Application Hybrid Laminar Flow Control to Global Range ... · Control to Global Range Military...

NASA Contractor Report 181638

Application of Hybrid Laminar Flow Control to Global Range Military Transport Aircraft

Roy H. Lange

Lockheed Aeronautical Systems Company A Division of Lockheed Corporation Marietta, Georgia 30063

Contract NAS 1-1 8036 April' 1088 ___ ~ . ~ __ -- - (NASA-CR-181638) APPLICATION OF H Y B R I D N88-21124

L A H I N A R FLOW C O N T R O L TO GLOBAL 6 A N G E HPLITARY TRANSPORT AIRCRAFT :Lockheed Aeronautical S y s t e n s C o . ) 101 p C S C L O 1 A Unclas

G3/02 8136058

National Aeronautics and Space Administration . La~lw Resoarch Center Harnpton, Virginia 23665-5225

https://ntrs.nasa.gov/search.jsp?R=19880011740 2018-05-31T23:17:49+00:00Z

FOREWORD

"he t e c h n i c a l work h e r e i n was accomplished under Task 2 of Con t rac t

NAS1-18036 s p o n s o r e d by NASA-Langley, L a n g l e y R e s e a r c h Center , Hampton, V i r g i n i a 23665-5225 and t h e Air Force F l i g h t Dynamics Labora to ry , WPAFB, Ohio. Task 2 go-ahead was i n i t i a t e d i n September 1986. Mr. Richard D. Wagner, Manager, Laminar Flow Cont ro l P r o j e c t O f f i c e , NASA-Langley, was t h e Task Manager , Mr. D. V. Maddalon was t h e T e c h n i c a l R e p r e s e n t a t i v e o f t h e Con t rac t ing O f f i c e , and Mr. Russ Osborn was t h e Technica l R e p r e s e n t a t i v e from t h e Air Force Wright Aeronau t i ca l L a b o r a t o r i e s , AFWAL/FIMM, WPAFB, Ohio.

Mr. Roy H. Lange, Lockheed Program manager, LFC P r o j e c t s , i s r e s p o n s i b l e

f o r t h e o v e r a l l gu idance and t e c h n i c a l d i r e c t i o n o f t h e Laminar Flow Enabl ing Technology Developnent C o n t r a c t , NAS1-18036 and a l s o se rved a s Task 2 Study

Manager. The t e c h n i c a l work f o r Task 2 was performed under t h e d i r e c t i o n o f t h e P re l imina ry Design D i v i s i o n (D/72-90) o f t h e Lockheed Aeronau t i ca l Systems

Company i n M a r i e t t a , Georgia . The key t e c h n i c a l people working du r ing t h i s r e p o r t i n g pe r iod i n c l u d e :

R. H. Lange

J. A . B e n n e t t

B. J. Hagemann J. F. Honrath R. W . P a t t e r s o n T. K. Randall

A. S. Woolf J. B. Clayton

Program and Study Manager

Aerodynamics Department P ropu l s ion & Acous t i c s k p a r t m e n t Aeronau t i ca l Systems Developnent Dept . Advanced S t r u c t u r e s Department

Aeronau t i ca l Systems Developnent Dept . Aeronau t i ca l Systems Developnent Dept . Engineering Technica l C o n t r a c t s Dept .

This r e p o r t has been g iven t h e Lockheed d e s i g n a t i o n LG87ER0145.

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TABLE OF CONTENTS

S e c t i o n

1.0

2 .0

3.0

4.0

5.0

FOREWORD

LIST OF FIGURES

LIST OF TABLES

SUMMARY

INTRODUC TI0 N

T i t l e

SYMBOLS A N D ABBREVIATIONS

STUDY APPROACH

4.1 Study O b j e c t i v e s

4.2 Study Plan

4 .3 Reference Technology Level

4.3.1 Aerodynamics

4.3.2 F l i g h t Cont ro ls

4.3.3 P ropu l s ion Systems

4.3.4 S t r u c t u r e s and M a t e r i a l s

4.3.5 HLFC Systems

4.3.6 A i r c r a f t Systems

BASELINE COW IGURATION DEVELOPMENT

5.1 Turbulent Flow A i r c r a f t S e l e c t i o n

5.2 HLFC Groundrules

5.3 HLFC A i r c r a f t S e l e c t i o n

5.4 Comparison o f ' h r b u l e n t and HLFC A i r c r a f t

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TABLE OF CONTENTS (CONT'D)

Section Title

5.5 HLFC A i r c r a f t D e f i n i t i o n

5.5.1 Conf igu ra t ion Design

5.5.2 HLFC Wing Aerodynamics Design

6.0 CONFIGURATION SENSITIVITY STUDIES

6.1 HLFC A i r c r a f t S e n s i t i v i t y S t u d i e s

6.1.1 I n c r e a s e of Cru i se Speed o r A l t i t u d e

6.1.2 E l i m i n a t i o n o f HLFC on h p e n n a g e

6. 1.3 E l imina t ion o f HLFC on bwer Wing S u r f a c e

6.1.4 Reduct ion of Aspect Rat io t o 10

6. 1.5 High Wing HLFC Conf igu ra t ion

6 .2 T u r b u l e n t Flow A i r c r a f t S e n s i t i v i t y S t u d i e s

6.2.1 I n c r e a s e o f A l t i t u d e t o 36,000 Fee t

6.2.2 I n c r e a s e o f Payload t o 212,000 Pomds

6.2.3 Reduction o f Aspect Ra t io t o 10

7.0 ASSESSMENT OF HLFC BENEFITS AND SELECTED CONF IGURATIONS

8.0 CO NC L US IONS AND REC OMM ENDAT IONS

REFERENCES

APPENDIX A GENERAL AIRCRAFT SIZING PROGRAM

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LLST OF FIGURES

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Figure

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1 OA

1OB

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1 9 B

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Title

Hybrid Laminar Flow Cont ro l Concept

Study Plan

Mission m a r a c t e r i s t i c s

Mission P r o f i l e

Wing Sec t ion Design Cbrves

Example o f Secondary Act ive T r a i l i n g Edge Flaps

S p e c i f i c F u e l Consumption (PLW STF686

Bare Engine Weight (PLW STF686)

Advanced Mate r i a l Technology f o r Hybrid LFC S tudy

Advanced M a t e r i a l U t i l i z a t i o n

Hybrid L F C Design

Hybrid LFC Design

Wing Leading Edge S l o t Loca t ions

S l o t and Ducting Cross Sec t ion

S u c t i o n System Schematic

Suct ion Punp

Wing Suc t ion System Schematic

Cleaning/ Ant i - ic ing System

Purge System Schematic

Nitrogen P r e s s u r i z a t i o n System

Turbu len t Baseline Pa rame t r i c Data; M = 0.77

Con c l ud ed

Turbu len t Flow B a s e l i n e Design Concept

HLFC A i r c r a f t Ground Rules

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F i g u r e

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LIST OF FIGURES (CONT'D)

T i t l e

Concluded

Paramet r ic S iz ing Data fo r HLFC A i r c r a f t ; M = 0.77, I n i t i a l Concepts

Co n c l ud ed

P a r m e t r i c S iz ing Data fo r HLFC A i r c r a f t ; Speed and A l t i t u d e Changes

Concluded

H L F C Base l ine Concept

Hor izonta l T a i l S iz ing Chart

Typica l HLFC Concept Payload-Range Curve

HLFC Concept Cargo P a l l e t s Arrangement

HLFC Concept Arrangement of Vehicles

Suct ion Pump/System General Arrangement

HLFC Concept Suct ion Systems Arrangement

Concluded

Leading Edge Design Concept

F u e l Tank Arrangement

Enpennage General Arrangement

Main Landing Gear

Wing T r a i l i n g Edge Design

General ized A i r c r a f t Si zing Program, LFC Subrout ine

Areas o f HLFC and Wing Lamina r i za t ion

Aerodynamic Design Procedure f o r LFC and ?I: Wings

Typical A i r f o i l Sec t ion on Base l ine HLFC Wing

R e p r e s e n t a t i v e Wing C D i s t r i b u t i o n From Wing S t a t i o n 'I = 0.488 P

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Figure

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LIST OF FIGURES (CONT'D)

Title

Mass Flow S u c t i o n Level

To ta l Suc t ion Flow Across h e Wing o f HLFC B a s e l i n e

Crossf low Dis turbance N F a c t o r , Wing S t a t i o n 7 = 0.488

S e n s i t i v i t y Data

Co ncl ud ed

S i z i n g Data f o r Aspect Ra t io 10 A i r c r a f t

Concluded

Comparison o f HLFC Base l ine and HLFC Aspect R a t i o 10 A i r c r a f t

Plan V i e w o f Laminar Flow Area Loss f o r Wing Mounted Engine Conf igu ra t ion ; Lower Wing S u r f a c e

S i z i n g Data f o r High Wing n r b u l e n t Flow and HLFC A i r c r a f t ; Sweepback 20

Concluded

General Arrangement Drawing o f High Wing HLFC A i r c r a f t

Turbu len t Flow A i r c r a f t S i z ing Data

Concluded

General Arrangement o f 21 2,000 l b Payload n r b u l e n t Flow A i r c r a f t

Comparison o f l b r b u l e n t Flow 21 2,000 l b Payload A i r c r a f t With C-5

Comparison o f l b r b u l e n t Flow B a s e l i n e and Aspect R a t i o 10 A i r c r a f t

Summary o f HLFC A i r c r a f t R e s u l t s R e l a t i v e t o Turbu len t Flow Base1 i n e

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Table - 1

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LIST OF TABLES

T i t l e

Upper Surface Design Data

Lower Surface Design Data

Leading Edge Metering System

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1.0 SUHMARY

T h i s r e p o r t summarizes t h e resul ts o f a s t u d y conducted by Lackheed under

NASA Cont rac t NAS1-18036 t o e v a l u a t e t h e a p p l i c a t i o n o f hybr id laminar f low

c o n t r o l (HLFC) t o g l o b a l r ange m i l i t a r y t r a n s p o r t a i r c r a f t .

By mutual agreement among NASA, t h e Air Force, and Lockheed t h e g l o b a l mi s s ion o f t h e a i r c r a f t i n t h i s s t u d y i n c l u d e d t h e c a p a b i l i t y t o t r a n s p o r t

132,500 pounds o f payload 6,500 n a u t i c a l miles, l a n d and d e l i v e r t h e payload and wi thout r e fue l ing r e t u r n 6,500 n a u t i c a l miles t o a f r i e n d l y a i r b a s e . The

d e s i g n c r u i s e Mach number f o r t h e miss ion i s M = 0.77. Bath t u r b u l e n t f low and h y b r i d l amina r flow c o n t r o l a i r c r a f t were s i z e d t o perform t h e g l o b a l

mission. A b a s e l i n e t u r b u l e n t f low a i r c r a f t was used a s t h e reference

a i r c r a f t f o r comparison w i t h t h e HLFC a i r c r a f t concep t s . The h y b r i d LFC

concep t restricts t h e a c t i v e s u c t i o n system t o t h e r e g i o n ahead o f t h e f r o n t

s p a r , i .e . , 15 percent wing chord and t h e remainder o f t h e a i r f o i l is t a i l o r e d ae rodynamica l ly t o a c h i e v e t h e maximum e x t e n t o f l amina r f low, which is

expected t o ex tend t o about 50 p e r c e n t chord .

O r i g i n a l l y i n t e n d e d a s a s i x month s t u d y , t h e scope was expanded a s

i n i t i a l r e s u l t s were o b t a i n e d t o i n c l u d e a d d i t i o n a l c o m p a r i s o n s and

s e n s i t i v i t y r u n s . This expans ion i n scope h a s provided f o r a better

unde r s t and ing o f t h e resul ts and f o r a more complete d a t a base o f a i r c r a f t d e s i g n and performance parameters. Thus, t h e p r e l i m i n a r y a i r c r a f t d e s i g n

s t u d y has g e n e r a t e d a c o n s i d e r a b l e m o u n t o f s i z i n g d a t a f o r bo th t u r b u l e n t flow and h y b r i d L F C a i r c r a f t c o n c e p t s .

P r e l i m i n a r y d e s i g n system s t u d i e s of t h e a p p l i c a t i o n o f hybr id laminar

f low c o n t r o l t o m i l i t a r y t r a n s p o r t s s i z e d . t o perform g l o b a l r a n g e m i s s i o n

c h a r a c t e r i s t i c s show s i g n i f i c a n t performance benefits o b t a i n e d f o r t h e hybr id

LFC a i r c r a f t as compared t o c o u n t e r p a r t t u r b u l e n t f low a i rc raf t . The s t u d y

results a t M = 0.77 show t h a t t h e l a r g e s t b e n e f i t s of HLFC a r e o b t a i n e d with a

h igh wing w i t h engines on t h e wing c o n f i g u r a t i o n . As compared t o t h e

t u r b u l e n t f l ow b a s e l i n e a i r c r a f t , t h e h i g h wing HLFC a i r c r a f t shows 17 percent

r e d u c t i o n i n fuel b u r n e d , 19.2 percent i n c r e a s e i n l i f t - t o - d r a g r a t i o , an i n s i g n i f i c a n t i n c r e a s e i n o p e r a t i n g weight, and 7.4 p e r c e n t r e d u c t i o n i n g r o s s

1

weight . For t h i s h i g h wing c o n f i g u r a t i o n , t h e performance d a t a a r e based on

t h e assumption t h a t t h e r e i s no loss i n l aminar flow on t h e upper wing s u r f a c e w i t h e n g i n e s mounted on t h e wings. It is f e l t t h a t t h i s is an o p t i m i s t i c

a s s u n p t i o n e s p e c i a l l y f o r t h e longe r laminar r u n s f o r t h e HLFC c o n d i t i o n s o f

t h i s s t u d y and f o r t h e mul t i -engine c o n f i g u r a t i o n s . The second bes t HLFC

c o n f i g u r a t i o n is t h e low wing, fuselage mounted arrangement w i t h no HLFC on t h e empennage. This c o n f i g u r a t i o n shows 13.7 p e r c e n t r e d u c t i o n i n f u e l burned, 18.2 p e r c e n t i n c r e a s e i n l i f t - t o - d r a g r a t i o , 5.4 p e r c e n t i n c r e a s e i n o p e r a t i n g we igh t , and 4 .2 p e r c e n t r e d u c t i o n i n g r o s s weight a s compared t o t h e t u r b u l e n t flow a i rc raf t .

S e n s i t i v i t y s t u d i e s i n c l u d e t h e d e t e r m i n a t i o n o f t h e e f f e c t s on per-

formance of i n c r e a s e i n c r u i s e Mach number from 0.77 t o 0.80, i n c r e a s e i n i n i t i a l c r u i s e a l t i t u d e t o 36,000 f e e t , and e l i m i n a t i o n o f HLFC on t h e lower

wing s u r f a c e . These changes g e n e r a l l y r e s u l t e d i n d e g r a d a t i o n i n performance a s compared t o t h e b a s e l i n e a i r c r a f t c h a r a c t e r i s t i c s . As expected, t h e

r e d u c t i o n i n a s p e c t r a t i o from t h e b a s e l i n e v a l u e o f about 13 t o a v a l u e o f 10

reduced t h e benef i t s f o r fuel consumption and l i f t - t o - d r a g r a t i o .

I n view o f t h e s u p e r i o r performance o f t h e h igh wing wi th e n g i n e s mounted on t h e wing HLFC c o n f i g u r a t i o n , it is recommended t h a t f u r t h e r r e s e a r c h and

deve lopnen t be conducted t o p rov ide t h e n e c e s s a r y d a t a base f o r v a l i d a t i o n o f

t h e effects o f e n g i n e o p e r a t i o n on l aminar boundary l a y e r t r a n s i t i o n f o r

f l i g h t Reynolds numbers co r re spond ing t o l a r g e , long r ange t r a n s p o r t a i r c r a f t .

Opera t ion a t h i g h e r a l t i t u d e s o f 36,000 f ee t and above a r e more f a v o r a b l e t o t h e a t t a i n m e n t and p r e s e r v a t i o n o f l a n i n a r flow. The d a t a i n t h i s s t u d y show

a moderate i n c r e a s e i n l i f t - t o - d r a g r a t i o f o r o p e r a t i o n a t i n i t i a l c r u i s e a l t i t u d e o f 36,000 feet , b u t w i th an a t t e n d a n t l a r g e i n c r e a s e i n e n g i n e t h r u s t

f o r t h e r e l a t i v e l y h i g h by p a s s r a t i o e n g i n e s used i n t h i s s t u d y . It i s recomnended t h a t a d d i t i o n a l s t u d i e s be made o f h igh a l t i t u d e o p e r a t i o n s wi th

lower by p a s s r a t i o e n g i n e s .

A l l HLFC a i r c r a f t i n t h i s s t u d y have been s i z e d w i t h o u t t h e u s e o f leading-edge h igh l i f t d e v i c e s on t h e wings. I n v i ew o f t h e f a v o r a b l e e f f e c t s

o f lead ing-edge h i g h l i f t sys t ems on t h e a i r f i e ld performance and t h e

2

s h i e l d i n g effects for HLFC o p e r a t i o n s , it is recomnended t h a t a d d i t i o n a l

s i z i n g s t u d i e s be conducted on t h e two b e s t HLFC c o n f i g u r a t i o n s o f t h i s s t u d y

wi th t h e a d d i t i o n of leading-edge h i g h l i f t sys tems .

3

2.0 INTRODUCTION

Among t h e many c o n c e p t s for a i r c ra f t d r a g r e d u c t i o n , l amina r flow

c o n t r o l , LFC, h a s i n d i c a t e d t h e g r e a t e s t p o t e n t i a l fo r s k i n f r i c t i o n d r a g r e d u c t i o n . A r e v i e w o f e a r l y p r o g r e s s s i n c e 1939 i n a n a l y t i c a l and

expe r imen ta l i n v e s t i g a t i o n s o f boundary l a y e r t r a n s i t i o n and methods fo r achievement o f l amina r flow is con ta ined i n a paper by Braslow and Muraca (Reference 1 ) .

Lockheed performed t h e i n i t i a l f e a s i b i l i t y s t u d y of advanced t echno logy

LFC a i r c r a f t beg inn ing i n October 1974 (References 2-5). lhe f a v o r a b l e r e s u l t s o f t h i s i n i t i a l s t u d y provided t h e impetus t o a d d i t i o n a l i n v e s t i g a -

t i o n s of LFC, and NASA, i n c o n c e r t w i th i n d u s t r y , h a s been sponsor ing LFC

t e c h n o l o g y deve lopnen t a c t i v i t i e s for t h e p a s t 12 y e a r s t o a c h i e v e LFC t echno logy r e a d i n e s s i n t h e 1 9 9 0 t s (References 6-9). Major Lockheed LFC

deve lopnen t programs funded i n 1980 under t h e N A S A ACEE program i n c l u d e wing s u r f a c e p a n e l s t r u c t u r a l deve lopnen t (References 10, 11) and t h e d e s i g n , f a b r i c a t i o n , and f l i g h t test o f leading-edge a r t i c l e s (Refe rence 12).

The Lockheed m o t i v a t i o n i n LFC a c t i v i t i e s has been directed t o t h e

e v e n t u a l a p p l i c a t i o n t o l o n g - r a n g e o r l o n g - e n d u r a n c e m i l i t a r y s t r a t e g i c a i rcraf t systems. During t h e time period of t h e i n t e n s i v e sys tem e v a l u a t i o n

s t u d i e s of commercial LFC t r a n s p o r t s t u d i e s under NASA c o n t r a c t s , Lockheed was c o n t i n u i n g i ts p r e l i m i n a r y d e s i g n studies of m i l i t a r y LFC a i r l i f t a i r c r a f t

under Independent Research and Developnent p r o j e c t s .

Encouraged by t h e p r o g r e s s made i n t h e deve lopnen t and v a l i d a t i o n of

lead ing-edge c l e a n i n g , a n t i - i c i n g , and s u c t i o n sys t ems so v i t a l t o t h e s u c c e s s of an LFC t r a n s p o r t , Lockheed and Douglas deve loped f l i g h t test ar t ic les wi th

NASA fund ing t h a t were i n s t a l l e d and tested on t h e NASA-Dryden F l i g h t Research F a c i l i t y JetStar a i r c ra f t . Ihe Lockheed a c t i v i t y is r e p o r t e d i n Reference 12.

An e a r l y r e v i e w of t h e t o t a l NASA program is given by Wagner and F i s c h e r i n Reference 9. A rev iew of t h e above a c t i v i t i e s since 1974 is g i v e n i n an A I A A

paper by Lange (Refe rence 13 ) .

4

Current a c t i v i t i e s i n t h e NASA/Lockheed Laminar Flow Enabl ing Technology

Development Cont rac t No. NAS1-18036 i n c l u d e t h e Task 1 deve lopnent o f a

s l o t t e d sur f a c e s t r u c t u r a l c o n c e p t u s i n g advanced aluminum m a t e r i a l s (Reference 14) and t h e Task 2 p re l imina ry concep tua l d e s i g n s tudy o f g l o b a l r a n g e m i l i t a r y HLFC t r a n s p o r t s r e p o r t e d herein.

This r e p o r t sumnar izes t h e r e s u l t s ob ta ined i n t h e Task 2 p r e l i m i n a r y

concep tua l d e s i g n s t u d i e s o f t h e b e n e f i t s de r ived from t h e u s e o f hybr id

laminar flow c o n t r o l ( H L F C ) f o r m i l i t a r y t r a n s p o r t s des igned t o a c h i e v e pay load / r ange r equ i r emen t s o f g l o b a l r a n g e a i r c r a f t . The Air Force P r o j e c t Fo recas t I1 e f f o r t has i d e n t i f i e d system PS-03 M u l t r i r o l e Global Range A i r c r a f t a s a s u b s o n i c element i n g l o b a l f o r c e p r o j e c t i o n . It i s a n t i c i p a t e d

t h a t t h i s g l o b a l r a n g e a i r c r a f t m u s t h a v e e x c e p t i o n a l a e r o d y n a m i c and p r o p u l s i v e e f f i c i e n c y t o a c h i e v e t h e mis s ion c h a r a c t e r i s t i c s . P rev ious

Lockheed p r e l i m i n a r y d e s i g n s t u d i e s h a v e shown s i g n i f i c a n t i n c r e a s e i n

aerodynamic e f f i c i e n c y by t h e a p p l i c a t i o n o f LFC t o m i l i t a r y t r a n s p o r t

a i r c r a f t . These resul ts were ob ta ined i n an Air Force c o n t r a c t s t u d y o f

T e c h n o l o g y A l t e r n a t i v e s f o r A i r l i f t Deployment (TAFAD) ( R e f e r e n c e 15 1. S e c t i o n 4.0 o f t h e r e p o r t c o n t a i n s background in fo rma t ion on t h e s t u d y

o b j e c t i v e s , s t u d y p l a n , assumpt ions b a s i c t o a l l s t u d y t a s k s , and t h e technology level a p p r o p r i a t e f o r t h e s t u d y . S e c t i o n 5.0 describes t h e m i s s i o n c h a r a c t e r i s t i c s and t h e b a s e l i n e c o n f i g u r a t i o n s t u d i e s o f t u r b u l e n t f l ow and

hybr id LFC a i r c r a f t sized t o perform t h e mis s ion r e q u i r e m e n t s o f t h e s tudy . Sect ion 6.0 con ta ins t h e r e s u l t s of s e n s i t i v i t y s t u d i e s o f s e v e r a l a i r c r a f t

performance p a r a n e t e r s f o r bo th hybr id LFC and t u r b u l e n t flow a i r c r a f t . I n S e c t i o n s 5.0 and 6 .0 comparisons a r e made between t h e t u r b u l e n t f l ow and

hybr id LFC c o n f i g u r a t i o n s i n o r d e r t o determine t h e b e n e f i t s a t t r i b u t a b l e t o

hybr id LFC. An o v e r a l l assessment o f hybr id LFC b e n e f i t s is c o n t a i n e d i n S e c t i o n 7.0, a long w i t h a p r e f e r e n c e f o r t h e best hybr id LFC c o n f i g u r a t i o n .

Conclusions and recommendations a r e provided i n S e c t i o n 8.0.

5

b

C

cL

CP

M

R N S t

t / c U

"S W X

x/ c

cy

rl

P

00

S

W

z

3.0 SYMBOLS AND ABBREVIATIONS

Symbols

wing s p a n , f t

l o c a l wing chord , f t

l i f t c o e f f i c i e n t

p r e s s u r e c o e f f i c i e n t

Mach nunber

Reynolds number

wing a r e a , f t

t h i c k n e s s , f t

2

wing th ickness- to-chord r a t i o

p o t e n t i a l f low v e l o c i t y , f t / s a r e a s u c t i o n v e l o c i t y , f t / s s l o t w i d t h , i n . s t reamwise c o o r d i n a t e , f t

chordwise l o c a t i o n

a n g l e o f a t t a c k , deg

c r u i s e power r a t i o , wing semispan l o c a t i o n

d e n s i t y , l b / f t 3

S u b s c r i p t s

f r e e s t ream

s l o t a t sur f a c e sucked h e i g h t o f boundary l a y e r

6

AR

CFL

HLF C

HP

H. P.

I G V

L /D LP MAC

O/ B

R P M

SF C

SL

STD

TOGW

A b b r e v i a t i o n s

aspe ct r a t i o

c r i t i c a l f i e l d l e n g t h , f t

hybrid l a n i n a r flow control

h i g h p r e s s u r e

horsepower

inlet g u i d e vane

1 i f t - t o - d r a g r a t i o

low p r e s s u r e

mean aerodynamic chord , f t

overboard v e n t

r e v o l u t i o n s per minute

s p e c i f i c f u e l c o n s u n p t i o n , lb&r

l b

s e a l e v e l

s tandard day

t a k e o f f g r o s s w e i g h t , l b

7

4.0 STUDY APPROACH

T h i s s e c t i o n o u t l i n e s t h e b a s i c a s s u n p t i o n s and c r i t e r i a which a r e

fundanen ta l t o a l l a s p e c t s o f t h e s t u d y . Included is a d e f i n i t i o n o f s t u d y o b j e c t i v e s , t h e o v e r a l l p l an employed t o a c h i e v e s t u d y o b j e c t i v e s , d e s i g n c r i t e r i a , and t h e assumed t echno logy level.

4.1 STUDY OBJECTIVES

The o b j e c t i v e o f t h i s t a s k i s t o de t e rmine by means o f p r e l i m i n a r y d e s i g n

s t u d i e s t h e b e n e f i t s d e r i v e d from t h e u s e o f hybr id laminar f low c o n t r o l f o r

m i l i t a r y t r a n s p o r t s designed t o a c h i e v e t h e payload/ range r equ i r emen t s o f t h e g l o b a l range a i rc raf t .

The Air Force P r o j e c t F o r e c a s t I1 e f f o r t has i d e n t i f i e d sys t em PS-03

M u l t i r o l e Global Range A i r c r a f t a s a s u b s o n i c element i n i ts g l o b a l f o r c e

p r o j e c t i o n . Although t h e sys t em c h a r a c t e r i s t i c s have n o t been f i n a l i z e d , it

i s a n t i c i p a t e d t h a t t h i s a i r c r a f t w i l l have t h e c a p a b i l i t y t o c a r r y l a r g e

pay loads f o r a d i s t a n c e o f 10,000 n a u t i c a l miles un re fue led a t h igh s u b s o n i c

cruise speeds . The a i r c r a f t w i l l l and and d e l i v e r t h e payload wi thou t s u p p o r t

a t t h e d e s t i n a t i o n a i r f i e l d and f l y back e i t h e r t o i t s b a s e i n t h e c o n t i n e n t a l U.S. o r t o a f r i e n d l y a i r b a s e u n r e f u e l e d . ?he a i r c r a f t m u s t have e x c e p t i o n a l

a e r o d y n a m i c and p r o p u l s i v e e f f i c i e n c y t o a c h i e v e t h e s e m i s s i o n c h a r a c t e r i s t i c s .

