V/STOL tilt rotor aircraft study. Volume 6: Preliminary design of a composite wing for tilt rotor research aircraft

NASA CR-114599

V/STOL TILT ROTOR AIRCRAFT STUDY

VOLUME V I

PRELIMINARY DESIGN

OF A COMPOSITE WING FOR TILT

ROTOR RESEARCH P.IRCRAFT

A A A Df'U 4 n 7 Q Y A S A - C R - 1 1 4 5 9 9 ) V/STOL T I L T F C T O F ~ 7 3 - 3130 1 0 B I h C R A F l 5iL'"Y. V O L O f i E 5 : E n i L I Y I H H a Y I CESIGtd O F A C S i I P O S I T E W I N G F C C TILT LiOTOR ( B o e i n q Vertol Co., P h r l a d t - l p t i i d , Il1,clds P a . ) 1 2 6 p HC 8 8 - 5 0 C S C L 01C' ( ; . 3 / 3 2 1 1 ~ t . 4 Distribution of this Report i3 provided in the interest of information e~chmnge. Rsponsibility for the contents resides in tho ruthor or oqmniution uha: prepared it.

Propard Unckr Contract No. NASZ-6598 by SOEING VERTOL COMPANY BOElNCi CENTER P. o. BOX 1 - 1 3 P)rikk@hia, Pmsybania 19142 NASA CR-114599

V/STOL TILT ROTOR AIRCRAFT STUDY

VOLUME VI

PRELIMINARY DESIGN

OF A COMPOSITE WING FOR TILT

R O T O R R E S E A R C H A I R C R A F T

MARCH 1973

Distribution of this Report i s provided in the interest of information exchange. Responsibility for the contents res~des in the author o r organization that prepared it.

Prepared Under Contract No. NAS2-6598 by BOEING VERTOL COMPANY BOEING CENTER P. 0. Box 16858 Philadelphia, Pennsylvania 19142 for Ames R-rch Center biifnui A e r o ~ u t i a and Space Administration and Unitd Smtr Army Air Mobility Rereerch and Dsvdopment Laboratory Amos Directorate REV LTR A

THE BCIEINE s o M r m J r

VERTOL DIVISION . PHILADELPHIA, PENNSYLVANIA

CODE IDENT. NO. 7 7 2 7 2 NUMBER D 2 2 2 - 1 0 0 6 0 - 2

TITLE V/STOL T I L T ROTOR AIRCRAFT STUDY -

PRELIMINARY DESIGN O F A COMPOSITE WING FOR T I L T ROTOR RESEARCH AIRCRAFT ORIWAL RELEASE D A T E - . . FOR THE RELEASE DATE OF SUBSEQUENT REVISIONS, SEE THE REVISION WEET. FOR LIMITATIONS IMPOSED ON THE DISTRIBUTION AND USE OF INFORMATION CONTAINED IN THIS DOCUMENT, SEE THE LIMITATIONS SHEET.

ISSUE NO. ISSUED TO: PREPARED BY DATE March, 1973 March, 1973 PREPAREDBY DATE

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APPROVEDBY DATE * APPROVEDBY DATE / I APPROVED BY W - Peck DATE CORM 4 @ 2 @ 0 (6/bOl SHEET I LIMITATIONS Ail revisions to this document sholl ba approved by the obove noted organization prior to release.

RIVUIOWS

LTR I DESCRIPTION I DATE I APPROVAL

A NASA comments incorporated 7/16/73 D222-10060- 2 FOREWORD REV. A T h i s r e p o r t i s r n c o f a series p r e p a r e d by The Boeing Vertol Company, P h i l a d e l p h i a , P e n n s y l v a n i a f o r t h e N a t i o n a l Aero- n a u t i c s and Space A d m i n i s t r a t i o n , Arnes R e s e a r c h C e n t e r , M o f f e t t F i e l d , C a l i f o r n i a u n d e r c o n t r a c t NAS2-6598. The s t u d i e s r e p o l ' t e d under Volumes I t h r o u g h I V and V I I I t h r o u g h X were j o i n t l y funded by NASA and t h e U.S. Army A i r M o b i l i t y Research and Development L a b o r a t o r y , Aves D i r e c t o r a t e . Volumes V t h r o u g h V I I were funded by t h e U.S. A i r F o r c e F l i g h t Dynamics L a b o r a t o r y , Wright P a t t e r s o n A i r F o r c e Base, Ohio.

T h i s c o n t r a c t was administered by t h e N a t i o n a l A e r o n a u t i c s an<[ Space A d i n i n i s t r a t i o n . M r . R i c h a r d J. A b k o t t was t h e C o n t r a c t A d m i n i s t r a t o r , M r . Gary B. C h u r c h i l l , T i l t R o t o r Research A i r c r a f t P r o j e c t O f f i c e , was t h e T e c h n i c a l E o n i t o r , and c o x - d i n a t i o n and l i a i s o n w i t h t h e U . S . A i r F o r c e F l i g h t Dynamics L a b o r a t o r y was t h r o u g h M r . D. F r a q a .

The complete l i s t o f r e p o r t s p u b l i s h e d u n d e r t h i s c o r t r a c t is as f o l l o w s :

Volume I -- C o n c e p t u a l Design o f U s e f u l M i l i t a r y and/or

Commercial A i r c r a f t , NASA CR-114437

Volume I1 -- P r e l i m i n a r y Design o f Research A i r c r a f t , NASA

CR-114438 Volume I11 -- O v e r a l l Research A i r c r a f t F r o j e c t P l a n , S c h e d u l e s , and E s t i m a t e d C o s t , NASA CR-114439

Volume I V -- Wind Tunnel I n v e s t i g a t i o n P l a n f o r a F u l l S c a l e

T i l t Rotor Research A i r c r a f t , CX-114430 Volume V -- D e f i n i t i o n o f Stowed Rotor Research A i r c r a f t , NASA CR-114598

Volume V I -- Pre1imir.ary Design of a Composite Wing f o r T i l t

Rotor A i r c r a f t , NASA CR-114599 -- T i l t R o t o r F l i g h t C o n t r o l Program Feedback s t u d i e s , Volume V I I NASA CR-114600

Volrune V I I I -- M a t h e m a t i c a l Medel f o r a Real Time S i m u l a t i o n o f

a T i l t R o t o r A i r c r a f t ( ~ o e i n g Vertol Model 2 2 2 ) .

NASA CR-114601

Volume I X -- P i l o t e d S i m u l a t o r E v a l u a t i o n of t h e Boeing Vertol

Model 222 Tilt Rotor A i r c r a f t , NASA CR-114602

Volume X -- Performance and S t a b i l i t y Test o f a 1/4.622 Froude

S c a l e d Boeing V e r t o l Model 222 T i l t Rotor A i r c r a f t (Phase I ) , NASA CR-114603 ABSTRACT

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This report presents the results of a study of the use of com- posite materials in the wing of a tilt rotor aircraft. An all- metal Search and Rescue (SAR) tilt rotor aircraft was first defined to porvide a basis for comparing composite with metal structure. A configuration study was then done in which the wing of the metal aircraft was replaced with composite wings of varying chord and thickness ratio. The results of this study defined the design and performance benefits obtainable with composite materials. Based on these results the aircraft was resized with a composite wing to extend the weight savinga to other parts of the aircraft. A wing design was then select- ed fur detailed structural analysis. A development plan including costs and schedules to develop this wing and incor-

porate it into ; proposed fligHt research tilt rotor vehicle

has been devised. ' VII REV. A TABLE OF CONTENTS

- -

2 . 0 INTRODUCTION. . . . .

3.0 PRELIMINARY DESIGN STUDIES .

. . . . . . . . . . . . 3.2 Design c o n s i d e r a t i o n s

3.2.1 Mission P r o f i l e . . . . . . . . . . . .

3.2.2 PropulsionSystem. . . . . . . . . . . .

3.2.3 A i r c r a f t Drag. . . . . . . . . . . . .

3.2.4 Criteria f o r S e l e c t i n g Design

. . . . . . . . . . . . P o i n t A i r c r a f t

3.2.5 wing S t r u c t u r a l Design criteria. . . .

3.2.6 composite weight F a c t o r s . . . . . . .

. . . . . . . . . . . . . 3.3 All-Metal A i r c r a f t .

. . . . . . . . . . . . . . . . 3.3.1 S i z i n g

. . . . . . . . . . . . . 3.3.2 Performance, . . . .

3.4 Composite Wing Configuration Studies.

Resized Compoaite Wing A i r c r a f t

. . . . . . . 3 ,q

3.5.1 S i z i n g . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . 3.5.2 Performance.

3.6 Design B o r n f i t s obtairrablo w i t h Compoaite8. .

4.1 St!ructural Dasign Requirements. . . . . . . .

. . . . . . . . . . . . . . 4.2 Goad D i r t r i b u t i o n .

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

4 , 3 Material S e l e c t i o n . 4-14

4.4 Wing Derign conmiderationr. . . . . . . . . . 4-21

VIII D222-10060-2 REV. A TABLE OF CONTENTS (continued) PAGE

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4.5 Stress Analysis . . . . . . . . . . . . . . . 4-36

4 . 6 Weig1.r::. S w 'ry - Composite Wing for SAR

A i r c f . n f t . . . . . . . . . . . . . . . . . . 4-62

5 . 0 DEVELOPMENT PLAN . . . . . . . . . . . . . . . . . 5-1

6 . 0 REFERENCES . . . . . . . . . . . . . . . . . . . . 5-1

LIST OF FIGURES FIGURE TITLE: PAGE

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3-1 DESIGN MISSION PROFILE, USAF SAR HI-HI

MISSION . . . . . . . . . . . . . . 3-4

DRAG TREND . . . . . . . . . . . . . 3-7

ALL-METAL SAR TILT ROTOR PARAMETRIC SIZING RESULTS. o . . . . . . . . . . . - 3 - 1 2

ALL-METAL SAR TILT ROTOR OGE HOVER CEILING . 3-13

ALL.YETAL SAR TILT ROTOR PPRF'ORXANCE SUMMARY 3-16 PLANFORM CHARACTERISTICS OF CONSTANT THICKNESS . . . . . . . . . .

RATIO WING SERIES 3-18 DRAG AREA INCREMENT BETWEEN COMPOSITE WINGS

AND BASELINE JUL-METAL WING . . 3-19

COMPOSITE WING WEIGHT FOR WING PARAMETER

. . . . . . . . . . . . . STUDY .3-21

EFFECTS OF WING THICKNESS RATIO (AT CONSTANT CHORD) ON TAKEOFP GROSS WEIGHT AND MAXIMUM

. . . . . . . . . RESCUECAPABILITY .3-24

EFFECTS OF WING 'YICXCNESS RATIO (CONSTANT CHORD)

ON MISSION R ' AND MIDPOINT HOVER ENDURANCE . 3-25

EFFECTS OF ' CIiORD (CONSTANT THICKNESS A T I O ) ON TAKEOFF .2 W I G m AND MAXIMUM RESCUE

CAPABILITY . . . . . . . . . . - 3 . 2 6

EFFECTS OF WING CHORD (CONSTANT TXICRWESS RATIO)

OK MISSION RADIUS AND MIDPOlNT HOVER ENDURANCE . 3-27

CQMBOGITE WING SAR TILT ROTOR PARAMETRIC

. . . . . . . . . . . SIZING RESC1LTS 3-28

COblPOSITE WING SAR TILT ROTOR OGE HOVER CEILING.3-33 C W O S X T E WING SAR TILT mTOR CRUISE PERFORM-

. . . . . . . . . . A N C E S U b J M A R Y . 3 - 3 4

WING TORQUE BCN ROOT CQHSTRUCTIOHAL DETAILS A T W S 3 0 . 0 4 - 2 . . . . . . . . . . .

WING TORQUE BOX COLOSTRUCTI- DETAILS AT

W S 1 8 0 . . . . . . . . . . . . 04-3

X REV. A L I S T O F FIGURES ( c o n t h u e d )

FIGURE - T I T L E - PAGE

4-3 SPANWISE SHEAR DISTRiBUTIClN COND. 1,

4 ' 9 ' VTO . . . . . . . . . - . . . . . . . . . . 4-10

SPANWISE BM D I S T R I B U T I O N COND. 1, 4 '

g o VTO . . . . . . . . . . . . . . . . . . . . -4-11

REQUIRED WING S T I F F N E S S D I S T R I B U T I O N . 0 -4-13

COMPOSITE MATERIALS PROVIDE D E S I G N

F'LEXIBILITY . . . . . . . . . . . , . . . . , . - 4 - 1 7

S P E C I F I C FATIGUE PROPERTIES . - . 4-19

RELATIVE WEIGHTS O F FAT IGUE-CRIT ICAL

STRUCTURES . . . . . . . . . . . . . . . . - 4 - 1 9

RESIDUAL STRENGTE O F COMPOSITES AND ALUMINUM

ALLOYS . . . . . . . . . . . . . . . . . - 4 - 2 0

COMPARISON O F MATERIAL PROPERTIES . . . . . . . . 4 - 2 0

MECHANICAL WING-FUSELAGE Z 3 I N T CONCEPT . . . . 4 - 2 9

T E S T SPECIMEN FOR HLH TRANSMISSION SUPPORT

F I T T I N G . . . . . . 4 - 3 0

D E S I G N O F T E S T SPECIMEN FOR KLH TRANSMISSION

S U P P O R T F I T T I N G . . . . . . . . - ? - 3 a

TRANSMISS ION SUPPOXT F I T T I N G CONCEPT FOR HLH

APPLICATION. . . ., . . . . . . 04-31

WING-NACELLE INTERFACE . . . . . . . . . . . . 4-32

CONSTRUCTION OP F I N A L T E h S ION J O I N T FOR HLH WITH S T A T I C T E N S I L E STRENGTH OF 14,400 P3UNDS. - 4-34

B A S I C GEOMETRY - WING DOX. . . . 4-37

P I S T O F FIGURES ( C o n t hued; PAGE T I T L E

-

4-19 ALIXrWABLE LOAD/IN. - COMPRESSION PANELS

(B,'E HONEYCOMB SANDWICH) 4-45 . . . . . . . . . . . .

