page ,

1 PACS C o n t r o l l e r I n p u t S i g n a l s . . . . . . . . . . . . . . . 7

I 2 F l i g h t C o n d i t i o n s Used f o r Advanced PACS C o n t r o l 1 2 S y s t e m A n a l y s i s . . . . . . . . . . . . . . . . . . . . . .

I 3

Feedback Gain Matrix (GI) . . . . . . . . . . . . . . . . . 15

i 4 PACS Gain Schedule Equation C o e f f i c i e n t s . . . . . . . . .

1 9 I 5 P i l o t e d F l i g h t Simulation Test C o n d i t i o n s . . . . . . . . .

6 PACS O p e r a t i n g C o n f i g u r a t i o n s Evaluated . . . . . . . . . .

7 P i l o t e d F l i g h t Simulation Evaluation Tasks . . . . . . . .

8 S m a l l H o r i z o n t a l T a i l Comparative Data . . . . . . . . . .

9 S m a l l H o r i z o n t a l T a i l Wind Tunnel T e s t s . . . . . . . . . .

4 8 ix L I S T OF SYMBOLS A S t a t e - s p a c e e q u a t i o n dynamic m a t r i x of aerodynamic d a t a

AAC s A i l e r o n a c t i v e c o n t r o l system

A i r c r a f t A / C ACEE A i r c r a f t Energy E f f i c i e n c y AFCS Automatic f l i g h t c o n t r o l system B S t a te-spac e e q u a t i o n i n p u t d i s t r i b u t i o n m a t r i x C S t a t e - s p a c e e q u a t i o n o u t p u t d i s t r i b u t i o n m a t r i x C* Blended normal-acceleration/pitch-rate parameter A i r c r a f t d r a g c o e f f i c i e n t CD L-1011 l o n g i t u d i n a l c o n t r o l s y s t e m f e e l s p r i n g CF A i r f o i l l i f t c o e f f i c i e n t CL Design a i r f o i l l i f t c o e f f i c i e n t QDES H o r i z o n t a l t a i l l i f t c o e f f i c i e n t CLH Maximum t a i l l i f t c o e f f i c i e n t ‘Lmax P i t c h i n g moment c o e f f i c i e n t Cm C C o e f f i c i e n t of t h e bank a n g l e term f o r secondary g a i n s c h e d u l i n g

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A i r c r a f t c e n t e r of g r a v i t y c . g .

c . p . Wing c e n t e r of p r e s s u r e D S t a t e - s p a c e e q u a t i o n feed-forward m a t r i x , o r d r a g Degrees deg F S t a t e - s p a c e e q u a t i o n feedback m a t r i x x i PRECEDING PAGE BLANK NOT FILMED Control column f o r c e FC FCES F l i g h t c o n t r o l e l e c t r o n i c system p a n e l f t Feet A c c e l e r a t i o n of g r a v i t y Computed feedback g a i n s J-curve compensated feedback g a i n s Computed feed-forward g a i n s Computed feedback g a i n s a f t e r e l i m i n a t i o n of v e l o c i t y s e n s o r s i g n a l Scheduled g a i n s (feed-forward and feedback) A 1 t i t u d e i n .

Inches J J-curve ( r e l a t i o n s h i p between c o n t r o l column d i s p l a c e m e n t and h o r i z o n t a l s t a b i l i z e r r o t a t i o n ) J' Space d e r i v a t i v e of J-curve Normalizing c o n s t a n t = 2.5 Feed-f orward g a i n Mach g a i n Normal a c c e l e r a t i o n g a i n Velocity g a i n Angle of a t t a c k augmented g a i n P i t c h a t t i t u d e g a i n P i t c h r a t e g a i n Bank a n g l e augmented g a i n Combined p i t c h - a t t i t u d e / v e l o c i t y g a i n Knots e q u i v a l e n t a i r speed x i i K t Knots A i r c r a f t Aerodynamic l i f t f o r c e L l b s Pounds L S W T Low speed wing t u n n e l L-1011 Lockheed wide body commercial t r a n s p o r t M Mach number mac Mean aerodynamic chord MC T Maximum c o n t i n u o u s t h r u s t MTO Maximum t a k e o f f t h r u s t Normal a c c e l e r a t i o n feedback s i g n a l N Z F i l t e r e d normal a c c e l e r a t i o n feedback s i g n a l Nz F Numerical c o n t r o l l e d N . C .

PAC S P i t c h a c t i v e c o n t r o l system PL F Power f o r l e v e l f l i g h t PT P r e s s u r e wind t u n n e l Dynamic p r e s s u r e q r a d Radian RSS Relaxed s t a t i c s t a b i l i t y S L a p l a c e t r a n s f o r m parameter S.L. Sea l e v e l sec Second S e r i a l number

s I N

s w s w i t c h

TP T T r a n s o n i c p r e s s u r e wind tunnel T / S T T r a n s o n i c / s u p e r s o n i c wind t u n n e l x i i i State-space e q u a t i o n i n p u t v e c t o r U A i r c r a f t e q u i v a l e n t a i r speed V A i r c r a f t maximum o p e r a t i n g speed Vmo A i r c r a f t s t a l l speed V S A i r c r a f t weight W W Input v e c t o r from p i l o t X S t a t e v e c t o r X Time d e r i v a t i v e of s t a t e v e c t o r E l e c t r o n i c i n p u t s i g n a l t o series s e r v o Control column displacement Output displacement of series s e r v o Output displacement of column t r i m I n i t i a l dynamic s t a t e of a i r c r a f t State-space e q u a t i o n feedback s i g n a l s Y a Angle of a t t a c k Angle of a t t a c k r e l a t i v e t o f u s e l a g e r e f e r e n c e l i n e a~~~ 6 P r e s s u r e of a i r a t a i r c r a f t a l t i t u d e / p r e s s u r e of a i r a t s e a level Outboard a i l e r o n symmetrical d e f l e c t i o n Elevator d e f l e c t i o n Wing f l a p d e f l e c t i o n H o r i z o n t a l s t a b i l i z e r a n g l e H o r i z o n t a l s t a b i l i z e r a n g u l a r v e l o c i t y H o r i z o n t a l s t a b i l i z e r command s i g n a l H o r i z o n t a l s t a b i l i z e r t r i m a n g l e x i v Modified h o r i z o n t a l s t a b i l i z e r feedback g a i n s i g n a l f o r secondary g a i n s c h e d u l i n g H o r i z o n t a l t a i l d r a g increment Mach t r i m compensation Mach t r i m s e r v o o f f s e t schedule Short-period and phugoid mode damping Short-period mode damping CSP Dynamic p r e s s u r e a t h o r i z o n t a l t a i l ‘71.

A i r c r a f t l o a d f a c t o r ‘7

e P i t c h a t t i t u d e feedback s i g n a l

P i t c h rate feedback s i g n a l F i l t e r e d p i t c h rate feedback s i g n a l S t a t e - s p a c e e q u a t i o n e i g e n v a l u e s S t a t e - s p a c e e q u a t i o n e i g e n v e c t o r s Force s e n s o r f i l t e r t i m e c o n s t a n t Mach t r i m compensation f i l t e r t i m e c o n s t a n t Power a c t u a t o r t i m e c o n s t a n t S e r i e s s e r v o t i m e c o n s t a n t Normal a c c e l e r a t i o n feedback s i g n a l f i l t e r t i m e c o n s t a n t P i t c h rate feedback s i g n a l f i l t e r t i m e c o n s t a n t Numerator t i m e c o n s t a n t of l a g - l e a d t r a n s f e r f u n c t i o n Denominator t i m e c o n s t a n t of l a g - l e a d t r a n s f e r f u n c t i o n Bank a n g l e o r phase a n g l e oEi Nyquist diagrams Short-period o r phugoid mode c i r c u l a r frequency S h o r t - p e r i o d mode c i r c u l a r frequency Reference s h o r t - p e r i o d mode c i r c u l a r f r e q u e n c y f o r t h e s p e c i f i c f l i g h t c o n d i t i o n xv I 1. INTRODUCTION

I 1.1 Background

J e t a i r c r a f t f u e l c o s t h a s i n c r e a s e d from a b o u t 1 2 c e n t s p e r g a l l o n i n I 1972 t o a b o u t $1.00 a g a l l o n i n 1983. A s a r e s u l t , t h e f u e l c o s t p o r t i o n o f a i r c r a f t d i r e c t o p e r a t i n g c o s t h a s i n c r e a s e d from 25 p e r c e n t t o n e a r l y 60 p e r - c e n t . T h i s t r e n d was recognized e a r l y by a i r c r a f t m a n u f a c t u r e r s and govern- ment l e a d e r s . T h e r e f o r e , i n 1975 t h e U.S. Congress r e q u e s t e d NASA t o implement a program t o develop f u e l s a v i n g technology f o r commercial t r a n s p o r t s .

I The NASA A i r c r a f t Energy E f f i c i e n c y (ACEE) program was i n i t i a t e d i n 1976.

S i n c e t h e Lockheed C a l i f o r n i a Company was m a n u f a c t u r i n g a wide body commer- ~ cia1 t r a n s p o r t (L-1011) and had been developing a c t i v e c o n t r o l technology 1972, Lockheed r e c e i v e d an ACEE program c o n t r a c t i n February 1977 f o r s i n c e "Development and F l i g h t E v a l u a t i o n of Active C o n t r o l Concepts f o r Subsonic

I

T r a n s p o r t A i r c r a f t " (NASA C o n t r a c t NAS1-14690). The c o n t r a c t r e s u l t e d i n t h e development of an a i l e r o n a c t i v e c o n t r o l system (AACS) which provided wing l o a d a l l e v i a t i o n . The AACS allowed a 5 . 8 p e r c e n t wing span i n c r e a s e f o r t h e L-1011-500 ( i n s e r v i c e d a t e 1980) which d e c r e a s e d f u e l consumption by a p p r o x i - m a t e l y t h r e e p e r c e n t (Reference 1 ) . Also, s t u d i e s were conducted under t h e c o n t r a c t t o e v a l u a t e b e n e f i t s of a p i t c h a c t i v e c o n t r o l system (PACS).