P rev ious Air Force/LASCGeorgia p r e l i m i n a r y d e s i g n s t u d i e s o f 1995 IOC m i l i t a r y t r a n s p o r t s have shown h igh performance and economic e f f i c i e n c y wi th c a p a b i l i t i e s o f Mach 0.80 cruise s p e e d , 212,000 pounds payload , and 3500 n a u t i c a l miles r a n g e (Reference 15). The a d d i t i o n o f laminar flow c o n t r o l on an a l t e r n a t e c o n f i g u r a t i o n provided a r a n g e i n c r e a s e t o 5800 n a u t i c a l miles

f o r t h e sane payload and cruise Mach number and a 26 p e r c e n t h i g h e r t a k e o f f

g r o s s weight . 'Ihe l amina r f low c o n t r o l concept u t i l i z e d a c t i v e s u c t i o n s l o t s

on t h e wing and empennage t o 70 p e r c e n t o f t h e chord .

I n o r d e r t o p rov ide f o r a near-term a p p l i c a t i o n o f laminar f low c o n t r o l ,

a more s i m p l i f i e d concept referred t o a s hybr id laminar flow c o n t r o l , HLFC,

8

has been used f o r t h e c u r r e n t s t u d y . The HLFC c o n c e p t , shown i n F igu re 1 h a s

t h e a c t i v e s u c t i o n system r e s t r i c t ed t o t h e r e g i o n ahead of t h e f r o n t s p a r o f

t h e wing. Aft o f t h e a c t i v e s u c t i o n r e g i o n t h e a i r f o i l shape is ta i lored t o a c h i e v e t h e maximum e x t e n t o f l amina r flow, and t h i s i s expec ted t o ex tend t o 50 p e r c e n t o r more o f t h e wing chord a s i n d i c a t e d by HLFC s t u d i e s b y b e i n g reported i n Reference 16. The HLFC concept a v o i d s a number of c o n c e r n s b y t h e

i n d u s t r y and t h e a i r l i n e s , i n p a r t i c u l a r , s u c t i o n s u r f a c e s and d u c t i n g are n o t

r e q u i r e d i n t h e main wing box a r e a s which a l s o c o n t a i n t h e f u e l for t h e a i r - c r a f t . Thus t h e weight and complex i ty of t h e s u c t i o n sys t ems is g r e a t l y reduced and t h e possible h a z a r d s wi th t h e f u e l are e l i m i n a t e d . 'Ihe s u c t i o n i n

t h e lead ing-edge r e g i o n can c o n t r o l t h e cross flow d i s t u r b a n c e s for swept wings and t h e a i r f o i l t a i l o r i n g ove r t h e wing box can s t a b i l i z e two-

d imens iona l d i s t u r b a n c e s .

4.2 STUDY PLAN

The p r e l i m i n a r y d e s i g n s t u d y c o n s i s t s of f i v e e l emen t s as shown i n F i g u r e

2. These e l emen t s c o n s i s t o f ( 1 ) Basic Data and Assumptions, (2) Mission C h a r a c t e r i s t i c s , (3) Conf igu ra t ion Developnent , (4) Conf igu ra t ion S e l e c t i o n ,

and (5) A n a l y s i s of Laminar Flow B e n e f i t s . These e l emen t s are b r i e f l y described i n t h e s e c t i o n s t h a t follow.

L E A D I N G EDGE T R E A T M E N T

A I R F O I L T A I L O R I N G T O M A I N T A I N N A T U R A L L A M I N A R FLOW

0 C L E A N I N G A N D A N T I - I C E S Y S T E M

0 S U C T I O N

Figure 1. Hybrid Laminar Flow Control Concept

9

BASIC DATA

ASSUMPTIONS MISSION CHARACTERISTICS

CONFIGURATION ANALYSIS OF SELECTION LAMINAR FLOW

BENEFITS

CONFlGURATlON DEVELOPMENT

TURBULENT FLOW LAMINAR FLOW

Figure 2 . Study Plan

Element 1

In t h e basic d a t a and assumption area and from c o n s i d e r a t i o n s of t h e

scope for t h i s s t u d y , t h e approach t aken was t o u t i l i z e t h e technology d a t a

base i n t h e Lockheed Generalized A i r c r a f t S i z i n g and Performance (GASP)

computer program t h a t was used i n t h e Air Force Technology A l t e r n a t i v e s for Airl if t Deployment (TAFAD) s t u d y mentioned p r e v i o u s l y (Refe rence 15). Modi-

f i c a t i o n has been made t o t h e d a t a b a s e t o account for t h e change t o t h e h y b r i d l a n i n a r flow c o n t r o l concep t and an upda te of t h e t echno logy da ta base i n c o r p o r a t e d . For t h i s s t u d y a technology r e a d i n e s s d a t e of 1994 is assuned a long wi th an i n i t i a l o p e r a t i o n a l c a p a b i l i t y IOC, d a t e of 2000.

Element 2

Mission c h a r a c t e r i s t i c s s u c h a s payload , r a n g e c r u i s e Mach nunber ,

a i r f i e l d performance, other sys tems o p e r a t i o n a l c o n c e p t s were e s t a b l i s h e d i n t h i s s t u d y element as t h e y a p p l y t o t h e Multrirole Global Range A i r c r a f t . As

noted p rev ious ly , t h e m i s s i o n characterist ics for System Ps-03 have n o t been f i n a l i z e d ; bu t as s u b s o n i c p a r t o f t h e Air Force Global Force P r o j e c t i o n , t h i s a i r c r a f t is t o p rov ide c o n v e n t i o n a l mass ive r e sponse and t a c t i c a l presence .

In t h e TAFAD s t u d y an in-depth miss ion a n a l y s i s o f t h e C o n g r e s s i o n a l l y

Mandated Mobi l i t y Study e s t a b l i s h e d an optimum payload of 212,000 pounds for

r a p i d deployment. In t h e Conf igu ra t ion I n t e g r a t i o n f o r Large Mul t ipurpose

10

A i r c r a f t s t u d y (Reference 17) t h e payloads v a r i e d from s m a l l 50,000 pounds f o r

AWACS t o medium l3O,000-150,000 pounds f o r ICBM l aunche r and c r u i s e missile

c a r r i e r and around 200,000 pounds f o r t h e a i r c r a f t c a r r i e r . A s p a c e v e h i c l e

launcher payload is about 275,000 pounds. U t i l i z i n g t h e s e r e s u l t s and o t h e r i n fo rma t ion , a set o f mis s ion c h a r a c t e r i s t i c s was m u t u a l l y agreed upon anong

NASA, t h e Air Force and Lockheed. These miss ion c h a r a c t e r i s t i c s summarized i n

F igu re 3A i n c l u d e a payload of 132,500 pounds, a c r u i s e speed of Mz0.77, and a r a n g e c a p a b i l i t y t o f l y o u t 6,500 n a u t i c a l miles w i t h f u l l pay load , l a n d and

o f f l o a d payload and f l y back 6,500 n a u t i c a l miles unre fue led . The t y p i c a l

mi s s ion p r o f i l e is provided i n F igu re 3B. A l l a i r c r a f t c o n f i g u r a t i o n s i n t h i s

s t u d y were s ized t o perform these mis s ion c h a r a c t e r i s t i c s . It should be noted o n t h e g e n e r a l a r r a n g e m e n t d r a w i n g s shown l a t e r , t h e l i s t e d a i r c r a f t pa rame te r s i n c l u d e an a s t e r i s k a f t e r "Range 6500 NM." Ihe a s t e r i s k d i r ec t s t h e

reader t o t h e mis s ion c h a r a c t e r i s t i c s o f F i g u r e 3 because t h e t o t a l c r u i s e

d i s t a n c e for t h e mis s ion i s no t a range or r a d i u s d i s t a n c e . Dev ia t ions from

these miss ion c h a r a c t e r i s t i c s were made i n t h e s e n s i t i v i t y s t u d i e s described

i n Section 6.0. For t h i s long r ange m i s s i o n , f u e l r e s e r v e s i n c l u d e 5 p e r c e n t

o f c r u i s e f u e l p l u s one h a l f hour .

Element 3

Conf igu ra t ion developnent u s ing p r e l i m i n a r y d e s i g n s t u d i e s were made of t u r b u l e n t flow and hybr id laminar flow control a i r c r a f t t o s a t i s f y t h e mis s ion charac te r i s t ics es tab l i shed i n Element 2. The Lockheed Genera l ized Aircraft S i z i n g and Performance (GASP) computer program used t o s i z e and d e f i n e t h e

0 PAYLOAD = 132,500 LB @ 2.5G

0 CRUISE SPEED = 0.77 MACH

0 I N I T I A L CRUISE ALTITUDE = FALLOUT VALUE

0 AIRFIELD (CFL) = 10,000 FT @ S.L. STD, DAY

0 FLYOUT 6,500 NM WITH FULL PAYLOAD, LAND AND RETURN 6,500 NM WITH ZERO PAYLOAD UNREFUELED

0 FIELD LENGTH @ MIDPOINT 5 8 , 0 0 0 FT @ S.L. STD, DAY

Figure 3A. Mission Characteristics

11

TAKEOFF TAKEOFF LAND 6 / L LANDING

ZERO PAYLOAD

f . , O O O . T i 6500 NM k10,OOO F T

6500 NM B 2.5G DESIGN PAYLOAD

10,000 FT L Figure 3B. Mission Profile

a i r c ra f t is described i n Appendix A. The a i r c r a f t c o n f i g u r a t i o n s were l i m i t e d

t o c o n v e n t i o n a l a r rangements i n t h i s s tudy . The o u t p u t of t h i s p r e l i m i n a r y d e s i g n Element 3 a c t i v i t y h a s provided t h e d a t a l i s t e d i n t h e fo l lowing f o r each c o n f i g u r a t i o n :

o General arrangement drawing

o Weight s t a t e m e n t i n c l u d i n g p ropu l s ion sys t ems o Geometric char ac t er i st i c s

o HLFC p e c u l i a r s t r u c t u r e s and c l e a n i n g ( a n t i - i c i n g f l u i d we igh t s )

o Payload-range c u r v e (some for a l l a i r c r a f t )

o Inboard p r o f i l e and cross s e c t i o n

In a d d i t i o n , sane s e n s i t i v i t y studies were performed i n t h e deve lopnen t of t h e b a s e l i n e t u r b u l e n t flow and h y b r i d LFC c o n f i g u r a t i o n s as described i n S e c t i o n 5.0.

Elements 4 and 5

The best t u r b u l e n t flow and hybr id LFC c o n f i g u r a t i o n s were selected i n

t h e s e e l emen t s of t h e s t u d y p l a n based on t h e r e s u l t s of Element 3 and addi -

t i o n a l s e n s i t i v i t y and c o n f i g u r a t i o n s t u d i e s performed i n Element 4 a s a

r e s u l t of a meet ing of NASA and Air Force t e c h n i c a l pe r sonne l a t t h e b c k h e e d

A e r o n a u t i c a l Systems Company i n Marietta, Georgia , on March 5, 1987. 'lhese a d d i t i o n a l studies extended t h e o v e r a l l scope of t h e s t u d y for b o t h t u r b u l e n t

12

f low and h y b r i d LFC c o n f i g u r a t i o n s and a r e d e s c r i b e d i n S e c t i o n 6.0. The

b e n e f i t s i n performance o f hybr id LFC a i r c r a f t a s compared wi th t u r b u l e n t flow

a i r c r a f t were determined from a d i r ec t comparison o f t h e best a i r c r a f t i n each

case .

4.3 REFERENCE TECHNOLOGY LEVEL

A s a p r e l i m i n a r y t o t h e p a r a m e t r i c c o n f i g u r a t i o n a n a l y s e s and subsequen t c o n f i g u r a t i o n a c t i v i t i e s l e a d i n g t o t h e d e f i n i t i o n o f selected a i r c r a f t , t h e level o f technology l i k e l y t o be a v a i l a b l e f o r a p p l i c a t i o n i n t h e e a r l y 1990

p e r i o d was e s t a b l i s h e d . T h i s s e c t i o n summarizes t h e r e f e r e n c e t echno logy level assumed f o r a l l c o n f i g u r a t i o n developnent a c t i v i t i e s .

4.3.1 Aerodynamics

4.3.1.1 Aerodynamics C r i t e r i a

The most comple te se t o f c r i t e r i a f o r t h e developnent o f e x t e r n a l aero-

dynamic c o n f i g u r a t i o n s compa t ib l e wi th LFC sys tems r equ i r emen t s was developed a s a p a r t of t h e X 2 1 program and is described i n Reference 5. The c r i t e r i a o f t h i s document were updated t o i n c l u d e r e s u l t s o f p e r t i n e n t r e c e n t i n v e s t i -

g a t i o n s . 'his upda t ing i n c l u d e d a c r i t i c a l review o f LFC s u c t i o n r e q u i r e m e n t s

and d u a l use o f a c t i v e t r a i l i n g - e d g e c o n t r o l f l a p s f o r g u s t a l l e v i a t i o n and min imiza t ion o f LFC s u c t i o n f low r a t e s in vary ing o p e r a t i o n a l c o n d i t i o n s .

Acous t ic e f fec ts on s u c t i o n requirements were addressed by i n c l u s i o n o f an

excess s u c t i o n system c a p a c i t y s i m i l a r t o t h e approach used f o r t h e X-21. A s

a result o f improvements i n aerodynamics d e s i g n and a n a l y s i s methods, aero-

dynamics c r i t e r i a used i n p r e v i o u s s t u d i e s , a s d e p i c t e d i n F i g u r e 3 of

Reference 4, were updated a p p r o p r i a t e l y f o r t h i s s tudy .

4.3.1 . 2 Airfoil Technology

The a i r c r a f t c o n f i g u r a t i o n s developed i n t h i s s t u d y i n c o r p o r a t e advanced

t echno logy s u p e r c r i t i c a l a i r f o i l s e c t i o n s c h a r a c t e r i z e d by an e x t e n s i v e r e g i o n o f s u p e r c r i t i c a l f l ow t e r m i n a t e d by a modera te -s t rength shock l o c a t e d f a i r l y

13

fa r a f t . Typ ica l wing s e c t i o n d e s i g n c u r v e s , which d e f i n e t h e technology

l e v e l o f t h e a i r f o i l t y p e , a r e shown i n F igu re 4. Some v a r i a t i o n i n a i r f o i l

t h i c k n e s s and form were examined t o maximize i n t e r n a l vo lune f o r f u e l and d u c t i n g and improve l e a d ing-edge boundary l a y e r c h a r a c t e r i s t i c s .

Advanced technology secondary a c t i v e t r a i l i n g - e d g e f l a p s of t h e t y p e

shown i n F igu re 5 were adopted as a means of a u t o m a t i c a l l y m a i n t a i n i n g des i r ed

p r e s s u r e g r a d i e n t s , c o n t r o l l i n g shock p o s i t i o n , and minimizing LFC s u c t i o n

r equ i r emen t s ove r a moderate r a n g e of o p e r a t i n g c o n d i t i o n s .

MACH NO.

.? =L

.6 CL .?

O L

SWEEP, DEC 0 15 25

Figure 4. Wing Section Design Curves

35

.15C 05:' .65C .9oc

LAMINAR AREA

Figure 5 . Example of Secondary Active Trailing Edge Flaps

14

4.3.1.3 High-Lift Device Technology

Design and a n a l y s i s s t u d i e s performed were compat ib le w i t h a c u r r e n t - t e c h n o l o g y mechanica l f l a p system w h i c h p r o v i d e s t h e r e q u i r e d a i r p o r t performance w i t h t h e smallest p e n a l t y t o d i r ec t o p e r a t i n g cost . S ingle- and

m u l t i p l e - s l o t t e d f l a p s were a s s e s s e d i n t h e s t u d y from t h e s t a n d p o i n t of

chordwise and spanwise extent, l i f t and d rag e f f e c t i v e n e s s , r e l a t i v e weight p e n a l t y , and h i g h - l i f t c o m p a t i b i l i t y wi th a i r f o i l s e c t i o n shapes desirable for

LFC. For t h i s s t u d y no l e a d i n g edge d e v i c e s are used i n o r d e r t o allow f o r HLFC on bo th upper and lower s u r f a c e s .

4.3.2 F l i n h t C o n t r o l s

The f l i g h t c o n t r o l sys tem inc luded i n t h e s i z i n g program i n c o r p o r a t e s t h e

e l emen t s of a c t i v e c o n t r o l t echno logy (ACT) which promise s i g n i f i c a n t improvements i n t h e e f f i c i e n c y of l a r g e transport a i r c r a f t .

The ACT system encompasses t h e fo l lowing modes of c o n t r o l :

o Relaxed S ta t i c S t a b i l i t y

o S t a b i l i t y Augmentation System

o Maneuver b a d Cont ro l

o Gust Load A l l e v i a t i o n o Flutter Mode Control

0 Ride Cont ro l

The major improvement offered by t h e above sys t ems are: min imiza t ion of a i r frame weight , i n c o r p o r a t i o n of au tomat i c t roub le - shoo t ing , and improved

r i d e c h a r a c t e r i s t i c s . These sys t ems were employed i n p r e v i o u s b c k h e e d LFC

a i r c r a f t studies and are d e s c r i b e d i n more d e t a i l i n Ref. 5 .

The four channel f l y - b y d i r e ( F B W ) system is c o n t r o l l e d on each channe l

b y an on-board d i g i t a l computer. A d i g i t a l sys tem is mandated by t h e exten-

s i v e complex s i g n a l p r o c e s s i n g , t h e f l e x i b i l i t y r e q u i r e d t o accommodate t h e

mul t i -mode control l o g i c laws, and t h e redundancy r e q u i r e d by an FBW system.

15

Geared e l e v a t o r s d r i v e n by t h e s t a b i l i z e r , a double hinged rudder, and

ou tboa rd a i l e r o n s provide low speed c o n t r o l . Ground-operable-only s p o i l e r s

a r e provided f o r deployment d u r i n g ground r o l l o u t o r rejected t a k e o f f . A l l

c o n t r o l s and i n s t r u m e n t a t i o n r e q u i r e d f o r t h e o p e r a t i o n o f t h e a i r p l a n e i n t h e a i r and on t h e ground a r e l o c a t e d i n t h e f l i g h t s t a t i o n . The on-board computers p rov ide feedback f o r two hydro-mechanica l u n i t s which p rov ide t h e p i l o t s w i t h a r t i f i c i a l f ee l i n a l l t h r e e c o n t r o l axes .

4.3.3 Propulsion Systems

The P r a t t & Whitney A i r c r a f t STF-686 s t u d y e n g i n e wa,s chosen a s t h e

pr imary p ropu l s ion u n i t f o r t h e b a s e l i n e a i r c r a f t . This eng ine i s a t w i n - spool, s e p a r a t e f low t u r b o f a n e n g i n e w i t h 19,350 pounds o f t a k e o f f t h r u s t . The p r e l i m i n a r y weight o f t h e eng ine is 3800 l b . The h igh p r e s s u r e spool is a

scaled v e r s i o n o f t h e STS-686 h i g h p r e s s u r e spool, made up of an 11 s t a g e h i g h

pressure compressor, a low emiss ions combustor , and a two stage h igh p r e s s u r e

tu rb ine . ?he low p r e s s u r e spool c o n s i s t s o f a s i n g l e s t a g e s h r o u d l e s s f a n , a

t h r e e stage low p r e s s u r e compressor and a f i v e s t a g e low p r e s s u r e t u r b i n e . An

a c t i v e c l e a r a n c e c o n t r o l system i s inco rpora t ed which c o n t r o l s t h e c l e a r a n c e s

o f s e v e r a l components i n o r d e r t o minimize t h e f u e l consunpt ion a t cruise.

Th i s system i s a c t i v a t e d a t a l l o p e r a t i n g c o n d i t i o n s a t a l t i t u d e s above 15,000

f e e t . The eng ine has a d e s i g n f a n p r e s s u r e r a t i o o f 1.66, a bypass r a t i o o f

6.97, and an o v e r a l l compression system p r e s s u r e r a t i o o f 37.2.

The performance and weight improvement s c h e d u l e s f o r t he STF-686 a r e

shown i n F igu res 6 and 7. F i g u r e 6 shows t h e p r o j e c t e d performance improve-

ment f o r t h e eng ine th rough t h e yea r 2005, u s i n g t h e PW2037 eng ine as a b a s e l i n e , by which time SFC w i l l have dec reased 1 3 - 5 p e r c e n t , F i g u r e 7 shows

t h e p r o j e c t e d weight improvement f o r t h e engine i n t h e sane time frane and

us ing t h e sane b a s e l i n e , a 13 p e r c e n t r e d u c t i o n i n weight by t h e y e a r 2005.

It is noted t h a t bo th c h a r a c t e r i s t i c s assune t h a t a g g r e s s i v e component and eng ine t echno logy programs w i l l be main ta ined du r ing t h i s time span .

16

SPECIFIC FUEL CONSUMPTION

- 1 4

-16

-1 8

F z w U

W n I

I- Z W E W > 0 p1 n r U L VI Q

-

- w

- - I I I I

0

- 2

- 4

-6

-8

-10

-1 2

- 1 4

-1 6

-18 I 1 fi I I I 1990 1995 2000 2005 1980 1985

YEAR OF TECHNOLOGY AVAILABILITY

Figure 6. Specific Fuel Consumption (PEW STF686)

BARE ENGINE WEIGHT

YEAR OF TECHNOLOGY AVAILABILITY

Figure 7. Bare Engine Weight (P&W STF686)

17

4.3.4 Structures and Materials

The a i r c r a f t s t r u c t u r e c o n t a i n s conven t iona l m a t e r i a l s and g r a p h i t e / p e e k

composi te m a t e r i a l s t o r e p r e s e n t a materials technology l e v e l o f approximate ly

1994. The p e r c e n t u t i l i z a t i o n of t h e advanced composi te m a t e r i a l is i l l u s -

t r a t e d i n F i g u r e 8. Craphi te /peek composite is composed of g r a p h i t e f ibers i n

a t h e r m o p l a s t i c resin and i t of fe rs t h e p o t e n t i a l o f h igh s t r e n g t h and s t i f f -

n e s s a long w i t h r e s i s t a n c e t o de lamina t ion and embr i t t l emen t .

F i g u r e 9 p r e s e n t s t h e weight r e d u c t i o n pe rcen tages t h a t have been de termined from d e t a i l e d a n a l y t i c a l t r a d e studies. These weight r e d u c t i o n

v a l u e s are inc luded i n t h e weight e s t i m a t i o n methods w i t h i n t h e v e h i c l e d e s i g n

s y n t h e s i s p rocess . Th i s s y n t h e s i s p rocess is per formed by MODGASP mere f u r t h e r weight decreases w i l l occur i n f u e l and o p e r a t i n g weight because o f

t h e advanced mater ia l w e i g h t r e d u c t i o n s .

The weight e s t i m a t i n g methods used for a i r c r a f t s i z i n g w i t h i n GASP are

those for c o n v e n t i o n a l t r a n s p o r t s p l u s a l lowances for LFC p e c u l i a r d e s i g n v a r i a b l e s and c o n s i d e r a t i o n of advanced technology.

H O R I Z O N T A L T A I L G R A P H I T E I PEEK 59% C O N V E N T I O N A L 41%

T O T A L S T R U C T U R E I G R A P H l T E l P E E K 63% C O N V E N T I O N A L 37% & - WING

G R A P H I T E / P E E K 69% C O N V E N T I O N A L 31%

C O N V E N T I O N A L 69% \\'s\n/ N A C E L L E G R A P H I T E I PEEK 31 %

F U S E L A G E G R A P H l T E l P E E K

CERTICAL TAIL G R A P H l T E l P E E K C O N V E N T I O N A L

71% >-- -.-==a C O N V E N T I O N A L 29%

LANDING GEAR G R A P H l T E l P E E K 27% C O N V E N T I O N A L 73%

Figure 8. Advanced Material Technology for Hybrid LFC Study

63% 37%

18 .

STRUCTURAL % GRAPHITE 8 CONVENTIONAL COMPONENT PEEK MATERIAL

WING 69 3 1

HOZ T A I L 5 9 4 1

VERT TAIL 63 3 7

FUSELAGE 7 1 2 9

LANDING GEAR 2 7 7 3

NACELLE 3 1 6 9

The methods for c o n v e n t i o n a l t r a n s p o r t s have been developed and improved

th rough pas t d e s i g n s t u d i e s . In these s t u d i e s t h e s e n s i t i v i t y o f t h e v a r i o u s

design parameters have been d e r i v e d from e x i s t i n g t r a n s p o r t d e s i g n d a t a w i t h

e x t r a p o l a t i o n s de r ived from v a r i o u s a n a l y t i c a l s t u d i e s . One o f these d e s i g n parameters i s wing a s p e c t r a t i o . Wing weight i n c r e a s e s a s a s p e c t r a t i o

i n c r e a s e s while drag decreases. The miss ion r equ i r emen t s of t h e v e h i c l e ,

t h e r e f o r e , i n f l u e n c e t h e s ize o f t h e optimum aspect r a t i o . O f c o u r s e , other f ac to r s must be cons ide red i n t h e f i n a l s e l e c t i o n of wing a s p e c t r a t i o . Same of these c o n s i d e r a t i o n s are f l u t t e r r equ i r emen t s , wing l o a d i n g and CLMAX,

runway w i d t h , f u e l c a p a c i t y , p r o d u c i b i l i t y , etc. The p r e s e n t wing weight

e s t i m a t i o n r e l a t i o n s h i p f o r t r a n s p o r t a i r c r a f t is t h e r e s u l t of aspect r a t i o s t u d i e s conducted d u r i n g 1984 and it p r o v i d e s higher aspect r a t i o d e s i g n s y n t h e s i s t h a n earlier methods.

% WEIGHT REDUCTION

2 9 . 5

2 2 . u

2 0 . u

1 9 . 1

1 1 . 4

2 1 . o

The we igh t e s t i m a t i n g methods for LFC p e c u l i a r d e s i g n v a r i a b l e s were

d e r i v e d d u r i n g t h e Air Force/Lockheed studies (Reference 15, 1982 t o 1985). These methods were r e v i s e d t o account for t h e hybr id L F C c o n c e p t where LFC is appl ied t o o n l y t h e wing l e a d i n g edge. The r e s u l t i n g weight i nc remen t s are d i s p l a y e d s e p a r a t e l y i n t h e v a r i o u s weight summaries for r eady i d e n t i f i c a t i o n .

Advanced technology weight a l lowances are p r o g r a m e d i n t h e weight

e s t i m a t i o n r o u t i n e s i n t h e form of i n p u t f a c t o r s f o r a p p l i c a t i o n t o c o n v e n t i o n a l t r a n s p o r t weight r e l a t i o n s h i p s . 'he i n p u t factors are d e r i v e d

f o r e a c h s t u d y program based upon t h e a d v a n c e d d e s i g n c o n c e p t s u n d e r

c o n s i d e r a t i o n for t h e p a r t i c u l a r v e h i c l e . A s p r e v i o u s l y descr ibed , g r a p h i t e

peek composite material is used i n t h i s s t u d y for basic and secondary

s t r u c t u r e i n t h e wing, f u s e l a g e , t a i l , l a n d i n g g e a r , and n a c e l l e .

19

Figure 10 i l l u s t r a t e s t h e Group Weight Summary f o r t h e Hybrid LFC

c o n f i g u r a t i o n a s produced by t h e G A S P weight l o g i c . On t h i s t ab l e , t h e weight

increments f o r L F C a re i d e n t i f i e d w i t h i n t h e S t r u c t u r e , P ropu l s ion , and Fixed

Equipnent c a t e g o r i e s . Th i s means t h a t t h e normal weight c a t e g o r i e s are

estimated from d e s i g n pa rame te r s u s ing conven t iona l t r a n s p o r t v a r i a t i o n s , and

t h e n t h e a i r c ra f t weight is modi f ied by t h e LFC weight increments . The d e s i g n

pa rame te r s , however, are in f luenced by t h e LFC performance q u a l i t i e s so t h a t

L F C b e n e f i t s must be e v a l u a t e d from o v e r a l l a i r c r a f t c o n f i g u r a t i o n pa rame te r s and n o t from t h e LFC weight i nc remen t s a lone . The WC weight i nc remen t s are

de r ived from a i r c r a f t d e s i g n pa rame te r s such a s l amina r i zed area, s u c t i o n c o e f f i c i e n t s , p r e s s u r e c o e f f i c i e n t s , source of a i r p r e s s u r e (LFC e n g i n e s ) , span of a i r d i s c h a r g e , and from t h e s u r f a c e c l e a n i n g p r o v i s i o n s and r e q u i r e - ments. The f u e l requi rement o f 2578 pounds f o r t h e s u c t i o n pumps is a l s o

l i s t ed i n S e c t i o n 4.3.5.3 on t h e s u c t i o n punp c h a r a c t e r i s t i c s . The a i r c r a f t b e n e f i t s are d e r i v e d from lower f u e l r equ i r emen t s , due t o d r a g r e d u c t i o n s , and its s c a l i n g effects on t h e v a r i o u s weight items are d e r i v e d from t h e reduced

geometry and g r o s s weight c o n d i t i o n s . This t y p e o f a n a l y s i s is handled w i t h i n G A S P t h rough i t e r a t i o n o f t h e a i r c r a f t weight q u a n t i t i e s a s compared wi th t h e q u a n t i t i e s r e q u i r e d f o r mission performance.