5-la COMPOSITE WING FOR T I L T ROTOR RESEARCH

. . . . . . . AIiiCIiArT PRCGRZM PLANNING COST. . 5 - 5

5-lb COMPOSITE WING FOR T I L T ROTOR RESEARCH

AIRCRAFT - PROGRAM PLANNING COPT. . 5-6

5 - l < COMPOSITE WING FOR T I L T ROTOR RESEARCH . . . . . . . .

AIRCRAFT PROGRAM PLANNING COST. 5 - 7

5-2 COMPCSITE WING FOR T I L T ROTOR AIRCRAFT -

SCHEDULE. . . . . . . . . . . . . . . . . . . - 5 - 8

I., D 2 2 2 - 1 0 0 6 0 - 2 REV. A L I S T OF TABLES PAGE

-

COMPOSITE HARDWARE DEVELOPMENT . . . . . . 3-10

SUMMARY WEIGHT STATEMENT - ALL-METAL SAR TILT

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

ROTOR 3 - 1 3

SUMMARY WEIGHT STATEMENT - RESIZED COMPOSITE

3-31 WIMS SAI: T I L T X T G R . . . . . . . . . .

DESIGN POINT AIRCRAFT COMPARISON. . . a * .

3-37

DESIGN POINT AIRCRAFT WEIGHT COMPARISON. . . . 3-3 8

COMPOSITE WING TORQUE BOX - MARGINS OF SAFETY .

4-5

USAF SAR T I L T ROTOR AIRCRAFT - SPANWISE BENDING

MOMENT AND SHEAR DISTRIBUTION COND. 1 (VTO)

4 ' 9 ' . . . . . . . . . . . . . . . . . . . . . . 4-9

PRELIMINARY DESIGN ALLOWABLES . . . . . . . . . 4 - 1 5

1 . 0 S U M M A R Y This r e p o r t p r e s e n t s t h e r e s u l t s of a s t u d y conducted by t h e Boeing V e r t o l Company t o d e f i n e t h e e f f e c t s of t h e u s e of composite m a t e r i a l s i n t h e wing of a tilt r o t o r a i r c r a f t .

The o b j e c t i v e s of t h e s t u d y were: 1. te d e f i n e t h e a e s i g n and performance b e n e f i t s o b t a i n e d w i t h composite m a t e r i a l s 2 . t o d e s i g n a composite wing f o r t h e tilt r o t o r r e s e a r c h a i r c r a f t 3 . t o e s t a b l i s h a development p l a n f o r a composite wing f o r t h e tilt r o t o r r e s e a r c h a i r c r a f t The USAF SAR a i r c r a f t d e s c r i b e d i n Reference 1 was used a s a b a s i s f o r t h e s t u d y . S i n c e t h i s a i r c r a f t was designed w i t h some composite s t r u c t u r e , it was r e s i z e d t o a n a l l - m e t a l con- f i g u r a t i o n . The all-metal tilt r o t o r a i r c r a f t was used a s a basj s f 3r comparison.

To d-etennine t h e optimum wing c o n f i g u r a t i o n , t w o p a r a m e t r i c t r a d e s t u d i e s were conducted. I n t h J f i r s t wing chord was h e l d c o n s t a n t and t h i c k n e s s was v a r i e d . I n t h e second t h i c k - n e s s r a t i o was h e l d and chord was v a r i e d . The& s t u d i e s were d e t a i l e d enough t o show t h e e f f e c t s of changing chord and t h i c k n e s s on t h e d r a g and weight of t.he wing and p a r t i c u l a r l y t o show t h e c r o s s o v e r p o i n t between. t h e s t r e n g t h c r i t i c a l and s t i f f n e s s c r i t i c a l d e s i g n c o n d i t i o n s .

~222-10060-2 Three d e s i g n p o i n t a i r c r a f t were d e f i n e d f o r purposes of com- p a r i s o n . The f i r s t is t h e a l l - m e t a l r e f e r e n c e a i r c r a f t de- s c r i b e d above. The second is a r e s i z e d composite wing a i r c r a f t i n which composite m a t e r i a l s were used f o r t h e e n t i r e wing and t h e r e s i z i n g was done t o e x t e n d t h e weight s a v i n g s tc o t h e r parts of t h e a i r c r a f t . The t h i r d d e s i g n p o i n t i n s l l . p l y t h e a l l - m e t a l a i r c r a f t w i t h a composite wing. I n t h i s mse t h e wing weight s a v i n g was t a k e n a s a payload o r p e r f o m . n c e b e n e f i t .

Gross w e i g h t s and r o t o r d i a m e t e r s of t h e s e t h r e e & i r c . r a f t a r e compared w i t h t h e a i r c r a f t o f Reference 1 i n t h e f o l l o w i n g Rotor D i a . Design GW A Weigirt A Weight % F e e t Lbs ?AS A i r c r a f t

- Y - , -

Reference 1 (Moderate u s e of composites throughout) 27.0 15,970 -1,055 -6.2

A l l - metal 28.9 18.025 l l l l l l l . l ---

A l l - metal p l u s max

use of composites i n 28.9 17,650 -3 75 - 2 . 1 wing o n l y i b

A l l - m e t a l p l u s cam-

p o s i t e wing - r e s i z e d 2 i . 1 17.242 -. 733 -4.3

I t may be n o t e d t h a t on t h e t h i r d a i r c r a f t (composite w i ~ ~ g o n l y ?

\ n o t r e s i z e d ) t h e o n l y weight s a v i n g is t h e 308 i ~ u u c t i o n i n wing weight f r o m 1,250 t o 875 pounds.

A wing c o n f i g u r a t i o n was chosen f o r f u r t h e r s t u d y and a d e s i g n and stress a n a l y s i s done. A simple two-spay :onfly- u r a t i o n was chosen f o r t h e wing t o r q u e box w i t h a rspanwiaie 1-2 D222-10060-2 well in the upper surface for the cross shaft. The torque box is a honeycomb shell consisting of Boron-Epoxy facings on a fiberglass honeycomb core.

A development plan has been devised which considers the desSgn, constructicn, and testing of a composite wing for the tilt rotor research aircraft. In order to arrive at a minimum cost program, only the main spar torque box ie built in composites for this program. The auxiliary surfaces (flaps, umbrellas, etc.) are existing metal components.

2 . 0 INTRODUCTION I n March 1972, t h e Boeing V e r t o l Company completed a s t u d y of tilt r o t o r a i r c r a f t under t h e j o i n t s p o n s o r s h i p of NASA and

t h e U.S. A m y (References 1 - 4 ) . P a r t of t h a t s t u d y (Refer-

ence 1) covered t h e conceptual d e s i g n of u s e f u l m i l i t a r y and c i v i l tilt r o t o r a i r c r a f t f o r t h e 1975-1980 t i m e p e r i o d .

Composite m a t e r i a l s were u t i l i z e d a s a means o f reducing a i r f r a m e s t r u c t u r a l weight. I n t h a t s t u d y t h e weight f a c t o r f o r composites was t a k e n a s 158. Design s t u d i e s and p r o t o t y p e t e s t d a t a have i n d i c a t e d , however, t h a t l a r g e r s a v i n g s could be r e a l i z e d w i t h p r e s e n t technology.

Consequently, t h e Boeing Company was asked by t h e U.S. A i r Force F l i g h t Dynamics Laboratory, through an add-on t o t h e N A S A c o n t r a c t , t o i n v e s t i g a t e t h e u s e of composite m a t e r i a l s i n t h e wing of a tilt r o t o r a i r c r a f t . The o b j e c t i v e s of t h i s s t u d y were a s follows: 1. t o d e f i n e t h e d e s i g n and performance improvements a composite wing p r o v i d e s f o r tilt r o t o r a i r c r a f t 2. d e s i g n a composite wing for t h e tilt r o t o r SAR a i r c r a f t 3 . e s t a b l i s h a development p l a n f o r a composite wing f o r t h e tilt r o t o r r e s e a r c h a i r c r a f t T h i s r e p o r t p r e s s n t s t h e r e s u l t s o f t h e study. 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 r e q u i r e d t o d e f i n e t h e optimum wing c o n f i g u r a t i o n a r e d e s c r i b e d i n S e c t i o n 3 , The advanced d e s i g n o f a compo- site wing i s d e s c r i b e d i n S e c t i o n 4 and t h e development p l a n i n S e c t i o n 5 .

D222-10060-2 REV. A 3.0 PRELIMINARY DESIGN STUDIES 3.1 INTRODUCTION A study was performed to show the potential benefits obtainable from the applicaticu; of composite mkterials to an advanced

type of VTOL aircraft - the tilt rotor. A promising op2ration-

a1 application for the tilt rotor configuration - a USAF search

and rescue (SAR) mission - was chosen for this study. This

application for the tilt rotor concept had previously been studied by Boeiag during 1971-72 in a NASA/Azmy sponsored "V/STOL Tilt Rotor Aircraft Study" (Reference 1). Moderate application of composites t o the fuselage, wing, and empennage was assumed for the aircraft defined in that study.

in more detail the application of The present study examines composites to the wing only with the following objectives: (1) To show the improvements in mission performance to the wing achievable by applying composite materials alone of an all-metal search and rescue tilt rotor aircraft.

(2) To show the overall weight and size benefits obtain- able by r-sizing the total aircraft structure to take

advantage of the reduced win3 weight - even though

composites were still applied only to the wing.

To provide a basis for comparison, the SAR aircraft of Reference 1 wae reeized to an all-metal structure.

D222-10060-2 REV. A (For reference, some of the characteristics of the Reference 1 aircraft are tabulated be1ow.j M222-1F SAR TILT ROTOR (REF. 1) Gross Weight (lb) 16970 Weight Empty (lb) 11500 Wing Area (sq. ft. ! 186 Wing Span (ft .)

34.4 Thickness ratio (t/c) 21% Rotor Diameter (ft.) 27.0 Solidity Ratio .I33 Power Plant (2) Lycoming PLT-27 Rated Power (Shp) In the remainder of this section, the criteria for aircraft sizing are first discussed (Section 3.2), followed by a description of the all-metal aircraft and its performance (Section 3.3). The effects of applying composite materials to the wing of the all-metal aircraft are then shown in Section 3.4. This study included the variation of wing geometry (thicknass and chord) to determine whether secondary benefits could be credited to the use of composites by making changes in the wing geometric design. Section 3.5 discusses the effect of resizing the remainder of the aircraft structure to take advantage of the lighter wing. Theee preliminary de- sign studies are then summarized in Section 3.6.

3.2 DESIGN CONSIDEFUiTiBNS - 3.2.1 Design Mission Profile All aircraft were sized to perform a 300 NM Eeatc;i arid Rescue (SARI mission (Figure 3-11. This is a "HI-HI'' mission consisting of a takeoff at SL/95OF, climb to optimum aititude, cruise out at NRP to the 500 NM radius, hover for i/2 hour at 5000ft/95OF and recover three (3) rescuees, and return without inflight refueling.

The optimum cruise altitude (based on minimum fuel) was found to be 20,000 ft.

The aircraft were assumed to carry a four-man crew consisting of two pilots, a crew chief, and a paramedic. The mission Load was specified at 150 lb of rescue equipment (litters, forest penetrator, rescue sling, et.), a i r b ~ r n e electronic3 and equipment required to locate the rescuee, and a 5.56mm machine gun and ammunition.

The engines, rotors, and drive system were sized by an alter- nate mission requirement. This was that the aircraft be cap- able of hovering at the mission midpoint at T/W=1.1 with a

total of seven rescuees - the additional four rescuees being

the crew of a downed sister ship. It was assumed that inflight refueling would be allowed under these conditions so that the mission fuel requirement ie determined by the basic mission ahown in Figure 3-1.