P i l o t e d f l i g h t s i m u l a t i o n tests were conducted on a moving b a s e s i m u l a t o r w i t h a n L-1011 cab. These tests showed t h a t a lagged p i t c h r a t e damper pro- v i d e d f l y i n g q u a l i t i e s w i t h s t a t i c l o n g i t u d i n a l s t a b i l i t y r e l a x e d t o n e a r n e u t r a l and i n heavy t u r b u l e n c e t h a t a r e e q u i v a l e n t t o t h o s e of t h e b a s e l i n e a i r c r a f t . The a f t c.g. s i m u l a t i o n r e s u l t s provided a s u f f i c i e n t b a s i s f o r p r o c e e d i n g t o a f l i g h t e v a l u a t i o n of t h e d e f i n e d augmentation c o n t r o l laws w i t h r e l a x e d s t a t i c s t a b i l i t y .

I n December 1978 Lockheed w a s awarded a second ACEE program c o n t r a c t I (NAS1-15326) f o r "Development and F l i g h t E v a l u a t i o n of a n Augmented S t a b i l i t y Active C o n t r o l s Concept w i t h a Small H o r i z o n t a l T a i l " . I n May 1980 t h e program w a s r e s t r u c t u r e d t o develop a PACS f o r improvement of f l y i n g q u a l i t i e s a t a f t c.g. f l i g h t c o n d i t i o n s u t i l i z i n g a s t a n d a r d L-1011 t a i l and t o c o n t i n u e s m a l l t a i l d r a g r e d u c t i o n e v a l u a t i o n by a n a l y s e s and wind-tunnel tests. The PACS

I

development program w a s d i v i d e d i n t o two p a r t s : a near-term PACS w i t h cap;lbll- i t y t o p r o v i d e s a t i s f a c t o r y f l y i n g q u a l i t i e s t o n e a r n e u t r a l s t a b i l i t y w i t h i n

!

t h e l i n e a r s t a b i l i t y f l i g h t r e g i o n , and a n advanced PACS w i t 5 c a p a b i l i t y t o p r o v i d e s a t i s f a c t o r y f l y i n g q u a l i t i e s f o r n e g a t i v e s t a t i c s t a b i l i t y margins o f 10 p e r c e n t mean aerodynamic chord ( m a c ) throughout t h e a i r c r a f t f l i g h t e n v e l o p e .

The near-term PACS p a r t of t h e program w a s s u c c e s s f u l l y completed and i s r e p o r t e d i n Reference 2 and Reference 3 Extended Work CR-172266. The advanced PACS p a r t of t h e program is documented i n Reference 4 and t h e s m a l l tsil program r e s u l t s i s documented i n Reference 5.

1.2 Program O b j e c t i v e s The program o v e r a l l o b j e c t i v e was t o d e v e l o p f u e l s a v i n g t e c h n o l o g y f o r commercial t r a n s p o r t a i r c r a f t . The g o a l s were t o a c h i e v e a f o u r p e r c e n t f u e l s a v i n g by c,g. management f o r a i r c r a f t w i t h advanced wing c o n f i g u r a t i o n s ( F i g - u r e 1) and t o a c h i e v e a t h r e e p e r c e n t f u e l s a v i n g s by r e d u c i n g t h e h o r i z o n t a l t a i l s i z e (Figure 2 ) .

The advanced PACS program o b j e c t i v e was t o develop a h i g h r e l i a b i l i t y p i t c h a c t i v e c o n t r o l system f o r a f u t u r e commercial t r a n s p o r t t h a t p r o v i d e s h a n d l i n g q u a l i t i e s a t a n e g a t i v e 10% s t a t i c s t a b i l i t y margin which are equiv- alent t o t h o s e of t h e b a s e l i n e L-1011 w i t h t h e c.g. a t 25 p e r c e n t mac (+15% The L-1011 h a s good h a n d l i n g q u a l i t i e s a t t h i s c . g . l o c a t i o n .

s t a t i c margin).

The small t a i l program o b j e c t i v e was t o d e t e r m i n e t h e maximum d r a g bene- f i t t h a t can b e a c h i e v e d by r e d u c i n g t h e h o r i z o n t a l t a i l area. The reduced area must be c o n s i s t e n t w i t h moving t h e c . g . r a n g e a f t and t h e c o n t r o l l a b i l i t y r e q u i r e m e n t 8 .

AFT c.g.

N O R M A L c.g.

F U E L SAVINGS 30 20 10 0 -10 -20

STATtC S T A B I L I T Y - % mec

F i g u r e 1 . - c.g. management system f u e l s a v i n g s .

800 ft2 FUEL SAVINGS 1,282 ft2 (REF) F i g u r e 2 . - Reduced a r e a h o r i z o n t a l t a i l f u e l s a v i n g s .

1.3 Scope of Program The advanced PACS program c o n s i s t e d of d e s i g n c r i t e r i a d e f i n i t i o n , c o n t r o l law s y n t h e s i s , f l y i n g q u a l i t y a n a l y s i s , p i l o t e d f l i g h t s i m u l a t i o n tests, and system a r c h i t e c t u r e s p e c i f i c a t i o n .

The s m a l l t a i l program c o n s i s t e d o f d e s i g n c r i t e r i a d e f i n i t i o n , d e s i g n and f a b r i c a t i o n of t h e reduced area h o r i z o n t a l t a i l models, and wind t u n n e l tests.

2. ADVANCED PACS DEVELOPMENT The Lockheed L-1011 house a i r p l a n e ( S / N 1001) w a s used as t h e b a s i s f o r I d e s i g n of t h e advanced PACS. T h i s a i r p l a n e ( F i g u r e 3 ) i s a n L-1011-1 model e x c e p t f o r t h e extended wing t i p s and a c t i v e c o n t r o l a i l e r o n s which are i n s t a l l e d on t h e L-1011-500 models. I The a i r p l a n e h a s a f l y i n g s t a b i l i z e r w i t h a g e a r e d e l e v a t o r which h a s been downrigged f i v e d e g r e e s t o p r o v i d e t h e I r e q u i r e d n o s e down a u t h o r i t y when t h e c . g . i s moved a f t t o a l l o w f l i g h t a t n e g a t i v e s t a t i c s t a b i l i t y margins t o t h r e e p e r c e n t mac. Center of g r a v i t y (c.g.) management i s provided by a n e l e c t r i c a l l y c o n t r o l l e d water b a l l a s t system f o r performance of f l i g h t tests (Reference 2 ) .

I I A b l o c k diagram which shows t h e major e l e m e n t s of t h e L-1011 l o n g i t u d i n a l c o n t r o l system w i t h t h e advanced PACS i n s t a l l e d i s g i v e n i n F i g u r e 4 . The

i

dashed l i n e s r e p r e s e n t t h e b a s e l i n e a i r c r a f t l o n g i t u d i n a l c o n t r o l s and t h e s o l i d l i n e s r e p r e s e n t t h e advanced PACS. I n p u t s i g n a l s t o t h e c o n t r o l l e r are d e f i n e d i n Table 1. The system h a s d u a l series s e r v o s f o r s a f e t y r e a s o n s . !

F a i l u r e of one s e r v o w i l l n o t induce a s w i f t e l e v a t o r d e f l e c t i o n ( h a r d o v e r ) which w i l l over stress t h e a i r c r a f t .

A schematic of t h e L-1011 c o n t r o l system

I

w i t h t h e series s e r v o s ( b l a c k l i n e s ) i n s t a l l e d i s shown i n F i g u r e 5.

2 . 1 C o n t r o l L a w Design O b j e c t i v e s Design o b j e c t i v e s f o r c o n t r o l law s y n t h e s i s i n c l u d e d dynamic s t a b i l i t y , maneuver s t a b i l i t y , and normal-acceleration/pitch-rate r e s p o n s e a s shown i n F i g u r e s 6 , 7 , and 8 r e s p e c t i v e l y . Speed s t a b i l i t y was n o t i n c l u d e d i n t h e I i n i t i a l d e s i g n o b j e c t i v e s . However, speed s t a b i l i t y a n a l y s i s ( S e c t i o n 2 . 3 . 3 ) I e s t a b l i s h e d a requirement f o r u s i n g t h i s d e s i g n o b j e c t i v e i n f u t u r e c o n t r o l law development programs.

I 2.2 C o n t r o l Law S y n t h e s i s The e n t i r e d i s c u s s i o n i n t h i s s e c t i o n i s based on t h e b l o c k diagram f o r c o n t r o l law s y n t h e s i s g i v e n i n Figure 9 .

The s y n t h e s i s p r o c e s s s t a r t e d w i t h a s e p a r a t e set o f aerodynamic d a t a t ( t r i m c o n d i t i o n s and s t a b i l i t y d e r i v a t i v e s ) f o r each of t h e f l i g h t cases l i s t e d i n T a b l e 2. The aerodynamic d a t a was i n p u t t o t h e b a s e l i n e a i r c r a f t I I model.

2 . 2 . 1 Feedback Loop Gain.- Development of t h e feedback l o o p g a i n s c o n s i s t e d of t h e f o l l o w i n g s t e p s .

Determining a r e f e r e n c e e i g e n s t r u c t u r e (A , v . ) f o r each f l i g h t c o n d i t i o n i i from t h e b a s e l i n e a i r c r a f t model 0 Computing t h e feedback g a i n s m a t r i x ( G ) by u t i l i z i n g modal c o n t r o l s y n t h e s i s PRECEDING PAGE BLANK NO FILMED'

&NlrWoBWU

ACTIVE CONTROL AILERONS EXTENDED WING GEARED ELEVATOR TIPS

c

5 DEGREES DOWNRIGGED / FLYING c.g. MANAGEMENT STAB I LI 2 E R SYSTEM PACS F i g u r e 3 . - F l i g h t test a i r p l a n e (L-1011 S/N 1 0 0 1 ) .

F i g u r e 4 . - L o n g i t u d i n a l c o n t r o l - s y s t e m w i t h t h e advanced PACS.

TABLE 1. - PACS CONTROLLER I N P U T SIGNALS SYMBOL SIGNAL TYPE USE Column force gradient Column force Feed-f orward FC N Z Normal acceleration Short period mode

e Feedback

Pitch rate Pitch attitude

I e

Dynamic pressure Compensation for Primary gain flight condition Horizontal stabilizer trim scheduling changes "T Angle of attack Compensation for Secondary gain pitch-up and AACS Bank angle scheduling outboard aileron Mach number operation a a , a , rl N U

r .-I

CT I- rl I O z I4 U a , Y Y U I 4-1 u t n V a d U

u

z U I u : I m { - 0.5 \ 0.8 mac - U A . SHORT PERIOD MOOE w 5 = 0.1

< = 0.5 I

- u 6. PHUGOID MOOE F i g u r e 6 . - PACS dynamic s t a b i l i t y d e s i g n o b j e c t i v e s .