4.3.5 HLFC Systems

4.3.5.1 Surface Design

The s u c t i o n s l o t d e s i g n must p r o v i d e s l o t s having flow c h a r a c t e r i s t i c s

t h a t are p r e d i c t a b l e , s t a b l e , uniform a long t h e l e n g t h of t h e s lo t , and free from s u r f a c e flow d i s t u r b a n c e s . Criteria and limits for s l o t d e s i g n were

developed t o meet t h e s e r e q u i r e m e n t s d u r i n g t h e X-21 program i n t h e e a r l y 1960,s by NORAIR and are summarized i n Reference 5. Unfor tuna te ly , s u p p o r t i n g

da t a are not well docunented i n t h e l i t e r a t u r e . When t h e s e c r i te r ia and limits are a p p l i e d t o t h e d e s i g n o f s l o t s for t h e c u r r e n t a i r f o i l r e q u i r e -

ments , m u t u a l l y e x c l u s i v e c o n f l i c t s exist between t h e cr i ter ia . A s t r i c t a p p l i c a t i o n o f t h e c r i t e r i a and limits t o d e f i n e t h e s u r f a c e s l o t conf igu ra -

t ion r e s u l t s i n s l o t w id ths and s p a c i n g s i n t h e leading-edge r e g i o n t h a t are impractical , i f n o t impossible, t o manufac ture on a p roduc t ion a i r p l a n e . For

t h e s e r e a s o n s , it is n e c e s s a r y t o a c c e p t some compromises i n these c r i t e r i a

20

ITEM

STRUCTURE WING T A I L CROUP (EMPENNAGE) BODY CROUP (FUSELAGE) ALIGHTING GEAR GROUP NACELLE/PYLON SPECIAL ITEMS

LFC WT. INCREMENT - WING LFC WT. INCREMENT - T A I L

PROPULSION PROPULSION GROUP

ENGINES FUEL SYSTEM THRUST REVERSERS MISCELLANEOUS

LFC WT, INCREMENT - LFC ENGINES LFC WT. INCREMENT - DUCTING, ETC.

SPECIAL ITEMS

WEIGHT (POUNt

5 6 3 1 7 1

2 0 , 2 8 2 6 , 5 9 8 3 , 4 0 7 2 , 0 0 0

536 1 , 2 3 9

Figure 10A. Hybrid LFC Design

ITEM

SYSTEMS AND EQUIPMENT SURFACE CONTROLS GROUP AUXIL IARY POWER PLANT CROUP INSTRUMENTS C NAV. EQ. GROUP HYDRAULICS AND PNEU. GROUP ELECTRICAL GROUP ELECTRONICS GROUP (AVIONICS) FURNISHING E EQUIPMENT CROUP AIR COND. C ANTI - ICE GROUP AUXIL IARY GEAR GROUP SPECIAL ITEMS

LFC W T . INCR. - CLEAN SYS.

WEIGHT EMPTY OPERATING EQUIPMENT

OPERATING WEIGHT CARGO

ZERO FUEL WEIGHT MISSION FUEL

LFC - L.E. CLEAN C ANTI - ICING FLUID

CROSS WEIGHT

6 8 , 8 8 8 6 , 4 5 0

4 2 , 9 9 9 1 6 . 8 1 5

4 . 0 9 9 7 3 4

3 2 . 2 8 7

1 , 7 7 0

1

1 3 9 , 9 8 5

3 9 , 0 5 7

WEIGHT (POUNDS)

9 5 9

4 , 2 8 0 8 4 2 9 1 2

1 , 9 9 5 3 , 4 2 3 2 . 3 8 1 5 , 8 5 0 2 , 9 5 6

2 3 . 5 9 8

959 I 1 9 7 , 6 4 0

5 , 0 0 6

2 0 2 , 6 4 6 1 3 2 , 5 0 0

3 3 5 , 1 4 6 2 5 2 , 2 1 6

4 . 2 5 8

5 9 1 , 6 2 0

Figure 10B. Hybrid LFC Design

21

and limits. However, t h e l a c k of s u f f i c i e n t s u p p o r t i n g d a t a p r e c l u d e s a sound

and c o n f i d e n t judgement of t h e s e compromises.

For t h e d e s i g n c r i t e r i a d i s c u s s e d i n Sec t ion 6.3.2.1 of Reference 5 and

us ing F igu re 1 1 i n Reference 5, p r e l i m i n a r y s l o t l o c a t i o n s for t h e s u c t i o n

sys tem were ach ieved . These s l o t l o c a t i o n s are shown i n F i g u r e 11 fo r wing s t a t i o n Y/b=0.832. The chordwise d e s i g n r e g i o n for both t h e upper and lower

wing s u r f a c e s s t a r t a t t h e f i r s t s l o t a f t of t h e leading-edge c l e a n i n g / de- ic ing sys tem r e g i o n located around X/C=.Ol. Since t h i s is an HLFC c o n f i g u r a t i o n , a c t i v e s u c t i o n e n d s a t t h e f r o n t s p a r (X/C=. 15); n a t u r a l l amina r flow is u t i l i z e d a f t of t h e f r o n t s p a r . S ince t h i s is o n l y a

p r e l i m i n a r y c o n f i g u r a t i o n , more a n a l y s i s w i l l be r e q u i r e d for a f i n a l d e s i g n .

I n t h e f i n a l d e s i g n , more s lo t s may have t o be added a t t h e inboa rd wing

s t a t i o n s , b u t these s l o t s w i l l n o t ex tend across t h e e n t i r e wing.

The p a r t i a l a i r f o i l s e c t i o n shown i n F igu re 11 shows t h e a i r f o i l a s a f u n c t i o n of X / C . The s u c t i o n s l o t s are r e p r e s e n t e d b y marks i n t e r n a l t o t h e

a i r f o i l o u t l i n e . The s l o t w i d t h and s l o t spac ing were f i n a l i z e d u s i n g t h e

d e s i g n c r i t e r i a i n S e c t i o n 6.3.2.1 of Reference 5. For manufac tu r ing r e a s o n s ,

- 20

10 -

0

10 -

- 20

Figure 11. Wing Leading Edge Slot Locations

22

t h e d i s t a n c e between s l o t s was k e p t t o a minimun of 0.65 i n c h . Table 1 shows t h e geometry and performance of t h e upper s u r f a c e o f t h e wing a t wing s t a t i o n

Y/b=0.488. These c a l c u l a t i o n s were made f o r t h e 0.77 Mach nunber c r u i s e

c o n d i t i o n a t 37,000 f e e t . The colunn head ings cor respond t o a s lo t number ( 1

being f a r thes t f o r w a r d ) , chordwise ( X / C ) l o c a t i o n , s l o t w i d t h ( W ) i n i n c h e s ,

s l o t s p a c i n g (CN) i n i n c h e s , s l o t Reynolds number ( R W ) , t h e r a t i o of s l o t w i d t h t o boundary l a y e r sucked h e i g h t ( W / Z ) , t h e r a t i o of sucked-height

v e l o c i t y t o boundary-layer edge v e l o c i t y ( U Z / U E ) , s l o t geometry and flow

parameter (BETA) , and s l o t p r e s s u r e loss c o e f f i c i e n t (CPS). S i m i l a r d a t a were g e n e r a t e d f o r a l l o f t h e v a r i a t i o n s i n spanwise l o c a t i o n and cruise f l i g h t

v a r i a t i o n s . The lower s u r f a c e geometry and performance are s i m i l a r l y i l l u s - t r a t e d on Figure 10 and l i s t e d for t h e des ign c r u i s e c o n d i t i o n s on Table 2.

As is shown, t h e performance pa rame te r s a r e a l l w i t h i n t h e limits described i n Reference 5 f o r optimum performance.

4.3.5.2 S u c t i o n meting System

The s u c t i o n d u c t i n g system i s composed of a combina t ion of d u c t s , l i n e s ,

and v a l v e s t o meter, c o l l e c t , and t r a n s p o r t t h e s u c t i o n flow from each s u r f a c e s l o t t o t h e s u c t i o n pump. This system w i l l p r o v i d e a s u c t i o n d i s t r i b u t i o n

compa t ib l e wi th boundary l a y e r l a m i n a r i z a t i o n ove r a range o f o p e r a t i n g

c o n d i t i o n s r e p r e s e n t i n g c r u i s e , cl imb and d e s c e n t w i th emphasis on c r u i s e .

TABLE 1. UPPER SURFACE DESIGN DATA

ALTITUDE - 37,000 FT MACH - 0.77 CHORD - 16.159 FT

> L O T

u t u2 u 3 u4 us U6 u7 U 8 u9 U I U u t 1 u1z U 1 3 U14 u15 U16 U17 U 1 8 U19 u20 u21

- X I C

0.5590E-02 0.816OE-02 0.1106E-01 0.1413E-01 0.1730E-01 O.ZOS7E-01 0.2385E-01 0.2722E-01 0.3058E-01 0.3395E-01 0.3764E-01 0.6159E-01 0.4581E-01 0.5033E-01 0.5512E-01 0.6018E-01 0.6554E-01 0.7143E-01 0.7787E-01 0.848SE-01 0.9263E-01

- U - 0.3000E-02 0.3000E-02 0.3000E-02 0.30OOE-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.300OE-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02

Y..'Z

0.8062E+00 0.16.9E+JI 0.14+3E+~ll 0.1309E+01

O.I213E+Ol O.l122E+01 O.IIIOE+OI O.L103E+OI O.L093E+01 O.IOLZE+OI O.l039E+OI 0.1019E+01 0.99f 2E+00 0.9 6:. 5 E+OO 0.93f3E+00 0.93; 5E+OO 0.914 9E+00 0.8949E+00 0.87t 8E+OO 0.8409E+00

--

O . I Z ~ ~ E + O I

B E T A - 0.8030E-01 0.3681E+OO 0.3774E+OO 0.4008E+00 0.42881+00 O.L447E+00 0 . 4 546E+00 0.4716E+OO 0.4957E+00 0.51966+00 O.S116E+00 0.513SE+00 0.5218E+00 0.5284E+00 0.5283E+OO 0.53S6E+00 0.5536E+00 0.5617E+00 0.5758E+00 0.5856E+OO 0.5725E+00

CBS

0.923OE+00 0.1438E+00 O.L536E+00 0.1539E+00 0.1476E+00 0.1419E+00 0.13885+00 0.1329E+OO O.I25SE+00 0.1189E+00 0.1208E+00 0. IZOZE+00 0.1179E+OO 0.1161E+00 0.1160E+00 O.l14lE+00 0.1097E+00 0.1077E+00 0 . 1 0 4 SE+OO O.I025E+00 0.1049E+00

-

TABLE 2 . LOWER SURFACE DESIGN DATA

A L T I T U D E - 37.000 FT HACll - 0 . 1 1 C H O R D * 16.159 FT

SLOT

L1 L2 L3 L4 L S L6 L7 L8 L9 LIO Ll I LIZ L13 LI4 L I S L16 L 1'7 L18 L19 LZO

- x / c

0.1001E-01 0.1296E-01 0.161&E-01 0.1937E-01 0.22728-01 0.2607E-01 0.2918E-01 0.32906-01 0.3633E-01 0.3980E-01 0.4354E-01 O.4754E-01 O.Sl83E-01 0.5640E-01 0.6124E-01 0.66638-01 0.72S6E-01 0.79318-01 0.86868-01 0.9550E-01

- U - 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3000E-02 0.3OOOE-02 0.3000E-02 0.3000E-02

U f Z

0.9045E+OO 0.1536E+01 0.1478E+01 0.1413EtOI 0.1371E+Ol 0.1323EtOI 0.1292EtOl 0.1269E+01 O.l239E+Ol O.I236E+OL 0.1192E+01 0.1153E+01 0 . I IZlE+OI 0.1091E+01 O.L067E+01 0.1035E+OI O.IOISE+OI 0.97878+00 0.9400E+00 0.8860E+r)0

- Ur / Ue

0.5776E+00 0.3086E+00 0.2792E+00 0.2553E+OO 0.2372E+OO 0.2197E+OO 0.2063E+OO 0.1939E+OO O.l820E+00 O.l74OE+00 0.1723E+00 O.I692E+OO 0.165OE+OO O.I598E+00 0.1531E+OO O,1504E+OO 0.1451E+OO 0.1413Et00 0.1366E+OO 0.1362E+00

- BETA - 0.1366E+OO 0.4190E+00 0.4383Et00 0.4532E+OO 0.4713EtOO 0.689 IE*00 0.5083E+00 0.5304E+00 0.5514E+00 0.5749E+00 0.5595E+OO 0.55 l O E + O O 0.5bI(.E+00 0.5516E+OO 0.5629E+OO 0.55S4E+00 0.56431+00 0.5586E+00 0.5548E+OO 0.5241E+00

~~ ~

CPS

O.I447E+OO 0.8923E-01 O.8726E-01 0.8617E-01 0.8334E-01 0.8083E-01 0.7787E-01 0.7454E-01 0.7167E-01 0.6857E-01 0.7088E-01 0.7230E-01 0.728 LE-01 0.7251E-01 0.7 107E-01 0.72271-01 0.1116L-01 0.7212E-01 0.7279E-01 0.7771E-01

-

I n d i v i d u a l s l o t flows w i l l be a d j u s t a b l e from t h e c o c k p i t , e n a b l i n g chordwise

s u c t i o n d i s t r i b u t i o n t o be v a r i e d i n f l i g h t . , This s e c t i o n w i l l o n l y dea l with

t h e s u c t i o n system i n t h e wing; t h e s u c t i o n pump w i l l be d i s c u s s e d i n S e c t i o n

6. 4.

Duct ing Concept - The d u c t i n g system for t h e b a s e l i n e a i r c r a f t evolved

ove r time based on bo th LFC system requ i r emen t s and s t r u c t u r a l c o n s i d e r a t i o n s . The r e s u l t i n g sys t em concept is compa t ib l e w i t h b o t h d i s c i p l i n e s and has

r e l a t i v e l y few compromises t o e i ther d i s c i p l i n e .

A t y p i c a l c r o s s - s e c t i o n of an i n d i v i d u a l s l o t is shown i n F igu re 12. The

slot d e s i g n and f a b r i c a t i o n are based on the Task 1 development efforts described i n Reference 14. Air is drawn th rough t h e s l o t i n t o t h e s l o t d u c t ,

t h rough t h e me te r ing h o l e s i n t o t h e collector d u c t , t h rough t h e collector d u c t o r i f i c e i n t o t h e s u c t i o n t u b e , and from t h e s u c t i o n t u b e th rough a n e e d l e v a l v e i n t o one of two main plenum d u c t s . There are two plenum d u c t s ex tend ing

t h e l e n g t h of t h e wing span: a h igh -p res su re d u c t for t h e lower s u r f a c e , and a

low-pressure d u c t for t h e upper s u r f a c e . These d u c t s lead t o t h e s u c t i o n

pump, which is d i s c u s s e d i n S e c t i o n 4.3.5.3. The schemat i c of t h e entire s u c t i o n system is shown i n F igu re 13. Dimensions of t h e s l o t s and me te r ing

h o l e s must be selected t o p r o v i d e as uniform s u c t i o n flow as p o s s i b l e t o a l l s lo t s wi th low p r e s s u r e l o s s e s . The n e e d l e v a l v e s w i l l be used t o m a i n t a i n

24

DIFFUSION BOND SLOT DUCT I

t d i T E R I N 4 /

ADHESIVE BOND

\ \ COLLECTOR DUCT / /

,

ADHESIVE BOND

ION

TO NEEDLE VALVE AND PLENUM DUCT

Figure 12. Slot and Ducting Cross Section

25

SURFACE - - - - - - - - - - - - - FROM CLEANING ------ SLOTS SLOT CLEANING DUCT LIQUID SHUTOFF VALVE

CLEANING >- LIQUID SUPPLY

SLOT TUBE FLOWMETER FLOWMETER

SHUTOFF VALVE

TO LOWER SURFACE SLOTS (SIMILAR TO

ISOLATION

TURBINE COMPRESSOR OVERBOARD OVERBOARD DISCHARGE DISCHARGE

Figure 13. Suction System Schematic

t h e proper d i s c h a r g e environment for t h e upstream s l o t meter ing holes and

provide proper flow m e t e r i n g downstream t o match t h e local p r e s s u r e s w i t h i n

t h e plenun d u c t s .

To p r o v i d e t h e c l e a n i n g and a n t i - i c i n g c a p a b i l i t y , t m s lo ts h a v e been

added on t h e upper s u r f a c e forward o f t h e first s u c t i o n s l o t s o l e l y dedica ted

t o e m i t t i n g t h e c l ean ing /de - i c ing f l u i d . On t h e lower s u r f a c e , t h e first f i v e s lo t s have t h e c a p a b i l i t y both for c l e a n i n g and s u c t i o n . A t low a l t i t u d e s

these seven s l o t s emit t h e c l ean ing /de - i c ing f l u i d t o keep t h e s l o t s open and t h e wing s u r f a c e c l e a n . Upon r e a c h i n g a c e r t a i n a l t i t u d e , t h e c l e a n i n g system w i l l be tu rned off and h igh -p res su re a i r from t h e a i r c r a f t ' s envi ronmenta l

c o n t r o l system w i l l be directed th rough t h e c l e a n i n g / s u c t i o n s lo t geometries. This airf low w i l l remove t h e c l e a n i n g f l u i d from t h e s l o t d u c t i n g s u r f a c e s t o p reven t d a m q e t o t h e v a l v e s and i n s t r u m e n t a t i o n . The c l e a n i n g system i s d i s c u s s e d i n more d e t a i l i n Sec t ion 4.3.5.5.

26

Ducting Design - The leading-edge me te r ing sys tem c o n f i g u r a t i o n h a s a l r e a d y been i l l u s t r a t e d i n F i g u r e 12. Tab le 3 shows t y p i c a l s l o t and m e t e r i n g geometry d imens ions used i n t h e performance c a l c u l a t i o n s for t h i s

p a r t i c u l a r s t u d y . These d imens ions were a r r i v e d a t t h r o u g h t h e parameter

g u i d e l i n e s d i s c u s s e d i n S e c t i o n 6.3.2.1 o f Reference 5 and a l s o i n Reference 18. However, there a r e i n d i c a t i o n s t h a t t h e v e r y low Reynolds nunber c h a r a c t e r i s t i c s of t h e leading-edge s l o t s , t o g e t h e r w i t h a v e r y f a v o r a b l e r a t i o of m e t e r i n g hole s p a c i n g t o slot d u c t d e p t h p e r m i t s some r e l a x a t i o n o f

t h e c r i t e r i a w i t h o u t any p e n a l t i e s t o performance.

TYPICAL NOMINAL DIMENSIONS

SLOT DUCT METERING COLLECTOR DUCT METERING . PLENUM SURFACE xlc DUCT SPACING ( I N ) DlAM ( I N ) SPACING (IN1 DlAM [ IN)

UPPER 0.017 HP 0.500 0.053 2.000 0.075

0.046 HP 0.500 0.072 2.000 0.095

0.100 HP 0.500 0.087 2.000 0.113

LOWER 0.019 LP 0.500 0.056 2.000 0.008

0.052 LP 0.500 0.076 2.000 0.069

0.087 LP 0.500 0.091 2.000 0.095 t

After e n t e r i n g t h e s l o t , flow p a s s e s th rough m e t e r i n g or i f ices which l e a d t o a col lector d u c t . The flow from these d u c t s is metered i n t o a s u c t i o n

tube . ?he s p a c i n g of t h e holes e x i t i n g t h e col lector d u c t is p r i m a r i l y d i c t a t e d b y t h e r equ i r emen t t o m a i n t a i n a uniform p r e s s u r e a long t h e collector

d u c t . The diameters of t h e m e t e r i n g h o l e s a r e p r i m a r i l y de te rmined b y t h e

r equ i r emen t t o c o n t r o l t h e p r e s s u r e w i t h i n t h e col lector d u c t t o a p r e d e t e r -

mined l e v e l , w h i l e matching t h e r e q u i r e d flow t o t h e l o c a l p r e s s u r e s w i t h i n t h e a p p r o p r i a t e s u c t i o n t u b e .

Connecting t h e s u c t i o n t u b e t o o n e of t h e t m plenun d u c t s are a series

of l i n e s across t h e span of t h e wing, each c o n t a i n i n g a c o c k p i t - c o n t r o l l e d n e e d l e v a l v e . This c o n f i g u r a t i o n p r o v i d e s i n - f l i g h t c a p a b i l i t y for chordwise

and spanwise s u c t i o n flow p r o f i l e ad jus tment . This v a l v e is t h e l a s t m e t e r i n g of t h e f l o w before t h e flow r e a c h e s t h e plenun d u c t , which has no m e t e r i n g

d e v i c e s , so it is v e r y c r i t i c a l t o s e t t h e n e e d l e v a l v e s c o r r e c t l y . I d e a l l y

TABLE 3. LEADING EDGE METERING SYSTEM

27

t h e flow from a l l t h e s l o t s w i l l p a s s t h rough t h e n e e d l e v a l v e s and e n t e r t h e

plenum d u c t a t t h e correct p r e s s u r e t o e n s u r e even flow t o t h e s u c t i o n pump.

I n s t r u n e n t a t i o n w i l l measure t h e flow th rough each need le v a l v e , and these

i n s t r u m e n t s w i l l h e l p e n s u r e t h a t n e i t h e r t oo much or too l i t t l e s u c t i o n is

a p p l i e d . If too much s u c t i o n is a p p l i e d , t h e boundary layer w i l l become too

t h i n wi th a co r re spond ing loss of performance because o f t h e i n c r e a s e d s e n s i t i v i t y of t h e boundary l a y e r t o g iven s u r f a c e i m p e r f e c t i o n s t h a t w i l l

r e s u l t i n i n c r e a s e d e x t e r n a l d r a g on t r a n s i t i o n . If too l i t t l e s u c t i o n is a p p l i e d , a i r could e n t e r t h e s l o t a t o n e p o i n t , t r a v e l spanwise , and e x i t t h e

s a n e s lo t , t r i p p i n g t h e boundary l a y e r .

4.3.5.3 S u c t i o n U n i t s

The s u c t i o n sys tem for t h e b a s e l i n e c o n f i g u r a t i o n i n c o r p o r a t e s t w i n t e r c h a n g e a b l e fuse lage-mounted s u c t i o n u n i t s , each powered by an independent g a s t u r b i n e power u n i t . Each u n i t i n c l u d e s flow and p r e s s u r e r a t i o c a p a c i t y

s u f f i c i e n t t o punp h a l f o f t h e flow from each s u r f a c e and d i s c h a r g e t h e t o t a l

pumped flow a t t h e freestream f l i g h t v e l o c i t y o f Mach 0.77 a t 37,000 f t

a l t i t u d e . S ince t h e v a r i o u s l amina r i zed s u r f a c e s have d i f f e r e n t s u r f a c e p r e s s u r e s , it is n e c e s s a r y t o d e s i g n t h e s u c t i o n pump t o accommodate t h e

v a r i o u s l e v e l s o f i n l e t p r e s s u r e whi le d i s c h a r g i n g a l l of t h e flow a t t h e same p r e s s u r e 1 ev e l .

S u c t i o n Requirements - The s u c t i o n r e q u i r e m e n t s and e x t e r n a l aerodynamics of t h e wing a i r f o i l are c o n s i s t e n t w i th t h e b a s e l i n e a i r f o i l

developed fo r t h i s a i r c r a f t . These i n c l u d e wing s u r f a c e C d i s t r i b u t i o n , d i s t r i b u t e d s u c t i o n r equ i r emen t s , and boundary l a y e r c h a r a c t e r i s t i c s for b o t h

t h e upper and lower s u r faces.

P

The l a m i n a r i z e d surfaces a r e as follows:

Wing upper Wing lower

*Hor izonta l t a i l - each s u r f a c e *Vertical t a i l - t o t a l

2126 f t 2

2074 f t 2

383 f t 2

675 f t 2

28

A l a n i n a r i z e d s u r f a c e is t h e t o t a l a r e a ove r which l a n i n a r boundary-layer

flow exists, and c o n s i s t s of t h e s l o t t e d s u r f a c e forward of t h e f r o n t spar and t h e a r e a a f t o f t h e f r o n t s p a r t o where t h e boundary layer becomes t u r b u l e n t .

The l amina r i zed wing areas a r e t h e t o t a l for t h e a i r p l a n e and i n c l u d e a d j u s t -

ments fo r 50 p e r c e n t nominal chord l a m i n a r i z a t i o n and a i r f o i l s u r f a c e curva- t u r e . The ( * I f i g u r e s are measurements for t h e empennage. mere w i l l be no a c t i v e s u c t i o n i n t h e empennage on t h e r e v i s e d HLFC b a s e l i n e c o n f i g u r a t i o n , b u t t h e c a p a b i l i t y exists i f a performance b e n e f i t s t u d y w a r r a n t s i t s add i ti on.

S u c t i o n Pump Characteristics .. The most pract ical s u c t i o n pump conf igu ra -

t i o n for meet ing t h e s u c t i o n r e q u i r e m e n t s o f t h e b a s e l i n e a i r c r a f t i s a compact a x i a l flow compressor which i n c o r p o r a t e s a h igh -p res su re compressor

punping t h e t o t a l flow w i t h a d d i t i o n a l lower flow boost u n i t s i n t e g r a l l y located on t h e i n l e t t o ra ise t h e p r e s s u r e of t h e low-pressure flows t o t h e

i n l e t p r e s s u r e of t h e h igh-pressure compressor. .

A r u d i m e n t a r y s u c t i o n u n i t d e s i g n was completed for t h e pu rpose of e s t a b l i s h i n g c o n c e p t u a l s i z e , weight, and g e n e r a l l a y o u t of t h e s u c t i o n u n i t .

While t h i s a n a l y s i s o b v i o u s l y l a c k s t h e r e f i n e m e n t s of an optimized d e s i g n , it

is r e a s o n a b l y a c c u r a t e for s a t i s f y i n g t h e p r e s e n t r e q u i r e m e n t s and s e r v e s t o i l l u s t r a t e some of t h e r e q u i r e d c o n s i d e r a t i o n s . The s u c t i o n pump for t h i s

u n i t is shown i n F igu re 14 , which i l l u s t r a t e s t h e u n i t p a r t i a l l y s e c t i o n e d . The s u c t i o n pump c o n s i s t s o f a forward frame, a two-stage low-pressure or boost e lement , a mid-frame, a fou r - s t age high-pressure e lement , and a scroll

d i f f u s o r . The forward frame s e r v e s as t h e a t t a c h n e n t for t h e a i r c r a f t s u c t i o n sys tem low-pressure d u c t and houses boost e lement v a r i a b l e i n l e t g u i d e vanes . These vanes p r o v i d e a c o n t r o l f o r matching t h e boost e lement flow c o n d i t i o n s

under v a r y i n g f l i g h t s u c t i o n r equ i r emen t s . 'he two-stage boost e l emen t is sized t o m e e t upper wing and empennage s u c t i o n flow and p r e s s u r e r a t i o r equ i r emen t s . The two stages operate a t a modest pressure r a t i o compatible

w i t h t h e d i s t o r t e d i n l e t c o n d i t i o n s t h a t w i l l undoub ted ly e x i s t wi th t h e

s u c t i o n flow. The mid-frame s e r v e s as t h e t r a n s i t i o n d u c t for t h e boost

e lement exhaus t flow t o t h e i n l e t o f t h e h i g h p r e s s u r e e lement . It a l so p r o v i d e s t h e i n t r o d u c t i o n of t h e h igh -p res su re s u c t i o n flow i n t o t h e high-

p r e s s u r e compressor. V a r i a b l e i n l e t g u i d e v a n e s are r e q u i r e d i n t h e high-

29

p r e s s u r e s u c t i o n flow i n t o t h e h igh-pressure compressor . Var i ab le i n l e t g u i d e vanes a r e r e q u i r e d i n t h e h igh-pressure s u c t i o n flow e n t r y pa th f o r o p e r a t i o n i n c o n j u n c t i o n with t h e boos t e lement v a r i a b l e i n l e t g u i d e vanes t o a s s u r e a

proper match between t h e boos t and h igh-pressure elements.