I L E - S . W HI-HI MISSION

3.2.2 P r o p u l s i o n System 3.2.2.1 Engine Cycle I n t h e R e f e r e n c e 1 s t u d y t h e Lycoming PLT-27 e n g i n e r a t e d a t 1950 horsepower was c h o s e n t o power t h e SAR tilt r o t o r a i r c r a f t .

T h i s e n g i n e met t h e m i d p o i n t hover r e q u i r e m e n t w i t h a r e a s o n - a b l e r o t o r d i a m e t e r . T h i s e n g i n e h a s been r e t a i n e d i n t h e p r e s e n t s t u d y .

3.2.2 T r a n s m i s s i o n and Rotor Design The t r a n s m i s s i o n s and r o t o r s were s t r u c t u r a l l y d e s i g n e d by t h e maxiinurn r a t e d horsepower o f t h e e n g i n e a t t h e hover rpm. That i s , no t r a n s m i s s i o n t o r q u e l i m i t s were a p p l i e d a t h o v e r rpm b u t power was l i m i t e d t o 70% of s e a l e v e l maximum a t c r u i s e rpm.

The r o t o r s c o n s i d e r e d I n t h e s t u d y were assumed t o b e of t h e same h i n g e l e s a d e s i g n as t h e r o t o r d e f i n e d f o r t h e T i l t Rotor Research A i r c r a f t i n NASA CR-114438, " P r e l i m i n a r y Design of Research A i r c r a f t " , R e f e r e n c e 2. The b l a d e s were assumed t o b e r e c t a n g u l a r i n planform and t o u t i l i z e t h e BV23010-1.58 a i r f o i l o u t b o a r d of t h e b i a d e c u f f . The sane b a s i c d e a i g n and t y p e of c o n s t . r u c t i o n was assumed and t h e aame w e i g h t f a c t o r s were u-sed.

3.2.3 A i r c r a f t Drag A s i m p l i f i e d d r a g model was u s e d f o r s i z i n g t h e d e s i g n p o i n t a i r c r a f t . The model r e p r e s e n t s t h e d r a g o f t h e a i r c r a f t a6 l i n e a r f u n c t i o n s of wing area. The methods o f Booing Document 0 8 - 2 1 9 4 -1, "Drag E s t i m a t i o n of V/STOL A i r c r a f t n , R e f e r e n c e 7 , D222-10060-2 were used t o c a l c u l a t e t h e i n t e r c e p t and s l o p e of t h e t r e n d c u r v e . The d r a g t r e n d used is shown i n F i g u r e 3-2. T h i s c u r v e i s i d e n t i c a l t o t h a t shown f o r t h e SAR a i r c r a f t i n R e f e r e n c e 1.

I n t h e composite wing t r a d e s t u d y p a r a m e t r i c v a r i a t i o n s of wing d r a g w i t h wing chord and t h i c k n e s s r a t i o were computed.

These v a l u e s were t h e n used t o i n c r e m e n t t h e b a s i c f , of t h e b a s e l i n e a i r c r a f t . The p r o c e d u r e and d r a g v a l u e s used a r e d i s c u s s e d i n m o r j d e t a i l i n S e c t i o n 3.4.1.

3 . 2 . 4 C r i t e r i a f ~ r S e l e c t i n g Design P o i n t A i r c r a f t The d e s i g n p o i n t a i r c r a f t were s i z e d t o t h e m 4 s s i o n r e q u i r e - ments d i s c u s s e d i n S e c t i o n 3.2.1. A d d i t i o n a l l y , t h e f o l l o w i n g d e s i g n c o n s t r a i n t s were imposed: Thrust-weight r a t i o c a p a b i l i t y a t t h e m i s s i o n 1.

mid-point of a t l e a s t 1.1 w i t h s e v e n ( 7 ) r e s c u e e s 2 . Maximum hover d i s k i o a d i n g of 1 5 p s f R o t r r s o l i d i t y g r e a t e r i h a n . 0 5 8 3 .

4 . Wing c h o r d t o r o t o r d i a m e t e r r a t i o of 0 . 2 I n g e n e r a l , t h e s e c o n s t r a i n t s a r e t h e r e s u l t of p r a c t i c a l con- s i d e r a t i o n s i n t h e d e s i g n of tilt r o t o r a i r c r a f t , The d i s k l o a d i n g l i m i t , f o r example, wag imposed t o a v o i d e x c e s s i v e downwaeh v e l o c i t i e s i n hover. Downwash v e l o c i t y i s d i r e c t l y r e l a t e d t o d i s k l o a d i n g . A t n i g h d i s k l o a d i n g s , t h e r e s u l t i n g h i g h downwaoh v e l o c i t i e s would terid t o hamper rescae o p e r a t i m e .

EQUIVALENT FLAT PLATE DRAG AREA F T 2 D222-10060-2 A itidximum thrust coefficient-to-solidity ratio, C~/o=0.135, was used, based rjfi stall ilutter cansiderations. !'.owever, in no event was the solidity permitted to go below a value of 0.058.

The rotor solidity limit is based on practical design and manu- facturing considerations related to blade torsional and flap- ping stiffness requiramcnts. As rstor blades become narrower and thinner at the lower solidities it becomes more and more difficult to tune them and still meet design fatigue life requirements.

The chord-diameter ratio value used is a nominal value selec- ted on the basis 2f previous design experience. It has been found that C/3=0.2 qives wing aspect ratios that provide ade- quate control of the static divergence and whirl flutter modes withoat excessive weight and performance penalties. The effect of chord-diameter ratio on mlssion performance has been investi- gated in this study and is discussed i n Section 3.4.2.

Fixing chord-diameter ratio fixes wing configuration because span has also been specified as a function of diameter. Thus wing loading is a function of disk loading and rotor diameter becomes the design parameter. TI- procedure for sizing the design point aircraft then become a matter of sizing aircraft for a series of rbtor diameters and determining the mininun weight configuration corresponding to the moat critical of the firsi three design constraints.

D222-10060-2 3.2.5 Wing Structural Design Criteria Wing structural design criteria for the study are based on t h w e established for the Model 222 tilt rotor research air- craft in Reference 5. In general the same loading criteria were applied except that the limit load factor was reduced to 2.67 to match that used in the Reference 1 design study. Stress allowables used are based on current doeing practice for com- posite materials.

3.2.6 Composite Weight Factors Analytical stcdies, complemented by actual hardware development, have established a 30 percent weight reduction potential for advaaced composite material. A survey paper, "Weight Predic- tion Techniques and Trends for Composite Material Structure", presented at the 30th annual SAWE me?ting in 1971 (Reference 6) identified 21 aerospace st::uctural components made from Further research was done to identify the advanced composites.

actual weight savings achieved compared to that predicted by the various analytical studies. The following table is re- produced from Reference 6.

Based on this analysis a weight reduction factor of 30% for a ccrnposite wing was used in this study. This reduction was agreed upon by the Air Force early in the study program.

Weight Component Material Saving hemarks C o n t r a c t o r - - F-4 Runder I n c l u d i n g Balance McDonnell A i r c r a f t Company Weights MCAIR IRAD S t a b i l a t o r 3 3 % S a v i n g s on McDonnell A i r c r a f t Company Refined Design

i F-111 S t a b i l a t o r 3 - P l y Minimum G e n e r a l Dynamit:~ Convair

Gage Aerospace-Fort Worth O p e r a t i o n F-111 Wing T i p U n p r e s s u r i z e d / Grumman A i r c r a f t E n g i n e e r i n g P r e s s u r i z e d C o r p o r a t i o n F-5 Landing G e a r D o o r Northrop C o r p o r a t i o n , N o r a i r D i v .

F-100D Wing S k i n s North Arrerican P.ockw'el1-LAD F-14 S t a b i l a t o r Grumman A i r c r a f t E n g i n e e r i n g C o r ~ A-4 F l a p Douglas A i r c r a f t Company A-4 F l a p Douglas A i l c r a f t Com,?any A - 4 s t a b i l a t o r Douglas A i r c r a f t Com:?any I n c l u d i n y End Royal A i r c r a f t E s t a b l i s h m e n t u ' V C - ~ G A i l e r o n S t r u t I F i t t i n g s I-' o C-5A Leadina Edge S l a t Lockheed Georgia Company T-39A Wing Box S e c t i o n North American Rockwell-LAD Advanced Composite Ning G e n e r a l Dynamics Convair Aero- S t r u c t u r e s p a c e - F o r t Worth Opera t i o n F-5 Leading Edge S e c t i o n Northrop C o r p o r a t i o n , N o r a i r Div.

F-111 F u s e l a g e G e n e r a l Dynamics Convair A e r o - s p a c e - F o r t Worth O p e r a t i o n Graphi te/Epoxy Boron/Aluminum CH-47 Rotor Blade Boron/Epoxy N/A Boeing V e r t o 1 S-61 T a i l Rotor BoronGlass/ S i k o r s k y A i r c r a f t EPOXY ICBM Reentry V e h i c l e G e n e r a l Electric Boron/Epoxy N/A North A r n e r i c a n Rockwell Space D i l Tubular S t r u t s Boron/Epoxy 3 0 % I n c l u d i n g End F i t t i n g s Missile Payload Adapter Boron/Aluminum 3 0 % G e n e r a l Dynamics Convair Aero- space-San Diego O p e r a t i o n TABLE 3-1 COMPOSITE HARDWARE DEVELOPMENT 3.3 ALL METAL AIRCRAFT 3.3.1 Description To provice a baseline agiinst which to measure the benefits obtainable with compositer,, the SAR a~rcraft of Reference I was resized to an all-ntet.il configuration. This was necessary because that aircraft had solle composite materials in it.

The parametric sizing results are shown in Figure 3-3. The data shown ase: disk loadi~g, midpoint thrust-weight ratio, and gross weight. As n o : e d , the aircraft is sized by ihe midpoint hover requirement. This gave a design gross weight of 18025 lb at a disk loading of 13.8 pef. The charasteristics of the aircraft qrz summarized in the following table.

Design Point All-Metal Aircraft characteristics Gross Weight (lb) Weight Empty (lb) Aspect Ratio Wing Area @t2) Wing Span (Ft) Wing Chord (Ft) Tapez Ratio ( A ) /Sweepback ( A ) Wing Thickness ( 8 ) Wing Loading ( ~ b / ~ t ~ ) Rotor Diameter (Ft) C h ~ r d to Diameter Ratio Rotor Solidity Ratio Disk Loading ( ~ b / ~ t ~ 1 Design CT/a Power Plant Rated Power @ SL/STD (SHP) A summary weight statement for the aircraft is presented in Table 3-2, 3.3.2 Performance The performance characteristics cf the all-metal aircraft are LIMIT DISK LOADING psF l2

i 10

MIDPOINT lW2 T

THRUST-WEIGHT RAT I0 MINIMUM T/W WITH 1.0 SEVEN RESCUEES MIDPOINT THRUST-WEIGHT ABOARD RATIO

-

i o n n n A - U Y Y Y GROSS WEIGHT -- LB 17000 GROSS WEIGHT ROTOR DIANETER FT F I G U R E 3-3. ALL-mA.L SAR TILT ROTOR PARAMETRIC SIZING RESULTS

TABLE 3-2: SUMMARY WEIGHT STATEMENT USAF -

SAR CURRENT MAT'L. & TECH.

ENG H . P . EA 1950 iiOTi3R DIA/C 2 8 . 9 7 ' 2 8 WING AREA 210 Fr2 ROTOR GROUP 1203 W I N G GROUP 1250 T A l L GROUP

BODY GROUP I

, AUX. POWER P L A N T I N S T R AND N A V . 1 3 5 H Y D R . AND PNEU. 130 L

E L E L T R I C A L GROUP 800 I

t E L E C T R O N I C S GROUP 1400 ARMAMENT GaOUP 17 5 F U R N . G E O U I P . GR?UP 3 50 P E R S O N . ACCOM.

M l S C . E O U I P M E N T -+- P H O T O G R A P H I C I I I I I

Y G€AR +-.__ I --

W E I G H T EMPTY 12380 F I X E D U S E F U L L O A D A (41 860 1 T R A P P E D L I OUI D S I 1 F N G L G O I L 1 1 I 1 I MISSION EQUIP. I 150 I I

FUEL ! 4450

5 I I I I GUN 6 m4MO 14 5 I I 1 I 0222-10060-2 REV. A summarized in Figures 3-4 and 3-5. The data are presented for 18025 lb gross weight and include hover ceiling, flight envel~pe, and climb characteristics.

The aircraft can hover at its design gross weight at about 3600 ft on a hot day (95OF) and over 8000 ft under standard day con- ditions. (Figure 3-4) These data are based on a thrust-

y ,-. ratio of 1.1 which allows 5% margin for download and 5 %

for maneuverability.

Cruise mode performance is summarized in Figure 3-5. The air- craft is capabla of 320 kt at normal power up to 5000 ft and cen exceed 300 KTAS up to 17000 ft. The aircraft has adequate climb performance and has absolute ceilings in excess of 25000 ft.