F L A P S UP C O N D I T I O N S 2 5 F(; ) I \ 1 5 IO 0 0 1 ,1 1 8 L O A D F A C T O R ' q Figure 7.- PACS column f o r c e g r a d i e n t d e s i g n o b j e c t i v e 2 5 NORMALIZED AT TIME

1 - - sec

2 0 FLAPS UP CONDITIONS 1 5 C' 1 0 0 5 I I I 0 I 2 3 4 TIME-SEC F i g u r e 8.- Blended normal-acceleration/pitch-rate (C*) r e s p o n s e o b j e c t i v e

ORlGlNAh PAGE €3

OF POOR QUALIW

I : I

z CCE?

3 : * O n I .

>; Y

-

W C D N N N N N 0 0 0 0 0 0 0 0 0 b b d L I I I I __ u u w E a u 1 2 0 Modifying the feedback gains (G1) to account for a nonlinear relationship between the L-1011 stabilizer rotation and control column displacement (J-curve). This is called J-curve compensation and provides a set of modified gains ( G 2 ) .

Modifying the compensated gain matrix ( G 2 ) to delete the requirement for a velocity sensor signal output. This provides the desired feedback gain matrix ( G q ) .

I The baseline aircraft model shown in Figure 9 was the open loop state-space equation (Equation 1) given in Figure 10.

I

I = [A] { x ) + [ B ] Iu)

( E q . 1)

I

The matrices are defined as follows.

A = Aerodynamic data x = state-space vector

& = derivative of the state-space vector

u = input vector B = input distribution matrix I Elements of the state-space vector and the input vector were: angle of attack increment pitch rate pitch attitude increment normalized airspeed increment {XI = filtered normal acceleration increment filtered pitch rate horizontal stabilizer angular velocity horizontal stabilizer angular increment horizontal stabilizer command signal Outboard aileron symmetrical deflection 1 3

i

I

ORlGlNAL PAGE e a

OF POOR QUALITY

FEED-FORWARD LOOP XO I

{;I= [ A I ( k ) t [ B l { u )

Compute Eigenvalues and Eigenvectors * F FEEDBACKLOOP = [ A t BFCl { , X \ Solve for Matrix F F i g u r e 10. - Advanced PACS c o n t r o l math model i n s t a t e - s p a c e form.

Equation 1 was used t o o b t a i n a set of e i g e n v a l u e s and e i g e n v e c t o r s ( l i , ~ i ) a t t h e 25 p e r c e n t mac c.g. p o s i t i o n For each of t h e 1 4 f l i g h t c o n d i t i o n s l i s t e d i n Table 2 . T h i s set of e i g e n v a l u e s and e i g e n v e c t o r s , c a l l e d t h e r e f e r e n c e e i g e n s t r u c t u r e , p r o v i d e s v a l u e s t h a t are used f o r t h e modal c o n t r o l s y n t h e s i s .

The modal c o n t r o l method of modern c o n t r o l t h e o r y w a s used t o d e t e r m i n e t h e feedback g a i n m a t r i x G 1 ( s e e Table 3 ) . T h i s method i s s u p e r i o r t o c l a s s i c a l c o n t r o l t h e o r y because t h e g a i n f o r several feedback l o o p s can b e computed simultaneously i n s t e a d of one a t a t i m e , and t h e p o l e placement i n t h e complex p l a n e i s s i g n i f i c a n t l y s i m p l i f i e d . Modal c o n t r o l s y n t h e s i s w a s accomplished by t h e s t a t e - s p a c e feedback l o o p e q u a t i o n (Equation 2 ) g i v e n i n F i g u r e 10 with t h e s w i t c h (SW1) c l o s e d .

Each element of Equation 2 i s known e x c e p t f o r t h e g a i n m a t r i x F. Thus, t h e e q u a t i o n is s o l v e d f o r F t o o b t a i n t h e feedback g a i n s f o r e a c h f l i g h t case t o p r o v i d e t h e feedback g a i n m a t r i x GI. The e l e m e n t s of t h e s t a t e - s p a c e equa- t i o n o u t p u t d i s t r i b u t i o n m a f r i x (c) were a r r a n g e d t o p r o v i d e t h e d e s i r e d f e e d - back l o o p s i g n a l s : N , and 0 f o r c o n t r o l of t h e s h o r t - p e r i o d mode, and u and 0 f o r c o n t r o l of t h e phugoid mode. Values of t h e e l e m e n t s f o r m a t r i c e s A, B , and C a r e d i f f e r e n t f o r each c.g. l o c a t i o n , whereas t h e r e f e r e n c e set of e i g e n v a l u e s (Xi) and e i g e n v e c t o r s ( v i ) remain unchanged f o r each f l i g h t c o n d i t i o n .

1 4

u~EGIFJAB PAGE 1 8

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TABLE 3 . - FEEDBACK G A I N MATRIX (GI) , 1 5 The relationship between the L-1011 stabilizer rotation (6 ) and the JJ control column displacement (XC) is a nonlinear function which is based on the stabilizer trim setting (~HT) as shown in Figure 11. Only the curves for 6HT = However, a family of 0 and ~ H T = -10 degrees are shown in the figure.

curves exist for points on the trim line. These curves are called the J-curve.

Figure 9 shows that the stabilizer trim signal is obtained from the aerodynamic data and supplied to the J-curve model. This model is a set of equations that was curve fitted to the family of curves of Figure 11. The output of the J-curve model which was used for J-curve compensation was the J-curve deriva- tive (J') corresponding to ~ H T for the specific flipht case being evaluated.

The slope of all members of the J-curve family is the same for any specified values of 6 ~ ~ .

The compensated feedback gain matrix (G2) was determined by application of Equation 3.

J' is a diagonal matrix of the J-curve derivatives corresponding to the 6 HT for the 56 flight cases (Table 1 ) .

Deletion of the velocity sensor is desirable because changes of trim conditions result in frequent velocity changes. Consequently a lag-lead circuit was devised to produce a signal component of derived incremental speed. Therefore, instead of using the velocity gain ( K , ) and the pitch attitude gain (KO), a new set of gains were determined in terms of a combined pitch-attitude/velocity gain (K3), a numerator time constant of the lag-lead circuit (TI), and a denominator time constant of the lag-lead circuit ( ~ 2 ) .

Thus the feedback gain matrix ( G 4 ) is expressed in terms of the gains KB, K N ~ , 1 1 ~ 2 , ~ 2 1 ~ 1 - 1, and K3 ~ 1 1 ~ 2 .

2.2.2 Feed-forward l o o p . - The feed-forward loop synthesis considered switches SW1 and SW2 o f Figure 9 to be closed. The control equation is expressed in terms of Equation 4 .

Matrix D is the state-space feed-forward matrix and w is the pilot input vector. This equation was solved by taking the Laplace transform of Equa- tion 4 and using Cramer's rule to obtain the force gradient (FC/Nz).

\There FC is the control column force and NZ is the aircraft normal acceleration at some specified location.

DEG -1 4 - 1 2 -10 -8 -6 - 4 -2 I 1 1 I I I 1 1 I I I I I I I 0 2 4 6 8 10 12 14

X c - in

F i g u r e 11. - L-1011 s t a b i l i z e r r o t a t : i o n ( 6 H ) r e l a t i v e t o c o n t r o l column displacement (X,).

1 7 2.2.3 Primary gain scheduling.- Primary gain scheduling refers to a procedure which expresses the feedback matrix (G4) and the feed-forward gain matrix (G3) in the form of a second degree polynomial equation with parameters stabilizer trim (6HT) and dynamic pressure ( q ) . Figure 9 shows that the ~ H T and q signals are provided from aerodynamic data for gain scheduling. The gain equation (Equation 5) is given below.

2 2

+ dCiHT +

K = a + bq + cq

(Eq. 5) Symbols in the equation are:

I

K = feedback or feed-forward signal gains a,b,c,d,e = Equation coefficients determined by using a least squares curve fit Equation 5 was applied for the flap-up and flap-down cases to determine the equation coefficients for the feedback gains (Ki), K N ~ , 1 / ~ 2 , T ~ / T ~ , K ? T ~ / T ~ ) and the feed-forward gain (KF ) which are given in Table 4. The flap-up

flight conditions gain schedu E ing curves for the feedback gain ( K i ) and the

feed-forward gain (KFF) are plotted in Figures 1 2 and 13 respectively. The other feedback gains (KN?, 1 / ~ 2 ,- i 2 / ~ 1 , T ~ / T ~ ) and the flap-down flight condition gains have a similar set of curves.

The primary gains such as those represented by the curves in Figures 1 2 and 13 are sufficient for gain scheduling of linear stability flight conditions.

However, for nonlinear stability flight conditions, a secondary gain scheduling is required as described in the following section.

2.2.4 Secondary gain scheduling.- Secondary gain scheduling is required to compensate for: 0 Pitch-up at high-Mach/high-g flight conditions 0 Symmetric activity of the aileron active control system outboard ailerons The pitch-up phenomena is caused by a l o s s of lift at the wing tips during the high-Mach/high-g flight conditions which causes the aerodynamic center of pressure (c.P.) to shift forward. Thus, the distance between the c.g. and the C.P. becomes less and the static stability margin is reduced in a manner similar to that when the c.p. is fixed and the c.g. is moved aft. Consequently, the gain scheduling curves already developed (e.g. Figures 1 2 and 13) can be used to stabilize the pitch-up conditions.

ORIGINAL PAGE 1 3 OF POOR QUALITY

TABLE 4 . - PACS GAIN SCHEDULE EQUATION COEFFICIENTS

FLAP

K F F - in/lb

SETTING COEF.

I

~~ -1.4295 9.7718 2.2322 x 10.’ 2.1328 x 10.’ 2.0571 x

k-

-4.0915 x FLAPS- 0 -4.1698 x I O UP 2.6975 x 10.’