I -

The h igh-pressure e lement is a four -s tage u n i t o f moderate stage l o a d i n g . A t h r e e - s t a g e u n i t would r e q u i r e s t a g e l o a d i n g s t h a t are h i g h b u t c o n s i s t e n t

wi th f o r e s e e a b l e p r a c t i c e f o r c o n v e n t i o n a l engine compressors . However t h e a n t i c i p a t e d i n l e t d i s t o r t i o n s and mismatch t o which t h i s u n i t w i l l be

s u b j e c t e d d i c t a t e t h e u s e of a more c o n s e r v a t i v e fou r - s t age c o n f i g u r a t i o n . Both t h e boos t and h igh -p res su re elements o p e r a t e on a common s h a f t .

PRIMARY COMPRESSOR

The exhaus t d i f f u s e r co l lec ts t h e d i s c h a r g e flow of t h e s u c t i o n pump and

t u r n s it th rough goo, w h i l e reducing t h e flow v e l o c i t y t o 0.3 Mach and al low- i n g t h e passage o f t h e s u c t i o n pump d r i v e s h a f t a x i a l l y th rough t h e c e n t e r of

t h e scrol l . The flow t h u s exi ts t h e s c ro l l i n a round d u c t a t a r i g h t a n g l e t o t h e axis of t h e s u c t i o n p m p . The d i f f u s o r / s c r o l l a lso p rov ides for a

DIFFUSER AND MIXER

/ BOOST COMPRESSOR

VARIABLE ICV

LOW PRESSURE IN LET

HIGH / PRESSURE INLET

-4

VARIABLE

3 / ICV I - DRIVE SHAFT

J

i

\ I I _

1 - DIFFUSER SCROLL

EXHAUST

Figure 14. Suction Pump

30

r i g i d mounting between t h e s u c t i o n pump and d r i v e u n i t . This i n c l u d e s a mounting for t h e d r i v e s h a f t hous ing a s well' as an e x t e r n a l t r u s s s t r u c t u r e t o

m a i n t a i n s h a f t a l i g m e n t and a b s o r p t i o n of t h e t o r q u e between t h e s u c t i o n pump and t h e power u n i t .

The s u c t i o n pumps a r e d r i v e n by independent power u n i t s provided wi th ram

i n l e t s exhaus t ing a t e s s e n t i a l l y f r e e s t r e a m v e l o c i t y . The independent d r i v e was adopted because i t h a s no impact on t h e p r imary p r o p u l s i o n u n i t s and can t h e r e f o r e be independen t ly s i z e d . In p rev ious studies, a l t e r n a t i v e sys t ems were cons ide red and i n c l u d e g e a r e d , b l eed , and b l eed /burn sys tems. The

p e n a l t i e s of t h e more complex sys t ems led t o t h e i r e l i m i n a t i o n . A conven- t i o n a l b u t advanced t echno logy s h a f t e n g i n e was adopted for t h i s s t u d y . The

t o t a l s u c t i o n u n i t weight was eva lua ted a t 536 l b ; t h i s f i g u r e i n c l u d e s b o t h t h e s u c t i o n pump and t h e power u n i t .

The performance charac te r i s t ics of t h e s u c t i o n pump fo r t h e b a s e l i n e HLFC t r a n s p o r t a r e provided i n t h e fo l lowing:

Power a t c r u i s e a l t i t u d e Mass flow a t c r u i s e a l t i t u d e

Pump speed

Punp p r e s s u r e r a t i o

Pump i n l e t p r e s s u r e Pump i n l e t temper a t u r e

C r u i s e a l t i t u d e Mission f u e l

Two s u c t i o n pumps are r e q u i r e d

200 H. P. 300 Lb/Min

50,000 RPM 4.0 4.1 p s i a

460' R 31,685 f t

2577.8 Lb

for e a c h HLFC t r a n s p o r t .

4.3.5.4 C o n t r o l s

C o n t r o l o f t h e LFC s u c t i o n sys tem p r e s e n t s a nunber of complex and un ique

problems. The r e q u i r e d s u c t i o n flow l e v e l s and d i s t r i b u t i o n s for re l iab le

l a n i n a r i z a t i o n are s u b j e c t t o t h e effects of p r o d u c t i o n t o l e r a n c e and de-

t e r i o r a t i o n , i n a d d i t i o n t o t h e v a r i a b l e f l i g h t c o n d i t i o n s . Ihe n e e d l e v a l v e s

and excess c a p a c i t y of t h e s u c t i o n pump a re required t o n e g a t e t h e effects of p roduc t ion t o l e r a n c e s and d e t e r i o r a t i o n , as well as t h e l i m i t e d v a r i a t i o n s i n

f l i g h t c o n d i t i o n s . It is a p p a r e n t t h a t s u f f i c i e n t r a n g e is n o t l i k e l y t o

31

e x i s t t o abso rb a l l o f these i n f l u e n c e s o v e r t h e e n t i r e f l i g h t s p e c t r u n and a c o n t r o l s y s t e m t o accomnodate these v a r i a b l e s m u l d become ext remely complex.

In t h e in te res t o f s i m p l i f y i n g t h e c o n t r o l sys t em, t h e r e b y improving t h e

r e l i a b i l i t y and r e d u c i n g t h e cost and main tenance , a c t i v e s u c t i o n is a p p l i e d o n l y d u r i n g t h e c r u i s e p o r t i o n of t h e mis s ion and a t h i g h e r a l t i t u d e s d u r i n g

climb and d e s c e n t . Th i s approach w i l l allow t h e c l ean ing /de - i c ing sys tem t o

f u n c t i o n on t h e ground and a t lower a l t i t u d e s , i n order t o e l i m i n a t e t h e

con tamina t ion i n c u r r e d d u r i n g o p e r a t i o n i n those c o n d i t i o n s and p reven t damage

t o t h e s u c t i o n system.

With t h i s approach , t h e major c o n t r o l of t h e s u c t i o n flow becomes t h a t of

o p t i m i z i n g t h e s u c t i o n r e q u i r e m e n t s between s t a r t c r u i s e and end c r u i s e . An

a p p r e c i a b l e change i n b o t h l e v e l and d i s t r i b u t i o n of wing s u r f a c e C v a l u e s

o c c u r s between these c o n d i t i o n s , p a r t i c u l a r l y i n t h e d i f f e r e n t i a l s between t h e

upper and lower s u r f a c e s . ?he i n t e r n a l s u c t i o n system p r e s s u r e s a r e d i c t a t e d

by these C v a l u e s . The s u c t i o n p m p uses v a r i a b l e i n l e t g u i d e vanes i n b o t h

t h e low p r e s s u r e and h i g h p r e s s u r e i n l e t s . These vanes a d j u s t t h e d u c t suc-

t i o n p r e s s u r e s t o t h e v a r y i n g upper and lower wing s u r f a c e C v a l u e s whi le

m a i n t a i n i n g desired s u c t i o n flows and an a c c e p t a b l e match between t h e p r imary

and boost e l e m e n t s o f t h e s u c t i o n p m p .

P

P

P

However, t h i s does n o t p r o v i d e discrete c o n t r o l t o accommodate t h e change i n t h e chordwise C d i s t r i b u t i o n s . The s u c t i o n sys tem is des igned t o minimize s e n s i t i v i t y t o t h i s change t h r o u g h correct s e l e c t i o n of me te r ing ho le s and

t h e i r s p a c i n g . To f u r t h e r e n s u r e t h a t l amina r flow is main ta ined th roughou t

changing c o n d i t i o n s , v a r i a b l e n e e d l e v a l v e s are i n s t a l l e d i n t h e l ines between

t h e s u c t i o n t u b e s and t h e main plenum d u c t s . These v a l v e s are c o n t r o l l e d from t h e c o c k p i t , and i n s t r u n e n t a t i o n w i l l be p r e s e n t t o d i s p l a y t h e amount of

s u c t i o n flow th rough each s l o t and t h e spanwise s lo t flow d i s t r i b u t i o n . An au tomat i c c o n t r o l sys tem w i l l mon i to r t h e s u c t i o n flow and a l t e r t h e n e e d l e

v a l v e s a s r e q u i r e d t o m a i n t a i n p r e s c r i b e d s u c t i o n l e v e l s .

P

The remain ing c o n t r o l problems are p r i m a r i l y o p e r a t i o n a l i n n a t u r e and c o n s i s t o f :

( 1 ) S u c t i o n u n i t s t a r t i n g a t b o t h sea l e v e l s t a t i c and a l t i t u d e .

32

( 2 ) Uni t f a i l u r e i n c r u i s e .

( 3 ) Atmospheric c o n d i t i o n s a t c r u i s e .

( 4 ) Sea l e v e l s t a t i c sys tem checkout .

S t a r t i n g t h e u n i t s a t a l t i t u d e w i l l p r e s e n t some problems because t h e

p r e s s u r e s a t t h e s u c t i o n pump i n l e t are a p p r e c i a b l y below ambient. I n t h e shutdown c o n d i t i o n , a s i g n i f i c a n t p r e s s u r e r a t i o e x i s t s across t h e s u c t i o n

pump. 'his p r e s s u r e r a t i o exceeds t h e c a p a b i l i t i e s o f t h e s u c t i o n pump u n t i l r o t a t i o n a l s p e e d s n e a r d e s i g n are achieved . This means t h a t t h e s u c t i o n pump

would be s t a l l e d th roughou t t h e s t a r t r a n g e u n t i l n e a r d e s i g n s p e e d s a r e

a t t a i n e d d u r i n g t h e s t a r t up. T h i s i s u n a c c e p t a b l e d u e t o power r e q u i r e m e n t s

and p o t e n t i a l danage t o t h e pump. To avoid t h i s problem, v a l v e s are provided i n t h e d u c t i n g sys tem nea r t h e pump i n l e t s t o i s o l a t e t h e s u c t i o n d u c t i n g and

v e n t t h e pump i n l e t t o ambient a i r d u r i n g t h e s t a r t c y c l e , a s shown on F i g u r e

15. When t h e s u c t i o n u n i t r e a c h e s a p r e s c r i b e d ro to r speed , t h e v e n t v a l v e

w i l l s l o w l y close w h i l e t h e i s o l a t i o n v a l v e s l o w l y opens acco rd ing t o a pre- scribed s c h e d u l e , T h i s o p e r a t i o n may be carried o u t ei ther b y an a u t o m a t i c

sys tem operated b y a " s t a r t - run" switch or manual ly by t h e f l i g h t e n g i n e e r .

VENT VALVES \ ,I RAM INLET /

DRIVE SHAFT

SUCTION JET EXHAUST

PART SPAN ISOLATION VALVES

L P SUCTION TRUNK DUCT CROSSOVER

DUCTS

ISOLATION VALVE

' \ I \ """' I/---- \

H P SUCTION TRUNK DUCT

POWER GENERATOR

Figure 15. Wing Suction System Schematic

33

In t h e even t of an i n f l i g h t s u c t i o n u n i t f a i l u r e , i n s t r u m e n t a t i o n a t the

s u c t i o n pump in l e t and d i scha rge w i l l immediately sense t h e f a i l u r e and s h u t

t h e u n i t down w h i l e s imu l t aneous ly c l o s i n g t h e i s o l a t i o n v a l v e s t o t h a t u n i t . After t h e i s o l a t i o n v a l v e s have c l o s e d , t h e v a l v e s l o c a t e d i n t h e low- and

h igh-pressure s u c t i o n plenum d u c t s near t h e wing mid-semispan, shown i n F igu re 14, w i l l a lso close, a l lowing o n l y inboa rd wing s u c t i o n .

In F igu re 15, there is an a l lowance f o r d u c t i n g back t o t h e empennage.

This f i g u r e i n c l u d e s t h e empennage d u c t t o show how it would t i e i n t o t h e

d u c t i n g network. In t h e empennage i t se l f , t h e s u c t i o n systems would be v e r y

s i m i l a r t o those i n t h e wings. In t h e even t of an i n f l i g h t s u c t i o n u n i t f a i l u r e ( t h e s i t u a t i o n mentioned above ) , t h e empennage v a l v e w i l l close

s i m u l t a n e o u s l y , e l i m i n a t i n g empennage s u c t i o n .

I n t h e e v e n t o f s l o t blockage, s u c h a s t h e i n c i d e n c e o f r a i n and ice

c r y s t a l s i n t h e c r u i s e mode and subsequent f a i l u r e of t h e s l o t c l ean ing /de - i c i n g sys t em, immediate shutdown of t h e s u c t i o n system would be r e q u i r e d t o p reven t pump s t a l l a s a r e s u l t of a i r f l o w s t a r v a t i o n . This w u l d be accom- p l i s h e d th rough s e n s i n g an a b r u p t i n c r e a s e i n pump p r e s s u r e r a t i o , s i g n a l i n g i n t e r f e r e n c e wi th t h e s u c t i o n f low i n g e s t i o n , and a u t o m a t i c a l l y s h u t t i n g t h e

system down. P r o v i s i o n could be made f o r au tomat i c re-start, or re-start cou ld be t h e r e s p o n s i b i l i t y of t h e f l i g h t eng inee r . An inc remen ta l i n c r e a s e

i n p r imary p r o p u l s i o n e n g i n e t h r u s t could be a u t o m a t i c a l l y accomplished t o compensate for t h e temporary d e l a m i n a r i zation.

A p re - f l i gh t s u c t i o n system checkout must be accomplished a t sea l e v e l

s t a t i c c o n d i t i o n s pr ior t o i n i t i a t i o n of t h e f l i g h t . This may be accomplished by t h e f l i g h t e n g i n e e r and w u l d c o n s i s t of a normal s t a r t w i t h t h e s u c t i o n

u n i t rotor speed l i m i t e d t o a low v a l u e t o minimize i n g e s t i o n of con taminan t s

t o t h e s u c t i o n sys tem and p r e v e n t e x c e s s i v e n o i s e i n t h e t e r m i n a l area. This

s t a r t would d u p l i c a t e t h e c r u i s e s t a r t and v a l v i n g sequence excep t for t h e

reduced rotor speed. S i n c e a l l wing s u r f a c e C v a l u e s are zero under s t a t i c

c o n d i t i o n s , t h e h i g h rotor speeds are not r e q u i r e d . When t h e s u c t i o n u n i t

reaches t h e prescr ibed speed , s u c t i o n system p r e s s u r e s and flows, pump

p r e s s u r e r a t i o , and punp and power u n i t o p e r a t i o n a l parameters (i.e., o i l p r e s s u r e , t u r b i n e t e m p e r a t u r e , f u e l flow, etc.) may be compared t o prescr ibed

P

34

limits. An adve r se a tmospher ic c o n d i t i o n m u l d be s imula t ed b y a s i g n a l selected b y t h e f l i g h t eng inee r t o a c t u a t e t h e au tomat i c shut-down sequence.

It is expected t h a t t h i s ground check may be normally accomplished i n a t o t a l

time o f less than 4 minutes .

4.3.5.5 Leading Edge Region Cleaning

S ince t h e ear l ies t c o n s i d e r a t i o n of app ly ing l a n i n a r flow c o n t r o l t o an

o p e r a t i o n a l a i r c r a f t , t h e p o t e n t i a l problems a t t e n d i n g leading-edge roughness due t o i n s e c t con tamina t ion and ice a c c u n u l a t i o n have been a c o n t i n u i n g con- c e r n . The i n s e c t a l l e v i a t i o n and a n t i - i c i n g sys tems des igned for t h i s LFC a i r c r a f t w i l l p r o v i d e p r o t e c t i o n from wing-sur face s l o t con tamina t ion due t o i n s e c t s a t low a l t i t u d e s and a method of de- ic ing t h e . . l e a d i n g edge a t a l l

a l t i t u d e s of o p e r a t i o n . The c l ean ing /de - i c ing sys t ems w i l l emit a f l u i d th rough t h e wing-surface s l o t s which w i l l c l e a n t h e wing s u r f a c e and p reven t

i c i n g , and t h e n pu rge t h e system of t h i s l i q u i d so t h a t s u c t i o n may be s a f e l y s ta r ted . The flow o f t h i s f l u i d w i l l be c o n t r o l l e d and monitored from t h e

c o c k p i t .

Cleaning/&-Icing System - The c l ean ing /de - i c ing sys tem, F i g u r e 16, w i l l

u s e selected s lo ts around t h e l e a d i n g edge t o d i s c h a r g e a sma l l flow of l i q u i d o n t o t h e s u r f a c e . This l i q u i d forms a p r o t e c t i v e f i l m ove r t h e s u r f a c e t o p reven t i n s e c t a c c r e t i o n and prevent/remove ice bu i ldup . The d e s i g n r e q u i r e - ments f o r t h i s l i q u i d flow have been demonstrated t o p r o v i d e a l i q u i d f i l m

s u f f i c i e n t t o prevent a c c r e t i o n i n a h igh-dens i ty i n s e c t environment . The l i q u i d used is a 60 p e r c e n t p ropy lene g lyco lme thy l e t h e r , PGME/4O p e r c e n t water mix tu re . A l l components used i n t h e system w i l l be s u i t a b l e for u s e

w i t h t h e PGME l i q u i d , and a l l components c o n t a i n i n g t h e PGME l i q u i d m i x t u r e w i l l be l o c a t e d i n a w e l l - v e n t i l a t e d area wi th s u f f i c i e n t d r a i n a g e p r o v i s i o n

t o p reven t entrapment o f any l i q u i d leakage . The system w i l l be des igned t o

p reven t a flamnable mixture of PGME and a i r from forming i n t h e v i c i n i t y of

any i g n i t i o n source i n t h e e v e n t o f any system l e a k s .

'he c leaning/de- ic ing system, i l l u s t r a t e d s c h e m a t i c a l l y i n F igu re 15

i n t e r f a c e s w i t h t h e s u c t i o n sys tem, t h e pu rge sys tem, and t h e n i t r o g e n

p r e s s u r i z a t i o n system. The major components of t h e system are t h e l i q u i d

35

N I TROCEN SUPPLY

TANK 7 RELIEF VALVE

SHUTOFF

LIQUID TRAP

OVERBOARD VENT

SURFACE

SLOT FLOW LIQUID ADJUSTMENT SHUTOFF

RESTRICTOR VALVE V A-LV E OR CHECKVALVE WITH RESTRICTOR

FLOWMETER

PRESSURE REGULATOR LIQUID

FLOW TO OTHER SLOTS

PRESSURIZATION :Liy SHUTOFF

I LIQUID I RESERVOIR 6 LIQUID

SYSTEM CONTROL VALVE

INSTRUMENTATION PRIMING

SUCTION SHUTOFF VALVE

TO SUCTION SYSTEM

F i g u r e 16. Cleaning/Ant i - Ic ing System

s u p p l y t a n k and t h e f l u i d d i s t r i b u t i o n system. "his system will d e l i v e r t h e

PGME m i x t u r e th rough t h e s i x 3-way v a l v e s a t t h e i n t e r f a c e w i t h t h e s u c t i o n

system t o t h e seven c l e a n i n g / a n t i - i c i n g s lots . These s lo ts ( t m dedica ted

c l e a n i n g s lo t s and f i v e dual-purpose s lo t s ) w i l l be l o c a t e d n e a r t h e l e a d i n g edge. Liquid f low c o n t r o l w i l l be provided by ad jus tmen t of t h e s u p p l y t a n k

n i t r o g e n p r e s s u r i z a t i o n sys t em p res su re . Flow d i s t r i b u t i o n w i l l be r e g u l a t e d by a d j u s t a b l e t h r o t t l i n g v a l v e s located upstream of each 3-way v a l v e .

The supp ly t a n k s w i l l be i n s t a l l e d so t h a t l i q u i d i s s u p p l i e d from t h e

bottom of t h e t a n k t h r o u g h t h e s i n g l e port located n e a r one end of t h e t a n k .

The two-port connec to r a t t h e o t h e r end of t h e t ank w i l l p rov ide for pres-

s u r i z a t i o n / v e n t i n g and f o r l i q u i d r e t u r n / s e r v i c i n g ove r flow as i l l u s t r a t e d i n

F i g u r e 16. The c l e a n i n g / a n t i - i c i n g system w i l l i n t e r f a c e w i t h t h e n i t r o g e n p r e s s u r i z a t i o n system upstream of t h e p r e s s u r e r e g u l a t o r . This n i t r o g e n

p r e s s u r e r e g u l a t o r w i l l be c o n t r o l l e d from t h e cockpit and w i l l connec t t o t h e

n i t r o g e n p r e s s u r i z a t i o n s h u t o f f va lve . Ibwnstream of t h e s h u t o f f v a l v e , t h e

v e n t / p r e s s u r i z a t i o n l i n e w i l l connec t t o an overboard ven t t h rough p a r a l l e l 1 i n e s .

A port on t h e t a n k connec t s t h rough t h e l i q u i d s h u t o f f v a l v e and f i l t e r

t o t h e flowmeter. The l i n e between t h e l i q u i d f i l t e r and flowneter is

36

connected t o t h e remain ing p o r t nea r t h e t o p o f t h e s u p p l y t a n k th rough a

f l u i d r e t u r n l i n e c o n t a i n i n g t h e l i q u i d v e n t s h u t o f f va lve . A t ank f i l l and d r a i n l i n e t ees i n t o t h e f l u i d s u p p l y l i n e between t h e t a n k o u t l e t and t h e

l i q u i d s h u t o f f v a l v e . T h i s l i n e p rov ides for s e r v i c i n g of t h e t a n k w i t h t h e PGMVwater s o l u t i o n for t a n k d r a i n i n g , and c o n t a i n s a manual s h u t o f f va lve .

The f l u i d r e t u r n l i n e l i k e w i s e c o n n e c t s t h rough a tee t o a manual s h u t o f f v a l v e t o p r o v i d e an o v e r f l o w for t a n k s e r v i c i n g . This c o n n e c t i o n is made

between t h e l i q u i d v e n t s h u t o f f v a l v e and t h e t ank . For both manual s h u t o f f v a l v e s descr ibed above, p r o v i s i o n is made for d r a i n i n g PGME f l u i d c lear of t h e

a i r c ra f t . Liquid i s plumbed from t h e flometer t o a man i fo ld plenum located

i n t h e wing l e a d i n g edge a r e a . The mani fo ld is provided w i t h ports for c o n n e c t i n g s i x s l o t c l e a n i n g l i n e s . These s i x s l o t l i n e s are r o u t e d from t h e

mani fo ld t o l i q u i d flow ad jus tmen t v a l v e s a l s o located i n t h e wing root area.

Each v a l v e i s connected t o co r re spond ing s l o t s u c t i o n l i n e a t t h e 3-way s u c t i o n / c l e a n i n g s e l e c t i o n v a l v e located i n t h e wing root. ?he 3+ay v a l v e s

p rov ide a mechanica l i n t e r l o c k t o avoid i n a d v e r t e n t d e l i v e r y o f c l e a n i n g l i q u i d t o t h e n e e d l e v a l v e s . A l l v a l v e s are r emote ly c o n t r o l l e d from t h e

c o c k p i t .

Purge System - The a i r pu rge system ( F i g u r e 17) w i l l be des igned t o remove l i q u i d from t h e c l e a n i n g / a n t i - i c i n g d u c t i n g and t o clear a l l s l o t s of l i q u i d before t h e i n i t i a t i o n of s u c t i o n . T h i s is n e c e s s a r y t o p r e v e n t

con tamina t ion of t h e s u c t i o n system wi th r e s i d u a l c l e a n i n g l i q u i d . Due t o t h e nunerous p o i n t s i n t h e sys t em bhere l i q u i d may be e n t r a p p e d , t h e sys tem is des igned so t h e c l e a n i n g / a n t i - i c i n g sys t em may be ven ted t o draw as much

l i q u i d as p o s s i b l e back i n t o t h e t a n k s and t h e n pu rge t h e r e s i d u a l l i q u i d o u t

t h r o u g h t h e s l o t s . T h i s sys tem i n t e r f a c e s w i t h t h e s u c t i o n and c l e a n i n g / a n t i - i c i n g sys tems.

The remote ly operated v a l v e s i n t h i s sys tem are c o n t r o l l e d from t h e

a i r c r a f t c o c k p i t . Electrical c o n t r o l i n t e r c o n n e c t i o n s are made t o p reven t ei ther t h e s u c t i o n or t h e pu rge v a l v e from be ing e n e r g i z e d t o t h e open

p o s i t i o n u n l e s s t h e o t h e r v a l v e is f u l l y closed.

A h igh-p res su re gaseous n i t r o g e n s u p p l y is used t o p r o v i d e p r e s s u r i z a t i o n

for t h e l i q u i d r e s e r v o i r s of t h e c l ean ing /de - i c ing sys tems, and in s t rumen ta -

37

-.- I WING ROOT ,

- c

1 -, NEEDLE c VALVES c

d

LETA z CONDITIONED AIR SUPPLY

PUMP COCKPIT ---

REGUL?TOR

AIR SUPPLY

Figure 17. Purge System Schematic

t i o n purge. The system a l s o p r o v i d e s pneuna t i c o p e r a t i o n of t h e pu rge system s h u t o f f v a l v e s and p r e s s u r e r e g u l a t o r . The n i t r o g e n source p r o v i d e s n i t r o g e n

a t a r e g u l a t e d p r e s s u r e o f abou t 350 p s i g . The system c o n f i g u r a t i o n is shown

s c h e m a t i c a l l y i n f i g u r e 18. To p rov ide p r o t e c t i o n from e x c e s s i v e r e g u l a t e d

p r e s s u r e , a l i n e is teed i n t o t h e n i t r o g e n l i n e downstream o f t h e p r e s s u r i z a - t i o n s h u t o f f v a l v e and connec t s t o an overboard v e n t t h rough a p r e s s u r e relief

va lve .

The normal a i r c r a f t sys t ems presumed t o be used i n t h e s t u d y a i r c r a f t are those g e n e r a l l y accep ted by i n d u s t r y as being v i a b l e c a n d i d a t e s for improve-

ment or upgrading d u r i n g t h e n e x t decade. Exanples o f s u c h improvement may be f u r t h e r m i n i a t u r i z a t i o n o f e l e c t r o n i c sys t ems , h i g h e r p r e s s u r e h y d r a u l i c

sys t ems t o reduce h y d r a u l i c a c t u a t o r s izes , and t h e major changes i n v o l v i n g fly-by-wire f l i g h t c o n t r o l sys tems i n c o r p o r a t i n g a c t i v e c o n t r o l s .

38

018 VENT To SUPPLY SHUTOFF t r T

MANUAL TANK SERVICING PRESSURE PORT

SHUTOFF VALVE

REGULATOR

RELIEF VALVE

N2 SYSTEM

INTERFACE I

/ - - - - - -

Figure

PRESSURE REGULATOR

SHUTOFF VALVE

t T O T O LIQUID INSTRUMENT SUPPLY PURGE TANK

18. Nitrogen Pressurization System

39

5.0 B A S E L I N E CONFIGURATION DEVELOPHENT

The p l a n developed f o r t h e r e a l i z a t i o n o f c o n t r a c t o b j e c t i v e s r e q u i r e s

t h e deve lopnent o f s t u d y b a s e l i n e a i r c r a f t t o be used a s v e h i c l e s f o r t h e e v a l u a t i o n o f a l t e r n a t i v e LFC sys t em c o n c e p t s d u r i n g subsequent s t u d y phases . Th i s s e c t i o n summarizes t h e a n a l y s e s conducted i n t h e p rocess o f deve lop ing

t h e b a s e l i n e c o n f i g u r a t i o n s . The b a s e l i n e c o n f i g u r a t i o n s c o n s i s t o f advanced t echno logy t u r b u l e n t f l ow a i r c r a f t , a s well a s hybr id L F C a i r c r a f t , a l l s i z e d

t o per form t h e mis s ion c h a r a c t e r i s t i c s d e s c r i b e d i n S e c t i o n 4.2. A s mentioned p r e v i o u s l y , t h e b c k h e e d GASP computer program h a s been used t o s i ze and d e f i n e a l l a i r c r a f t i n t h i s s tudy .