ALTITUDE

- FT

DESIGN GROSS WEIGHT

HOVER GROSS WEIGHT - LB

NOTES t 1, T/w = 1.1 2, Military Power 3, R o t o r Tip Speed: 750 FPS 4, Design Gross Weight: 18025 LB FIGURE 3-4. ALL-METAL SAR TILT ROTOR OGE HOVER CEILING 3-15 I 'Y Altitude - Ft.

30000 SPEED ENVELOPE S P E C I F I C RANGE 25000 20000 15000 10000 True Airspeed -- Kt.

Specific Range - i

Figure 3-5. All-Metal SAR Tilt Sotor Cruise E Summary MAXIMUM RATE Or' CLIMB

(lo@/, RPM)

Notes : 1. Standard Day 2. VTIp = 525 FPS (70% max.) except as noted 3 . Zero n a c e l l e incidence ( c r u i s e f l i s h t mode)

' i c Rancre - N M P P Maximum Rate of Climb -- FPM

FRAME

D222-10060-2 REV. A 3.4 COMPOSITE WING CONFIGURATI3N STUDIES The first objective of the study was to determine the perform- ance benefits obtainable with composites as affected by wing geometry. his was done by replacing the metal wing of the all- metal aircraft (Section 3.3) with composite wings and varying geometry and then computing the mission performance of the modified aircraft.

Wing geometry was varied in two ways: in the first the base- line wing planform was used and thickness ratio was varied from

159 to 24% - the characteristics of these wings

are tabulated as follows:

chozd - 5.78 ft

- 36.33 ft

span

aspect ratio - 6.29

- 210.1 ft2

area

thdskness - 15 to 2 4 % ~

In the second series, the baseline thickness ratio (21%) was used and chord was varied from 4 to 8 feet. The planform characteristics of t h e m wings are ohown in F?gure 3-6.

The drag of the wing8 was estimate;! using the methods of Reference 7. The resulting drag increments for the composite There values were added to the wings are given in Figure 3-7.

f , of the baseline aircraft.

NOTES : 1. Span = 3 6 . 3 3 Ft.

PLAN FORM 2 . t/c = 21% AREA F T ~ ASPECT *TI0

/'

PmFORM AREA 4

CHORD DIAMET RAT I0 4 5 6 7 8 WING CHORD - : T .

FIGURE 3-6: PLANFORM CHARACTERISTICS OF CONSTANT THICKNESS RATIO WING SERIES DRAG AREA INCREMENT ffe) CHORD VARYING (t/c = 21%)

- CHORD 1.'3NSTANT

1 1 I I IA l 6 I 8 20 22 24

THICKNESS-CHOU RATIO, t/c -

I

5 6 7 8

CHORD - Fr.

FIGURE 3-7: DWAG AREA INCREPIENT BETWEEN COMPOSITE WINGS AND BASELINE ALL-METAL WnTG fi222-10G60-2 REV. A A s t r u c t u r a l a n a l y s i s was done t o oLta.in t h e w e i g h t s of t h e composite wings. T h i s was done s o t h a t t h e w e i g h t s would re- f l e c t t h e e f f e c t s of wing d e s i g n ground r u l e s p a r t i c u l a r l y w i t h r e g a r d t o s t r e n a t h and s t i f f ~ e s s r e q u i r e m e n t s (Sect.:on 3 . 2 . 5 ) . The r e s u l t i n g wing w e i g h t 8 a r e shown i n F i g u r e 3-8.

1; w i l l b e n o t e d t h a t t \ e wings become s t i f f n e s s : . . t i c a l a t t h l c k n e a s e s below 17.15% and chortle below 4 . 5 5 f t i n t h e t h i c k n e s s and chord t r a d e s , r e s p c t i v e l y . A t h i g h e r t h i c k n e s s an: zhord v a l u e s t h e wings a r e s t r e n g t ; . c r i t i c a l .

The VASCOY? program ( R e f e r e n c e 9 ) was u s e d t o compute t h e p e r - formance of t h e a i r c r a f t w i t h t h e d i f f e r e n t wings. T h e s e re- s u l t s a r e p r e s e n t e d i n F i g u r e s 3-3 t h r o u g h 3-12.

The performance b e n e f i t s o b t a i n a b l e w i t h cornposited a r e ex- p r e s s e d i n two Lays: i n terms o f t h e improvement i n m~ximum r e s c u e w e i g h t c a p a b i l i t y and i n t e r m s of t h e a d d i t i o n a l r a d i u s o r m i d p o i n t h o v e r t i m e o b t a i n e d a t a g i v e n r e s c u e c a p a b i l i t y .

I n t h e f i r s t c a s e t a k e o f f g r o s s w e i g h t was r e d u c e d t o t a k e a d v a n t a g e of t h e r e d u c e d w e i g h t a£ t h e c o m p o s i t e wing F u e l r e q u i r e d was computed f o r t h e sAR m i s s i o n p r o f i l e ( S e c t i o r 3 . 2 . 1 ) . (VASCOMP h a s ?, p r o c e d u r e th;.t s o l v e s f o r TOGW when OWE and payload a r e g i v e n . ) he r e d u c e d g r c s s w e i g h t a t t h e m i d p o i n t a l l o w e d a n i n c r e a o e i n t h e maximum r e s c u e w e l g h t .

These b e n e f i t s a r e i n d i c a t e d by t h e cu:ves l a b e l e d " C o n c t a n t M i s s i o n C a p a b i l i t y " i n F i g u r e s 3-9 and 3--11.

\

ALL-METAL SARP

\

TILT ROTOR WING \ .

WEIGHT

- LB

STRENGTH

CRITICAL + CRITICAL

CONSTANT CHORD - C = 5.78 F T

1 1 I I I 1'4 1 ' 6 n'8 2'0 2'2

THICKNESS-CHORD RATIO, t/c - %

WING WF! IGHT

- LB

STIFFNESS CRITICAL CRITICAL CONSTANT T H I C X N E S S - ~ C H O R D W T I Q

- t/c = 21%

1 1

4 5 6 7 --

CHORD - f i .

FIGURE 3-88 COMPOSITE WING WEIGHT FOR WING PiWMETER STUDY D222-10060-2 REV. A In the second case takeoff gross weight was specified as 18025 lb. and the benefits due to the composite wings were put into an increased fuel load. 'his allowed either the mission radius to be increased over the basic 500 NM or the midpoint hover time to be increased over the basic 30 min. The maximum resuce capability in this case was maintained at 1400 lb or seven rescuees. The radius and hover time improvements are indicated by the curves labeled "Const TOGW" in Figures 3-10 and 3-12.

Also shown for refe~ence are the all-metal and resized compo- site wing aircraft described in Sections 3.3 and 3.5, re- spectively, and the design point aircraft obtained by replacing the metal wing with a composite wing.

It w i i ~ be noted in general that the curves reflect the strength

critical - stizfness critical crossover resulting from the struc-

tural analysis. The benef .'.ts due to composites decrease rapidly as thickness or chord is decreased below the crossover because of the rapid increase in wing weight.

- .~.

When constant mission capability is specified, replacing the metal wing with one of composite construction increases the rescue capability by more than 350 lb. With thicknesses greater than 19% or chords greater than 5.4 ft, two additionai men can be picked up with a small margin in capability. The reduction in takeoff gross weight in these cases is about 500 lb.

When takeoff gross weight is fixed the composite wing will give mission radius increases ' 50 miles or more or midpoint hover time increases in the 15 to 20 minute range. Note that these cases are mutually exclusive. The additional fuel can be put into additional range or additional hover endurance but not both. Of course, both radius and endurance could be increased simultaneously, but not to the maxima shown.

The effect of composite construction in the wing is further iliustrated by the design point aircraft shown in Figures 3-9 to 3-12. Replacing the mete1 wing with a composite wing re- duced gross weight by about 500 lb. The reduced wing weight is reflected in empty weight and m~ssion fuel. Resizing the aircraft with composites in the wing gave gross wsight reduc- tion of about 750 lb and decreased the rotor diameter from 28.9 ft to 27.1 ft. These comparisons are discussed in greater detail in Section 3.6.

The all-metal aircraft was used as a basis for the wing struc- tural analysis presented in Section 4. The design gross weight used in the calculations shown reflects just the reduction in empty weight due to the reducticn in wing weight.

Maximum fuel for the all-metal aircraft was used in the struc- tural acalysis.

D 2 2 2 - 1 0 0 6 0 - 2 REV. A TAKE OFF GROSS WEIGHT

---

2o 1 TOTES:

'"zoo 1 , [CONSTANT - 1 8 0 2 5 LB

lLB ALL-METAL T I L T ROTOR 18 V ALL-METAL w/coMP. W ING I

com .WING T/R

I 14 1 6 18 2 0 3 2 24 2 6

THICKNESS-CHORD RATIO, t / c - %

MAXIMUM RESCUE C A P A B I L I W ALE-METAL W/COMP.WING NO. O F

I --

RESCUEES M I S S I O N C A P A B I L I T Y (TOGW VARYING) I /

I

TANT TOGW - 1 8 0 2 5 LB

/ RESCUE

I

- 100 1

ALL-METAL AND RES I ZED

I

COMPOSITE WING T/R I S T I F F N E S S STRENGTH C R I T I C A L +CRITIC&

t

FIGURE 3-9: EFFECT OF WING THICICNESS RATIO (AT CONSTANT CHORD) ON TAKE OFF GROSS WEIGHT AND MAXIMUM RESCUE C A P A B I L I T Y .

D222-10060-2 REV. A MISS102 W I U S (30 M N. HOVER @I M.P.)

CONSTANT TOGW - 18025 LR

r

t

RADIUS

I

- NM

/- I ALL-METAL TILT ROTOR

-. / - - I -

I

-t CONSTANT MISSION CAPABILITY

I (TOGW VARYING ) STIFFNESS %TmNGTH

400 .- CRIT ICAL CRITICAL

TFIICKNESS-CHORD RATIO, C / C - "/, HOVER ENDURANCE AT MIDPOINT

-

(500 NM R A D I U S ~

CONSTANT TOGW - 18025 LB

?

- -

CONSTANT MISSION CAPABILITY (TOGW VARYING ) CRITICAL CRITICAL

L . 1

I 14 16 18 10 1 2 24

THICKNESS-CHORD RATIO, t/c - %

FIGURE 3-10: EFFECT OF WING THICKNESS RATIO (AT CONSTANT CHORD) ON MISSION RADIUS AND MIDPOINT HOVER ENDURANCE D 2 2 2 - 1 0 0 6 0 - 2 REV. A M I S S ION RADIUS NOTE: t/c = 21%

COYSTANT TOGW - 1 8 0 2 5 LB

T I L T ROTOR

- --

CONSTANT M I S S I O N C A P A B I L I T Y

I

(TOGW VARYING ) S T I F F N E S S C R I T I C A L I 1 I I 1 I 3 4 5 6 7 8 9

CHORD - FT

HOVER ENDURANCE AT MIDPOIJTI' - - -.

HOVER ENDUR- ANCE

- MIN

ALL-METAL T I L T ROTOR

' - + , -

I CONSTANT M I S S I O N C A P A B I L I T Y

I (TOGW VARYING)

S T I F F N E S S ' STRENGTH

20 C R I T I C A L -+* C R I T I C A L

I I 4 5 7 8 9 6 FIGURE 3-12 : EFFECT O F WING CHORD (AT CONSTANT THICKNESS R A T J 9 ) ON M I S S I O N RAPI','5 AND MIDPOINT HOVER ENDURANCE D222-10060-2 .- - REV. A

3.5 RESIZED COMPOSITE WING AIRCRAFT

3.5.1 Description Another way in which the benefits resulting from composite construction can be shown is in their effect on overall air- craft size and weight. To obtain these results the all-metal SAR tilt rotor (Section 3.3) was rasized with the wing weight coefficient reduced by 3 0 % to reflect composite construction.

The 3 0 % reduction factor has been agreed upon with USAF as being the weight saving obtainable with the use of composites (Section 3.2.6).

The parametric sizing results are shown in Figure 3-13. The data shown are: disk loading, midpoint thrust-weight ratio, and gross weight. In this instance the aircraft is sized by the disk loading limit (W/A=15 psf) and is just over the thrust-weight ratio requirement (T/W=1.1 with 7 rescuees). The aircraft therefore has nearly matched hover and cruise power requirements.