-5.0386 x 10-1 I 4.5098

-4.2834 x 10.‘

I 0

n 3.8428 x 10.’ -7.4588 x los3 8.7222 x 10.’ -1.577 1 -1.0483 3.2829 x 10-1 n 2.4014 x lo-’

r-r -3.5719 x loT5

FLAPS -1.4232 x l 0-3 6.8201 x 10‘’ -

, -3.9313 x 10.’

4.3358 x 10.l 0 -9.7027 x los 4.0296 x

OOWN F

-0.00 -0.08 -0.16 -0.24 -0.32

KC - sec

-0.4[ FLIGHT DYNAMIC -0.4t CONDITIONS PRESSURE 0 3. 254.

A 4. 461.

0 5. 461.

-0.51 203.

0 6.

218. 0 7.

+ 8. 234.

-0.61 0 9. 258.

A 10. 266.

0 11. 285.

-0.1

I 1 1 1 1 1 1

-0.81 -2.8 -2.4 -2.0 - 1 . 6 0 -1.20 -0.8 -0.4 -3.6 -3.2

~ H T - deg

F i g u r e 1 2 . - Scheduled p i t c h r a t e feedback g a i n c u r v e s , f l a p s - u p c o n d i t i o n s .

ORIGIIMAL PAGE I$

OF POOR QUALITY

.5t F L I G H T DYNAMIC

-

.41 CONDITIONS PRESSURE 254.

4 :: 451.

451.

0 5.

- .40 209.

0 6 .

@ 7.

218.

234.

8- 258.

0 9.

A 10.

255.

- .32 0 11. 285.

.24

K F F - inllb

.16 - .08 - . 1 6 - . 2 4 - 3.60 - 3 . 2 0 . - 2 . 8 0 -2.40 - 2 . 0 0 - 1.60 - 1 . 2 0 - 0.80 - 0 . 4 0 ~ H T deg Figure 13. - Scheduled feed-forward gain curves f l a p s - u p conditions.

2 1 The v a l u e of ~ H T i s incremented by t h e secondary g a i n c o n t r o l l e r ( F i g u r e 1 4 ) t o p r o v i d e a modified v a l u e ~ H T * ( s e e F i g u r e 4 ) .

S i g n a l s r e q u i r e d t o p r o v i d e t h e c o r r e c t 6 H ~ * v a l u e are a n g l e o f a t t a c k (a), bank a n g l e ( I $ ) , and Mach number (M) as shown i n F i g u r e 1 4 . The modified v a l u e 6 ~ ~ " changes t h e feedback g a i n s t o p r o v i d e t h e i n c r e a s e d c o n t r o l command f o r t h e h o r i z o n t a l s t a b i l i z e r and changes t h e feed-forward g a i n s t o p r o v i d e t h e d e s i r e d column f o r c e g r a d i e n t s .

The AACS o p e r a t e s t h e outboard a i l e r o n s i n a symmetric mode i n r e s p o n s e t o normal a c c e l e r a t i o n of t h e a i r c r a f t c.g. and wing t i p s . T h i s symmetric mode produces a c . p . s h i f t t h a t i s e q u i v a l e n t t o a n a f t c.g. s h i f t of about f i v e p e r - c e n t mac.

The change i n p i t c h i n g moment can b e c o r r e c t e d i n t h e same way a s f o r t h e pitch-up c o n d i t i o n . T h i s i s accomplished by s w i t c h SW1 i n F i g u r e 1 4 .

Also, a switch S W 2 p r o v i d e s f o r flap-up and flap-down g a i n changes. Switch SW1 i s a u t o m a t i c a l l y c l o s e d when t h e AACS i s engaged and S W 2 i s a u t o m a t i c a l l y o p e r a t e d t o correspond t o t h e f l a p s e t t i n g .

2.2.5 Advanced PACS c o n t r o l l a w . - The advanced PACS c o n t r o l l a w b l o c k The diagram i s c o n s i d e r e d t o b e d i v i d e d i n t o diagram i s given i n F i g u r e 15.

t h r e e p a r t s f o r t h e purpose of d i s c u s s i o n .

0 Control column and a c t u a t o r system: c o n t r o l column, column t r i m , s e r i e s s e r v o s , J-curve, s t a b i l i z e r t r i m , and power a c t u a t o r

0 Feedback l o o p s : p i t c h rate ( e ) , normal a c c e l e r a t i o n ( N Z ) , and

p i t c h a t t i t u d e ( 0 ) 0 Feed-forward l o o p : column f o r c e (FC) The c o n t r o l column and a c t u a t o r system i s c o n s i d e r e d t o start w i t h t h e c o n t r o l column displacement ( X , ) and t h e c o n t r o l column t r i m (XT). These The n o n l i n e a r i z e r i n p u t s are summed w i t h t h e series s e r v o o u t p u t s (Xs).

( s e e F i g u r e 5) r e p r e s e n t e d by t h e J-curve ( J ) p r o v i d e s t h e J-curve compensa- t i o n . The s t a b i l i z e r t r i m ( ~ H T ) i s t h e n s u b t r a c t e d t o p r o v i d e t h e s t a b i l i z e r s e r v o command s i g n a l ( ~ H c ) . The s t a b i l i z e r s e r v o c o n t r o l s t h e h y d r a u l i c flow t o t h e s t a b i l i z e r power a c t u a t o r s t o p r o v i d e t h e d e s i r e d s t a b i l i z e r a n g l e (6~).

The power a c t u a t o r l a g c h a r a c t e r i s t i c s a r e shown i n t h e f i g u r e t o b e l / ( ~ ~ s + l ) .

Where T~ i s t h e power a c t u a t o r t i m e c o n s t a n t and s i s t h e L a p l a c e Transform parameter.

The feedback l o o p u s e s t h e e and N Z feedback s i g n a l s f o r c o n t r o l of t h e s h o r t - p e r i o d modes. These s i g n a l s are f i l t e r e d t h r u t h e f i r s t o r d e r low-pass f i l t e r s , l/(.rZs + 1) and l / ( ~ ; t s + l), shown i n F i g u r e 15. The f i l t e r t i m e c o n s t a n t s T~ and ~6 are e q u a l t o 0.03 seconds. The f i l t e r e d s i g n a l s 6~ and N Z F are s u b j e c t e d t o t h e scheduled g a i n s K i and K N ~ r e s p e c t i v e l y .

A n o r m a l i z i n g c o n s t a n t l / K c i s used i n each feedback l o o p s i g n a l so t h a t g a i n s c h e d u l e s t h r u t h e J-curves p r o v i d e a 6HT v a l u e of 10 d e g r e e s t o produce a - 0 + + k aJ CT Y 2 h k rd a L a i : F: u a J m 2 3

ORIGINAL PAGE 8 8

OF POOR QUALlVV

s i g n a l v a l u e of 1. The e feedback s i g n a l i s used t o c o n t r o l t h e phugoid mode.

T h i s s i g n a l i s p r o c e s s e d through a p i t c h s y n c h r o n i z e r , a l a g - l e a d , c i r c u i t , and a g a i n a m p l i f i e r . The p i t c h s y n c h r o n i z e r s u p p r e s s e s t h e a t t i t u d e h o l d d u r i n g maneuvers and sets a new a t t i t u d e r e f e r e n c e a t t h e s y n c h r o n i z e r o u t p u t when a c o n t r o l column f o r c e i s a p p l i e d . The l a g - l e a d c i r c u i t eliminates t h e need f o r a v e l o c i t y s i g n a l t h a t would be r e q u i r e d f o r phugoid mode c o n t r o l . Thus, t h e r e s u l t i n g s e t of g a i n s t o b e scheduled are: ( ~ 2 1 - r ~ - 11, 9 and K 3 ( ~ 1 / ~ 2 ) .

The feed-forward l o o p i s used t o p r o v i d e t h e d e s i r e d c o n t r o l column f o r c e g r a d i e n t s . (C,) c o n v e r t s t h e c o n t r o l column d i s p l a c e m e n t The f e e l - s p r i n g (X,) t o column f o r c e (Fc).

The f o r c e s e n s o r c o n v e r t s FC t o a n e l e c t r i c a l s i g n a l . A flaps-up/flaps-down b i a s s i g n a l s w i t c h t h e low-pass f i l t e r t i m e c o n s t a n t (rc) from 0.66 sec f o r flaps-up t o 1.06 sec f o r flaps-down.

These t i m e c o n s t a n t s are r e l a t e d t o t h e s h o r t - p e r i o d mode o f t h e b a s e l i n e a i r c r a f t .

The s i g n a l i s t h e n passed through t h e scheduled feed-forward g a i n (K ) and i s FF summed w i t h t h e feedback s i g n a l s t o provide t h e series s e r v o i n p u t s i g n a l (XA).

The dashed l i n e s i n F i g u r e 15 r e p r e s e n t a Mach compensation c i r c u i t and a feed-forward g a i n lower l i m i t (KFF = 0) which w a s added t o p r o v i d e t h e r e q u i r e d speed s t a b i l i t y ( s e e S e c t i o n 2.3.3) t h a t w a s n o t accounted f o r i n t h e i n i t i a l c o n t r o l l a w development. The Mach compensation c i r c u i t c o n s i s t s of two c i r c u i t s t h a t o p e r a t e through t h e Mach t r i m system. The Mach t r i m com- p e n s a t i o n A6, and c o r r e s p o n d i n g f i l t e r w a s p a r t of t h e b a s e l i n e a i r c r a f t con- t r o l system. The Mach t r i m s e r v o o f f s e t s c h e d u l e ( A & , ) and t h e l o o p g a i n s c h e d u l e (KM) p r o v i d e t h e d e s i r e d c o n t r o l column f o r c e g r a d i e n t f o r speed s t a b i l i t y of t h e PACS c o n f i g u r e d A i r c r a f t . The s t a b i l i z e r g a i n f i l t e r h a s a 20 second t i m e c o n s t a n t and t h e o f f s e t s c h e d u l e f i l t e r h a s a 10 second t i m e c o n s t a n t t o l i m i t series s e r v o o f f s e t g a i n o v e r s h o o t .

2.3 FLYING QUALITY ANALYSIS The v a l i d i t y of t h e c o n t r o l l a w w a s e v a l u a t e d f o r t h e f l i g h t c o n d i t i o n s t h a t w e r e t o b e used f o r t h e p i l o t e d f l i g h t s i m u l a t i o n test (Table 5 ) . The a n a l y s i s i n c l u d e d dynamic s t a b i l i t y , maneuver s t a b i l i t y , speed s t a b i l i t y , and t r i m m a b i l i t y .