5 . 1 TURBULENT FLOW AIRCRAFT SELECTION

Paramet r i c d a t a f o r t u r b u l e n t f low a i r c r a f t are p r e s e n t e d i n F igu re 19

for wing sweep a n g l e s va ry ing from 25' t o 40' and a t a cruise Mach number of 0.77. The d a t a i n F igu re 19 summarize t h e o u t p u t of t h e GASP and i n c l u d e mis s ion c h a r a c t e r i s t i c s ; w e i g h t s ; wing d a t a ; mi sce l l aneous d a t a such a s l i f t -

to -drag r a t i o , eng ine t h r u s t , and HLFC d a t a when a p p r o p r i a t e . These d a t a a r e a r r anged i n columns s t a r t i n g wi th a reference c o n f i g u r a t i o n w i t h wing sweep o f

25' fol lowed by Options 1 th rough 3 f o r wing sweep a n g l e s o f 30°, 3S0, and 40°, r e s p e c t i v e l y . I n columns D-1 t h r o u g h D-3 t h e p e r c e n t a g e change i n each

a i r c r a f t d e s i g n parameter a s compared t o t h a t f o r t h e r e f e r e n c e c o n f i g u r a t i o n is d i s p l a y e d .

Pa rame t r i c d a t a f o r t u r b u l e n t f low a i r c r a f t p re sen ted i n F igu re 19 show a

s l i g h t s u p e r i o r i t y o f t h e Option 1, 30' sweep c o n f i g u r a t i o n based on an over -

a l l comparison o f m i n i m u m f u e l burned , maximum l i f t - t o - d r a g r a t i o , L/D, and

m i n i m u m g r o s s weight . Accordingly, t h e Option 1 c o n f i g u r a t i o n was s e l e c t e d a s

t h e b a s e l i n e t u r b u l e n t f l ow a i r c r a f t . A g e n e r a l arrangement drawing is pre- s e n t e d i n F igu re 20 wi th a l i s t i n g o f p e r t i n e n t d e s i g n and aerodynamic para-

meters. These pa rame te r s i n c l u d e a t a k e o f f g r o s s weight o f 616,125 l b s , l i f t

t o d rag r a t i o o f 26, t h r u s t per eng ine o f 30,195 l b s , a wing a s p e c t r a t i o o f 13.54, wing span o f 255.9 feet , and a c r i t i c a l f i e l d l e n g t h o f 7,558 feet.

40

.-. I 1 I I I 1 I I I I I1 I1 I 1 I I I I I1 I 1 I I I1 I I II I1

!!

TURBULENT BASELINE SWEEP (25,30,35, & 40 DEGREES) STUDY 19-IUV-IY1b

............................................................................................................................ 1 1 COIVb1ISOI - ICOCL*T CMbNLL FCU1 UbSkLIhC I 1 ............................................................................................................................ 1 1

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I I1 I I I I I I I I I 1 I 1 I 25.91 I 1 2 5 . 9 1 I 1 as.11 I I 2 5 . b 2 I I 0 . 0 1 I 1 -0.17 I I -1.11 I I

I I

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I 1 ~ h N U ~ ~ / M & I C M l I 0.13 I 1 0.19 I I 0.19 1 1 0.19 I 1 - U . I O I I u.21 I 1 0.7s 1 1 I 1 V L Y T b N k b - LO i T I 471.a1 I I 4 1 9 . I 2 I I au1.00 I 1 114.11 1 1 -1.12 I I -11.99 I I -fu.ma 1 1 1 1 11u11Z b1Lb - IO fT I 41b.12 I 1 bI1.64 I1 429.56 I1 442 .01 I1 -1.10 I1 U . 1 0 1 1 1.1, 1 1 . _ . . ........................................ , I .-..--

I 1 abSIC SUCCI - OCG I 15.00 I I 10.00 1 1 15.00 1 1 40.00 I 1 20.00 I I I U . O ( I

II L o i n t i c - w s a rr I 11a.bb 1 1 I 2 4 . b l 1 1 110.01 1 1 I 1 b . U I 1 4.15 I 1 * . S b I 1 1 1 . 4 8 1 1

I 1

1 1 TU.UST(CYCINO - L I I [email protected] I 1 10~195.00 I1 J0.1fi9.00 I1 10.IY9.00 I I -1 .58 I I -1.27 I1 * 0 . 2 b ( 1

Figure 19A. Turbulent Baseline Parametric Data; M=O .77

IW-MUU-I91b 1__--_1.-.1-.-._..__..-..11.--..--------------......--...--.-.-..----1-----111-..-1----.--- l-_.l.-.-.-.------.-..-- 1 ~ ~ , a m ~ ~ c I u w i n n I I I UVTIIIII a I I UYIIUII 1 I I Y - I I I o - i I I n - 1 I I II ............................................................................................................................ I I I 1 D l K L L L b I E O U I ICON’TI I I I I 1 I 1 I 1 I I I 1 I I

I I 1 I 1 I 1 I 1 I I I 1 I 1 II UJU.LII LLTTIIC 0.11 I I 0.11 I I 0.11 I I 0.11 1 1 -0.11 I 1 - * . I 1 I 1 - 0 . 2 1 1 1 I t I I IbJIC U I N C blkb I 4,llS.W I I 4,119.11 I I 1,911.24 I I lr1lb.S) I 1 - 5 . 1 5 . I 1 -1u.11 I 1 -12.11 I 1 I I TUTbL UtIC bRkb I Sr l lb .O1 I I 4 , O S . 1 9 I1 4 , b l S a 1 9 I I 4,419.50 I 1 - $ . I 5 I I -1U.II I1 -12.1- I 1 I I U T bltb - t TOTAL I 14.19 I 1 14.19 I 1 14.19 I I 14.79 I I 0.00 I I U.OB I1 0.OJ I1 I I TOTbLIRbIIC I 1.11 I I 1.11 I1 1.11 I 1 1.17 1 1 0.00 I 1 u.00 I 1 0.UY I 1 I I 1 . C . LUCTIOU t C I I 1 I I I 1 I I I I I I I I 1 1 L . ~ . T * b I ; I I T I O N 8 C I I 1 I 1 I I I I I I I 1 I 1 I 1 C l l i LUC * 8 1bC I I1 I I I I 1 1 II I I I 1 I I r U L U b C C 11/10 V I I 4.Ub I1 4.05 1 1 4.U5 I 1 4.05 I I -0.24 I I -u.d4 I I -0.24 I I

I 1 ............................................................................................................................ ....................................................................................................................................

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

I I I 1 I1 I I 1 1

I1 I I I1 I I I 1

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I I I 1 I 1 I I I 1

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I U T U I

I . U l l U u u m TUIIULCIIT IIbCYbfT, 25 U I I C I1LPY 1. O V T l U I I m TUUIULCNZ b I I C 1 b f T , 10 1bIIC SULLY I . U I T I U U 2 0 T U I W L C 1 1 b I I C 1 b r T r I5 IbIIC S U U ? 4. M I 0 8 1 T U R W L U T U R C I U T , b0 IIIIC IIIU

Figure 19B. Concluded

41

PAYLOAD RANGE MACH NO. ALTITUDE TOCW FUEL L I D MAC SPAN AR C14 SWEEP

132,500 LB 6 , 5 0 0 NM' 0.77 32,119 F T 616,125 LE 291,401 LB 25.99

255.91 FT 13.54 30 DEC

22.88 FT

I

*SEE FIGURE 3

Figure 20. Turbulent Flow Baseline Design Concept

Other f e a t u r e s of t h e a i r c r a f t i n c l u d e nose loading c a p a b i l i t y o n l y ,

l a n d i n g gear f l o t a t i o n f o r hard s u r f a c e runways, f u l l span wing f u e l t a n k s , no

lead ing edge h i g h l i f t devices, 25 percent chord t r a i l i n g edge f l a p s , accom-

modat ions f o r 3 p i l o t s , one loadmaster, and two bunks for t h e l o n g r ange mis s ion .

5.2 HLFC GROUNDRULES

"he ground r u l e s for t h e conduct o f t h e 21 paranetric s i z i n g studies for hybr id LFC a i r c r a f t a r e l i s t e d i n f i g u r e 21 for t h e b a s e l i n e c o n f i g u r a t i o n s .

H i g h l i g h t s of these g r o u n d r u l e s i n c l u d e p r o v i s i o n s for a c t i v e s u c t i o n on t h e

wing and empennage from t h e l e a d i n g edge t o 15 p e r c e n t of t h e chord and

a c t i v a t i o n of t h e HLFC system o n l y a t i n i t i a l c r u i s e a l t i t u d e . Turbu len t flow is assumed t o occur d u r i n g 6 p e r c e n t of c r u i s e f l i g h t time t o a s s u r e mis s ion

comple t ion should atmospheric c o n d i t i o n s p r e c l u d e t h e u s e of HLFC for shor t

periods du r ing c r u i s e . The 12 p e r c e n t excess c r u i s e t h r u s t p r o v i d e s t h e

c a p a b i l i t y t o m a i n t a i n c r u i s e a l t i t u d e and/or speed wi th t h e HLFC sys t em

i n a c t i v e . A low w i n g , a f t f u s e l a g e mounted e n g i n e c o n f i g u r a t i o n s i m i l a r t o

t h a t of t he TAFAD s t u d y c o n s t i t u t e s t h e b a s e l i n e a i r c r a f t . Inc luded i n t h e

42

s i z i n g program i s a l i m i t i n g t r a n s i t i o n Reynolds number based on t h e d i s t a n c e

from t h e wing l e a d i n g edge t o t h e d e s i r e d p e r c e n t chord f o r laminar flow.

T h i s f u n c t i o n p r e v e n t s u n r e a l i s t i c p a r a m e t r i c o p t i m i z a t i o n s o f HLFC

c o n f i g u r a t i o n s i n t h e si z ing process .

5.3 HLFC AIRCRAFT SELECTION

I n i t i a l p a r a m e t r i c s i z i n g d a t a f o r t h e HLFC a i r c r a f t are p resen ted i n

F i g u r e 22 i n t h e same g e n e r a l format as t h a t for t h e t u r b u l e n t flow a i r c r a f t

i n F igu re 19 b u t w i t h t h e a d d i t i o n of HLFC p e c u l i a r da t a . These HLFC p e c u l i a r d a t a i n c l u d e weight of t h e l e a d i n g edge c l e a n i n g f l u i d ( F i g u r e 22A), chordwise

e x t e n t of l e a d i n g edge s u c t i o n , chordwise l o c a t i o n of t r a n s i t i o n p o i n t from

l a n i n a r t o t u r b u l e n t f low, and a l l o t h e r system weight a d d i t i o n s ( F i g u r e 228).

0 WING & EMPENNAGE ACTIVE SUCTION = 15% CHORD

0 WING FRONT AND REAR BEAM @ 15 & 65 8 CHORD

0 HLFC ACTIVATED ONLY UPON REACHING INITIAL CRUISE ALTITUDE

0 TURBULENT FLOW = 6% CRUISE TIME

12% MINIMUM EXCESS CRUISE THRUST AVAILABLE

0 WING T.E. FLAPS = 25% WING CHORD

0 INDEPENDENT HLFC SUCTION POWER SYSTEM

0 ACCOMMODATIONS = 3 PILOTS, 1 LOADMASTER, AND 2 BUNKS

Figure 21A. HLFC Aircraft Ground Rules

LOW-WING CONFIGURATION

PEW STF-686 AFT FUSELAGE MOUNTED ENGINES

0 NOSE LOADING CAPABILITY ONLY

0 HARD SURFACE LANDING GEAR

0 AERO SURFACE L.E. HOT AIR ANTI - ICE SYSTEM DELETED

0 FULL SPAN WING FUEL TANKS

0 L.E. DEVICE DELETED

I Figure 21 B. Concluded

43

OE POOR QUALITY

.. .. .. II LLEA SO f? 1 4.uJb.JY I ) 4 , U b l . 1 1 ) I 4 ,491 .80 I 1 4 , 1 4 4 . 0 0 II $ . e l II -0 .9b II 0 . 1 1 II U L I L N t - LO I 5c ,911 .00 I I b 9 , 7 b 5 . 0 0 l l bS,bIO.00 l l bY.451 .00 I I 18.1b II 11.07 I 1 17.10 II I L I C I I ? - L O I S 0 i T I 12.19 II 1 4 . 1 1 II 14.56 II 14 .J1 I 1 11.50 l l I Y . 4 1 II 11.5b l l I S P L C I Y b t l U I 13.54 II 11.86 I 1 1 4 . I I I 1 I J . 9 5 II 2 - 1 6 l l 4 .11 I 1 1 .01 1 1 I I USIC SWLEP - nEc I 10.00 ll 1 0 . 0 0 II 15.00 I I 1O.UO I 1 -1J .J ) II -1b .6b II -11.11 II I I ah? SULCC - DkC I 1 5 . 5 J II 1 5 . 0 0 l l 10.03 I 1 1b.00 1 1 -19 .63 1 1 *I5.5b 1 4 -d9 .bJ II II LObOIhG - LBlsO fl I 1 7 4 . 4 1 II LlU.YO I l 116.69 I 1 I I Y . I ~ I I - 4 . 4 3 II 1 . 1 ) II - 4 . 2 1 II II fUCL WOL. U T I 0 I 1.00 I 1 1.00 I 1 1.00 II 1.00 II 0.00 II u.00 II 0.BU I 1 II W A N - FT I 155 .91 l l 1 5 9 . 7 4 l l 1 5 1 . 5 b I I 1bU.00 II 1 .49 II -1.30 II 1 . b ~ II II e.? LLlSPAN TT I 0.00 II l b . 4 9 II 14.99 II 16.14 II LYW II LYO II &*a II I I I b C - rT I 11.10 II 11.bU l l 11.U9 I I n . 6 ~ I I -U.UI II 0.04 II - 1 . 1 ~ lI I I t / e - t I 11 .49 l l 11.14 I I 11.b1 II 11.11 II -11.97 II - 8 . 0 1 II -11.11 II I, __.-_1-_1_-_11_--__1.-..-...---.-.----------.--.-.---.----.---..---..-.----.---..--.--.---.--.------------------------------- I 1 * I ICLLLbIIFOUb I 1 1 I 1 I 1 II I 1 II I 1 I I I I 1 II I 1 II II II II

I 10.91 1 1 10.18 II 10 .92 I 1 1 I . V J II 18 .41 II 1 8 . ~ 0 II 1 5 . 9 9 II II L i n II MLlb I 1 0 . 0 1 I 1 l J . 8 0 I / 31.10 l l 1 3 . 8 0 II I U a I II 11 .4J 1 1 11.9- II I I CLUISE SFC I 0 . 5 b II 0 . 5 1 I I 0.57 I 1 0 . b l 1 1 2 . d 5 II 1.15 II ~ . u a I I II CL R b A T.U. I 2.bO I I 1 . 5 ) I 1 2 . 4 4 I 1 1 . 5 ) II - 1 s . 9 II -*.IS 1 1 - 2 . 0 , 1 1 ll CYUlSL CL I 0.32 I 1 0 . 4 9 1 1 0 .49 I 1 0 . 4 1 II -..dl II - a . i o II - 1 . 0 ~ II

II T ~ U U S t l n L l C n ? I 0 . 1 9 l l 0.11 I 1 0 . 1 1 1 1 0.81 1 1 - a . I J II - 9 . J I 11 - ~ . J J II

I 1

1 1 ~ ~ M u ~ T I L N I . I ~ I L - LB I J O . I Y 5 . 0 0 I 1 17 .1d1 .00 I I 16,999.l lO II 1b .9Jb .00 l l -9 .51 II -11 .69 1 1 -1U . lY 1 1

II VLYT bYkA - SO )‘I I 4 J 9 . > 1 I 1 fb1 .50 I 1 5J3.J. I I 541 .00 I 1 d9.11 II dU.6b I 1 d 4 . 4 3 1 1 b l u . 0 0 I 1 51.09 II 48 .4b 1 1 4 1 . 0 4 II I I W U S l Z b M b A - S O ) 1 I 411.u4 II 0 9 . 5 b I ) 6 1 1 . 9 U II

I I ~__~I_________._~~~_~~~~~~~--~~~~~~~~-~.----.--*.-------.~.---..-.--*------...--.--.---..-------.--~--------~~.--~~-~~~~.~-~ I #

F igure 22A. Parametric Sizing Data fo r HLFC Ai rcraf t ; M=0.77, In i t ia l Concepts

The comparison o f p a r a m e t r i c d a t a fo r H L F C c o n f i g u r a t i o n s presented i n F i g u r e 22, i . e . , Opt ions 1 - 3, shows mixed r e s u l t s . F i r s t , a comparison o f

HLFC Opt ions 1 and 2 f o r wing sweep e f f e c t s w i t h t h e same engine l o c a t i o n

shows s l i g h t l y h ighe r l i f t t o drag r a t i o f o r t he Option 1 c a s e b u t s l i g h t l y

lower g r o s s weight and f u e l burned f o r t h e @ t i o n 2 case . Option 1 was deemed

s u p e r i o r because it was f e l t t h a t l e s s l e a d i n g edge c r o s s flow e f f e c t s m u l d

be encountered f o r t h e lower wing sweep o f @ t i o n 1, 20°, a s compared t o t h e

Option 2 h ighe r wing sweep c a s e , 25’. The comparison o f d a t a f o r H L F C Option 1 and Option 3 f o r t h e same wing sweep geometry b u t w i t h t h e eng ines moved forward 5.6 f e e t on t h e f u s e l a g e shows i d e n t i c a l l i f t - t o - d r a g r a t i o f o r bo th c a s e s and e s s e n t i a l l y n e g l i g i b l e b u t lower g r o s s weight and f u e l burned f o r Option 3. The HLFC Option 1 was deemed s u p e r i o r because it was f e l t t h a t more d i f f i cu l t i e s i n ma in ta in ing laminar f low would be encountered f o r Option 3

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

44

I , - - - - - - ~ ~ ~

I1 UIACLLLAULOUS ICbh'T) I I 1 I 1 I1 I1 0.11 1 1 0.11 1 1 1 1 4,149.11 1 1 J.5U1.11 I1

1 1 14.1b 1 1 10.~1 I I I1 4 , b b 7 . @ 1 I 1 4,491.#0 I 1

I 1 1.11 I 1 1.13 I 1 ... I P . 0 0 I1 15.00 I 1

I 1 ' I 5o.ao t i 50.00 I I I ( LOP.00 I I 100.00 I 1 I1 ( . a i I I 4.70 1 1

I1 I1 I 1 11 I I I1 I 1 11

U.11 I I -11.59 I 1 -11.47 1 1 - 1 1 . 4 1 I1 4 .111 .00 1 1 0.11 I 1 -18.01 I1 0.01 I 1 4 , 1 4 4 . 0 0 I1 0.bl I I - * . W 1 1 4.11 I 1

14.90 I 1 - U . l J I 1 Jl.10 1 1 J.11 I 1 1.11 I 1 -0.04 I 1 1.00 I 1 0.12 I 1

1S.00 1 1 I1 I1 11 $O.UO I 1 I1 11 1 1 11.00 II u r n I I w h II u r n I I

4.20 I1 1.1) I 1 4.59 I 1 1.11 I 1 1 I

I 1 U l l C - L l I I MODI2 - L I I 1 UE.? I L l

11 11 11 11 11 1 1 11 11 11 11 11 1 1


with t h e e n g i n e s i n c l o s e proximi ty w i t h t h e wing upper s u r f a c e t h a n for

Option 1. Accord ingly , t h e Option 1 HLFC c o n f i g u r a t i o n was selected as t h e i n i t i a l b a s e l i n e HLFC a i r c r a f t . P e r t i n e n t d e s i g n and aerodynamic pa rame te r s i n c l u d e a t a k e o f f g r o s s weight o f 594,548 l b s , l i f t - t o - d r a g r a t i o o f 30.9,

t h r u s t per e n g i n e of 27,321 l b s , a wing a s p e c t r a t i o o f 13.86, wing span of

259.7 f e e t , and a c r i t i c a l f i e l d l e n g t h o f 8 ,267 feet .

Refinements were made i n t h e aerodynamic and s t r u c t u r a l i n p u t s t o the

HLFC i n i t i a l b a s e l i n e c o n f i g u r a t i o n s i z i n g p r o c e s s t o i n c l u d e ( 1 ) a r e d u c t i o n

i n d u c t weights from a p rev ious s t u d y (Ref. 15) and (2) a change i n l amina r f low time l o s s i n c r u i s e f l i g h t c o n d i t i o n from 10 p e r c e n t t o t h e desired 6

p e r c e n t . These i n p u t changes t o t h e s i z i n g program r e s u l t e d i n v e r y sma l l

changes i n t h e we igh t s and performance o f t h e HLFC r e v i s e d b a s e l i n e a i r c r a f t .

The r e v i s e d p a r a m e t r i c s i z i n g d a t a f o r t h e baseline HLFC a i r c r a f t is c o n t a i n e d

i n Option 1 o f F i g u r e 23. For exanp le , a s compared t o t h e i n i t i a l HLFC

45

b a s e l i n e a i r c r a f t t h e revised HLFC b a s e l i n e a i r c r a f t t a k e o f f g r o s s weight is

591,636 I b s , versus 594,548 l b s , t h e l i f t - t o - d r a g r a t i o is 30.8 v e r s u s 30.9,

t h r u s t per engine is 26,990 l b s , versus 27,231 l b s , and t h e wing a s p e c t r a t i o

is 13.87 v e r s u s 13.86. A g e n e r a l arrangement drawing o f t h e b a s e l i n e HLFC a i r c r a f t is p resen ted i n F igu re 24, a long w i t h o t h e r p e r t i n e n t d e s i g n and aerodynamic pa rame te r s i n c l u d i n g a wing span o f 258.85 fee t and a c r i t i c a l

f i e l d l e n g t h o f 8 ,383 f e e t .

5 . 4 COnPARISON OF TURBULENT AND HLFC AIRCRAFT

%e d a t a i n F igure 23 e n a b l e a comparison t o be made between t h e revised

b a s e l i n e HLFC a i r c r a f t (Opt ion 1 ) and t h e i n i t i a l b a s e l i n e t u r b u l e n t f low a i r -

c r a f t l i s t e d a s t h e r e f e r e n c e a i r c r a f t i n t h e f i r s t co lunn . This comparison shows f o r t h e HLFC a i r c r a f t a r e d u c t i o n o f 13.4 p e r c e n t i n f u e l bu rned , an

I I II STRUCTURAL ULICMT I I URCPdI.SIflN SYSTEM I I SYST1*f L L O U I P . II OPLlATINC COUIP.

U P C E A T I Y G .CIGhT LLlO FULL bEICHT L.C. CLCAN rLUID r U c L PAKLOAO U l C r U L LOA0 .-l_---.-.__....l_...

WING D A I A

I I 1 I I I I I I I I

I .-I - - *. .

128,013.00 15.b15.00 i i , a 5 3 . 0 0 5,311.uo

19a.a1~.00 324.711.00

0.00

l l I I II II II II I I l l I 1 I t l l

l l #....

I I 153.651.00 I 11.101.00 I 14 .441 .00 I 5,llb.OO I 110,425.00 I 352,915.00 I 5,049.00 I 11*,5*1.00

I ~Ilr091.00 I II~,SOO.OO

154,161.00 3#.119,00 14,011.00

111r919.00 354 ,419 .00

4.604.00 155,301.00 iia,soo.oo 111.101.00

5.030.00

*-----.-----. I l l II 1 II II I 5,574.b6 l l 1,211,Ob I 78,594.00 II 11,536.00 I 14.09 I 11.15 I 20.00 I 1 5 . 0 0 I 111.11 I 1.00

l l 15.51 II I3.0 I I 20.00 I I 25.00 I I 114 .14 II 1.00

I 1 6 0 ~ 1 0 1 1 260;16 I I).)* I I 17.11 I 25.17 II 23.51 I 10.10 I I 10.51

i i LISCFLLANCOUS II I I L I D I I * L I D I I CRUISL S f C I I CL *A1 t.0. II C R U I S E CL II TWYUST/LNCINC * L I I I T H R U S I l U C I G H T I 1 YCRT AREA - SO IT l l M C i I I Z A R f A - SO f?

I

I I I I 1 3 , I I I

I I II II II II ll

15.99 I 1 10.11 II 29.51 II 11.90 1 0 . 0 1 l l 11.61 II 11.bO I I 24.56 015b I I 0.Sl II 0.59 II 0.5b 2.00 I 1 1.51 II 1.53 I 1 1 . 5 1 0.11 1 1 d .48 l l 0.4 i II 0.51

,195.oo I I 1i.99t.00 I I ~a.111.00 I I 11 ,001 .00 0.19 I 1 0.18 I I 0.11 I 1 0 . 1 0

419.51 I 1 5 C J . 0 1 II 634.70 I 1 b14.11 411md4 I I 641.15 l l 710.81 l l 611.11

I 1 ¶.I4 I 00.02 I I 10.11 I I

-4 .40 I 4 - 1 4 II 1.16 I 1

-1 .14 II -5.39 I I -::: I 14.61 l l 15.41 I 1 3.11 I 1.b1 I I 9.16 I I

1 . 4 1 I 1 . 1 0 I I 1.11 I I

- I 3 , 4 4 O.EO

-9 .14

I -0.06 I 16.10 I 16.11 I 1.41 I -13.31 I -19.bI I - 4 - 2 a . .~- . I 0.00 I 1 - 1 4 1 can

I I I I I I I I I I I 15.11 I I 1.11 II I 13.15 1 1 11.14 II I 15.51 II 27.19 II I -10.16 I I 1.01 II I -13.11 II -11.11 1 1 I -19.63 I 1 -29.63 1 1 I -10.12 II -1.11 II I 0.00 I 1 0.00 I I I 1.71 II 5.06 I 1 I cnll I 1 111 I I

I I I I I I I I I I

Ia.35 11.19

1 .15 . l e * * -7.u

-10.61 -6.91 11.10 51.12

II I I I I II I 1

Y I I I 1 I I II

11.14 17.9b 5.15

-7.61 -1l.bl

0.11 -1.34 44.40 Ib.41

II II I I I 1 I I II I I I I II II

aa.11 11.71 1.42

- 1 . 0 10.26 5.91 b.11

41.05

II II I 1 I 1 I 1 I I l l l l I I I I

Figure 23A. Parametric Sizing Data for HLFC Aircraft; Speed and Altitude Changes

46 ORIGSNAL PAGE IS

WOR QUALITY

.................................................................................................................................... 11-rra-1917 .................................................................................................................................

I1 --.-.----.---..---..---- I*-.YCI l l U?lILl* I I I O P I I O * 1 I I OP110* 1 I I 0 . I I I 0 - 2 I1 0 - 3 1 I ............................................................................................................................ I 1 UISCCLLAICEUI (COM'II I I 1 I I I1 . I I I I I I I 1 I I I 1 1 I1 I I I I I I I I *O.tL StTIl*C I 0 .11 I1 0.11 I I 0.11 I I 0 . 1 6 I I -11.31 I1 -10.19 1 1 -11.95

I 1 ?OTAL UI*C A R C A I 4,115.39 I I 4.~31.11 1 1 5.514.b4 l l 5,211.0b I1 -0 .Ob I I 15.11 I I 8.18 I I 111 AlCA 8 TUIAL I 14.19 I I 14.11 I I 11.11 I I 14.10 I I -0,Ib I I -11.11 I I -0.65 I I ?O?bL/IAIIC I 1.17 I I 1.11 I I 1.11 I I 1.11 I I -0 .01 I 1 -1.91 I I -0.11 I 1 L.9. I U C I I C I - t c I I I 15.00 I I 15.00 I 1 15.00 I 1 I I I I I1 L . r . I L A I a I I 1 0 * . 8 E I I I 10.00 I1 50.00 I1 50.00 I I I I I1 II ENC LOC - b M A E I I I ioo.oo I I io0.00 I I io0.00 I I t i n I I cnn I I m a

I1 LASIC . I N C A 1 C A I 4.1l9.11 I 1 4 . I l 1 ~ 3 1 I I 4.195.14 I 1 4 . 4 b l . 9 1 I I -0.01 I1 11.11 I1 1.30

I1 I U S t L L C C 11/10 I T I 4.05 I I 6.10 I1 4 . 3 1 1 1 4 . 1 1 I I 3.14 I I * .ab I I 4.18 I , ............................................................................................................................