The composite wing tilt rotor aircraft has a design gross weight of 17242 pounds and 27.1 ft diameter rotors. The design . ..

character-istics-of the dircraft are summarized as follows : Gross Weight (lb) 17242 Weight Empty (lb) 11747 Aspect Ratio 6.37 Wing Area (f t2) 187.7 Wing Span (f t ) 34.55 Wing Chord (ft) 5.42 REV. A Taper R a t i o (A)/Sweep Angle ( A ) 1.0/0 Wing Thickness Wing ~ o a d i n g ( p s f ) Rotor Diameter (f t ) Chord t o Diameter R a t i o 0.2 S o l i d i t y R a t i o ,087 Disk Loading ( p s f ) 15.0 Design CT/a .135 Power P l a n t ( 2 ) Lycoming PLY-27 Rated Power @ SL/STD (SHP) 1950 A summary weight s t a t e m e n t f o r t h e a i r c r a f t i s g i v e n i n Table I t i s noted t h a t t h i s a i r c r a f t i s v e r y n e a r l y t h e same s i z e a s t h e tilt r o t o r SAR a i r c r a f t d o s c r i b e d i n Reference 1. That a i r c r a f t had a d e s i g n g r o s s weight of 16970 l b , an empty weight of 11500 l b , and a r o t o r diameter of 27 f t . Although t h e weight r e d u c t i o n f a c t o r used t o a.ccount f o r composites i s l a r g e r , t h e composite wing a i r c r a f t of t h i s s t u d y i s h e a v i e r because t h e f u s e l a g e and empennage weights do n o t i n c l u d e composites. I n a d d i t i o n d e t a i i d e s i g n s t u d i e s have i n d i c a t e d t h a t a c t u a l wing weights a r e g r e a t e r t h a n t h o s e i n d i c a t e d by the weight t r e n d s o r i g i n a l l y used. The wing weight t r e n d s used i n t h i s study t h e r e f o r e r e f l e c t t h e r e s u l t s of f u r t h e r e t u d i e s i n t h e d e s i g n of t h e wing.

D 2 2 2 - 1 0 0 6 0 - 2 REV. A DISK LOADING DISK LOADING LIMIT MIDPOINT THRUST-WEIGHT RATIO 1.2 MIDPOINT THRUST-WEIGHT RATIO 1 . 0 SEVEN RESCVEES ABOARD DESIGN GROSS WEIGHT

ROTOR DIAMETER - FT

FIGURE 3-13 : COMPOSITE WING SAR TILT ROTOR PARAMETRIC S IZ ING RESULTS V E ~ T O L D I V I ~ I O W t I I I I K I I, I I Y , - - & L a.Vb , , n A T t i u -,,. REV. A

- -

TABLE 3-3 : SUMMARY WEIGHT STATEMENT - R E S I Z E D

COMPOS ITE WING AIRCRAFT ENG RATING @ SL&D R o T 9 R D I A / ~ NING AREA ROTOR GROUP W I N G GROUP

----

T A I L GROUP BODY-- B A q I C SECONDARY SECOND. - D O O R S . E TC.

A L I G H T I N G GEAR F L l G H T P h - 1 ON - - - P R O P U L S I ON GROUP . - -_ E N G I N C S L S ) --- A I R I N D U C T I O N E X t i A U S T S Y S T E M L U R R l C A T l NG S Y S T F h l . . .-a ---.

c u p - - + . - F U E L S Y S T E M P R O D L c L E R A T .

-- I t - -

-

A L X , P(r*ER F L A N T _ . -- t - - - - - INSTR, A N D N A V . 1 3 5 I

HYDP. AND PNCU. 1 3 0 1

E L E C T R I C A L GROJP E l ECTRON I C S GROUP ARMAMENT ~ ' I O U P FURN. t1 L Q U I P . G R O U P 3 5 0 PERSON. ACCOM. .-- ' d l sc. E O l l I ~ ~ y ~ ~ . - - . ~ t- I - F U H - 5 --- - Eh:ERG. E O k I PS'EhT - PHOTOGR A P H I C 1 1 0 1 ~ ---- ~ . - - - - - W U U N U L t L A R -.

W E I G H T E M P T Y F I X E D U S E F U L L V A D

I b

, CARGO 1

- . l : L = + - -- --

P A F F - GUN & A M I U 145 I G R O S S W E I G H T 1 7 2 4 2 fi 3.5.2 Performance The performance characte~istics of the rosized composite wing aircraft are summarized in Figures 3-14 and 3-i5. The data, presented for 17242 lb gross weight, inciude hover ceiling, flight envel~pe, and climb characteristics.

The aircraft can hover at design gross weight at 3600 ft c-1 a 9S°F d; 1 and at 8300 ft on a standard day (Figure 3-14). This performance is based on a thrust to weight ratio of 1.1 which allows 5% margin for download and 5% for maneuvarability.

Cruise mode performance is summarized in Figure 3-15. Tha aircraft has a sea level normal power speed of 326 kt. and can exceed 300 ::TAS up to almost 19000 ft. Climb performance is good with absolute ceilings in excese of 25000 ft.

D222-10360-2 REV. A ALTITUDE

- rn

SllANDARD DAY

HOVER G R O S S WEIGHT - LB

MOTES : 1. T/W = 1.1 2 . Militarv Power

3. Rotor Tip Speed - 750 E S

". Design Gross Weight = 17242 LB

FIGURE 3-14: COMPOSITE WING SAR TILT hOTOR OGE HOLLR CEILING Jtitude -- Ft.

30304, SPEED EN'.:'ELOPE S P E C I F I C RANGE - - - True Airspeed -- Kt.

S p e c i f i c Range -- N W P

Figure 3-15. C w p o s i t e Wing SAR T i l t R. . . . C ~

Performance S o u n a r y D222-10060-2 REV. A 'ECIFIC RANGE MAXIMUM RATE OF CLIMB

(lop/, RPM)

Notes : 1. Standard Day 2. VTIp = 525 FPS (7@4 max.) except as noted 3 . Zero n a c e l l e incidence ( c r u i s e f l i g h t mode) 4 . l7,242 l b . GW Maximum Rate of Climb -- FPM ~ l t Rotor Cruist D222-10060-2 REV. A

3.6 DESIGN BENEFITS OBTAINABLE WITH COMPOSITES -

From a configuration design point of view the chief benefits resulting from the use of composites are the reductions in air- craft size and weight that can be obtained. From a structural .,,.I ccmpzsites sffer supsri~r corrosion resis- design poiat zf **: tance, greater fatigue strength a1.d reduced notch sensitivity (hence greater damage tolerance). (These are discussed in greater detail in Section 4.3).

Three separate design point aircraft are shown in Figures 3-9 and 3-11. These are the all-metal aircraft, the all-metal aircraft with composite wing, and the reslzed composite wing aircraft. The first twc are identizal except for wing con- struction. The third has been resized to extend the wing weight benefits iato other components of the 3ircraft (rotors, drive system, otc.). Physical and performance characteristics of the three are summarized for co~nparison in Table 3-4. A weight comparison is given in Table 3-5.

Re2iacing the metal wing of the all-metal aircraft with a composite wing gave a reduction in 3ross weight of 506 lb.

Of LL: - ..--...- L

c A l r a a u u u A A L . 375 I b is a t t r i b l i i - ; ? h i e to the W ~ Z G and ths Resizing with composites rest to a reduction in fuel required.

reduced the physical size of the aircraft as well as its weight.

Rotor diameter, for example, dropped to 27.1 ft rrom 28.9 ft for the all-metal aircraft. Wing area was reduced to 187.7 sq.

D222-10060-2 REV. A f t . The e f f e c t s of t h e composite wing a r e a l s o s e e n i n t h e r e d u c t i o n s i n group w e i g h t s down t h e l i n e f o r t h e r e s i z e d a i r - c r a f t . The t o t a l r e d u c t i o n i n empty w e i g h t between t h e a l l - m e t a l and r e s i z e d a i r c r a f t i s 6 3 3 l b which i s 5.1% of t h e a l l - m e t a l v a l u e . Tne t o t a l r e d x t i o n i n g r o s s w e i g h t is 783 l b o r 4.3%. The t o t a l r e d u c t i o n i n wing w e i g h t i s 4 3 8 l b o r 35% of t h e a l l - m e t a l wing w e i g h t . T h i s i n c l u d e s t h e i n h e r e n t weight r e d u c t i o n due t o composites and t h e e f f e c t s of s m a l l e r s i z e .

The e f f e c t of composite c o n s t r c c t i o n i s a l s o s e e n i n t h e f r a c - tior1 of empty w a i g h t a t L , r i b u t a b l e t o t h e wing. The m e t a l wing i s 1 0 . 1 % of t h e empty weight w h i l e t h e composite wing i s o n l y 6 . 9 % of it. T h i s f a c t o r would h e l p t o o f f s e t t h e i n c r e a s e d c o s t of composite c o n s t r u c t i o n .

D222-10060-2 REV* A TABLE 3-4. DESIGN POINT AIRCRAFT COMPARISON All-Metal Resized All-Metal w/Composite Comp.Wing Physical characteristics Aircraft Wing Aircraft Gross Weight (lb) 18025 17519 17242 Empty Weight (lb) Wing Span (ft) Wing Area (sq.ft.1 210.1 210.1 187.7 Wing Thickness 21% 21% 21% 85.8 83.4 91.8 Wing Loading 9psf) Rotor Diameter (f t) 28.9 28.9 27.1 Disk Loading (psf) 13.8 13.4 15.0 Max. Hover Gross Weight: (lb) SL/STD 22870 22870 21850 Forward Flight Performance (SLJSTD, Design Gross Weight) Max. Speed (Mil Pwr) (kt) 342 342 348 Best Range Speed (kt) 225 222 228 Specific Range @ VBR- ,-..4PP) ,271 ,279 ,276 Max. Rate of Climb (fpm) 4010 4120 4050 D l 1 1 r M O D t L N O .

.

TABLE 3-5: D E S I G N P O I N T A I R C R A F T WEIGHT COMPARISON 28.9/. 0 8 28.9/. 08 2 7 . u.087 LL-METX ALL- A S E L I N E A/C

I R C R A F T m1m

h

ROTOR GROUP W I N G GPOUP - T A I L GROUP P t i A S I C SECONDARY SECOND.-DOORS. E T C .

A L l GHT l NG GEAR F I I G H T CONTfiQL S - ENGINE- P R O P U L S I O N GROUP - E V G I N E S ( S ) A I R I N D U C T 1 ON--- L I t B R I C A T I NC S Y S T C h l F U E L S Y S T E h l P R O P t L L E R I N S T . 1 1427 1427 ! 1355

I

' D R I V F S Y S T E M I

. I , I I

I

AUX. POWER P L A N T I- \

I

I N S T R . A h D N A V . 3 5 1

H Y D P . AND PNEU. b 3 0 \ -

E L E C T R I C A L GROLJP 0 0 1 i

E L E C T R O N I C S GRO " P b4001

A R M A M E N T GROUP h 7 5 I F U R Y , L E Q U I P . GROUP D50 ) 3 2 0 @ 3200 3200 P T R S O N . ACCOM. I I 1 1 I hl l SC. EQU I-PVENT 1- B B I 1 I I

W E I G H T EMPTY 12380 12005 11747 /

, I F l X E D USE1 d L L O A D C R f W (4 ) T R A P P E D L I Q U I D S MISS ION E Q U I P .

OP S I 1 1 I

GUN & AMMO - +145 - j 145 1 145 1

G R O S S b~ I G H T 18025 1 7 5 1 9 17242 4.0 ADVANCED DESIGN STUDIES This section presents the design and s t r e s s analysis of a composite wing torque box f o r a tilt r o t o r a i r c r a f t f o r t h e USAF-SAR r o l e .

TWO concepts f o r the wing torque box configuration were inves- tigated, namely, a multi-spar and a two-spar torque box. For reasons discussed i n Section 4.4, the two-spar configuration has been chosen f o r the torque box, a s shown i n Figures 4.1 and 4 . 2 . A well i n t h e upper surface provides space f o r cross shafting. The torque-box s h e l l is a honeycomb sandwich with boron-epoxy facings on a f i b e r g l a s s honeycomb core. ~ 1 1 corners a r e gusseted using Xv25i-S g l a s s cross ply t o provide shear t r a n s f e r c a p a b i l i t y and increase s t a b i l i t y of the skin panels.

Although graphite-epoxy construction would r e s u l t i n a s l i g h t decrease i n weight, boron epc..; was selected f o r t h i s design f o r its superior impact r e s i s t a n c e over graphite. This w i l l provide the rugc dness required under normal service condi- t i o n s and reduce maintenance costs.

The estimated weight is 626 lbs. f o r the boron torque box.

Total weight f o r the composite wing is estimated a t 875 lbs.

An equivalent all-metal wing w i l l weigh approximately 1250 lbs.

D222-10060-2 Thus, the composite design represents a weight saving of about 375 lbs. or 30"k of the metal wing weight.

A summary of margins of safety is shawn in Table 4-1.

TABLE 4-1. COMPOSITE WING TORQUE BOX - bIARGINS OF SAFEW Element Locat i o n P r i n c i p a l Margin o f (winq S t a . ) Load Condition S a f e t y

Upper Cover 30 Compn. + Shear

0.02 Forward 180 Shear 0.33

Compn. + Shear

Upper Cover 30 0.03 A f t

180 Compn. + Shzar 0.34

Tension + Shear

Lower Cover 30 .01 Forward I

180 Tension + Shear -18

L o w e r Cover 30 Compn. + Shear

.07 A f t

180 compn. + Shear .32

F r o n t S p a r 30 Shear .28 180 Shear .23 Rear Spar 30 Shear .28 180 Shear - 5 4 . 1 STRUCTURAL D E S I G N REQUIREMENTS 4.1.1 B a s i c Data G. W t . W = 17,650 Lbs.