2 . 3 . 1 Dynamic s t a b i l i t y . - Dynamic s t a b i l i t y a n a l y s i s i n c l u d e d l i n e a r a n a l y s i s and n o n l i n e a r a n a l y s i s a p p l i c a t i o n s .

The l i n e a r a n a l y s i s w a s performed t o show t h a t t h e c h a r a c t e r i s t i c r o o t s o f t h e c o n t r o l system m e t t h e d e s i g n o b j e c t i v e s of F i g u r e 6 . T h i s w a s accomplished by p l o t t i n g t h e short-period and phugoid mode r o o t l o c i o f e a c h l o c a t i o n s from 25 t o 50 p e r c e n t m a c i n t h e complex(s) f l i g h t c o n d i t i o n f o r c.g.

p l a n e .

TABLE 5. - PILOTED FLIGHT SIMULATION TEST CONDITIONS

V Weight c.g. Altitude KEAS Flight Condition 1000 Ibs % mac 1000 ft 408 25 to 50 37 7. Cruise 254

W/6 = 1 . 9 x lo6 Ibs (M = 0.83)

10. Cruise 360 25 to 50 33

W/6 = 1 . 4 x lo6 Ibs (M = 0.83)

15. Cruise 360 25 to 50 36

W/S = 1 . 6 x lo6 Ibs ( M = 0.83)

16. M m o / V m o 350 25 to 50 25 357 17. Holding 335 25 to 50 10 250 18. Landing 330 25 to 50 2 135 (SF = 33 deg) (1.3V , ) 19. Takeoff 380 25 to 50 2 137 26 deg) ( 6 F ( 1 -2V , ) The nonlinear analysis was performed to determine time histories of the longitudinal dynamic response for control column step inputs and for discrete Figure 1 6 shows the aircraft response for angle of attack, vertical gusts.

pitch rate, and load factor with the PACS on and off for Flight Condition 7 (Table 5 ) with the c.g. at 5 0 percent rnac. The baseline aircraft diverges quickly from its trim condition for any constant force input until it reaches a region of increased stability at high angle of attack. Engagement of the PACS reduces the angle of attack and load factor excursions significantly.

Figure 17 gives a comparison of the aircraft response with the PACS off and on for a discrete severe vertical gust with a peak of -54 ft/sec. This gust is representative of a severe disturbance €or a heavy thunderstorm. For this severe disturbance the baseline airplane with c.g. at 25 percent r n a c will return to its initial trim condition. If the c.g. is aft of 25 percent rnac, the aircraft diverges from its trim condition and seeks a new equilibrium at high angle of attack. The PACS configured aircraft had well-behaved and stable response characteristics for all of the c.g. positions from 25 to 5 0 percent rnac.

Figure 18 shows that the blended normal-acceleration/pitch rate response (C*) for flight condition 7 with the c.g. at 25 percent r n a c was in compliance with the design objective (Figure 8).

ORiGlMAL PAGE 1 3 OF POOR QUALITY FLIGHT CONDITION 7 AACS ON, c.g. AT 50% mac BASELINE AlRCRAFl PACS CONFIGURED AIRCRAFT -

, F c - 20 Ibs

- - 8 - 8 - D 6 - 6 - B I 4 - 4 - I I 1 1 1 1 I L I I I I I I 0 4 8 12 16 20 24 8 12 16 20 2 4 0 4 TIME * sec

TIME - sec

1 F C - 20 Ibr

- 2 t

- z - 4 t

- 4 L u c I I , I 4 I l , , , 0 4 8 12 16 20 24 0 4 8 12 16 20 24 TIME * sec TIME * sec

2 ' 5 c

FC - 20 Ibs

2.0 - u s

0.5 t

OL L I I 1 I I I I 0 4 8 12 16 20 24 0 4 8 12 16 20 24 TIME n, rec TIME * rec F i g u r e 16.

- Comparison of a i r c r a f t r e s p o n s e w i t h and w i t h o u t PACS engaged f o r v a r i o u s l e v e l s of c o n t r o l column s t e p i n p u t s .

FLIGTH CONDITION 7

W ~ E A K GUST - - 5 4 FTISEC

nLsEuwL AIRCRAFT PACS CONFIGURED AIRCRAFT -

-

- -

- 8 - .I

c.0. - 25%

t 8 -

I

4 - 8:+ 2 - 2 0 - 0 - I I 1 I I I I o 4 E 12 18 20 24

I

. s -2 Q 4 E 12 18 20 24

TIME - uc

, c.g. = 26% mac

2 ' 5 1 2 . 0 2'51

-

2.0

-

1.5 - 1.0 - 0.5

0'5 t

O L I 1 I I I O L 0 4 E 12 1E 20 24 I I 1 1 0 4 8 12 18 20 24

TIME - u c

TIME * uc

Figure 17. - Comparison of aircraft response with and without

PACS engaged for a severe vertical gust.

ORIGINAL PAGE 8 8

OF POOR QilALlTY FLIGHT CONDITION 7 c.g. = 50% mac 2.0 1.5 NORMALIZED C* 1 .o 0.5 0 1 2 3 4 TIME sec F i g u r e 18. - Blended n o r m a l - a c c e l e r a t i o n / p i t c h - r a t e r e s p o n s e .

2.3.2 Maneuver s t a b i l i t y . - The maneuver s t a b i l i t y a n a l y s i s w a s performed t o d e t e r m i n e t h e c o n t r o l column f o r c e g r a d i e n t as a f u n c t i o n of l o a d f a c t o r .

The b a s e l i n e a i r c r a f t maneuver s t a b i l i t y c o n t r o l column c h a r a c t e r i s t i c s f o r c r u i s e a r e shown i n F i g u r e 19. These d a t a show t h a t t h e f o r c e g r a d i e n t s f o r e a c h c . g . l o c a t i o n does n o t comply with t h e MIL-F-8785C c r i t e r i a i n t h e n o n l i n e a r s t a b i l i t y f l i g h t c o n d i t i o n range.

Engagement of t h e PACS improved t h e column f o r c e g r a d i e n t s s i g n i f i c a n t l y as shown i n F i g u r e 20. T h i s improvement complies w i t h t h e d e s i g n o b j e c t i v e e x c e p t a t h i g h l o a d f a c t o r s and i s p r i m a r i l y due t o t h e secondary g a i n s c h e d u l i n g . T o e v a l u a t e t h e secondary g a i n s c h e d u l e , t h e s i x PACS o p e r a t i n g conf g u r a t i o n s l i s t e d i n Table 6 were analyzed. C o n f i g u r a t i o n number two ( t h e PAC S w i t h f u l l g a i n ) w a s shown t o provide t h e b e s t c o n t r o l column g r a d i e n t .

2 . 3 . 3 Speed s t a b i l i t y . - The speed s t a b i l i t y a n a l y s i s d e t e r m i n e s t h e con- t r o l column f o r c e r e q u i r e d t o maintain t h e a i r c r a f t a t s p e e d s d i f f e r e n t from t h e r i m speed. FAR P a r t 25 requirements were used as t h e speed s t a b i l i t y c r i t e r i a . FAR P a r t 25 d e f i n e s s a t i s f a c t o r y column f o r c e c h a r a c t e r i s t i c s as f o l l o w s .

a A p u l l f o r c e s h a l l b e r e q u i r e d t o m a i n t a i n speed below t r i m speed and a push f o r c e s h a l l be r e q u i r e d t o m a i n t a i n speed above t r i m speed.

80 PULI FLIGHT CONDITION 7 AACS ON

A’

FC Ibr I 4[ I I I I I I 1 1 2 1 4 1 6 I 8 2 0 2 2 I O LOAD FACTOA * g Figure 19. - Baseline aircraft maneuver stability column force gradients, cruise.

FLIGHT CONDITION 7

I I I I I

1 .o 1.2 1.4 1.6 1.8 2.0

LOAD FACTOR - g

Figure 20. - PACS configured aircraft maneuver stability column force gradients, cruise.

It d Y U U U U I Y Y II It L L U U U Y x Y u W W U

t

s E

u a a a

-

a L?

W w z -1 .

W v) v) O a n e Stick forces shall vary monotonically with speed.

e The average stick force gradient shall be at least -1 lb per six KEAS increase throughout the speed range.

I Speed stability had not been included in the design criteria for control law development. Consequently, the speed stability analysis showed an abrupt 1 I column force reversal for the takeoff condition with c.g. at 25 percent mac and 50 percent c.g. location. These prob- unstable column force gradients for the

i

lems were corrected by limiting the lower bound of the feed-forward gain (KFF) to zero and by adding a Mach compensation circuit that operates through the Mach trim system (Figure 1 5 ) . The speed stability column forces for the PACS with Mach compensation are shown in Figure 21 for c.g. positions from 25 percent to 50 percent rnac. As shown in the figure, the speed stability satisfies the FAR Part 25 criteria.

2.3.4 Trimmabi1ity.- This analysis was performed t o determine changes of the baseline aircraft control system that were required for the advanced PACS configured aircraft. Control system characteristics that were evaluated included: stabilizer/elevator deflection range, trim servo range, elevator versus stabilizer gearing relationship, control column limits, and pitch feel- spring rate. The control system design criteria were: e Capability must be provided to trim the aircraft for all flight conditions FLIGHT CONDITION 7 AACS ON v)

rr

LL - 40 PUSH - 80 c I I 1 I I I 180 200 220 240 260 280 300 2 :O

v - KEAS

Figure 21. - Speed stability column forces, cruise.

Sufficient control power must be provided to give a minimum pitch angular acceleration of -5.73 degjsec2 for stall recovery from any flight condition Control power must be provided to recover from maneuvers in high angle-of-attack regions.

The baseline L-1011 c.g. range is 12 to 35 percent rnac. The corresponding stabilizer/elevator deflection range is from -14 deg/-25 deg nose-up to +1 deg/0 deg nose-down. The stabilizer deflection trim range is -10 deg to 0 deg. The advanced PACS configured c.g. range is 25 to 50 percent rnac. Anal- ysis showed that the corresponding stabilizer/elevator deflection range should be from -14 deg/-20 deg nose-up to +4 deg/+5 deg nose-down, and the trim range should be -10 deg to +1 deg.