.I--

I1 I I I I II I I I I I1 I I I 1 I I I I I 1 I 1 I 1 ....................................................................................................................................

I I I I W C I?. COUAllOD I 11.00 1 1 I J .00 I I 13.00 1 1 13.00 I I I 1 I 1 I I 1 ............................................................................................................................ I I

W U I C l l ....................................................................................................................................

l . U * . . C Y m TULBULLU? AllCLAl? 1. OIllOl 1 m MLlC BAILLIIIC, W 0 . 1 1 3. 011101I 2 m YLrc BAICLInc, aa0.10 b. O I I l O l 3 m N L l C IAICLlwL. ll.0.11. Y.3b.000 f T

I I atnucwnc I I I I I *:IC - LI I 0.00 I I 163.00 I 1 *0112 - 1.8 I 0.00 I I 91.00 ii r ir i - - ~ n - I 0;oo I 1 10.00 I I TOTAL L I I 0.00 I I 134.00

I I 1 J 0.00 I 1 11b.00 I 0.00 I I 1b9.00 I 0.00 I1 b15.00 I 0.00 I I t . i i o . a o

.. I I i i rr io i rLuii - E i I 0;oo I I 4.111.00

I I 10TAL DCLtA l?. - La I 0.00 I I i.iai.oo

I 0.00 II 5,aii.oo I ............ 1 1 ............ II TOTAL - L I

I I

I I I I1

w c m r t i c m i AOOIIIOIIS-----------

I I I I I I 111.00 I I *14.00 I I I l l I . 0 0 I I 95.00 I1 I 90.00 I 1 11.00 I I I 811.00 I 1 1 9 1 . 0 0 I I

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

I I1 I1 I I1 I I I 5b3.00 I 1 14b.00 I I I 578.00 I 1 591.00 I I I **1.00 I1 bl1,OO ll I 1,819.00 I1 I , I I 4 * 0 0 I 1 I I I I I

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

I I DASC I I I A S C I I I A S C l l 8AbC I I

1 1 - I I I I1

I bI4.00 I I b1J.00 I I I 5.049.00 I1 4.bOb.00 I I BAS9 I I I b,185.00 l l 5.642.00 I I IASC I I 1 ............ [ I ............ 1 1 ........ 1 1 I a , i i a .oo I I I,ISI,OO I I I A ~ C I I I I1

I ' u a . o o I I ao3.00 I I

I!

I1 I*.b9 I I 9.05 11-97 1 1 4.39 11.50 I 1 10.00 1*.19 I I 8.11

I 1 I I

5.01 I 1 1.U 1.11 I I 1.1b 4.9b I 1 L.10

~~ . . ~~

I I I RW..WCL I1 O??lOl 1 I I OPIIOl 1 I1 0?110* 1 I I 0 - 1 - y 0 - 1 I I 0 -1 I I I I ............................................................................................................................ 1 I .............................

ll-Is~.l9al


PAY LOAD RANGE MACH NO ALTITUDE TOCW FUEL L / D MAC SPAN AR L.E. SWEEP

132,500 L8 6,500 NM* 0.77

31,361 FT 591,636 LE 252,216 LB

30.8 22.6 FT

258.9 F f 13.87 20 DEC

+

L+!L, -

'SEE FIGURE 3

rb \\\

Figure 24 . HLFC Baseline Concept

47

i n c r e a s e of 18.4 p e r c e n t i n l i f t - t o - d r a g r a t i o , a r e d u c t i o n o f 10.6 p e r c e n t i n

eng ine t h r u s t , and a r e d u c t i o n o f 4.0 p e r c e n t i n t a k e o f f g r o s s weight . The

d a t a a l s o shows a 5.4 p e r c e n t i n c r e a s e i n t h e o p e r a t i n g weight empty o f t h e

HLFC a i r c r a f t over t h a t f o r t h e t u r b u l e n t f low a i r c r a f t . This i n c r e a s e i n

o p e r a t i n g weight f o r t h e HLFC a i r c r a f t i s d u e p r i m a r i l y t o t h e 7,721 pounds of

HLFC p e c u l i a r s t r u c t u r a l weight a d d i t i o n s and a l s o t o t h e 53 percent i n c r e a s e

i n h o r i z o n t a l t a i l a r e a a s compared t o t h e t u r b u l e n t flow a i r c r a f t .

Exanina t ion o f t h e t u r b u l e n t f low and HLFC a i r c r a f t pa rame te r s i n d i c a t e s

t h a t t h e t a i l volume c o e f f i c i e n t s f o r bo th c o n f i g u r a t i o n s a r e r e a s o n a b l e . How- ever, t h e approx ima te ly 50 p e r c e n t l a r g e r h o r i z o n t a l t a i l volume c o e f f i c i e n t f o r t h e fuselage-mounted e n g i n e s c o n f i g u r a t i o n r e s u l t s from t h e r equ i r emen t

f o r a s u b s t a n t i a l l y g r e a t e r c e n t e r o f g r a v i t y r ange , 37 p e r c e n t MAC, a s

compared w i t h t h a t f o r t h e wing mounted eng ine c o n f i g u r a t i o n o f 26 p e r c e n t a s

shown i n F igu re 25.

WING-MOUNTED ENGINES

--- FUSELAGE-MOUNTED ENGINES

' *or

COEFFICIENT

INC-MOUNTED

-0.2 0.0 0.2 0.4 0.6 0.8

CENTER OF GRAVITY - MAC

Figure 2 5 . Horizontat Tail Sizing Chart

48

A f t f u s e l a g e mounted e n g i n e c o n f i g u r a t i o n s e x h i b i t wider center o f

g r a v i t y t r a v e l because t h e wing (and i ts f u e l ) are d i s p l a c e d a f t w i th r e s p e c t

t o t h e c e n t r o i d o f t h e payload compartment. "his effect is d i s c u s s e d on page 299, F i g u r e 8-10, o f Reference 19.

The a f t e n g i n e sys t em, because o f t h e r ea rward s h i f t o f we igh t , r e s u l t s

i n nose wheel l i f t - o f f b e i n g much more c r i t i c a l for t h i s c o n f i g u r a t i o n t h a n f o r t h e wing mounted e n g i n e c o n f i g u r a t i o n , a s also shown i n F igu re 25.

5.5 HLFC AIRCRAFT D E F I N I T I O N

5.5.1 C o n f i g u r a t i o n Design

The b a s e l i n e L F C c o n f i g u r a t i o n shown i n F i g u r e 24 i s a wide-body t r a n s -

port c o n f i g u r a t i o n des igned t o c a r r y a 132,500 pound payload a t a r a n g e of

6500 m a t M = 0.77 w i t h adequa te f u e l t o accoun t f o r a d v e r s e winds , i n t e r -

m i t t e n t L F C d i s r u p t i o n s due t o a tmosphe r i c c o n d i t i o n s a t c r u i s e a l t i t u d e , and

normal i n t e r n a t i o n a l f u e l r e s e r v e s . A t y p i c a l payload-range c u r v e is g iven i n F igu re 26.

A t y p i c a l arrangement o f 38x108 ca rgo p a l l e t s is shown i n F i g u r e 27. The

c a r g o compartment is 19.5' w i d e , 13.5' h igh and 110.75' long . In a d d i t i o n t o c a r g o , v a r i o u s v e h i c l e s and o ther equipment may be accommodated i n t h e c a r g o compartment (See F igure 28).

Two s u c t i o n pumps are r e q u i r e d f o r t h e HLFC t r a n s p o r t a i r c r a f t . A g e n e r a l a r rangement of t h e s u c t i o n system which is r e q u i r e d for t h e H L F C T r a n s p o r t is shown i n F igu re 29.

The power u n i t which i s r e q u i r e d t o d r i v e t h e s u c t i o n pump is mounted be-

hind t h e s u c t i o n punp. A d r i v e s h a f t c o u p l e s t h e power u n i t and s u c t i o n pump

t o g e t h e r . This ar rangement is used t o minimize t h e s ize and weight of t h e d u c t i n g n e c e s s a r y t o l a n i n a r i z e t h e l e a d i n g edge. An S-shaped i n l e t d u c t pro-

v i d e s a i r t o t h e power u n i t . The power u n i t and t h e s u c t i o n pump b o t h exhaus t t h r o u g h d u c t s i n t h e s i d e of t h e pod used t o house t h e s u c t i o n system. The

ar rangement o f t h e s u c t i o n pumps and d u c t s is shown i n F i g u r e s 30A and 30A.

49

HLFC PAYLOAD - RANGE

40’000Em3M 0 - 0 P.000 8.000 12,000 16,000 20.000

RANGE (NM)

Figure 26. Typical HLFC Concept Payload-Range Curve

25 PALLETS A T 5300 LBS EACH

Figure 27. HLFC Concept Cargo Pallets Arrangement

50

I I 1- 19.5 FT-

I 13.5 FT

Figure 28. HLFC Concept Arrangement of Vehicles

. .

VIEW A

Figure 29. Suction Pump/System General Arrangement

51

SUCTION PUMP POWER GENERATOR

LP SUCTION

HP SUCTION

Figure 30A. HLFC Concept Suction Systems Arrangement

GENERATOR AIR INLET SUCTION PUMP POWER GENERATOR

I HIFH PRESSURE DUCTS

SECT B - B SECT C - C SECT A - A

Figure 308. Concluded

52

A l l o f t h e 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 i n t h i s s t u d y are f u e l vo lune

c r i t i c a l because o f t h e l a r g e un re fue led r a n g e r equ i r emen t , and therefore u s e

bo th wing and center s e c t i o n f u e l . S ince f u e l volume s izes t h e wing, wing s ize and a i r c r a f t weight and d rag could be reduced i f f u e l volume could be

ob ta ined elsewhere, as for example, under t h e f u s e l a g e cab in floor.

5.5.1.1 Leading Edge

L F C s u c t i o n c a p a b i l i t y i s provided o n l y i n t h e l e a d i n g edge of t h e wing. The l e a d i n g edge i s removable and it c o n t a i n s a system of chordwise s lo t s wi th s u b s u r f a c e compartments which are used t o c o n t r o l t h e p r e s s u r e g r a d i e n t s i n s i d e t h e l e a d i n g edge ( F i g u r e 31).

The l e a d i n g edge is f a b r i c a t e d i n t w o s e c t i o n s . The upper s e c t i o n is a

f i x e d nose pane l and t h e lower s e c t i o n is removable t o p rov ide access for maintenance and ad jus tment o f t h e LFC s u c t i o n and s l o t c l e a n i n g equipment.

Two f u l l l e n g t h diaphragms p rov ide s u b s t r u c t u r e s u p p o r t . These members

I

i

Figure 31. Leading Edge Design Concept

53

p r o v i d e s u p p o r t for t h e upper and lower p a n e l s and form t h e boundar i e s of t h e

upper and lower s u r f a c e d u c t s . A l l l e a d i n g edge components are of sandwich

c o n s t r u c t i o n and f e a t u r e g raph i t e / epoxy sheets and a c o r r o s i o n r e s i s t a n t

aluminum honeycomb core. The s u c t i o n s l o t s are c u t i n t o t h e t h i n guage o u t e r

t i t a n i u m s k i n which is bonded t o t h e o u t e r pane l f a c e s h e e t . T h i s t i t a n i u m s k i n a l so p rov ides envi ronmenta l p r o t e c t i o n for t h e s t r u c t u r e .

5.5.1.2 Fuel Tanks

An a u x i l i a r y f u e l t a n k and two main f u e l t a n k s i n each wing a r e t h e bas i s

for t h e fuel sys tem (See F igure 32). In a d d i t i o n , an a u x i l i a r y f u e l t a n k i s

located i n t h e c e n t e r wing box w i t h i n t h e f u s e l a g e . The main t a n k s are used

t o s u p p l y f u e l t o t h e main eng ines . The f u e l for t h e LFC s u c t i o n pump power

u n i t s w i l l be s u p p l i e d from t h e a u x i l i a r y c e n t e r t a n k s . Fue l t r a n s f e r between t h e a u x i l i a r y and main f u e l t a n k s is accomplished a s r e q u i r e d .

I n

I

7 TANK-MAIN NO. 1

P 7 \ TANK-M'AIN NO. 2

TANK-CENTER AUX Y Figure 32. Fuel Tank Arrangement

54

Access door s i n t h e lower s u r f a c e o f t h e wing a r e provided f o r inspec-

t i o n , main tenance , and r e p a i r o f t h e wing s t r u c t u r e . All components i n c l u d i n g boos t pumps, fuel p r o b e s , and f u e l level c o n t r o l v a l v e s a r e removable from t h e

e x t e r i o r o f t h e lower wing s u r f a c e . S i n g l e p o i n t ground p r e s s u r e f u e l i n g i s accomplished from t h e main l a n d i n g gea r wheel well.

5.5.1 . 3 Empennage

The Ehpennage is a T t a i l arrangement a s shown i n F igu re 33.

Cons t ruc t ion o f t h e t a i l is s i m i l a r t o t h e wing w i t h t h e e x c e p t i o n t h a t

t h e l e a d i n g edges o f t h e h o r i z o n t a l and v e r t i c a l t a i l s do no t have t h e s l o t s o r plumbing f o r LFC s u c t i o n .

The h o r i z o n t a l t a i l i n c o r p o r a t e s e l e v a t o r s a s p a r t o f t h e t r a i l i n g edge.

?he A double hinged r u d d e r is a t t a c h e d t o t h e a f t s p a r o f t h e v e r t i c a l t a i l .

e l e v a t o r and rudder a r e p a r t o f t h e FBW ( f l y by wire) c o n t r o l sys tem.

INCE POINTS

Figure 33. Empennage General Arrangement

5.5.1.4 Landing Gear

A t w o wheel nose g e a r and a two s t r u t twelve wheel main gea r ( F i g u r e 3 4 )

comprise t h e l a n d i n g g e a r sys tem.

The main l a n d i n g g e a r is f u s e l a g e mounted. Each strut r e t r a c t s i n t o t h e

f u s e l a g e . ?he g e a r r o t a t e s abou t a l o n g i t u d i n a l axis 90' from f u l l y r e t r a c t e d t o f u l l y extended. A t o t a l o f s i x wheels is mounted on each s t r u t . The wheels

a r e a r r anged i n t h r e e p a i r s o f two on each strut . No form o f d i r e c t i o n a l s t e e r i n g or s w i v e l l i n g is employed on t h e main g e a r s .

The nose g e a r r e t r a c t s forward i n t o t h e b e l l y o f t h e a i r c r a f t . Nosewheel s t e e r i n g ope ra t ed by a handwheel c o n t r o l l e d by t h e p i l o t is p rov ided .

Figure 34. Main Landing Gear

56

5.5.1.5 High L i f t Systems

The h i g h l i f t system c o n s i s t s o f s i n g l e s l o t , h i n g e d , f u l l span f l a p s

( F i g u r e 35). The ou tboa rd p o r t i o n o f t h e f l a p s a r e provided by drooping t h e a i l e r o n s . Con t ro l o f t h e hydro-mechanical ly ope ra t ed f l a p s is accomplished by a s i n g l e p i l o t i n p u t . Systems t o p reven t a s y m n e t r i c a l f l a p o p e r a t i o n and f l a p p o s i t i o n i n d i c a t i o n a r e a l s o provided .

Secondary f l a p s o f 9.5% chord a r e b u i l t i n t o t h e entire span o f t h e main f l a p s . These secondary f l a p s a r e used a long wi th t h e pr imary f l a p s i n t h e

h i g h l i f t mode o r s e p a r a t e l y a s p a r t of t h e a c t i v e c o n t r o l system. The onboard fly-by-wire computer system a c t u a t e s t h e secondary f l a p s i n t h e a c t i v e c o n t r o l mode by means o f e l e c t r o - h y d r a u l i c s e r v o u n i t s .

SPOILER PANEL

REAR SP

L \

Figure 35. Wing Trailing Edge Design

57

5.5.2 HLFC Wing Aerodynamics Design

S i n c e 1974, a n m b e r of automated d e s i g n and a n a l y s i s methods have been

developed by Lockheed for wings having e x t e n s i v e r e g i o n s of l amina r flow.

These methods are a p p l i c a b l e t o Natura l Laminar Flow (NLF), Laminar Flow

Con t ro l (LFC), and Hybrid Laminar Flow Cont ro l (HLFC) wing des ign . The d e s i g n

methodologies are d e p i c t e d i n b l o c k diagram form i n F igu re 36. Aerodynamics e f fo r t i n t h i s c o n t r a c t was c o n c e n t r a t e d on u s e of these a v a i l a b l e H L F C d e s i g n

methods t o d e v e l o p an i n i t i a l b a s e l i n e HLFC wing c o n f i g u r a t i o n which is close t o t h e f i n a l p a r a n e t r i c wing d e f i n e d i n t h e s i z i n g and o p t i m i z a t i o n studies.

T h i s o b j e c t i v e was a c h i e v e d , however some a d d i t i o n a l r e f inemen t of wing

geometry and a p p l i e d s u c t i o n m u l d be r e q u i r e d t o deve lop a t r u e t lp roduct iont t

c o n f i g u r a t i o n w i t h near-minimal s u c t i o n . The r e s u l t i n g f i n a l c o n f i g u r a t i o n

might t h e n r e q u i r e r e s i z i n g t o f i n a l i z e t h e b a s e l i n e . A summary of t h e aerodynamic d e s i g n and d e r i v a t i o n of n e c e s s a r y i n p u t s t o t h e s u c t i o n system

d e s i g n i s provided i n t h e s e c t i o n s t h a t follow.

Figure 36. Generalized Aircraft Sizing Program, LFC Subroutine

58

5.5.2.1 Parametric S i z i n g of Aircraft I n c l u d i n g Suc t ion Requirements and

Suc t ion System

As p r e v i o u s l y o u t l i n e d i n Reference 2 and repeated herein, t h e HLFC system i s c h a r a c t e r i z e d i n t h e parametric s i z i n g code w i t h t h e fo l lowing inpu t s :

( 1 Type of HLFC s u c t i o n power

Opt ion 1: Independently-powered s u c t i o n u n i t s (used i n t h i s con-

t r a c t )

Option 2: S u c t i o n u n i t s in tegra ted wi th pr imary p ropu l s ion

Opt ion 3: S u c t i o n u n i t s i n t e g r a t e d w i t h other a i r c r a f t systems

Number and l o c a t i o n of HLFC s u c t i o n u n i t s

P a r a m e t r i c geometric d e s c r i p t i o n o f WC g l o v e and s u c t i o n d u c t s

Parametric s u c t i o n e x t e r n a l p r e s s u r e s and s u c t i o n d i s t r i b u t i o n s

Ex ten t o f l a n i n a r flow provided by e x t e n t of HLFC s u c t i o n

Pa rame t r i c i nc remen ta l costs of HLFC systems ( o m i t t e d i n t h i s s t u d y )

Determina t ion of B a s e l i n e Wing Detailed E x t e r n a l Contours

Bas - l ine c o n f i g u r a t i o n paranetric code o u t p u t s were used as a s t a r t i n g place for d e t a i l e d d e s i g n of t h e b a s e l i n e wing e x t e r n a l c o n t o u r s (and s u r f a c e

p r e s s u r e s ) . The d e t a i l e d d e s i g n s t a r t i n g p o i n t b a s e l i n e wing geometry d e r i v e d from parametric s t u d i e s is i l l u s t r a t e d i n F igu re 37. As p r e v i o u s l y exp la ined

i n Reference 2, t h e subsequen t aerodynanic d e s i g n followed t h e procedure o u t l i n e d i n F igu re 38.

59

EAR EXTENT OF AMINAR FLOW

RONT SPAR

BELLY POD LINE

I , -- 1

Figure 37. Areas of HLFC and Wing Laminarization

60

- CHOICE OF

SUCTION EXTENT DETERMINATION AND TYPE OF AIRFOIL PRESSURE GEOMETRY

DISTRIBUTION

t

I INCORPORATION

LAMINAR BOUNDARY TURBULENT LAYER AND SUCTION 0 BOUNDARY

REQUIREMENTS LAYER

INTO 3-DIMENSIONAL

WING

t CHECK OF RESULTS A I RFOl L AI RFOl L

GEOMETRY WITH VISCOUS GEOMET RY FOR USE ON AIRFOIL CORRECTION LFC WING PRESSURE

PROGRAM

-

FINAL WING AND SUCTION

REQUIREMENTS H MODIFICATION OF WING SECTIONS

TO PRODUCE STRAIGHT ISOBARS

1 1 L I

Figure 38. Aerodynamic Design Procedure for LFC and TF Wings

Using t h e above-mentioned methods, t h e b a s e l i n e wing geometry was d e r i v e d

w i t h a t y p i c a l a i r f o i l shape as i l l u s t r a t e d i n F i g u r e 39. The shape cor- r e sponds t o t h a t used a t t h e wing c o n t r o l s t a t i o n a t 0.350 nondimensional

semispan p o s i t i o n . Wing s e c t i o n c h o r d l i n e i n c i d e n c e s and t h i c k n e s s r a t i o s a t t h e f u s e l a g e s i d e , b reak s t a t i o n , and t i p a re i n d i c a t e d below:

P o s i t i o n I n c i d e n c e ( d e g r e e s ) T h i c k n e s s ( t / c )

Side of F u s e l a g e 0.75 Wing Break S t a t i o n 1.25 Wing T ip -0.25

5.5.2.3 Baseline Wing Surface Pressure R e s u l t s

0.130

0.118

0.118

F i g u r e 40 i l l u s t r a t e s a t y p i c a l s u r f a c e p r e s s u r e r e s u l t on t h e d e r i v e d

b a s e l i n e wing geometry. T h i s p r e s s u r e d i s t r i b u t i o n shape is fo r t h e 0.488

nondimensional wing s t a t i o n and is t h e r e s u l t of v i s c o u s t h r e e - d i m e n s i o n a l

61

0 UPPER SURFACE 0 LOWER SURFACE

D Y I D X (DEC)

20.

0 .

-40. I

F i g u r e 39. Typica l Airfoil Sec t ion on Base l ine HLFC Wing

f low computa t ions a t t h e b a s i c d e s i g n p o i n t a t s t a r t - c r u i s e c o n d i t i o n s .

Su r face p r e s s u r e r e s u l t s from t h i s s t a t i o n and o t h e r s ove r t h e semispan of t h e

wing were used t o determine wing boundary l a y e r and boundary l a y e r s t a b i l i t y

r e s u l t s . Discuss ion o f these results is provided i n t h e s e c t i o n s t h a t fo l low.

5.5.2.4 B a s e l i n e S u c t i o n and Boundary Layer S t a b i l i t y Results

Using t h e wing s u r f a c e pressure r e s u l t s descr ibed above, a s u c t i o n

d i s t r i b u t i o n was developed u s i n g t h e parametric d e s c r i p t i o n a s t h e s t a r t i n g

p o i n t . Changes t o a i r f o i l p r e s s u r e s and s u c t i o n were made t o produce better results from t h e boundary l a y e r and boundary l a y e r s t a b i l i t y p r e d i c t i o n s . The

s u c t i o n d i s t r i b u t i o n from these d e s i g n i t e r a t i o n s , shown i n F igure 41, i s s i m i l a r i n shape and t o t a l s u c t i o n mass f low t o t h a t used i n p a r a m e t r i c

studies, Add i t iona l d e s i g n work is warranted t o a r r i v e a t an lloptimumtl wing

d e s i g n . Any a d d i t i o n a l work i s , h o w e v e r , o u t s i d e t h e t ime and c o s t

c o n s t r a i n t s of t h e c u r r e n t c o n t r a c t .

62

n :oo

CL=O. 5 5 1 M=0.77 RN=26x106

-1.2

1 .a0

1.30 LOCAL MACH

1.20

1.10

1 .oo

0.90

0.80

1.2 I I

0.4 0.6 0.8 1 0.0 0.2

X I C

REPRESENTATIVE WING Cp DISTRIBUTION FROM WING STATION I)- 0.488

Figure 40. Representative Wing C Distribution from Wing Station q= 0.488 P

The d a t a i n F igu re 42 shows t h e amount o f flow r e q u i r e d t o be s u c t i o n e d o f f t h e boundary l a y e r t o m a i n t a i n l a m i n a r i z e d flow on each wing. This flow

is t y p i c a l f o r t h e b a s e l i n e a i r f o i l o p e r a t i n g a t Mach 0.77 and 37,000 feet

a l t i t u d e . The s u c t i o n pump used on t h e a i r c r a f t has t h e c a p a c i t y t o punp 150

percent o f t h e f low o f bo th wings , so there should be no d i f f i c u l t y i n

o p e r a t i n g t h e pump a t o f f -poin t c o n d i t i o n s and s t i l l m a i n t a i n l a m i n a r i zed

flow.

Results o f boundary l a y e r c r o s s f l o w s t a b i l i t y c a l c u l a t i o n s u s i n g t h e

SALLY code are shown i n F igu re 43. Note t h a t results i n d i c a t e t h a t t r a n s i t i o n

is l i k e l y t o occur near midchord p o s i t i o n f o r bo th t h e upper and lower

63

-. 0005

TOTAL SUCTION FLOW FOR ONE WING =

-.20 -.15 - . lo - .05 0 .05 .10 .IS .20

LOWER SURFACE X I C UPPER SURFACE

105.51 L B l M l N

Figure 41. Mass Flow Suction Level

T O T A L SUCTION FLOW REQUIRED FOR ONE WING

SECTION 1.

SECTION 2.

SECTION 3.

SECTION 0 .

SECTION 1.

SECTION 2.

SECTION 3.

SECTION 9.

UPPER SURFACE, FROM WING ROOT TO Y l b = .35

UPPER SURFACE, FROM Y l b = .35 TO Y l b * .56

UPPER SURFACE, FROM Y l b = .56 TO Y l b = .78

UPPER SURFACE, FROM Y l b = .78 TO WING T I P

LOWER SURFACE, FROM WING ROOT TO Y l b = .35

LOWER SURFACE, FROM Y l b = .35 TO Y l b = .56

LOWER SURFACE, FROM Y l b = .56 TO Y l b = .78

LOWER SURFACE, FROM Y l b = .78 T O WING T I P

SUCTION FLOW

17.66 LB lMIN

13.43 LB~MIN

15.36 L B l M l N

12.04 LBMN

3.08 LB lM lN

1.02 LB lM lN

1.36 L B l M l N

1.56 L B l M l N

Figure 42. Total Suction Flow Across One Wing of HLFC Baseline

64

I I I

I I

I

1 I

I I I

I 0 - 1

I

- LOWER SURFACE I 1 1

s u r f a c e s w i t h a c t i v e s u c t i o n no fa r ther a f t t h a n 15 p e r c e n t chord. F u r t h e r

d e s i g n r e f inemen t should improve t h e i n i t i a l b a s e l i n e r e s u l t s i l l u s t r a t e d .

Tol lmien-Schl ich t ing i n s t a b i l i t y c a l c u l a t i o n s do n o t p i c k up i n s t a b i l i t y on

ei ther t h e upper or t h e lower s u r f a c e back t o 50 p e r c e n t chord.