Wing Span b = 3 6 . 3 F t .

Wing Chord c = 5.78 F t . ( C o n s t a n t ) Thickness R a t i o = 0 . 2 1 ~ F r o n t Spar a t 0 . 1 5 ~ Rear Spar a t 0 . 7 5 ~ Weight 9f T i l t i n g and Fixed Nacelles = 2260 Lbs./Side U l t i m a t e Load F a c t o r = 4.09 F u e l ( A l l i n Wing) = 4450 L b s .

Wing Root Attachment a t W.S. 30 4.1.2 C r i t i c a l Design Condition O Based on 51adei 222 s t r u c t u r a l a n a l y s i s , d e s i g n wing t o r q u e

box t o l o a d s f o r f l i g h t c o n d i t i o n 1 - vTO a t 49 u l t i m a t e

O Check lower s k i n f o r compression l o a d s d u r i n g l a n d i n g and ground t a x i o p e r a t i o n s 4.2 LOAD DISTRIBUTION 4.2.1 Wing Mass Distribution 10% Fuel 15 Lbs.

r-7

I -I-- 1 I Structure

1 50 100 150 200 1

Wing Sta. In.

I 'W/S 217.8

w/S 30

E Rotor

4.2.2 Rotor Loads 4.0g Condition (Ultimate) a) 100% Fuel G.W. = 17,650 Lbs.

Rotor Download = 5% Rotor Thrust . * . Thrust T = 17650 1

- - x - x 4 = 37200 Lbs. (Ult.)

2 .95 b) 10% Fuel G . W . = 17650 -4450 +445 = 13645 Lbs.

. * . Thrust T = 13645 1

- x - x 4 = 28750 Lbs. (Ult.)

2 .95 Assume Wing Torsion = 500000 In.-Lbr. (Ult.)

LSpanwise R M and Shear Distribution 4 . 2 . 3 RM a t W/S X,

3 0 < X < 180, is g i v e n by

( K = F r a c t i o n o f F u e l ~ e m a i n i n g ) b The bending moments and s h e a r s a l o n g the span a r e computed i n Table 4-2 and shown g r a p h i c a l l y i n Figures 4 . 3 and 4 . 4 respectively.

Table 4-2 USAF SAR T i l t Rotor Aircraft Spanwise Rendin3 Moment and Shear Di.

Cond. 1 (m) 4 ' g 1

- -

1 K = Fraction of F u e l Rema li.lg 1.00 2 T = Rotor Thrust X Wing S t a . I n .

2 1 7 . 8 - @

I n .

1 - .000114 @

In. Lb.

0 0 0

9200 @ In. a.

200 - 9

In.

2 @2 In. Lb.

a3/150 In. Lb.

180 - @

I n .

30 00'

In. Lb.

@ + @ + @ + @

In. Lb.

In. Lb.

-

Lb .

Lb.

Lb .

Lb.

Lb .

Lb .

10% Fuel Aircraft Shear Distribution #'g' 20 40 60 80 100 120 140 160 180 200 220

P

E A/C WING STA. INCH

FIGUBF 4 2 . SPANWISE SHEFR DISTRIBUTION COND. 1, 4 ' g ' VTO 0 20 40 60 80 100 120 140 160 180 200 2 I

g A/C WING STA. - INCH E ROTORN

FIGURE 4-4, SPANWISE BM DISTRIBUTION COND. 1, 4'g' VTO D222-10060-2 4.2.4 E s t i m a t e of S t i f f n e s s Requirements rt is assumed t h a t t h e s t i f f n e s s d i s t r i b u t i o n should match Model 222 wing s t i f f n e s s and :hat t h e wins f r e q u e n c i e s should be t h e same a s f o r Model 222 wing designed f o r a n u l t i m a t e \

SAR Wing Bending S t i f f n e s s - - 4.2' '

Model 222 Wing Bending S t i f f n e s s .3712 X 10.32 X lo6 and SAR Wing T o r s i o n a l S t i f f n e s s Model 222 Wing T o r s i o n a l S t i f f n e s s The r e s u l t i n q E I and GJ d i s t r i b u t i o n a r e shown i n F i g u r e 4.5.

WING STA. - INCH FIGURE 4 3 . REQUIRED WING STIFFNESS DISTRIBUTIOJS 4 . 3 MATERIAL SELECTION 4.3.1 FILAMENT-MATRIX SYSTEMS The selection of the basic composite fiber has a major impact on the ~verall cost and performance of the system. Four basic filament-matrix systems (two of which are stats-of-the-art and the others considered advanced) were evaluated for application in the wing structure.

Representative values of the basic material properties are presented in Table 4-3. The values shown are design allowables, .L, % statistically reduced, based on component fatigue experience and extensive coupon testing. For graphitt epoxies, consid- erable data are currently being generated in support of the HLH program.

Design allowables for composite materials are based on tests conducted under Army and Air Force sponsorship, as well as Boeing research. The desig,~ properties are derived from over 1,000 tests of boron/epoxy compositee, 3,003 tests of glass/ epoxy compositee, and 350 tests of mixtures of glass m d high-.nodulw. ecmpssites. The data fnolude effects of notches, temperature, humidity, load eequencing, effect o f mean load, D222-10060-2 TABLE 4-3. PRELIMINARY DESIGN ALLOWABLES ' I q n - - Material L U U L - s Gra- XP-251-S Gra- or on/ Glass Glass phite/ phite/ Epoxy E P O X Y E P O X Y HT HM 175 14 3 F tu O0

(ks i) - + 4S0

28.2 10 90 O 2.98 7.5 126 90 F cu O0

(ksi) - + 45" 31

90 O 25 O0 7.1 8 F su

(ksi) - + 4 S 0 27 24

90° 30.1 E 0° 7.15

( p s i x - + 4 s 0 1.8

90 O 1.74 .058 .010 Thickness P e r Cured .010 Lap (In.)

D222-1006C-2 and f a i l u r e modes. Although d d t a r e g a r d i n g m a t e r i a l proper- t i e s a r e a v a i l a b l e f o r t h e most p a r t , t h e r e a r e gaps which l i m i t t h e a p p l i c a t i o n of t h e s e f i l a m e n t - m a t r i x systems.

The above f o u r b a s i c f ilament-matrix systems were s e l e c t e d because of t h e i r range of c o s t , s t r e n g t h , s t i f f n e s s , e s t a b l i s h e d performance c o n f i d e n c e , and r e l a t e d e x p e r i e n c e e x i s t i n g w i t h i n t h e Boeing V e r t o l . E-glass and S-glass have been used f o r s e v e r a l y e a r s and t h e i r b a s i c p r o p e r t i e s a r e g e n e r a l l y w e l l - known.

Since most of t h e composite m a t e r i a l s a v a i l - a b l e are n o n m e t a l l i c , t h e i r s u s c e p t i b i l i t y t o c o r r o s i o n as it i s commonfy understood is n e g l i g i b l e , (Galvanic c o r r o s i o n s h o u l d b e c o n s i d e r e d when c e r t a i n composites a r e i n c o n t a c t w i t h metals. E s p e c i a l l y s u s c e p t i b l e i s an aluminum/graphite i n t e r f a c e ) . C u r r e n t l y a v a i l a b l e epoxy m a t r i x systems a r e a l s o h i g h l y r e s i s t a n t t o e w i r o n m e n t a l e f f e c t s .

T h e p l o t o f e x t e n s i o n a l modulus d i v i d e d by d e n s i t y ( E / p ) and t o r s i o n a l modulus d i v i d e d by d e n s i t y (G/p) i n F i g u r e 4.6 i s an i n d i c a t i o n o f t h e f l e x i b i l i t y a v a i l a b l e t o t h e d e s i g n e r i n a c h i e v i n g a match o f dynamic c h a r a c t e r i s t i c s r e q u i r e d f o r d y n a m i c - c r i t i c a l wings w h i l e a t t h e same time a c h i e v i n g weight savings.

Much of t h e primary s t r u c t u r e of a t y p i c a l m e t a l V/STOL a i r p l a n e is designed by f a t i g u e c o n s i d e r a t i o n s . The h i g h r a t i o o f UNIDIRECTIONAL GRAPHITE

- UNIDIRECTIONAL --

, TORSIONAL M O D L l M DENSITY

Figure 4-6. . Composite Matcrids Provide Design Flexibility

f a t i g u e s t r e n g t h t o u l t i m a t e s t r e n g t h o f advanced c o m p o s i t e s e x h i b i t e d by boron and g r a p h i t e o f f e r s a major a d v a n t a g e i n i n c r e a s i n g t h e f a t i g u e s t r e n g t h o f t h e a i r c r a f t . Not o n l y is w e i g h t s a v e d , b u t reduced maintenance costs a r e a n t i c i - p a t e d due t o a s i g n i f i c a n t r e d u c t i o n of i n - s e r v i c e f a t i g u e problems. The a d v a n t a g e of composite m a t e r i a l s o v e r aluminum f o r f a t i g u e i s shown i n F i g u r e 4.7. A d i s p l a y o f r e l a t i v e w e i g h t s o f f a t i g u e - c r i t i c a l s t r u c t u r e s i s p r e s e n t e d i n F i g u r e 4 . . For a g i v e n d e s i g n l i m i t l o a d f a c t o r , it is e x p e c t e d t h a t most of t h e primary wing s t r u c t u r e can b e d e s i g n e d f o r l i m i t and u l t i m a t e c o n d i t i o n s i f advanced composites a r e used, w h i l e s t i l l p r o v i d i n g a f a t i g u e l i f e i n e x c e s s o f t h a t uscd f o r t h e d e s i g n o f c o r r e s p o n d i n g m e t a l s t r u c t u r e s .

Damage t o l e r a n c e i s an i m p o r t a n t c o n s i d e r a t i o n a l o n g w i t h ' a : P' s p e c i f i c p r o p e r t i e s o f s r r u c t u r a l m a t e r i a l s . The s u p e r i o r f r a c t u r e t o u g h n e s s o f composite m a t e r i a l s r e l a t i v e t o aluminum a l l o y s i s c l e a r l y i n d i c a t e d i n F i g u r e 4.9. It s h o u l d b e n o t e d however t h a t exposed g r a p h i t e / e p o x y s u r f a c e s a r e e x t r e m e l y v u l n e r a b l e t o impact damage u n d e r normal s e r v i c e c o n d i t i o n s .

The boron/epoxy h a s i m p r e s s i v e compressive q u a l i t i e s f o r u s e i n combination w i t h o t h e r a p p r o p r i a t e f i l a m e n t - m a t r i x m a t e r i a l s i n p r i m a r i l y compression l o a d e d e l e m e n t s .

" - - . .

, , . I.. , -7- . . .. - - -. . . . . - - - , Figure 4 4 . Specific Fatigue Properties ALUMINUM TITANIUM STEEL 4340 BORON 8-GWSS 2024 6AL-4V COMPOSITE COMPOSITE 150 KSIprO MATERIAL Relative Welghte of F~tigue-C ritlcal Structures Figure 4 3 .

( : S% z, ;~:.r,"~:;---~mw~

' 1-1,. .U,',.Y I*,.. I!.,* I,,1, z y

""' '1 ; z: g :::=:::;:::: , , * m r . . . . . . , m , . . 1 . ;::::::=.::, Figure 4-9.

Residual Strength of Composites and Aluminum Alloys ' J STIFFNESS/WEIGHT COMPARISON H Figure 4-10. Comparison of Mntcriol Propcrtics A f t e r a c a r e f u l r e v i e w o f t h e a v a i l a b l e d a t a , boron/epoxy was s e l e c t e d a s t h e b a s i c f i l a m e n t m a t r i x f o r u s e i n t h e wing s t r u c t u r e .

4 . 3 . 2 CORE MATERI.,X For sandwich p a n e l s , aluminum, g l a s s , and Nomex ( a n y i o n v a r i a n t ) were examined a s p o s s i b l e c o r e m a t e r i a l s . Aluminum c o r e , i s e x t r e m e l y v a l n e r a h l e t o m a j o r damage by l i g h t n i n g when combined w i t h e i t h e r g r a p h i t e or. boron f a c e s h e e t s .

T h i s s i t u a t i o n can r e a d i l y be a l l e v i a t e d w i t h c u r r e n t d s s i g n approaches, n u t c f f o r t was f o c u s e d on r e p l a c i n g alumini..,~ a s a prime c a n d i d a t e f o r t h e c o r e m a t e r i a l . Nomex i s t h e primary c a n d i d a t e from an e n v i r o n m e n t a l r e s i s t a n c e c o n s i d e r a t i o n ; however, i t s p h e n o l i c b i n d i n g is e x t r e m e l y v u l n e r a b l e t o f u e l exposure. Hence f o r t h i s s t u d y , a g l a s s c o r e h a s been s e l e c t e d .