I

Thus, modifications of the stabilizer/elevator gearing curve, the J-curve, the trim servo, and the feel-spring rate would be required for the L-1011 in I order to flight test an advanced PACS to a three percent negative stability ~ margin. This three percent negative stability margin represents the aft c.g.

limits for which the L-1011 can be tested without significant modifications.

2.4 Piloted Flight Simulation Test I The advanced PACS piloted flight simulation test was performed to identify pilotfcontrol interface problems and to evaluate flying qualities of the air- craft.

The test was performed at the NASA Langley Flight Simulation Facilities.

Setup of the simulator included a check of the simulation computer program, motion system interface, cockpit controls, and instrumentation. Two Lockheed and three NASA pilots performed the flight simulation tests.

I 2.4.1 Flight Simulator.- The NASA flight sinulator is a visual motion simulator with a two-man cockpit mounted on a six degree-of-freedom synergistic motion base. A collimated visual display provides a 60 degrees out-the-window color display which was activated during the landing approach task.

2.4.2 Simulation Computer Program.- The simulation mathematical model represents the L-1011 S/N 1001 Aircraft. Engine characteristics were repre- sented by the installed thrust for three Rolls Royce R.B.211-22B high-bypass ratio turbofan engines.

2.4.3 Simulation Test Conditions.- The pilot.ed flight simulation test conditions are listed in Table 5 and designated in Figure 22. The simulation - tests were performed for calm air and moderate turbulence atmospheric condi- tions. Evaluation tasks performed for the different flight conditions are listed in Table 7.

FLIGHT CONDITIONS IDENTIFIED BY NUMBERS: SEE TABLE 5

c -

T 30

x Y =I a 18, 1 / I I I I I I 1 1 I I 0 0 320. 3 6 0 4 0 0 4 4 0 120 160 200 2 4 0 2 8 0 V, Kts F i g u r e 22. - Piloted flight simulation test conditions.

TABLE 7. - PILOTED FLIGHT SIMULATION EVALUATION TASKS Flight Conditions ~~

Evaluation Task Cruise Max. Oper. Speed Landing Holding I Take off

Wind-up turns S-pattern turns Airline operational turns X Trimmabil ity Pitch attitude change Power effects Emergency descent Short-period mode stability Phugoid mode stability I LS approach Heading change X 2.4.4 S i m u l a t i o n Test R e s u l t s . - The p i l o t used t h e Cooper-Harper r a t i n g s c a l e (NASA TND-5163) g i v e n i n F i g u r e 23 t o e v a l u a t e t h e a i r c r a f t f l y i n g q u a l i t i e s w i t h t h e PACS on and o f f . Ratings f o r each p i l o t were p l o t t e d a s shown i n F i g u r e 24. A summary of t h e test r e s u l t s i s shown i n F i g u r e 25.

With t h e PACS o f f t h e p i l o t r a t i n g s became u n a c c e p t a b l e (Cooper-Harper r a t i n g of 6.5) when t h e c.g. was n e a r t h e neutral p o i n t . However, engagement of t h e advanced PACS r e s u l t e d i n good handling q u a l i t i e s o v e r t h e e n t i r e c . g . r a n g e of from 25 t o 60 p e r c e n t rnac.

2.5 PACS System A r c h i t e c t u r e T h i s s e c t i o n shows how t h e c o n t r o l l a w s were mechanized t o p r o v i d e an advanced PACS which i s s u i t a b l e f o r performing f l i g h t tests w i t h t h e Lockheed house L-1011 ( S / N 1 0 0 1 ) . A f t c . g . l o c a t i o n c o n s t r a i n t s do n o t permit f l i g h t a t n e g a t i v e s t a t i c s t a b i l i t y margins g r e a t e r t h a n t h r e e p e r c e n t rnac.

The advanced PACS i n t e r f a c e block diagram i s shown i n F i g u r e 26. The c o n t r o l l e r i n p u t s i g n a l s from each s e n s o r element are shown on t h e l e f t s i d e of t h e f i g u r e a l o n g w i t h t h e e l e c t r i c a l power. Output s i g n a l s t o t h e series s e r v o c h a n n e l s and f a i l u r e s i g n a l s t o t h e F l i g h t C o n t r o l E l e c t r o n i c System (FCES) p a n e l are shown on t h e r i g h t s i d e of t h e i l l u s t r a t i o n .

S a f e t y p r o v i s i o n s c r i t e r i a were a s f o l l o w s . S i n g l e f a i l u r e s a r e bound t o occur and i t i s i m p o s s i b l e t o p r e d i c t e x a c t l y when t h e y w i l l happen.

T h e r e f o r e , t h e d e s i g n a i m i s t o i n c o r p o r a t e s a f e t y p r o v i s i o n s t h a t w i l l p r o t e c t t h e system a g a i n s t c r i t i c a l e f f e c t s f o r any s i n g l e f a i l u r e . A l s o , t h e f l i g h t crew needs t o be warned o f any f a i l u r e , c r i t i c a l o r n o t , so t h a t exposure t i m e f o r b u i l d up of p o s s i b l e hazardous m u l t i p l e f a i l u r e s and t h e p r o b a b i l i t y of p o t e n t i a l hazardous s i n g l e f a i l u r e s a r e a c c e p t a b l y remote.

The redundancy of t h e advanced PACS components t o comply w i t h t h i s c r i t e r i a i s shown i n F i g u r e 2 7 .

a 000 0, z w I& U z m P 0 I-

=

U n w a u- L1 z z a x P w m U w a a a 0 0 0

a a a

I E I I

v

- 0 PILOT 1

0 PILOT 4

- v PILOT 5

t

H PACS OFF 0 PACSON

-

v

v

C J

z 6 t

e a z W E 5 U T a W n E 4 u

-

0 0 0 0

I I 1 I

c.g. - % mac

Figure 2 4 . - Cooper-Harper ratings for flight condition 7, moderate turbulence.

GfflGBiVAh PAGE S S

OF POOR QUALITY

v) C .rl I a c l x a m rl C O P e a Vl- aJ

= =

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I 3 . REDUCED AREA HORIZONTAL T A I L The s m a l l t a i l program o b j e c t i v e was t o d e t e r m i n e t h e maximum d r a g b e n e f i t t h a t can b e achieved by r e d u c i n g t h e h o r i z o n t a l t a i l area which i s c o n s i s t e n t The w i t h moving t h e c . g . r a n g e a f t and t h e c o n t r o l l a b i l i t y r e q u i r e m e n t s .

reduced area h o r i z o n t a l t a i l development t a s k s were: d e s i g n c r i t e r i a d e f i n i - t i o n , r e f i n e m e n t of a n a l y s i s methods, and wind t u n n e l tests.

3.1 Design C r i t e r i a Design c r i t e r i a f o r a reduced a r e a h o r i z o n t a l t a i l i n c l u d e high-speed and low-speed c h a r a c t e r i s t i c s . The o b j e c t i v e of t h e t a i l d e s i g n d e f i n i t i o n w a s t o select a n a i r f o i l which is a s a t i s f a c t o r y compromise between t h e high-speed and low-speed o b j e c t i v e s .

3.1.1 High-speed d e s i g n c r i t e r i a . - The h i g h speed d e s i g n r e q u i r e m e n t s w e r e : e S e c t i o n d r a g c h a r a c t e r i s t i c s s h a l l b e similar t o t h e s t a n d a r d L-1011 T a i l d r a g o No t a i l drag rise s h a l l o c c u r w i t h i n t h e c r u i s e Mach number r a n g e 0 Maximum l i f t - t o - d r a g r a t i o of t h e t a i l s h a l l o c c u r a nominal c r u i s e t r i m l o a d .

3.1.2 Low-Speed d e s i p n c r i t e r i a . - The low speed d e s i g n r e q u i r e m e n t s w e r e : e Achieve n o s e wheel l i f t o f f a t forward c.g. f o r p r e s c r i b e d n o s e wheel l i f t o f f speeds.

e Have s u f f i c i e n t c o n t r o l power t o s t a l l a t forward c . g .

e Have s u f f i c i e n t c o n t r o l power f o r s t a l l r e c o v e r y a t a f t c . g .

3 . 1 . 3 S p e c i f i c d e s i g n requirements.- S p e c i f i c s m a l l h o r i z o n t a l t a i l d e s i g n r e q u i r e m e n t s were: e Takeoff nose wheel l i f t o f f a t forward c.g. w i t h maximum t a k e o f f f l a p s a t t h e lesser of 1.05 minimum c o n t r o l speed o r t h e FAA s t a l l speed e C o n t r o l - t o - s t a l l a t forward c.g. w i t h maximum l a n d i n g f l a p s ( 4 2 deg) and i d l e t h r u s t A t l e a s t 4.58 d e g / s e c 2 nose-down p i t c h a c c e l e r a t i o n f o r s t a l l recovery a t a f t c . g . and a t t h e FAA s t a l l speed f o r maximum l a n d i n g weight A t l e a s t n e u t r a l s t a b i l i t y a t a f t c . g .

A c.g. r a n g e e q u a l t o t h a t of t h e s t a n d a r d L-1011 (12 t o 35 per- c e n t m a c ) which i s 6 7 . 5 i n c h e s f o r a i r c r a f t w e i g h t s of 338,000 pounds.

3.2 Small T a i l C o n f i g u r a t i o n s Evaluated The small t a i l c o n f i g u r a t i o n s t h a t were e v a l u a t e d a r e l i s t e d below.

S m a l l t a i l , s t a n d a r d f u s e l a g e (L-1011-1) l e n g t h , Wortmann A i r f o i l

H16 - (MFX 69-H-098-090-1)

H i 7 - S m a l l t a i l , s h o r t f u s e l a g e (L-1011-500) l e n g t h , RSS2 A i r f o i l c o n f i g u r a t i o n Hi8 - S m a l l t a i l , s h o r t f u s e l a g e (L-1011-500) l e n g t h , RSS2 A i r f o i l c o n f i g u r a t i o n Hi9 - S m a l l t a i l , NASA c o n f i g u r a t i o n

H ~ c - Reference t a i l , s t a n d a r d L-1011

F i g u r e 2 8 shows p l a n views of t h e h o r i z o n t a l t a i l s and Table 8 g i v e s compara- t i v e geometric d a t a .