T h e b a s i c t r a n s i t i o n N F a c t o r l e v e l i s c o m p a t i b l e w i t h l e v e l s demonst ra ted t o be adequa te from ear l ier Lockheed LFC work for NASA as

o u t l i n e d i n Reference 12. Based on t h i s work and o t h e r Lockheed e x p e r i e n c e , a n a l y s i s work was c o n c e n t r a t e d on crossflow s t a b i l i t y v e r i f i c a t i o n . For t h e

c o n f i g u r a t i o n s o f t h i s s t u d y , t h e Tollmien-Schl i c h t e n s t a b i l i t y mode was much less c r i t i c a l t h a n crossflow and was n o t examined a t g r e a t l eng th . The

leading-edge a t t a c h e n t l i n e momentum t h i c k n e s s Reynolds n m b e r , Rea . l . , was a l s o not s t u d i e d a t g r e a t l e n g t h s i n c e t h e c o n f i g u r a t i o n s o f t h e s t u d y have

leading-edge r a d i i , p r e s s u r e g r a d i e n t s , and sweep s imilar t o p r e v i o u s l y -

s t u d i e d wings and should have s imilar , s a t i s f a c t o r y leading-edge con tamina t ion

charac te r i s t ics .

a.o.00

Mnor=0.770 RN=26x106

DISTURBANCE FREQUENCY = 0.5 Hz l6 1 1 2

z

---- UPPER SURFACE

1 2

z

I !/ ---- UPPER SURFACE

X I C

CROSSFLOW DISTURBANCE N FACTOR, WING STATION 7Ir.488

Figure 43. Crossflow Disturbance N Factor, Wing Station q= 0.488

65

The Gaster bump, lead ing-edge n o t c h , and i n i t i a t i o n of s u c t i o n n e a r e r t h e

l e a d i n g edge p r o v i d e s e v e r a l p r o v e n means o f a v o i d i n g u n s a t i s f a c t o r y

charac te r i s t ics should more d e t a i l e d d e s i g n i n d i c a t e t h a t a problem e x i s t s , l e r e s u l t i n g e f fec ts o f any such changes on a i r c r a f t weights wbuld be v e r y

s m a l l shou ld t h e y be n e c e s s a r y .

5.5.2.5 Possible Future Wing Refinements

The i n i t i a l b a s e l i n e wing can p robab ly be f u r t h e r r e f i n e d t o r educe t h e

l i k e l i h o o d of t r a n s i t i o n forward of 50 p e r c e n t chord wi th some changes t o ( a )

geomet r i e s and r e s u l t a n t p r e s s u r e , ( b ) s u c t i o n l e v e l s and d i s t r i b u t i o n s , and ( c ) s u c t i o n s u r f a c e c o n f i g u r a t i o n . Data from t h e wing e x t e r n a l aerodynamics

d e s i g n were used i n b a s e l i n e s u c t i o n s u r f a c e d e t e r m i n a t i o n , i n t e r n a l HLFC geometry d e s i g n , and v e r i f i c a t i o n of s u c t i o n powerplan t s ize r e q u i r e d . D e t a i l s .

of t h i s p a r t o f t h e HLFC d e s i g n are con ta ined i n S e c t i o n 4.4.5 of t h i s report.

66

6.0 CONFIGURATION SENSITIVITY STUDIES

S e n s i t i v i t y s t u d i e s r e l a t i v e t o performance and c o n f i g u r a t i o n parameters

a r e reviewed i n t h i s s e c t i o n f o r both hybr id LFC and t u r b u l e n t f low a i r c r a f t .

6.1 HLFC AIRCRAFT SENSITIVITY STUDIES

6.1.1 Increase of C r u i s e Sbeed or A l t i t u d e

Pre l iminary s e n s i t i v i t y s t u d i e s were performed on t h e b a s e l i n e HLFC

a i r c r a f t (Option 1, F igure 23) t o i n c l u d e t h e e f f e c t s on performance o f ( 1 )

i n c r e a s i n g t h e c r u i s e Mach number from 0.77 t o 0.80 (Option 2, F igure 23) and

( 2 ) i n c r e a s i n g t h e i n i t i a l c r u i s e a l t i t u d e from 32,361 f e e t t o 36,000 f e e t

(Option 3, F igure 23 ) .

As compared t o t h e HLFC a i r c r a f t a t M = 0.77 c r u i s e speed , t h e HLFC

a i r c r a f t a t M = 0.80 shows an i n c r e a s e i n f u e l burned o f 10.9 p e r c e n t , a

r e d u c t i o n i n l i f t - t o - d r a g r a t i o o f 4.1 p e r c e n t , an i n c r e a s e i n engine t h r u s t

o f 12 p e r c e n t , an i n c r e a s e i n g r o s s weight of 7.8 p e r c e n t , and a r e d u c t i o n i n

ML/D o f 0.3 of one p e r c e n t .

The e f f e c t s o f an i n c r e a s e i n i n i t i a l cruise a l t i t u d e from 31,361 f e e t t o

36,000 f e e t a t M = 0.77 cruise speed shows an i n c r e a s e i n f u e l burned o f 1.2 percent, an increase i n l i f t - to-drag r a t i o o f 3.7 p e r c e n t , an i n c r e a s e i n

engine t h r u s t o f 18.6 p e r c e n t , and an i n c r e a s e i n g r o s s weight o f 3.8 percent.

Nei ther o f t h e t w o p t i o n s appea r s t o be b e n e f i c i a l t o t h e performance o f

t h e HLFC a i r c r a f t . Of t h e t w o , t h e i n c r e a s e i n i n i t i a l cruise a l t i t u d e t o 36,000 f e e t has t h e smaller d e g r a d a t i o n i n t h e o v e r a l l a i r c r a f t performance.

6.1.2 E l i m i n a t i o n of HLFC on Empennage

The e f f e c t s on performance o f d e l e t i n g HLFC from t h e empennage o f t h e

HLFC a i r c r a f t is provided i n t h e d a t a f o r Option 4 i n F igu re 44. As compared

t o t h e b a s e l i n e t u r b u l e n t f low a i r c r a f t , t h e HLFC a i r c r a f t w i t h no HLFC on t h e

67

empennage shows a r e d u c t i o n i n fuel burned o f 13.7 p e r c e n t , an i n c r e a s e i n l i f t - t o - d r a g r a t i o o f 18.2 p e r c e n t , a d e c r e a s e i n engine t h r u s t o f 11.9 percent, and a d e c r e a s e i n g r o s s weight o f 4.2 p e r c e n t . Except f o r t h e i n s i g n i -

f i c a n t r e d u c t i o n i n l i f t - t o - d r a g r a t i o , t h e s e improvements i n performance a s compared t o t h e t u r b u l e n t f low a i r c r a f t are s l i g h t l y b e t t e r t han t h o s e f o r the

b a s e l i n e HLFC a i r c r a f t . It appea r s from t h e s e d a t a t h a t e l i m i n a t i o n o f HLFC

from t h e empennage has a f a v o r a b l e o v e r a l l effect on HLFC a i r c r a f t performance and r educes t h e complexi ty o f t h e systems a s well.

6.1.3 Elimination of HLFC on Lower Wing Surface

The e f f e c t s on HLFC a i r c r a f t performance o f d e l e t i n g H L F C on t h e lower

s u r f a c e o f t h e wing is provided i n t h e d a t a f o r @ t i o n 5 i n F igu re 44. As

compared t o t h e b a s e l i n e t u r b u l e n t flow a i r c r a f t , t h e HLFC a i r c r a f t wi th no

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I I

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Figure 44A. Sensi t iv i ty Data

ORIGINAI: PAGE IS OF POOR QUALITY

68

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

QIUGINAU PAGE Is OF POOR QUALITY ......................................................................................

I.4I k.IC.1 .................................................................................................................................... I I ............................................................................................................................ I I II LIACILI.AILFUL ICOY'l) I I I I I I I I I I1 1 1 I I

I I I I I I I I I I f I I 1 1 II ?own ATITI*C I U.*B I I 0.lB I I 0.?@ I I 0 . 0 I I -11.1b I I - 1 U . W I I -b.M I I

la*** I I -0.1b I I 1.13 I I -0.*1 I 1 I I TDlAL/bASIC I 1.11 I I 1.17 II 1.11 I I 1.11 1 1 -0.02 I t 0.S1 1 1 -0.1s 11

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I I t I t I t t l I 4lb.00 I I DASC I I -10.01 II -11.18 I I

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1 auu.iWr I I o c i i u ~ i I I o i i i o n 4 I I OWION s I I o - 1 II D - a I I o - 1 I I ............................................................................................................................ I I ................................................................................................................................. I2.ICD.1*@1


HLFC on t h e lower wing s u r f a c e shows a d e c r e a s e i n f u e l burned o f 7.9 percent, an i n c r e a s e i n l i f t - t o - d r a g r a t i o o f 12.5 p e r c e n t , a d e c r e a s e i n engine t h r u s t of 6.4 percent, and a decrease i n gross weight of 0.6 percent. These improve-

ments i n performance are c o n s i d e r a b l y less t h a n t h o s e shown f o r t h e b a s e l i n e

HLFC a i r c r a f t a s compared t o t h e t u r b u l e n t f l o w a i r c r a f t , showing 41 p e r c e n t more f u e l burned , 32 p e r c e n t lower l i f t - t o - d r a g r a t i o , and 40 p e r c e n t h i g h e r t h r u s t required. It a p p e a r s from these d a t a t h a t e l i m i n a t i o n of t h e IUFC from t h e lower wing s u r f a c e has an o v e r a l l u n f a v o r a b l e effect on t h e performance o f t h e HLFC a i r c r a f t .

6.1.4 Reduction of Aspect Ratio to 10

S i z i n g runs were made i n o r d e r t o determine t h e effects o f a r e d u c t i o n i n wing a s p e c t r a t i o from t h e b a s e l i n e v a l u e s o f over 1 3 t o a more modera te

a s p e c t r a t i o o f 10. The s i z i n g d a t a p re sen ted i n F igu re 45 i n c l u d e t h e HLFC

69 .

b a s e l i n e a i r c r a f t f o r a s p e c t r a t i o 13.87 and a s p e c t r a t i o 10 d a t a f o r t h e HLFC a i r c r a f t and a t u r b u l e n t f l ow a i r c r a f t .

Results i n d i c a t e t h a t t h e HLFC concep t w i t h a s p e c t r a t i o 10 compared t o t h e HLFC b a s e l i n e has a 2.9 pe rcen t h ighe r g r o s s weight , 11.5 p e r c e n t more

f u e l burned , 12.2 p e r c e n t decrease i n L/D, and 11.7 p e r c e n t more e n g i n e t h r u s t . I n a d d i t i o n , t h e a s p e c t r a t i o 10 HLFC a i r c r a f t has a r e d u c t i o n i n

wing span from 258.85 f e e t t o 21 9.35 f e e t , or 15.3 p e r c e n t , a s compared t o t h e

baseline HLFC a i r c r a f t . ?he wing span comparison is shown i n F igu re 46. Com- p a r i n g t h e t u r b u l e n t f low and HLFC a s p e c t r a t i o 10 r u n s i n F i g u r e 4 5 shows t h a t t h e HLFC c o n f i g u r a t i o n has a 4 .1 p e r c e n t r e d u c t i o n i n g r o s s we igh t , 12.1

p e r c e n t less fue l burned , 15.1 p e r c e n t h ighe r L/D, and a 9 p e r c e n t r e d u c t i o n

i n e n g i n e t h r u s t . These d i f f e r e n c e s between t h e a s p e c t r a t i o 10 a i r c r a f t a r e

s l i g h t l y s m a l l e r , y e t v e r y similar t o t h e d i f f e r e n c e s found between t h e h i g h e r a s p e c t r a t i o t u r b u l e n t and HLFC b a s e l i n e a i r c r a f t r e p o r t e d i n S e c t i o n 5.4.

I.

a II I I II I 1 I1 I 1 II II I I 1 1 ..

,*..*.

II II I I II I I II II I I I I

1 W .

II II II II I I II

I I I I I1 I I I I I I I I I I I I I I I I I I

I 1 I 1 II I I I I II 'I I1 II ' I # I II II I

n

).

.. II II II II

II I 1 II II II II II II I 1 II

Figure 45A. Sizing Data for Aspect Ratio 10 Aircraft

70 QRIGINk& PAGE IS OE POOR QUALITY

.............................. NLIC SITED ICICMT AOO1TlOR0 ............................. 1 1 .................................................................................................................................... I I S4a.00 I 91.00 I 00.00 I 1a4.00 I I I SJ4.00 I Sb9.09 I 4bLL.00 I 1,710.bt

li I I 11. I I s41.00 I I 0 .00 I I II 494.00 I I 0.00 I I II 410.00 I 1 0.00 I I I I 1.101.00 II 0.00 I 1

I t 1 1 I I J I II II I 1 I I 1 U C A I K D C 1 1 1 T l D I I I II I I II II II I I II 018TtLD - L I I 113.00 II 19a.00 I I 0.00 I I II 8.10 II 100 1 1 ERR 1 1 I 1 ?0A?PCO ILUKO LD I S0b.M II 411.00 I I 0.00 I I 1 1 1.10 1 1 ne0 1 1 801 1 1 )I R l U I M C U I 1 0 I 0 I b,lSl.OO I 1 40b90.00 I I 0.00 I I I I s.os 11 C R I I I 111 1 1

~~ai1.00 II s,s00.00 II 0.00 I I II %,*a II 111: II 118 II ! ............................................................. U. .............. o..... .. i


HLFC BASELINE: SPAN = 258.85 F T ASPECT RATIO = 10: SPAN = 219.35 F T

BASELINE

HLFC AR = 10

- ---.

Figure 46. Comparison of HLFC Baseline and HLFC Aspect Ratio 10 Aircraft

71

6.1.5 High Wing HLFC C o n f i g u r a t i o n

S i z i n g runs fo r hybr id LFC t r a n s p o r t s for t h e more conven t iona l a r range-

ment of h igh wings w i t h t h e eng ines mounted on t h e wings were made i n order t o p r o v i d e t h e d a t a fo r an assessment o f t h e wing-mounted v e r s u s t h e f u s e l a g e -

mounted eng ine a r rangement of t h e hybr id L F C b a s e l i n e c o n f i g u r a t i o n . I n order

t o account for t h e i n t e r f e r e n c e o f t h e pylon-mounted eng ines under t h e h i g h wing c o n f i g u r a t i o n , it was dec ided t h a t t h e r e w u l d be a l o s s of l amina r flow

on t h e lower wing s u r f a c e i n a streamwise d i r e c t i o n w i t h an area of loss c o n s i s t i n g of t h e maximun w i d t h o f t h e eng ine a t t h e wing l e a d i n g edge p l u s a 7' i n c r e a s e i n area ove r t h e wing s u r f a c e t o t h e t r a i l i n g edge. A p lan v i ew

of t h i s loss i n l amina r flow is shown i n F i g u r e 47. No upper s u r f a c e loss of

l a n i n a r flow is assumed f o r these s i z i n g r u n s .

LOSS OF LAMINAR FLOW I - d

Figure 47. Plan View of Laminar Flow Area Loss for Wing Mounted Engine Configuration; Lower Wing Surface

72

S i z i n g d a t a fo r t h e h igh wing c o n f i g u r a t i o n s a r e g iven i n F i g u r e 48 f o r a

c r u i s e Mach nunber o f 0.77. It should be noted t h a t t h e h i g h wing t u r b u l e n t flow a i r c r a f t u sed fo r r e f e r e n c e h e r e has 20' o f wing sweep a s c o n t r a s t e d t o 30' wing sweep f o r t h e p r e v i o u s l y r e p o r t e d t u r b u l e n t f l ow b a s e l i n e a i r c r a f t .

As noted i n F i g u r e 48, t h e s i z i n g r u n s c o n s i s t o f a r e f e r e n c e t u r b u l e n t f low a i r c r a f t , an HLFC a i r c r a f t w i t h l a n i n a r f low on bo th upper and lower s u r f a c e s ,

and an HLFC a i r c r a f t w i t h l aminar flow on t h e upper s u r f a c e on ly .

The results o f F i g u r e 48 i n d i c a t e t h a t t h e HLFC h i g h wing a i r c r a f t w i t h

l amina r flow on bo th upper and lower s u r f a c e s as compared t o t h e t u r b u l e n t

f low a i r c r a f t has 10.4 p e r c e n t lower g r o s s we igh t , 19.6 p e r c e n t lower f u e l

bu rned , 18.1 p e r c e n t i n c r e a s e i n L/D, 15.3 p e r c e n t d e c r e a s e i n e n g i n e t h r u s t , and 4.8 p e r c e n t r e d u c t i o n i n wing span . These r e s u l t s f o r t h e HLFC c o n f i g u r a -

t i o n a r e t h e best o v e r a l l performance d a t a o b t a i n e d f o r an HLFC a i r c r a f t i n

Figure 48A. Sizing Data for High Wing Turbulent Flow and HLFC Aircraft; Sweepback 20'

73

WWC 110?CW U C I W I bWI11O.b .. ........................... ................................................................ 1 II I t 1 0.00 II ss0.00 II I 0.00 II sa.oo II I 0.00 II b4.00 II I 0.00 I 1 W6.08 I 1

I 1 I II I II I I

0.00 I I sa0.00 II I 0.00 I I SSl.00 I 1 1

I 0.00 II b44.00 II I

.II II II II II

II 100 II I D 0 II II 111 I 1 101 I 1 I 1 1 I0 II 110 II

II m e II m II

II I 1

I 1 I 1 II II I 1 II II I 1 II II II II

F i g u r e 48B. Concluded

t h i s s t u d y t a s k . As compared t o t h e HLFC b a s e l i n e a i r c r a f t wi th f u s e l a g e - mounted engines, t h i s HLFC h i g h wing a i r c r a f t has 3.5 percent lower g r o s s weight, 5.0 percent lower o p e r a t i n g weight empty, 4.1 percent lower f u e l

burned , a s l i g h t i n c r e a s e i n L/D, 2.7 p e r c e n t decrease i n e n g i n e t h r u s t , and

1 .1 p e r c e n t r e d u c t i o n i n wing span. In t h e a u t h o r l s o p i n i o n , however, these

h i g h wing HLFC r e s u l t s a r e o p t i m i s t i c because o f t h e assumption o f no loss o f

laminar flow on t h e wing upper s u r f a c e f o r t h i s wing-mounted eng ine con-

f i g u r a t i o n . A g e n e r a l arrangement drawing o f t h e h i g h wing HLFC a i r c ra f t is presented i n F i g u r e 49.

The r e s u l t s f o r t h e HLFC h i g h wing a i r c r a f t w i th l a n i n a r f low on t h e

upper s u r f a c e o n l y , a s compared t o t h e t u r b u l e n t flow a i r c r a f t , show 7.5

percent lower g r o s s w e i g h t , 14.6 percent lower fue l burned, 12.5 percent

h i g h e r L/D, 12.8 percent d e c r e a s e i n engine t h r u s t , and 3.1 percent r e d u c t i o n

i n wing span f o r t h e HLFC c o n f i g u r a t i o n . As compared t o t h e HLFC b a s e l i n e

74

I a i r c r a f t wi th fuse l age -moun ted e n g i n e s , t h i s HLFC h igh wing a i r c r a f t i s s l i g h t l y i n f e r i o r i n performance w i t h a n e g l i g i b l e d i f f e r e n c e i n g r o s s we igh t ,

1.8 p e r c e n t i n c r e a s e i n f u e l burned , 4.0 p e r c e n t r e d u c t i o n i n L/D, and a

s l i g h t i n c r e a s e i n e n g i n e t h r u s t . As c i t e d p r e v i o u s l y , t h e r e s u l t s for t h e

I

I

I high wing HLFC c o n f i g u r a t i o n s are cons ide red t o be o p t i m i s t i c because o f t h e ~ assumption of no loss i n l a n i n a r flow on t h e upper wing s u r f a c e .

- I

P A Y L O A D RANGE M A C H NO. A L T I TU D E TOCW F U E L L I D SPAN AR L.E. SWEEP

132,500 L B 6,500 NM' 0.77 31,950 FT 570.734 L B 241,833 LB 30.99 255.9 FT 13.86 20 DEG

*SEE F I G U R E 3

Figure 49. General Arrangement Drawing of High Wins HLFC Aircraft

6.2 TURBULENT FLOW AIRCRAFT SENSITIVITY STUDIES

6.2.1 I n c r e a s e of A l t i t u d e t o 36,000 F e e t

S i z i n g c a l c u l a t i o n s f o r a d d i t i o n a l s e n s i t i v i t y s t u d i e s o f t h e t u r b u l e n t

f low a i r c r a f t were made, and t h e d a t a a r e presented i n Figure 50. The @ t i o n

1 s i z i n g r u n was performed t o de te rmine t h e e f f e c t o f i n c r e a s e i n i n i t i a l

c r u i s e a l t i t u d e from 32,119 f e e t t o 36,000 f e e t . As compared t o t h e lower cruise a l t i t u d e performance, t h e i n c r e a s e i n i n i t i a l c r u i s e a l t i t u d e results i n 2.1 p e r c e n t i n c r e a s e i n f u e l b u r n e d , 6.1 pe rcen t i n c r e a s e i n l i f t - t o - d r a g r a t i o , 11.7 pe rcen t i n c r e a s e i n engine t h r u s t , and 1.6 pe rcen t i n c r e a s e i n g r o s s weight . These r e s u l t s a r e s i m i l a r t o t h o s e f o r t h e HLFC a i r c r a f t when

t h e i n i t i a l c r u i s e a l t i t u d e is inc reased t o 36,000 f e e t .

‘Figure 50A. Turbulent Flow Aircraft Sizing Data

76


6.2.2 I n c r e a s e of Payload to 212,000 Pounds

I n t h e s t u d y o f these advanced t echno logy m i l i t a r y t r a n s p o r t s , it was

desired t o determine t h e performance d a t a f o r a t u r b u l e n t f low a i r c r a f t s i z e d f o r t h e payload d e r i v e d i n t h e TAFAD s t u d y r e s u l t i n g from an in-depth mis s ion

a n a l y s i s o f t h e Congres s iona l ly Mandated M o b i l i t y S tudy . This s t u d y e s t a b - l i s h e d an optimum payload o f 212,000 pounds f o r r a p i d deployment.

F i g u r e 50 i n c l u d e s d a t a f o r t h e t u r b u l e n t f low a i r l i f t e r c a p a b l e o f

t r a n s p o r t i n g 212,000 pound pay loads over t h e s a n e g l o b a l r ange mis s ion a s t h e

baseline. This t u r b u l e n t f l ow a i r c r a f t , a s expected, is v e r y l a r g e a s compared t o t h e t u r b u l e n t b a s e l i n e . The 60 percent increase i n payload

r e su l t s i n a 48.7 percent i n c r e a s e i n g r o s s weight , 45.4 p e r c e n t more f u e l

burned , and 44 percent h i g h e r t h r u s t per engine a s compared t o t h e t u r b u l e n t b a s e l i n e a i r c r a f t . A g e n e r a l arrangement drawing of t h e a i r c r a f t is presented i n F i g u r e 51.

This l a r g e a i r c r a f t h a s a g r o s s weight of 916,333 pounds, m i s s i o n f u e l r equ i r emen t o f 423,769 pounds, wing a s p e c t r a t i o o f 13.4, t h r u s t per eng ine o f

43,313 pounds, and a l i f t - t o - d r a g r a t i o o f 26.4. 'he wing span of 293.78 feet

as compared t o 222.8 fee t f o r t h e C-5 a i r c ra f t a s shown i n F i g u r e 52.

77

PAY LOAD RANGE MACH NO. ALTITUDE TOGW FUEL L I D MAC SPAN AR L.E. SWEEP

212,000 LB 6,500 NM' 0.77 31,198 F T 916,333 LB 923,769 LB 26.38 26.527 193.78 FT 15.79 30 DEG

i k !. 293.78' I L 212.,;1 -, I

*SEE FIGURE 3

Figure 51. General Arrangement of 212,000 LB Payload Turbulent Flow Aircraft

WING SPANS I

TRANSPORT

Figure 52. Comparison of Turbulent Flow 212,000 LB Payload Aircraft With C-5

78

6.2.3 Reduction of Aspect Ratio to 10

A comparison o f performance c h a r a c t e r i s t i c s o f t h e t u r b u l e n t b a s e l i n e

a i r c r a f t and t h e a s p e c t r a t i o 10 t u r b u l e n t a i r c r a f t a r e o b t a i n e d from t h e d a t a o f F i g u r e s 50 and 45, Option 2. The r e s u l t s show t h a t t h e t u r b u l e n t f l o w a i r c r a f t with a s p e c t r a t i o 10, as compared t o t h e t u r b u l e n t b a s e l i n e a i r c r a f t , h a s a 1.5 percent i n c r e a s e i n g r o s s weight , 9.8 percent more f u e l burned , a

9.7 pe rcen t d e c r e a s e i n L/D, a 9.7 pe rcen t i n c r e a s e i n t h r u s t required, and a 14.9 percent r e d u c t i o n i n wing span . The difference i n wing span o f 255.9

feet f o r t h e b a s e l i n e t u r b u l e n t a i r c r a f t and 217.86 feet f o r t h e a s p e c t r a t i o 10 t u r b u l e n t a i r c r a f t is shown i n F igu re 53.

TURB. BASELINE: SPAN = 255.91 FT

ASPECT RATIO = 10: SPAN = 217.86 FT

-BASELINE

-----TURB. A b 1 0

Figure 53. Comparison of Turbulent Flow Baseline and Aspect Ratio 10 Aircraft

79

7.0 A S S E S S l E N T OF HLFC B E N E F I T S AND SELECTED CONFIGURATIONS

I n t h i s p r e l i m i n a r y d e s i g n system s t u d y a c o n s i d e r a b l e m o u n t o f a i r c r a f t

s i z i n g d a t a h a s been gene ra t ed t o e s t a b l i s h b a s e l i n e t u r b u l e n t f l ow and h y b r i d

LFC a i r c r a f t c o n f i g u r a t i o n s which perform t h e g l o b a l range mis s ion r e q u i r e - ments. lhese results have been d i s c u s s e d i n Sec t ion 5. In a d d i t i o n , c e r t a i n

s e n s i t i v i t y studies have been made i n c l u d i n g changes i n performance pa rame te r s and a l s o changes t o t h e a i r c r a f t c o n f i g u r a t i o n concep t a s described i n S e c t i o n

6. It i s t h e purpose o f t h i s section o f t h e r e p o r t t o provide an o v e r a l l assessment o f t h e results o f t h e s t u d y i n c l u d i n g comments on f i n a l selected HLFC c o n f i g u r a t i o n s .

A summary o f t h e s i g n i f i c a n t s t u d y results a r e provided i n F igu re 54 showing changes i n HLFC performance parameters r e l a t i v e t o t h e baseline

t u r b u l e n t f l ow a i r c r a f t . Ihese r e s u l t s have been d i s c u s s e d i n more d e t a i l

p r e v i o u s l y i n S e c t i o n s 5 and 6. A review o f t h e d a t a i n F i g u r e 54 shows t h a t t h e l a r g e s t benefits o f HLFC i n r e d u c t i o n of fuel consumption and w i t h t h e

l e a s t i n c r e a s e i n a i r c r a f t o p e r a t i n g weight a r e ob ta ined wi th t h e h igh wing ,

e n g i n e s on wing HLFC c o n f i g u r a t i o n . As compared t o t h e t u r b u l e n t flow base-

l ine a i r c r a f t , t h e high wing H L F C a i r c r a f t shows 17 percent r e d u c t i o n i n f u e l bu rned , 19.2 p e r c e n t i n c r e a s e i n l i f t - t o - d r a g r a t i o , an i n s i g n i f i c a n t i n c r e a s e i n o p e r a t i n g weight, and 7.4 p e r c e n t r e d u c t i o n i n g r o s s weight. The second

b e s t HLFC c o n f i g u r a t i o n is t h e low wing, f u s e l a g e mounted arrangement w i t h no H L F C on t h e empennage. This c o n f i g u r a t i o n shows 13.7 p e r c e n t r e d u c t i o n i n

f u e l burned , 18.2 p e r c e n t i n c r e a s e i n l i f t - t o - d r a g r a t i o , 5.4 p e r c e n t i n c r e a s e

i n o p e r a t i n g weight , and 4.2 percent r e d u c t i o n i n g r o s s weight . This con-

f i g u r a t i o n w i t h no HLFC on t h e empennage is favored over t h e low wing HLFC i n i t i a l b a s e l i n e a i r c r a f t because o f t h e reduced complexi ty o f e l i m i n a t i o n of

HLFC p e c u l i a r equipment f o r t h e empennage. The c a n d i d a t e f o r t h e f o u r t h

selected c o n f i g u r a t i o n was de te rmined from t h e r e s u l t s o f t h e high wing HLFC

w i t h no lower s u r f a c e l a n i n a r flow and t h e low wing b a s e l i n e a i r c ra f t o p e r a t e d a t i n i t i a l c r u i s e a l t i t u d e o f 36,000 feet. The r e s u l t s show approx ima te ly t h e

sane p e r c e n t a g e r e d u c t i o n i n f u e l burned, b u t t h e c o n s i d e r a b l y lower o p e r a t i n g weight and g r o s s weight o f t h e h igh wing c o n f i g u r a t i o n is considered more

f a v o r a b l e as compared t o t h e h i g h e r a l t i t u d e o p e r a t i o n case . As expected, t h e

80

a s p e c t r a t i o 10 HLFC c o n f i g u r a t i o n r e s u l t e d i n reduced o p e r a t i n g weight and t h e b e n e f i t s i n f u e l consunp t ion and l i f t - t o - d r a g r a t i o were ve ry low as compared t o t h e h ighe r a s p e c t r a t i o t u r b u l e n t flow b a s e l i n e a i r c r a f t .