4 . 4 W I N G DESIGN CONSIDERATIONS 4 . 4 . 1 ENGINEERING APPROACH C u r r e n t manufacturing c a p a b i l i t i e s and p r o c e s s e s i n t h e f i e l d o f advanced composites g i v e t h e d e s i g n e r c o n s i d e r a b l e l a t i t u d e i n a r r i v i n g a t a n a p p r o p r i a t e s t r u c t u r a l c o n f i g u r a t i o n f o r t h e b a s i c l o a d - c a r y i n g e l e m e n t o f t h e wing, t h e t o r q u e box.

I n g e n e r a l , t h e a d v a n t a g e s t r a n s l a t e i n t o fewer p a r t s , a reduced number o f mechdnical f a s t e n e r s , and an a s s o c i a t e d w e i g h t r e d u c t i o n , and t h u s , r e d u c e d m a n u f a c t u r i n g ma3hours.

The wing p r o v i d e s s u p p o x t f o r t h e rotc)r,'transmission/engine co&inat.ion a t its extreme ends. The t o t a l f u e l c a p a c i t y o f t h e tilt r o t o r a i r c r a f t i s c a r r i e d i n t h e wings o u t b o a r d of t h e f u s e l a g e . The p r o p u l s i o n u n i t s a t t h e extreme e n d s o f t h e wing are c o n n e c t e d by a c r o s s s h a f t r u n n i n g t h r o u g h t h e upper c e n t e r p o r t i o n o f t h e t o r q w box.

I n a d d i t i o n t o t h e b a s i c prohlem o f c o n f i g u r i n g t h e w i a 1 g box s e c t i o n , t h e t.ilt r o t o r h a s j o i n t d e s i g n r e q u i r e m e n t s which The e:lcompass b o t h fjxed-wing and rotary-w!nq t e c h n o l o g y .

d e s i ~ n e f f o r t bas f o c u s e d on t h r e e main a r e a s i n t h e wirg: o F a s i 2 s t r u c t u r a l s h e l l ( t o r q a e box) o Wing-fuselage j o i n t o J o i n t . 3 ( h a x d p o i n t s ) Emphasis was d i r e c t e d toward: o Reducina t h e number of p a r t s and t o o l s o Reduction i n machining operatio:s Theze a r e a c h i e v e d respectivc. ly by: D222-10060-2 O A d i s c r e t e application of composite filaments/laminates O Use of adhesives 0 pressure-molding techniques 4.4.2 BASIC SZ'RUCTURAL SHELL (TORQUE BOX) Two concepts f o r the basic wing torque box were considered (see Figure 4.11). One is a four-spar configuration and the other -i two-spar configuration; tb.ese w i l l be r e f e r r e d t o a s concepts A and B r e s p x t i v e l y .

Concept A - The four-spar configuration is oriented toward min-

imizing the number of heat cycles during the manufacturing pro- cess. The primary aim i s t o achieve a co-cured zssembly; i . e . , a one-cycle heat exposure operation. The inclusior. c f r i b s , howet-er, prevents t h i s goal from being attained.

Since the inclusion of r i b s presents not only manufacturing d i f f i c u l t i e s but a l s o design problem, r i b s w i l l be provided oniy a t f l a p hinges and leading-edge umbrella hinges. I n t e r - mediate r i b s t o r e a c t panel crushing loads w i l l be eliminated.

The decision t o eliminate intermediate r i b s was based on: Relatively short span of the wing O dequired panei con.pression strength obtainable for r e l a t i v e l y minor weight penalties Ihsic. Wing 'l'orclut* I;ox Figure 4-11; 7 --.-- D222-10060-2 O No external loads being applied at the intermediate rib locations ~lthough the reduction in the number of curing cycles so as to approach co-curing in the manufacturing process is highly desirable, the concept A configuration has design considera- tions for which extensive development is required: Too Many Access Holes O Access required for inspection at three locations chordwise (one in each spa1 bay) at about 36 inches on center for length of span. Each hole has tc be at least 5 inches in diameter. Providing for lightening holes in the lower portion (below cross shaft channel) of the center two spars or using a trussed configura- tion could eliminate access holes in the forward an2 aft bays, requiring them only in the center bay.

O Access hole atinboard tank end rib has to be big enough to allow fur the installation of a fuel boost O Access cutouts limit area f .r locating chordwise fil- aments, if these are required. The wing is primarily loaded in spanwise bending, spanwise shear, and tor- sion: chordwise loads are negligible.

Ribs a r e d i f f i c u l t t o i n s t a l l , especially the tank end r i b a t the inboard location.

The d e l e t i c n of intermediate r i b s may require S a f f l e s t o reduce f u e l slosh.

I f the center two spars do not have lightening holes o r a r e not of a t r u s s configuration, cutouts a r e needed between t h e forward, center, and a f t bays f o r f u e l drainage and a i r venting. This means interrupting spar chords and webs in t h e spanwise d i r e c t i o n (high a x i a l load i n members attaching t~ the spar could cause peel- ing problems) .

C o n c e ~ t s - The two-spar , m u l t i r i b (25-inch spacing) conf igur-

- .-

' . + .)

a t i m (Figure 4.11), i n general, exhibits: O Ease of assembly O A provision f o r good dimensional control and tolerance washout O Good access i.s provided f o r inspection of s t r u c t u r e and maintenance of systems inside the wing * A l l s t r u c t u r e is used e f f i c i e n t l y and is multipurpose.

Ribs a r e used f o r f u e l b a f f l e s and t o carry s t r u c t u r a l loads, spars carry s t r u c t u r a l loads and serve a s i n t e - g r a l f u e l tank walls, etc.

D222-10060-2 Although the method f o r the design and f a b r i c a t i o n of d e t a i l p a r t s and assembly a r e within t h e s t a t e of the a r t , the incor- poratlon of t h l s technology in a wing design has not y e t been demonstrated on a flightworthy vehicle.

Further development is needed t o : O Reduce t h e number of heat cycles required i n construc- t i o n of d e t a i l p a r t s and assembly O Reduce t h e number of p a r t s requiring hand layup. Use of pulkruded components f o r spar-to-rib and spar-to- panel j o i n t s should be investigated.

Based on t h e above discussion concept B has been chosen a s a conservative approach t o t h e design of t h e composite wing.

4.4.3 WING-FUSELAGE JOINT The wing-fuselage joint depends on the configuration of the Assembly and field replaceability individual components.

requirements virtually eliminate adhesive bonding.

A mechanical fastener design is shown in Figure 4.12. This ..

< I . .

concept is currently being deveioped for Boeing Vertol's Heavy Lift Helicopt .r. Figure 4.13 is a photograph of the fitting and Figure 4.14 a general arrangement drawing. Figure 4.15 shows the HLH application. This method offers high strength capability with minimum weight and relatively simple tooling requirements. The barrel nut installation eliminates eccen- tricities by placing the load path directly on the centerline of the sandwich fuselage structure. Four attachment locations using four bolts, or eight for fail safety, could provide the load paths for all the wing-to-fuselage loads, 4 . 4 . 3 WING-NACELLE INTERFACE A design concept of the wing tip fitting structure is shown in Figure 4.16. This design envisions compression molding of a basic chopped-fiber el,-vant j truss j reinforced with uni- directional tape/fiber eiements, BA~REL NUT Figure 4-12. Mechanical Wing-Fuselage J o i n t Concept Figure 4-15. ' Transmission Support Fitting Concept for H1.B Applicat,bn 4.4.5 HARDPOINTS 4.4.5.1 Tension attachment o f wing t o t h e I n t h e e v e n t t h a t a m u l t i b o l t f u s e l a g e is emplcyed (more e f f i c i e n t l o a d t r a n s f e r i n a hcneycomb s t r u c t u r e ) , s h e a r and t e n s i o n b o l t j o i n t s need t o be i n v e s t i g a t e d .

F i g u r e 4.17 d e p i c t s a t e n s i o n b o l t concept which i s capable of t r a n s m i t t i n g wing bending l o a d s ( c o n v e r t e d t o a x i a l loads-lb.,/in.) f o r b o t h c o n c e n t r a t e d and uniform l o a d p a t h s . The j o i n t c o n f i g u r a t i o n shown is now b e i n g e v a l - u a t e d f o r t h e HLH f o r a major f i e l d s p l i c e . These f i t t i n g s w i l l be exposed t o a f a t i g u e environment i n a d d i t i o n t o e s t a b l i s h i n g an u l t i m a t e l o a d t r a n s f e r c a p a b i l i t y .

4.4.5.2 Shear A s p r e v i o u s l y s t a t e d , m e t a l s a r e f a t i g u e - c r i t i c a l while advanced composites ( g r a p h i t e and boron/epoxy) a r e n o t . The q u a n t i t a t i v e e x t e n t of t h i s advantace h a s y e t t o b e e s t a b - l i s h e d and demonstrated. . I c u r r e n t Boeing-Vertol program has y i e l d e d p r e l i m i n a r y r e s u l t s f o r a s h e a r j o i n t conpept which i n d i c a t e t h a t t h e advantage is ccnaiilerable when comparing a parameters of s t e e l and g r a p h i t e .

- F I L L E R BLOCK

/I /

Figure 4-17 .' Conrrtruction of Final 'I'ension Joint for HLH With Static Tensile Strelv'5 of 14,400 Pounds D222-10066-2 A preliminary test specimen of 0.4-inch thickness fabricated from Hercules 2002T Gr/E (HTS/BP901) and subjected to a fatigue environment has outperformed its matint, ,tee! clevis of equal width and c total thjskness of 0.7 inch. The steel clevis failed at the net tension s~ction. When comparing tx parameters, graphite/epoxy laminate has a weight advantag<> P over the steel in the order of 10.

At present the number of tests is statistically insufficient for determination of joint design allowables. Hnwever, the tests do indicat the magnitude of the impact that the use of advanced colposites will have in fatigue-critical structures.

4 . 5 STRESS ANALYSJS 4 . 5 . 1 SECTION . ?OPERTIES Assume u n i f o r r t h i c k n e s s o f material f o r t h e s h e l l , f u , l y e f f e c t i v e i n b e n d i n g .

S e e F i g u r e 4.18 f o r g e o m e t q .

Y

I n . In.

-

12.5 4.87 5.33 7.5 3 . 5 4 . 7 5 5 . 8 8 22.25 6 . 5 4.10 2.08 2.78 -1.39 22.25 - 4 . 5 7.5 -5.62 12.5 -5.58 4.86 -2 - 4 3 6.95 3.475

-

112.8'1 L . 1 8 1

--

. 1/t = 3447.1 - 112.87 ( 1 , 1 8 1 ) ~

= 3447.1 - 1 4 6 . 1 = 3303.0 I n .

A r e a f o r T o r s i o n = 4 7 1 n - ~ / t = - 4~~ = 4 ( 4 7 1 1 2 = 7060 In.

.

L Q 112.87

FIGURE 4J.8. BASIC GEOMEThJ - WING BOX

D222-10060-2 4.5.1.1 Preliminary Check f o r S t i f f n e s s S t i f f n e s s Required a t Root:

EI = 7960 X l o 6 Lb In2

GJ = 1600 X l o 6 Lb In2

Using the E and G v a l u e s f o r boron assuming X = f r a c t i o n of u n i d i r e c t i o n a l boron

l o 6 I30 2 + 3 (1 - X ) 1 3303t = 7960 X l o 6

i , e . (27 X + 3 ) t = 2.41 . . . . . @

and l o 6 (8.8 (1 - Z ) + 1 (1 ) 1 7860t = 1600 X l o 6

i.e, (8.8 - 7 . 8 1 ) t = .204 . . . . @

Solving above equations y i e l d s X = . 8 4 6 t = .093511 = .079" t o t+45 .0145

-

As one l a y e r of boron is .007" t h i c k and t h e c r o s s p l y h a s t o 90 i n p a i r a , assume 11 p l i e s of u n i and 4 p l i e s a t +MO

-

for p r a c t i c a l design then X = .734; t = .105"

E I = {(27 X .734) + 3 1 (3303) (.105) X l o 6

= 7900 X lo6 Lb. 1n.*

GJ (8.8 - 7.8 X ,734) (7860 X .105) X lo6

= 2540 X lo6 Lb. m e 2

S t i f f n e s s Required a t T i p S t a . 1 8 0 : EI = 6450 x l o 6 ~ b . I n .

6 2 G J = 1600 X 10 Lb. I n .

Proceeding a s b e f o r e 9 p l i e s a t O 0 4 p l i e s a t +45 -

. '. EI = {(27 X .692) + 3 ) (3303) ( .091) X l o 6 = 6520 X l o 6 Lb. In. 2

GJ = C8.8 - ( 7 . 8 X .692)] (7860) ( . 0 9 1 ) x l o 6 =

2430 X lo6 Lb. 1 n e 2 4.5.2 PANEL INSTABILITY Compression Panels The critical axial loading for compression panels with different lay-ups are computed below and shown in ~ i g u r e 4.19: ~ a y u p A n at 0 = 12 = ,667 n at +45 = 4 = .222

-

n a?. 90 = 2 = .111 tf = 0.5 X 18 X .007 = .063" . .

c = core thickness ~ o - ' E , = (30 X .666 + 3 X .222=) 20.67 Lb./~n.