3.3 S m a l l T a i l Design Procedures A c o n v e n t i o n a l a i r p l a n e h o r i z o n t a l t a i l i s s i z e d t o p r o v i d e a s p e c i f i e d margin of s t a t i c s t a b i l i t y and t h e r e q u i r e d l o n g i t u d i n a l c o n t r o l by having a n a d e q u a t e C h a x i n down l i f t . The requirement f o r a s p e c i f i c p o s i t i v e margin of s t a t i c s t a b i l i t y r e s u l t s i n a l a r g e s t a b i l i z e r s u r f a c e and forward c e n t e r - o f - g r a v i t y range which p e n a l i z e performance i n terms of d r a g and w e i g h t .

I f a p i t c h a c t i v e c o n t r o l system i s i n c o r p o r a t e d i n t o t h e a i r p l a n e t o p r o v i d e s t a t i c s t a b i l i t y a r t i f i c i a l l y , t h e n t h e h o r i z o n t a l t a i l can be s i z e d t o provide t h e r e q u i r e d p i t c h c o n t r o l by u s i n g t h e t a i l maximum l i f t c a p a b i l i t y i n b o t h t h e up and down d i r e c t i o n s .

Analysis methods used f o r a n a l y z i n g t h e d i f f e r e n t s m a l l h o r i z o n t a l t a i l c o n f i g u r a t i o n s are d i s c u s s e d i n t h e f o l l o w i n g s e c t i o n s .

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I 4 3 TABLE 8. - SMALL HORIZONTAL T A I L COMPARATIVE DATA H17 H18 19 H8c H16 Small Small Small Standard Small 4.5 4 4 4.5 Aspect ratio 4 0.33 0.33 0.33 0.33 Taper ratio 0.33 0.013 mac 0.016 rnac 0.016 rnac Camber 0 0.015 rnac 0.0089 rnac 0.006 mac 0.036 mac 0.036 mac Leading-edge radius 0.10 Thickness ratio 0.09 0.09 0.1045 0.1045 35 deg 25 deg 25 deg Quarter chord sweep 35 deg 28 deg 800 898 Total Area ft2 1282 800 552 552 652 Exposed Area f t 960 0.72 0.69 0.73 0.75 0.69 Exposed/Total 0.3 mac 0.3 mac 0.3 mac 0.25 rnac 0.3 mac Elevator chord ratio 17 deg 20 deg 17 deg Stabilizer throw 15 deg 2 0 deg 35 deg 4 0 deg 4 0 deg 35 deg Elevator throw 25 deg 3.3.1 H o r i z o n t a l t a i l a n a l y s i s . - The Hi6 t a i l development was t h e i n i t i a l e f f o r t t o develop a s m a l l L-1011 h o r i z o n t a l t a i l and w a s accomplished w i t h Lockheed funds (Reference 6 ) . T h i s work i s summarized i n t h i s r e p o r t t o p r o v i d e a f o u n d a t i o n of t h e work t h a t was accomplished under t h e ACEE c o n t r a c t .

A h e l i c o p t e r r o t o r b l a d e a i r f o i l (Wortmann MFX 69-H-098-090-1, Reference 7 ) was s e l e c t e d and s l i g h t l y modified f o r t h e sma1.1 t a i l a i r f o i l becadse of i t s good high-speed and h i g h - l i f t c h a r a c t e r i s t i c s . The r e d u c t i o n i n aerodynamic drag w a s e s t i m a t e d by u s i n g s t a n d a r d handbook methods f o r l i f t i n g s u r f a c e s .

A form f a c t o r f o r t h e a i r f o i l t h i c k n e s s w a s a p p l i e d t o t h e p l a n a r s u r f a c e c o m p r e s s i b l e s k i n f r i c t i o n drag which w a s computed by t h e Sommer and S h o r t T ' method. The form f a c t o r w a s determined by a s p e c i a l Lockheed c o r r e l a t i o n of a i r f o i l drag w i t h t h i c k n e s s r a t i o .

The a n a l y s i s showed t h a t t h e s m a l l h o r i z o n t a l t a i l (800 f t 2 a r e a ) d r a g r e d u c t i o n was 11 d r a g c o u n t s a t wind t u n n e l test c o n d i t i o n s and 7 c o u ? t s a t f u l l - s c a l e c r u i s e c o n d i t i o n s , T h i s r e p r e s e n t e d a p o t e n t i a l d r a g s a v i n g s of 2.7 percent a t nominal c r u i s e c o n d i t i o n s . The n e t improvement i n c r u i s e e f f i c i e n c y would b e approximately t h r e e p e r c e n t due t o t h e weight r e d u c t i o n of t h e s m a l l e r h o r i z o n t a l t a i l .

1117 H o r i z o n t a l t a i l a n a l y s i s . - Advanced a i r f o i l technology was

3.3.2 used t o d e s i g n a r e l a t i v e t h i c k a i r f o i l w i t h l a r g e leading-edge r a d i u s t o p r o v i d e l i f t a t low speed f o r c o n t r o l l a b i l i t y w i t h o u t degrading h i g h speed d r a g c h a r a c t e r i s t i c s . P r e s s u r e d i s t r i b u t i o n of t h e Hi7 a i r f o i l ( d e s i g n a t e d RSS2 A i r f o i l ) w a s determined by u s i n g t h e Jameson-Caughey e x a c t p o t e n t i a l i n v i s c i d flow a n a l y s i s code, Flow 22 (References 6 and 7 ) . A f u l l y automated c u r v a t u r e a i r f o i l shaping design system ( F i g u r e 29) was used t o o p t i m i z e t h e s m a l l t a i l a i r f o i l .

3.3.3 a8 H o r i z o n t a l t a i l a n a l y s i s . - T h i s s m a l l h o r i z o n t a l t a i l (898 f t 2 a r e a ) f o r both short-body and long-body L-1011 i s a common s i z e t a i l d e r i v a t i v e s .

The d e s i g n g o a l was t o provide a C L ~ ~ ~ = -1.4 f o r f u l l s c a l e f l i g h t c o n d i t i o n s . The handbook methods used f o r p r e d i c t i n g d r a g of t h e HI6 t a i l w e r e 27 p e r c e n t lower t h a n measured i n t h e wind t u n n e l even though t h e same method had p r e d i c t e d d r a g of t h e s t a n d a r d I,-1011 t a i l (Hac) c o r r e c t l y .

T h e r e f o r e , d r a g of t h e H i 8 t a i l was e s t i m a t e d by a p p l y i n g a n exposed a r e a cor- r e c t i o n t o p r e v i o u s wind-tunnel t e s t r e s u l t s of t h e Hi6 t a i l .

t h a t t h e L-1011 d r a g would be reduced by about The e s t i m a t i o n showed 6 c o u n t s a t wind t u n n e l test c o n d i t i o n s .

Thus, a two p e r c e n t L/D b e n e f i t was p r e d i c t e d based on reduced a r e a and l o w e r weight of t h e t a i l .

3.3.4 1 1 9 H o r i z o n t a l t a i l a n a l y s i s . - A review of independent develop- ment of s m a l l h o r i z o n t a l t a i l d e s i g n s by NASA and Lockheed r e v e a l e d s i g n i - f i c a n t d i f f e r e n c e s . The f o l l o w i n g d i f f e r e n c e s w e r e i d e n t i f i e d : Lockheed NASA Sweep Angle (mac/4) 32.5 deg 25 deg Aspect R a t i o 3 . 0 4.5 A i r f o i l S e c t i o n s NASA developed Lockheed Developed i n v e r s e camber i n v e r s e camber & symmetrical A i r f o i l Thickness 10% 10.5% A f t e r reviewing t h e d a t a it w a s concluded t h a t d r a g c r e e p problem e n c o u n t e r e d w i t h t h e Lockheed c o n f i g u r a t i o n c o u l d be e l i m i n a t e d by: 1) i n c r e a s i n g t h e sweep a n g l e of t h e t a i l , 2) u s i n g one of t h e NASA a i r f o i l s e c t i o n s which i s one-half p e r c e n t t h i n n e r , and 3) r e t a i n i n g t h e 4 . 5 a s p e c t r a t i o planform. Of t h e two NASA a i r f o i l s , t h e symmetrical s e c t i o n w a s s e l e c t e d because of i t s good low-speed p r o p e r t i e s .

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It w a s f u r t h e r decided t o determine a n a p p r o p r i a t e sweep a n g l e f o r t h e t a i l by means of v i s c o u s Jameson-Caughey a n a l y s i s , and based on h o r i z o n t a l t h e r e s u l t s of t h a t a n a l y s i s t o proceed w i t h high-speed model c o n s t r u c t i o n and t e s t i n g .

3.4 Wind Tunnel Tests S m a l l t a i l wind t u n n e l tests which were performed are l i s t e d i n T a b l e 9.

The Hi6 tests were Lockheed funded, t h e H 1 7 tests were funded by t h e f i r s t ACEE a c t i v e c o n t r o l technology c o n t r a c t awarded t o Lockheed (NASI-14690) .

The remainder of tests were funded under t h e c u r r e n t c o n t r a c t (NASI-15326) - The wind t u n n e l test r e s u l t s i s d i s c u s s e d i n t h e next s e c t i o n .

3.5 Wind Tunnel T e s t R e s u l t s The s m a l l h o r i z o n t a l drag b e n e f i t s and h i g h l i f t c h a r a c t e r i s t i c s are p r e s e n t e d i n F i g u r e s 30 and 31 r e s p e c t i v e l y .

The HI6 t a i l shows a d r a g c o u n t r e d u c t i o n of 7 ( F i g u r e 30) o v e r t h e c r u i s e r a n g e of Mach .8 t o .83. However, t h e low-speed h i g h - l i f t c h a r a c t e r - i s t i c s d i d n o t m e e t t h e C h = -1.4 t a r g e t ( F i g u r e 31).

ax The H17 t a i l low-speed h i g h - l i f t c h a r a c t e r i s t i c s d i d n o t meet t h e C L ~ ~ t a r g e t . Consequently, h i g h speed t e s t s w e r e n o t performed.

The H 1 8 t a i l d r a g count r e d u c t i o n a t Mach .8 w a s a b o u t 5 d r a g c o u n t s , The C L ~ ~ ~ t a r g e t w a s b u t a d r a g - c r e e p occured throughout t h e Mach r a n g e .

a c h i e v e d .