In sumnary, t h e f i n a l selected HLFC c o n f i g u r a t i o n s ob ta ined from t h e

p a r a m e t r i c s i z i n g s t u d i e s o f M = 0.77 HLFC g l o b a l r a n g e m i l i t a r y a i r c r a f t are l i s t e d below i n order of h i g h e s t p r i o r i t y :

<

HIGH WING

HLFC ON SURFACE ALTITUDE RATIO ENGINES ON NO LOWER NO HLFC NO LOWER ASPECT HIGH WING ENGINES ON WING;

BASELINE EMPENNAGE HLFC 36,000 F T 10 WING, HLFC SURFACE HLFC

WEIGHTS

OPERATING EMPTY 5.4 5.9 7.9 15.5 -0.7 0.2 1.9

GROSS - 9.0 - Y.2 -0.6 - 0.3 -1.1 - 7.1 - 9.9

FUEL CONSUMPTION -13.4 -13.7 -7.9 -12.4 -3.5 -17.1 -11.9

L IFT TO DRAG RATIO 18.9 18.2 12.5 22.7 4.0 19.2 13.6

1.

2.

3.

4.

5.

High wing wi th wing mounted eng ines and no HLFC on empennage.

Low w i n g wi th f u s e l a g e mounted eng ines and no HLFC on empennage.

Low wing w i t h f u s e l a g e mounted eng ines and i n c l u d i n g HLFC on empennage.

High wing w i t h wing mounted eng ines and no HLFC on lower wing

s u r f ace.

Low wing wi th f u s e l a g e mounted eng ines and i n c r e a s e d i n i t i a l c r u i s e

a l t i t u d e t o 36,000 feet .

Figure 54. Summary of HLFC Aircraft Results Relative to Turbulent Flow Baseline

81

is an o p t i m i s t i c assumption which canno t be v a l i d a t e d a t t h i s time f o r an HLFC

a i r c r a f t w i t h wing mounted eng ines . P re l imina ry f l i g h t tests o f a n a t u r a l

l a n i n a r f low gloved wing s e c t i o n j u s t ou tboa rd o f t h e eng ine on a b e i n g 757

a i r c r a f t have been made t o determine t h e e f f e c t s o f engine n o i s e on boundary

l a y e r t r a n s i t i o n as reported i n Reference 20.. These f l i g h t test results showed f o r t h e c o n d i t i o n s o f t h e tests a n e g l i g i b l e effect o f engine power on t r a n s i t i o n on t h e upper s u r f a c e and s m a l l e f f e c t s on t h e lower s u r f a c e . It is

f e l t , however, f o r c o n d i t i o n s i n v o l v i n g longe r laminar r u n s such a s those f o r t h e HLFC c o n f i g u r a t i o n s o f t h i s s t u d y t h a t t h e e f f e c t s o f t h e e n g i n e on laminar boundary l a y e r t r a n s i t i o n are incomplete . Furthermore, it i s expec ted

t h a t e n g i n e effects on t r a n s i t i o n could be more pronounced f o r t h e mul t i -

engine a r rangements o f t h i s s t u d y . Fu r the r r e s e a r c h and development work i n

t h i s impor t an t a r e a is war ran ted .

In t h e ranking o f selected c o n f i g u r a t i o n s listed above, t h e low wing w i t h

f u s e l a g e mounted engines and i n c l u d i n g HLFC on t h e empennage r e c e i v e d a h igh

r ank ing . It is f e l t , however, t h a t a p r e f e r r e d c o n f i g u r a t i o n w i l l n o t have HLFC on t h e empennage i n o r d e r t o s i m p l i f y t h e d e s i g n concep t and p rov ide a more p r a c t i c a l a i r c r a f t .

The o p e r a t i o n o f t h e HLFC a i r c r a f t a t 36,000 fee t i n i t i a l cruise a l t i t u d e

p rov ides a n o t a b l e i n c r e a s e , 3.7 p e r c e n t , i n l i f t - t o - d r a g r a t i o a s compared t o

t h e b a s e l i n e HLFC a i r c r a f t o p e r a t i n g a t 31,361 feet . T h i s improvement i n l i f t - t o - d r a g r a t i o , however, is accompanied by an 1 8 p e r c e n t i n c r e a s e i n

e n g i n e t h r u s t . Th i s l a r g e i n c r e a s e i n e n g i n e t h r u s t is a t t r i b u t a b l e t o t h e

r e l a t i v e l y h igh by p a s s r a t i o o f 6.97 o f t h e STF-686 eng ine used i n t h i s

s t u d y . S ince o p e r a t i o n a t h i g h e r a l t i t u d e s is more f a v o r a b l e t o t h e a t ta in- ment and p r e s e r v a t i o n o f laminar flow, a d d i t i o n a l i n v e s t i g a t i o n s of h i g h a l t i t u d e o p e r a t i o n s w i t h lower by p a s s r a t i o e n g i n e s is war ran ted .

82

8.0 CONCLUSIONS AND RECOIO(ENDATI0NS

P r e l i m i n a r y d e s i g n system studies o f t h e a p p l i c a t i o n o f hybr id l amina r

f low c o n t r o l t o m i l i t a r y t r a n s p o r t s s i z e d t o perform g l o b a l r a n g e mis s ion c h a r a c t e r i s t i c s show s i g n i f i c a n t performance benefits ob ta ined f o r t h e h y b r i d LFC a i r c r a f t a s compared t o c o u n t e r p a r t t u r b u l e n t f l ow a i r c r a f t . The parametric a i r c r a f t s i z i n g studies inc luded t h e development o f a c o n s i d e r a b l e d a t a base cove r ing bo th s e n s i t i v i t y s t u d i e s o f b a s e l i n e a i r c r a f t as well a s changes i n t h e o v e r a l l des ign c o n c e p t s o f t h e a i r c r a f t c o n f i g u r a t i o n s . The s t u d y r e s u l t s a t M = 0.77 show t h a t t h e l a r g e s t b e n e f i t s of HLFC a r e ob ta ined wi th a high wing with engines on t h e wing c o n f i g u r a t i o n . As compared t o t h e tu rbu- lent f low b a s e l i n e a i r c r a f t , t h e h i g h wing HLFC a i r c r a f t shows 17 percent

r e d u c t i o n i n fuel bu rned , 19.2 percent i n c r e a s e i n l i f t - t o - d r a g r a t i o , an

i n s i g n i f i c a n t i n c r e a s e i n o p e r a t i n g we igh t , and 7 .4 r e d u c t i o n i n g r o s s weight . For t h i s high wing c o n f i g u r a t i o n t h e performance d a t a are based on t h e assump- t i o n t h a t there i s no upper s u r f a c e l o s s i n l a n i n a r f low w i t h e n g i n e s mounted

on t h e wings . It is f e l t t h a t t h i s is an o p t i m i s t i c a s s u n p t i o n e s p e c i a l l y f o r t h e l o n g e r l a n i n a r r u n s f o r t h e HLFC c o n d i t i o n s o f t h i s s t u d y and f o r t h e

mul t i - eng ine c o n f i g u r a t i o n s . The second best H L F C c o n f i g u r a t i o n i s t h e low

wing, f u s e l a g e mounted arrangement w i t h no HLFC on t h e empennage. This

c o n f i g u r a t i o n shows 13.7 percent r e d u c t i o n i n f u e l burned , 18.2 p e r c e n t i n c r e a s e i n l i f t - t o - d r a g r a t i o , 5.4 p e r c e n t increase i n o p e r a t i n g weight , and

4.2 p e r c e n t r e d u c t i o n i n g r o s s weight a s compared t o t h e t u r b u l e n t f low a i r c r a f t .

S e n s i t i v i t y s t u d i e s inc luded t h e d e t e r m i n a t i o n o f t h e effects on p e r f o r -

mance o f i n c r e a s e i n c r u i s e Mach number from 0.77 t o 0.80, i n c r e a s e i n i n i t i a l c r u i s e a l t i t u d e t o 36,000 f e e t , and e l i m i n a t i o n of HLFC on t h e lower wing

s u r f a c e . These changes g e n e r a l l y r e s u l t e d i n d e g r a d a t i o n i n performance as

compared t o t h e b a s e l i n e a i r c r a f t c h a r a c t e r i s t i c s . As expected, t h e r e d u c t i o n

i n aspect r a t i o from t h e b a s e l i n e v a l u e o f about 13 t o a v a l u e o f 10 resulted i n v e r y low improvements i n f u e l consumption and l i f t - t o - d r a g r a t i o .

I n v i ew of t h e s u p e r i o r performance o f t h e h igh wing wi th e n g i n e s mounted on t h e wing HLFC c o n f i g u r a t i o n , it is recommended t h a t f u r t h e r r e s e a r c h and

83

developnent be conducted t o p r o v i d e t h e n e c e s s a r y d a t a base for v a l i d a t i o n of

t h e effects o f engine o p e r a t i o n on laminar boundary l a y e r t r a n s i t i o n for

f l i g h t Reynolds nunbers cor responding t o l a r g e , long r a n g e t r a n s p o r t a i r c r a f t .

Opera t ion a t h ighe r a l t i t u d e s of 36,000 feet and above are more f a v o r a b l e t o . . t h e a t t a i n m e n t and p r e s e r v a t i o n of l a n i n a r flow. 'Ihe d a t a i n t h i s s t u d y show

a moderate i n c r e a s e i n l i f t - t o - d r a g r a t io for o p e r a t i o n a t i n i t i a l c r u i s e

a l t i t u d e o f 36,000 f e e t , b u t w i t h an a t t e n d a n t l a r g e increase i n e n g i n e t h r u s t for t h e r e l a t i v e l y h igh by p a s s r a t i o eng ines used i n t h i s s tudy . It i s

recommended t h a t a d d i t i o n a l s t u d i e s be made o f h i g h a l t i t u d e o p e r a t i o n s w i t h

lower by p a s s r a t i o engines .

A l l HLFC a i r c r a f t i n t h i s s t u d y have been sized w i t h o u t t h e u s e o f

leading-edge h igh l i f t d e v i c e s on t h e wings. In view of t h e f a v o r a b l e e f fec ts

of leading-edge h i g h l i f t sys t ems on t h e a i r f i e l d performance and t h e

s h i e l d i n g effects for HLFC o p e r a t i o n s , i t is recommended t h a t a d d i t i o n a l s i z i n g s t u d i e s be conducted on t h e t w best HLFC c o n f i g u r a t i o n s o f t h i s s t u d y

wi th t h e a d d i t i o n of leading-edge h igh l i f t systems.

84

REFERENCES

1. Braslow, Alber t L. and Ralph J. Muraca, ' lA P e r s p e c t i v e o f Laminar-Flow

Control,11 A I A A Paper 78-1528, Los Angeles , CA, August 1978.

2. S turgeon, R.F., e t a l , "Study o f t h e App l i ca t ion o f Advanced Technologies

t o Laminar-Flow Con t ro l Systems f o r Subsonic lkanspor t s , ' l NASA CR

133949, prepared by Lockheed-Georgia Company, May 1976.

3. S tu rgeon , R.F., Developnent and Eva lua t ion o f Advanced Technology

Laminar-Flow-Control Subsonic Transpor t A i r c r a f t A I A A Paper 79-96,

H u n t s v i l l e , Alabama, January 1978.

4. S turgeon, R.F., llToward a Laminar-Flow Cont ro l T ranspor t ,11 CTOL Transpor t

Technology 1978, NASA Conference P u b l i c a t i o n 2036, Februa ry 28 - March 3, 1978.

5. S tu rgeon , R.F., et a l , l lEva lua t ion o f LFC System Concepts f o r S b s o n i c

Commerc ia l T r a n s p o r t A i r c r a f t , I 1 NASA C R 159253 , p r e p a r e d b y Lockheed-Georgia Company, S e p t . 1980.

6. Kraner, J. J., "Planning a New Era i n Air T r a n s p o r t a t i o n E f f i c i e n c y , "

A s t r o n a u t i c s and A e r o n a u t i c s , July/August 1978, pp. 26-28.

7. Conner , D.W., "CTOL Concepts and Technology Developnent ," A s t r o n a u t i c s l and A e r o n a u t i c s , July/August 1978, pp. 29-37.

t I

I

8. Leonard, R.W., l1 A i r f r a n e s and Aerodynamics ," A s t r o n a u t i c s and Aeronau-

- t i c s , Juiy/August 1978, pp. 28-46.

9. Wagner, Richard D. and Michael C. F i s c h e r , "Developnents i n t h e NASA I T r a n s p o r t A i r c r a f t Laminar Flow Rogram," A I A A Paper 83-0090, Reno, I Nevada, January 1983.

85

10. L i n e b e r g e r , L.B., e t a l . , llDevelopnent o f Laminar Flow Cont ro l Wing Sur face Composite S t ruc tu res ,11 NASA C R 172330, prepared by t h e Lockheed-Georgia Company, May 1984.

11. L inebe rge r , L.B., e t a l . , l l S t r u c t u r a l Tests and Developnent o f a Laminar Flow Con t ro l Wing S u r f a c e Composite Chordwise Joint ,(I NASA CR

172462, prepared by t h e b c k h e e d - G e o r g i a Company, December 1984.

12. E tchbe rge r , F.R., e t a l . , "LFC Leading Edge Glove F l i g h t - A i r c r a f t Modi- f i c a t i o n & s i g n , Test Article Developnent, and Systems I n t e g r a t i o n , 1 1 NASA CR 172136, prepared by Lockheed-Georgia Company, Nov. 1983.

13. Lange, R.H., "Design I n t e g r a t i o n o f Laminar Flow Con t ro l f o r T ranspor t A i r c r a f t , " J o u r n a l o f A i r c r a f t , Vol. 21, No. 8, August 1984, pp. 61 2-61 7.

14. Goodyear, M.D., " A p p l i c a t i o n o f S u p e r p l a s t i c a l l y Formed and D i f f u s i o n

Bonded Aluminum t o a Laminar Flow Can t ro l Leading Edge," NASA CR

1783 1 6, p repared by Loc kheed-Georg i a Company, J u l y 1987.

15. Moore, J.W., e t a l . , "Technology A l t e r n a t i v e s for Airlift Deployment ,11

Air Force Wright Aeronau t i ca l t a b o r a t o r y Report AFWAL-TR-85-3001, prepared by Lockheed-Georgia Company, A p r i l 1985.

16. Anon, "Hybrid Laminar Flow Con t ro l Study," NASA CR-165930, p repa red by

b e i n g C o m e r c i a l A i r c r a f t Company, October 1982.

17. Smethers , Rol lo G., and John F. Honra th , l lConf igu ra t ion I n t e g r a t i o n for L a r g e Mul t i -Pur p o s e A i r c r a f t ,'l A i r F o r c e W r i g h t Aeronaut ical Labora to ry Re p o r t AFWAL-T R-85-3075, p repa red by Lmkheed-Georg i a Company, September 1985.

18. S t a f f , LFC Engineer ing S e c t i o n : " F i n a l Report on LFC Aircraft Design

Data, Laminar Flow Cont ro l Demonstration Progran ,11 NOR 67-1 36, Nor throp Corpora t ion , Nora i r D i v i s i o n , June 1967.

86

I 19. Torenbeek, Egbert , l *Synthes i s o f Subsonic Airplane Des ign", Delft

University Press, Holland, 1982, p . 299.

20. Anon, "Flight Survey of t h e 757 Wing Noise F i e l d and Its Effects on

Laminar Boundary Layer Transition ,lt NASA CR 178216, 178217, prepared

by the b e i n g Commercial Airplane Company, March 1987.

87

APPENDIX A

GENERAL AIRCRAFT SIZING PROGRAM (GASP)

The Lockheed Genera l ized Aircraft Si zing and Performance (GASP) computer

program is used t o s ize and d e f i n e t h e a i r c r a f t i n t h i s s tudy . l'he method-

o l o g y of t h i s program is o u t l i n e d i n F igure A-1.

GASP c o n t r o l s t h e i n t e r a c t i o n of t h e program modules provided by t h e

var ious t e c h n i c a l d i s c i p l i n e s and t h e i n p u t s provided for t h e s p e c i f i c con- f i g u r a t i o n . GASP t h e n g e n e r a t e s a component b u i l d u p of d rag and weight , and in tegra tes t h e s e resul ts i n t o t o t a l a i r c r a f t d r a g and weight . P ropu l s ion

sys tem s ize is d e f i n e d b y matching c r u i s e t h r u s t r e q u i r e m e n t s o r , if r e q u i r e d ,

by mismatching t h e s e r equ i r emen t s so as to o v e r s i z e t h e e n g i n e a t c r u i s e t o p rov ide a d d i t i o n a l t a k e o f f t h r u s t . ?he c a p a b i l i t y of s i z i n g a c o n f i g u r a t i o n w i t h a f i x e d - s i z e p r o p u l s i o n system is also a v a i l a b l e . The a i r c r a f t s ize r e q u i r e d f o r t h e mission is d e f i n e d by an automated i t e r a t i v e p rocess . GASP h a s been used i n a nunber o f p r e v i o u s s t u d i e s (References 3, 5, 6, 8, 11, 12,

-1

AIKRAFT E31 T A U O F F

- 1 I

w w r I NSRVCMNS

1 I

Figure A-1 . Generalized Aircraft Sizing and Performance (GASP) Program

88

13, 14, 15) t o s y n t h e s i z e a i r c r a f t f o r d e s i g n v a r i a b l e s such a s wing l o a d i n g , a s p e c t r a t i o , cruise power s e t t i n g , Mach nunber, r ange , pay load , and f i e l d

performance; and t o d e f i n e a i r c r a f t op t imized t o f i g u r e s o f merit such a s m i n i m w n d i r e c t o p e r a t i n g c o s t , g r o s s w e i g h t , a c q u i s i t i o n c o s t , fuel usage , and

l i f e cycle c o s t . These studies have encompassed c o n v e n t i o n a l and a s s a u l t t r a n s p o r t s a s well a s l o i t e r / e n d u r a n c e mis s ions . Turbofan and propfan pro-

p u l s i o n systems were examined and v a r i o u s advanced m a t e r i a l s were eva lua ted .

The g e n e r a l method o f p a r a m e t r i c a n a l y s i s t o be used is i l l u s t r a t e d i n

F igu re A-2. A i r c r a f t c h a r a c t e r i s t i c s a r e g e n e r a t e d by GASP f o r p a r a m e t r i c

v a r i a t i o n s o f s i z i n g v a r i a b l e s . For t h e s e d a t a , performance c o n s t r a i n t s such a s one-engine-out climb g r a d i e n t c a p a b i l i t y , f i e l d l e n g t h r e q u i r e m e n t s , f u e l volume a v a i l a b i l i t y , and l a n d i n g approach speed can be gene ra t ed and s u i t a b l e

c o n s t r a i n t s imposed on t h e r e s u l t i n g c o n f i g u r a t i o n . The two c a r p e t p l o t s i n F i g u r e A-2 p r o v i d e a p a r a m e t r i c e v a l u a t i o n i n which t h e c o n s t r a i n i n g p e r f o r -

mance v a r i a b l e s a r e t a k e o f f f i e l d l e n g t h and second-segment climb g r a d i e n t . These pa rame te r s a r e p re sen ted a s a f u n c t i o n o f a s p e c t r a t i o ( A R ) and i n i t i a l cruise power s e t t i n g ( t h e percent o f a v a i l a b l e cruise t h r u s t r e q u i r e d a t t h e i n i t i a l cruise p o i n t ) f o r a g i v e n wing l o a d i n g .

S p e c i f i c f i e l d l e n g t h and grad i e n t c a p a b i l i t i e s , a s determined from t h e s e c a r p e t p l o t s , a long wi th a r equ i r emen t t h a t t h e e n g i n e s have f i v e p e r c e n t excess a v a i l a b l e c r u i s e t h r u s t ( = 0.951, d e f i n e t h e group o f a c c e p t a b l e c o n f i g u r a t i o n s i n s i d e t h e hatched a r e a i n t h e m i d d l e s e c t i o n i n F igu re A-2. If d e s i r e d , t h e comple te c a r p e t p l o t s o f a i r c r a f t c h a r a c t e r i s t i c s such as

g r o s s weight and l i f e cycle c o s t s can be g e n e r a t e d t o show t h e impact o f s i z i n g c o n s t r a i n t s on t h e v a r i o u s f i g u r e s of m e r i t . However, t h e optimun v a l u e s o f t h e f i g u r e s o f merit w i l l be d e f i n e d b y t h e envelope of c o n s t r a i n t l ines shown i n t h i s p l o t .

Ano the r method o f p a r a m e t r i c a n a l y s i s i s p r o v i d e d b y a n u m e r i c a l

o p t i m i z e r t h a t h a s been coupled wi th t h e s i z i n g progran . Th i s p r o v i d e s t h e

c a p a b i l i t y o f a u t o m a t i c a l l y s e l e c t i n g a i r c r a f t t h a t minimize a g iven f i g u r e o f

merit wh i l e s i m u l t a n e o u s l y meet ing a d e f i n e d set o f c o n s t r a i n t s .

89

'I

t 1

Figure A-2. Typical Parametric Se lec t ion Procedure

I n p u t r e q u i r e m e n t s t o t h e GASP program t h a t a r e of p a r t i c u l a r s i g n i f i -

cance i n c l u d e : ( a ) m i s s i o n d e f i n i t i o n , ( b ) d e s i g n payload and speed, ( c ) c o n c e p t and t e c h n o l o g y d e f i n i t i o n , and ( d ) economic ground r u l e s . Other

i n p u t s , i n c l u d i n g atmospheric d a t a , geometric c h a r a c t e r i s t i c s , and o p t i m i z e r i n p u t s and c o n s t r a i n t s a r e a l s o r e q u i r e d .

I n s t a l l e d e n g i n e performance d a t a i n t h e form of t h r u s t and SFC are p r o v i d e d b y t h e p r o p u l s i o n o r g a n i z a t i o n fo r b o t h t u r b o f a n and p r o p f a n

i n s t a l l a t i o n s . For p r o p f a n s , t h i s r e q u i r e s coupl ing of a specific t u r b o s h a f t

engine w i t h a specific propeller d e s i g n . For a g iven propel le r d e s i g n , an

optimun propeller d i s c l o a d i n g can be d e f i n e d for a g iven a l t i t u d e ( d e n s i t y

r a t i o ) , c r u i s e Mach number, number of b l a d e s , and t i p speed. S ince t h e

a l t i t u d e and speed r e q u i r e m e n t s a r e v a r i a b l e , a i rc raf t a r e sized for s e v e r a l

d i s c l o a d i n g combina t ions .

I n a d d i t i o n , r o u t i n e s a r e i n c o r p o r a t e d i n GASP t o c a l c u l a t e t h e wing

scrubbing d r a g d u e t o t h e propwash of wing-mounted propellers and t o i n c l u d e

90 c -L

t h i s i n t h e a i r f r a m e d r a g b u i l d u p . Based on r e c e n t NASA/industry s t u d i e s , our

c u r r e n t propfan studies do no t i n c l u d e swirl d r a g p e n a l t i e s ; however, a method

t o i n c o r p o r a t e t h i s e f f ec t is a v a i l a b l e i f l a t e r d a t a i n d i c a t e s t h i s p e n a l t y

e x i s t s .

Concept and t e c h n o l o g y d e f i n i t i o n s are r e q u i r e d i n p u t t o t h e GASP program

i n terms of weight , performance, and/or cost ad jus tmen t factors. lhe weight,

performance, and cost r e l a t i o n s h i p s upon which t h e CLASP program operates are based on c o n v e n t i o n a l concep t and c u r r e n t technology d e f i n i t i o n s . Therefore, for each unique concep t and/or t echno logy , factors t h a t reflect t h e i r in -

f l u e n c e on weight , performance, and other charac te r i s t ics w i l l be developed t o a degree t h a t i s c o n s i s t e n t w i t h t h e c o n c e p t u a l d e s i g n o f each

concep t / tech nology i n t e g r a t i o n .

I

91

rn utwi & ~ ~ J U I C S a-c Report Documentation Page Yxe %-~SVJIW

1. Report No.

NASA CR-181638 2. Government Accession No.

4. Title and Subtitle A p p l i c a t i o n of Hybrid Laminar Flow Con t ro l t o Global Range M i l i t a r y Transpor t A i r c r a f t

7. Authork)

Roy H. Lange

9. Performing Organization Name and Address

Lockheed A e r o n a u t i c a l Systems Company 86 South Cobb Dr ive Narietta, GA 30063

12. Sponsoring Agency Name and Address

Nat iona l Aeronau t i c s and Space Admin i s t r a t ion Langley Research Center Hampton, VA 23665-5225

15. Supplementary Notes

3. Recipient's Catalog No.

5. Report Date

A p r i l 1988 6. Performing Organization Code

8. Performing Organization Report No.

LG87ER0145

10. Work Unit No.

11. Contract or Grant No.

NAS1-18036

13. Type of Report and Period Covered C o n t r a c t o r Report Sep t . 1986 - S e p t . 1987

14. Sponsoring hgency Code

Langley Techn ica l Monitor: D a l V . Naddalon F i n a l Report

16.Abrtract A Study w a s conducted t o e v a l u a t e t h e a p p l i c a t i o n of hybr id l amina r f low c o n t r o l (HLFC) t o g l o b a l r ange m i l i t a r y t r a n s p o r t a i r c r a f t . The g l o b a l mi s s ion

inc luded t h e c a p a b i l i t y t o t r a n s p o r t 132,500 pounds of payload 6,500 n a u t i c a l miles, l a n d and d e l i v e r t h e payload and wi thou t r e f u e l i n g r e t u r n t o 6,500 n a u t i c a l

m i l e s t o a f r i e n d l y a i r b a s e .

performance b e n e f i t s o b t a i n e d f o r t h e HLFC a i r c r a f t as compared

t u r b u l e n t f l ow a i r c r a f t . b e n e f i t s of HLFC are o b t a i n e d w i t h a h i g h wing w i t h eng ines on t h e wing

c o n f i g u r a t i o n .

HLFC a i r c r a f t shows 17 p e r c e n t r e d u c t i o n i n f u e l burned, 19.2 p e r c e n t i n c r e a s e i n l i f t - t o - d r a g r a t i o , an i n s i g n i f i c a n t i n c r e a s e i n o p e r a t i n g we igh t , and 7.4 r e d u c t i o n i n g r o s s weight .

The p r e l i m i n a r y d e s i g n s t u d i e s show s i g n i f i c a n t

t o c o u n t e r p a r t The s t u d y r e s u l t s a t M = 0.77 show t h a t t h e l a r g e s t

A s compared t o t h e t u r b u l e n t f low b a s e l i n e a i r c r a f t , t h e h i g h wing

17. Key Words (Suggested by Author(sJJ 18. Dirtribution Statement Laminar f low c o n t r o l Drag r e d u c t i o n Hybrid l a m i n a r f low c o n t r o l Long r a n g e t r a n s p o r t s Fue l c o n s e r v a t i o n

19. Secuw Cbssif. (of thir report)

U n c l a s s i f i e d U n c l a s s i f i e d 20. Security Classif. (of this page) 21. No. of paw 22. Price

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