The The H19 t a i l shows a d r a g count r e d u c t i o n of 7 and a C h a x = - 1 . 2 .

d r a g r e d u c t i o n and corresponding weight of t h e Hi9 t a i l p r o v i d e s a n e s t i m a t e d U s e of the H 1 9 t a i l on a n L-1011 r e q u i r e s t h e L / D b e n e f i t of two p e r c e n t .

forward c.g. l i m i t b e r e s t r i c t e d , Use o f t h i s t a i l on a n e x t g e n e r a t i o n t r a n s p o r t may be f e a s i b l e i f t h e a i r c r a f t i s equipped w i t h a PACS and h a s a p r o p e r l y d e s i g n e d c.g. range.

TABLE 9. - SMALL HORIZONTAL T A I L WIND TUNNEL TESTS Date Test No. Wind Tunnel Type of Test H8C "16 H17 H18 H19 Apr 76" N-307 Calspan 8 f t Limited High-speed Force Data X TPT in Cruise-No Elevator X X Nov 76" La04 Calac LSWT LowSpeed Force Data a t Low Reynolds Number Nov 78"" L429 Calac LSWT LowSpeed Force & Pressure X Data a t Low Reynolds Number Mar 79"" L-442 Calac LSWT Complete LowSpeed Force X Data a t Low Reynolds Number July 79 N-336 Langley 8 f t Complete High-speed Force X X TPT & H.T. Pressure Data Sept 79 N-340 Calspan 8 f t Complete High-speed Force & X X TPT Data H .T. Pressure Jan 80 N-337 Ames 12 f t Complete Low-Speed Force X X PT Data a t High Reynolds Number Feb 80 S.387 Calac 4 f t Horizontal Tail Drag a t X X X X TIST Cruise Mach Number Aug 81 N-369 Calspan 8 f t Complete High-speed Force X TPT Data a t Cruise Mach Number Jan 82 N-337 Ames 12 f t Low-Speed Force Data a t X PT High Reynolds Number "Lockheed Funded "*NASA Contract NAS1 -1 4690 H ~ c STANDARD TAIL

-

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I I 1 18 18 4 -24 d. HIB TAIL d. HIB TAIL 0.8 c. H,, TAIL Figure 31. - Comparison of small horizontal tail high-lift characteristics.

CONCLUSIONS T h i s program h a s demonstrated by p i l o t e d f l i g h t s i m u l a t i o n tests (based PACS w i l l p r o v i d e h a n d l i n g q u a l i - on L - 1 0 1 1 aerodynamic d a t a ) t h a t a n advanced t i e s a t n e g a t i v e s t a b i l i t y margins up t o 20 p e r c e n t which are e q u i v a l e n t t o t h e h a n d l i n g q u a l i t i e s o f t h e b a s e l i n e a i r p l a n e a t a p o s i t i v e 15 p e r c e n t sta- b i l i t y margin. Also, i t h a s shown t h a t t h e modal c o n t r o l method of modern c o n t r o l t h e o r y can b e used f o r c o n t r o l l a w s y n t h e s i s o f a m u l t i p l e feedback l o o p PACS and p r o v i d e s a v a l i d c o n t r o l l a w t o c o n t r o l t h e dynamic s t a b i l i t y .

The p i l o t e d f l i g h t s i m u l a t i o n t e s t s demonstrated t h a t t h e PACS r e q u i r e s a feed-forward l o o p t o p r o v i d e t h e d e s i r e d c o n t r o l column g r a d i e n t s f o r maneu- v e r s t a b i l i t y and a Mach compensation loop t o p r o v i d e t h e d e s i r e d column f o r c e s f o r speed s t a b i l i t y .

The s m a l l h o r i z o n t a l t a i l program has demonstrated by wind t u n n e l tests t h a t a 30 p e r c e n t t a i l area r e d u c t i o n ( r e l a t i v e t o s t a n d a r d L - 1 0 1 1 t a i l ) pro- v i d e s a n i n c r e a s e i n c r u i s e e f f i c i e n c y of about two p e r c e n t and t h a t a 38 per- c e n t t a i l area r e d u c t i o n p r o v i d e s a n i n c r e a s e i n c r u i s e e f f i c i e n c y of about t h r e e p e r c e n t . However, forward c.g. l i m i t a t i o n s would have t o b e imposed on t h e a i r c r a f t because t h e maximum h o r i z o n t a l t a i l l i f t g o a l w a s n o t achieved and s u f f i c i e n t a i r c r a f t nose-up c o n t r o l a u t h o r i t y w a s n o t a v a i l a b l e . T h i s l i m i t a t i o n would n o t be r e q u i r e d f o r a p r o p e r l y d e s i g n e d new a i r c r a f t .

P o t e n t i a l f u e l s a v i n g s f o r a f u t u r e t r a n s p o r t a i r c r a f t t h a t h a s a s m a l l f l i e s a t n e g a t i v e s t a t i c s t a b i l i t y margins would b e about h o r i z o n t a l t a i l and s i x p e r c e n t . T h i s a r i p l a n e would have t o b e equipped w i t h a h i g h - r e l i a b i l i t y PACS t o p r o v i d 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 .

i'KECEDING PAGE BLeNK NOT FILMED REFERENCES 1. Urie, D.M., "Acceleration Development and Flight Evaluation of Active

Controls Concepts for Subsonic Transport Aircraft - Volume 11: Aft

CG Simulation and Analysis," NASA CR 159098, September 1979.

2.

Guinn, W.A., "Development and Flight Evaluation of an Augmented Stability Active Controls Concept," NASA CR 165951, September 1, 1982.

3. Guinn, W.A., Willey, C.S., and Chong, M.G., "Extended Flight Evaluation of a Near-Term Pitch Active Controls System," NASA CR 17226, December 21, 1983.

4 .

Rising, Jerry J., "Development of a Reduced Area Horizontal Tail for a Wide Body Jet Aircraft," NASA CR-172278, February 1, 1984.

5. Urie, D.M., and Passer, J.S., "Aerodynamic Development of a Small Horizontal Tail for an Active Control Released Stability Transport Application," A I M Paper 79-1653, August 6, 1979.

6 . Bingham, Gene J., and Noonan, Kevin W., "Low Speed Aerodynamic Charac- teristics of Five Helicopter Blade Sections at Reynolds Numbers from 2.4 x 106 to 8.4 x 106, ' I NASA TMX-2467, 1972.

PRECEDING PAGE BL4NK NOT FKmD

2. Government Accession No. 3. Recipient's Catalog No.

I. Report No.

NASA CR - 172283 5. Report Date 1. Title and Subtitle Development o f a Advanced P i t c h Active C o n t r o l System and February 1, 1984 - a Reduced Area H o r i z o n t a l T a i l f o r a Wide-Body Jet. A i r c r a f t 6. performing Organization Code E x e c u t i v e Summary 8. Performing Organization Report NO.

7. Author(s) LR 30463 Wiley A . Guinn 10. Work Unit No.

9. Performing Organization Name and Address 11. Contract or Grant NO.

Lockheed C a l i f o r n i a Company NAS 1-15326 Burbank. CA 13. T pe of Report and Period Covered s o n t r a ct o r Rep0 r t 2. Sponsoring Agency Name and Address Dec 1978-April 1983 N a t i o n a l Aeronautics and Space A d m i n i s t r a t i o n 14. Sponsoring Agency Code Washington, D . C . 20546

I

I 5. Supplementary Notes Langley Technical Monitor: Dennis W. Bartlett 6. Abstract T h i s r e p o r t documents work t h a t w a s accomplished toward development o f a advanced p i t c h a c t i v e c o n t r o l system (PACS) and a reduced area h o r i z o n t a l t a i l f o r a wide-body j e t t r a n s p o r t (L-1011) w i t h a f l y i n g h o r i z o n t a l s t a b i l i z e r . The advanced PACS c o n t r o l l a w d e s i g n o b j e c t i v e s were t o p r o v i d 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 f o r a f t c . g .

f l i g h t c o n d i t i o n s t o n e g a t i v e s t a t i c s t a b i l i t y margins of 10 p e r c e n t and t o p r o v i d e goo1 maneuver c o n t r o l column f o r c e g r a d i e n t s f o r n o n l i n e a r s t a b i l i t y f l i g h t c o n d i t i o n s .

V a l i d i t y o f t h e c o n t r o l l a w s w e r e demonstrated by p i l o t e d f l i g h t s i m u l a t i o n tests on thi NASA Langley V i s u a l Motion S i m u l a t o r . 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 were a c t u a l l y demonstrated t o a n e g a t i v e 20 p e r c e n t s t a t i c s t a b i l i t y margin. The PACS c o n t r o l l a w s f o r a n L-1011 w e r e mechanized t o p r o v i d e t h e system a r c h i t e c t u r e t h a t would b e s u i t a b l e f l i g h t test program t o a n e g a t i v e s t a b i l i t y margin of 3 p e r c e n t which r e p r e s e n t s t h e a f c . g . l i m i t s of t h e a i r c r a f t . Reduced area h o r i z o n t a l t a i l d e s i g n s of 30 and 38 p e r c e n t w i t h r e s p e c t t o t h e L-1011 s t a n d a r d t a i l w e r e d e s i g n e d f a b r i c a t e d and wind t u n n e l t e s t - ed. Drag r e d u c t i o n s and weight s a v i n g s of t h e 30 p e r c e n t smaller t a i l would p r o v i d e a n L I D b e n e f i t of about 2 p e r c e n t and t h e 38 p e r c e n t s m a l l t a i l L I D b e n e f i t would b e a b o u t 3 p e r c e n t . However, forward c . g . l i m i t a t i o n s would have t o b e imposed on t h e a i r c r a f t because t h e maximum h o r i z o n t a l t a i l l i f t g o a l w a s n o t a c h i e v e d and s u f f i c i e n t a i r c r a f t nose-up c o n t r o l a u t h o r i t y w a s n o t a v a i l a b l e . T h i s l i m i t a t i o n would n o t be r e q u i r e d f o r a p r o p e r l y designed new a i r c r a f t .

~ ~~~ ~ ~~ 17. Key Words (Suggested by Author(s)) 18. Distribution Statement A c t i v e C o n t r o l S y s t e m , C o n t r o l System, P i t c h C o n t r o l , L o n g i t u d i n a l C o n t r o l , A i r c r a f t F u e l Savings 22. Price' 21. No. of Pages 19. Security Classif. (of this report) 20. Security Classif. (of this pap)

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