= (30 X .I11 + 3 X .222=) 4 . 0 0 ~b./~n.

1 0 - ~ 0 = ( .666 + .111 + 8 . 8 X .222=) 2.73 Lb./~n.

Assume p, = .4

The allowable load/in. is given by For a rib spacing of 25" K,= 2 . 5 Evaluating as function of b and c b (inch) 10 20 30 40 c = .3 7850 1960 872 491 (inch) .5 20300 5060 2250 1266

Layup B (Layup A + 2 Lails at 9 0 ' )

Evaluating NXCR b (inch) 10 20 N~~~ c = - 3 10980 2740 ( inch)

. 5 28200 7 040

.6 39700 9920 .7 53200 13390 \ \

Layup C (Layup A + 4 Laps at 90°)

3222-10060-2 E v a l u a t i n g NXCR a s before b ( i n c h ) 1 0 2 0 30 4 0 c = . 3 13720 3430 1525 858 ( i n c h ) .5 35000 8 7 5 0 3885 2 188 Shear B u c k l i n g Assume V = 0 . 1 -.

E '

Then FS = -

4 h

For b > a use value of a for b in equation

b (inch) b/a KM = K c = .3; h = .363 10950 3320 3 190 (inch) .5; (inch) .563 26400 8000 7660 .6; .663 36500 11000 10620 9690 .7; .763 48500 14700 14100 12850 Layup B = 11.42 X lo6 ~ b . / ~ n . ~ X FSCR = 28.2 X lo6 K K Values As For Layup A = .14 X lo3 (9X.6+9X.2+67X.2)= 2880 ~b./~n.

ULT

i

L a P P C = 12.78 X lo6 ~ b . / ~ n . ~ h

I

tf = -077

= 4.86 X lo6 K (h)2

9CR \b K Values As For Layup A

%R > % T for c > 0.3 and b ,L 40"

I

5 10 15 20 25 30 35 b - i n .

Figure 4-19. Allowable ~ o a d / ~ n . - Compression Panels

( ~ 1 % Honeycomb Sandwich) 4 . 5 . 3 STRENGTH CHECK Root (Wing S t a . 30)

-

1. Axial Loads BM = 4 . 3 2 X l o 6 In. Lb.

1/t = 3303 1 n O 3 Distance From Front Spar (Inch)

t

YU In.

1 5 - 7 6 9

yL 1n. 6.031 NXU L b . / ~ n . -7560 NXL ~b./~n. 7910 F (Shear Lag 1.35 Factor) Nm ~ b . / ~ n . -10200 NXL Lb./In. 10690 - - - 2 . Shear Loads V = 16200 Lb.

T = 2200,000 In. Lb. ( ~ s s u m e d ) 2A = 942 1nO2 Basic Shear Flow q = g, + q l + ey ~ b . / ~ n .

(without Shear ~ a g ) where q, = shear flow i n c u t s t r u c t u r e q1 = balancing shear flow

ep = shear flw due to torque

4 4 6 Basic Shear Distribution V = 16200 Lb.

X K X Value of 2 x Area

q, Distribution

E F ~ = 250+454-2760+3620-4?40-4450-5860-3230+253+347+16200

= 21124 -21040 = (84) Taking Moments About PT 'A'

M = 295X137+564X57+669X58 - 771X87+1155X348+lu91X105

+1480~152+1126+676~5 = 40400+32150+38800-67200+402000+156900+223500+71000+3480 = 901030 In. Lb.

2AEFF = 137+57+58+87+348+105+152+63+5-70 = 942 1 n a 2 q ( w i t h o u t s h e a r lag e f f e c t s ) = Q o - 9 1 2 ! : ~ ~ V a l u e s for R . H . Wing Shown i n P a r e n t h e s i s Shear Lag E f f e c t s (Average Shear Flows) --- Shear Lag % (Due t o V N o Shear Lag) F a c t o r G - 2 . 7 5 1 . 7 5 1 . 4 3 1 . 2 5 1 . 1 5 1 . 5 2 . 7 5 1 . 5 1 . 2 5 1 - 7 5 Tip (Sta. 180) BM = 1.05 X lo6 In. Lb.

Shear = 27,500 Lb.

Torque = - +500,000 In. Lb. (Assumed)

1. Axial Loads (Geometry and 1/t as at root)

-

Distance From Front Spar (Inch)

I

(R.S.)

10 15 20 30 41.7 Lb./In. -1835 -2315 -740 -2230 -1790 -946

: z Lb./1n. 1 1920 2160 2160 2160 1860

--

. 1 2 , 0 -I

2. Shear Distribution Panel BC A t root select layup 3; c = 0 . 5 " b = 12:0 e f f e c t i v e M . S . = 1

- -1 = 0.02

,982 ~t t i p select layup A; c = 0.5" Panels CD, DE, and EF Layup A will be satisfactory by compariscn with panel BC Panel FG b = 22" effective A t tip select layup A wit3 c = 0.6" M.S. = 1

- = 0.34

.743 Lower Cover To a l l m for taxi conditions, assume design compression load = . 4 X Design Tension Load Case Assume beff = 4 0 " Compression Case Select layup C w i t h c = 0.7" N X~~ ~ C R R~ RS

.'. M,b

At tip; layup A with C = 0.7" Check T e n s i o n C a s e A t root N

= , 1 5 4 ( 1 7 8 X , 5 4 5 + 22 X , 1 8 2 + 1 0 X . 2 7 3 ) X 10'

XALL

L a y u p C i s i n a d e q u a t e o v e r f o r w a r d p o r t i o n - u s e 2

a d d i t i o n a l l a y e r s a t o0 over f o r w a r d 2 0 " t h e n

-

At t i p N = . 1 2 6 ( 1 7 8 X , 6 6 6 + 22 X , 2 2 2 + 1 0 X ,111) X l o 3 X A L ~ = 1 5 5 3 0 l b / i n .

RS = . 7 1

.'. M.S. -

- -1 = ' 1 8

Front Spar A t Root = 2572 l b / i n .

qMAx A s t h e s p a r i s assumed t o b e e f f e c t i v e i n bending, d e s i g n for an average a x i a l l o a d of + 5000 l b / i n .

-

Layup D n @ 0 = 1 2 = .545 = . 1 5 4 ( 3 0 0 X - 5 4 5 + 30 X -364 + 25 X .091)X 10 N ~ u L T ( F u l l y = 27203 l b / i n .

S t a b l e ) Allowing for u l t i m a t e c a p a b i l i t y and e f f e c t i v e modulus e f f e c t s use N = 25000 l b / i n .

' C R w i t h V = 0 . 1 a = 25" b = 12" b/a = . 4 8 K = K M = 3 and M.S.- -1 = 0 . 2 8

-178

A t tip

i

cl = 2 1 2 5 lb/'in.

N = + 1000 l b / i n .

'AV -

For layup A with C = 0 . 7 = 2760 l b / i n .

ALL

NxALL= 24000 l b / i n .

+ '041

RS = . 7 7 Page 4-55 , * . M . S . = 1 - 1 ~ 0 . 2 3

m

Rear S p a r A t root Use l a y u p A w i t h C = . 6 " ALL = 2 7 6 0 l b / i n .

= 1 9 0 0 l b / i n , N~~~~ RC = . I 8 5

.*. M.S. - -1 = 0 . 2 8

- 3 3 7 A t t i p q - = 1 4 5 5 l b / i n .

N = + 6 0 0 l b / i n .

-

*AV Layup E n @ 0 = 6 = . 5 0 n @ + 4 5 = 4 = . 3 3 3

-

n @ 90 = 2 = - 1 6 7 C = 0 . 6 " tf = 0 . 5 X 1 2 X.007 = - 0 4 2 Assume uxy = 0 . 4 Use N = 1 2 0 0 0 l b / i n .

XALL With V = 0 - 1 a = 25" b = 1 2 " b/a = . 4 8 Corner G u s s e t s G u s s e t s a r e r e q u i r e d a t a l l c o r n e r s t o c a r r y t o s s i o n a l s h e a r and a t C , D, E and F t o s t a b i l i z e p a n e l .

Maximum s h e a r flow f o r d e s i g n = 1300 l b s / i n .

w i t h (XP251-'S' Glass) = 48000 p s i X-Ply .*. Av, 't' r e q u i r e d = -027"

Minimum p r a c t i c a l t h i c k n e s s = 4 X ,010 - .040n

-

(2, + 4 5 O l a y e r s e a c h s i d e o f j o i n t )

But for e f f i c i e n t d e s i g n t h e 'AG' f o r t h e g u s s e t s h o u l d b e about aqua1 t o 'AG' o f s h e l l Use p r o p e r t i e s of l a y u p B w i t h o u t c o r e

'AG' s h e l l / i n . = 2.56 X l o 6 X .14

= . 3 5 8 X l o 6 l b / i n .

G f o r XP251-S, X-PLY = 2.4 X l o 6 l b / i n ?

. . tREQ = - '358 = .15 i.e. 15 l a y e r s

2 . 4 For p r a c t i c a l d e i s g n u s e 16 l a y e r s , i.e. -16" t h i c k .

4 . 5 . 4 WING STIFFNESS The e f f e c t i v e v a l u e s of E I and G J are g i v e n by a r e show.1 b e l o w .

EX = 1 8 . 0 1 G = 2 . 2 9 t = .168"

T i p ( ~ t a 180) I

(Note: EX and G i n l o 6 i b / i n 2 u n i t s )

The stiffness parameters at wing Station 30 and Station 180 are computed below.

I I 1 1

Y-Y F r o m t h e t a b u l a t e d d a t a , s t i f f n e s s a t root s e c t i o n ( s t a t i o n 3 0 )

( E I ) X EFFECTIVE = ( 8 7 7 3 . 8 + 1 7 2 . 2 ) X l o 6

( E l ) REQUIRED = 7 9 6 0 X l o 6 LB IN.^

4 X 4 7 1 2

( G J ) EFFECTIVE = -

2 9 6 . 2 7 7 X 1 0 ' ~

= 2 9 9 5 X l o 6 LB I N . 2

( G J ) REQUIRED = 1 6 0 0 LB IN.

S i m i l a r l y a t t i p s e c t i o n ( S t a t i o n 1 8 0 ) ( E I ) X EFFECTIVE = 8 5 3 6 . 5 X !06 LB I N .

( E l ) REQUIRED = 6 4 5 0 X l o 6 LB I N .

4712= 2 6 8 2 ~ 1 0 ~ LB I N .

( G J ) EFFECTIVE

=m

( G J ) REQUIRED = 1 6 0 0 X l o 6 LB I N .

H e n c e , d e s i g n i s s a t i s f a c t o r y .

4 . 6 WEIGHTS SUMMARY - COMPOSITE WING FOR SAR AIRCRAFT

ITEMS WEIGHT (LD)

-

COVERS 415 CORE 50

. JOINTS & GUSSETS 6 7

. RIBS

. ADHESIVES

. NACELLE ATTACHMENT STRUCTURE 8 6

LEADING & TRAILING EDGES, FAIRINGS 1 6 9

. SPLICES, FASTENERS, MISC. 3 0

TOTAL WING WEIGHT PER AIRCRAFT 8 7 5 D222-10060-2 5.0 DEVELOPMENT' kLAN Presente? j p ' h i s s e c t '*in is a d e v e l ~ p m e ~ t p l a n fox t h e i n c o r - tilt r o t o r p o r a t i o n o f a composite wing box on an e x i s t i n g NASA r e s e a r c h a i r c r a f t . Cost and s c h e d u l e s w e r e d w e l o p e d on t h e b a s i s o f modifying an e x i s t i n g a i r c r a f t followi?g completion o f i t s f l i g h t test program.

5.1 D E S I G N he c o n c e p t u a l d e s i g n o f t h e main s p a r t o r q u e box would be i d e n t i c a l t o t h a t described i n t h t preceding s e c t i o n f o r t h e SAR a i r c r a f t . For t h e r e s e a r c h ~ i r c r a f t , t h e wing could b e r e s i z e d t o meet t h e e x i s t i n g NASA a i r c r a f t , and t o minimize c o s t o n l y t h e main t o r q u e box woulfl b e b u i l t o f composites.

The e x i s t i n g metal a u x i l i a r y s u r f a c e s would be a t t a c h e d t o t h e composite s p a r . F a b r i c a t i n g a u x i l i a r y s u r f a c e s from com- p o s i t e s h a s been demonstrated on p r e v i o u s p r o j e c t s s o t h a t demonstration o f t h e s p a r box i t s e l f w o u 3 b e t h e prime o b j e c - t i v e of t h e program.

The d e s i g n e f f o r t would mainly c o n s i s t o f t h e e s t a b l i s h m e n t of a f i n a l d e s i g n f o r t h e composite wing box wl,h p r o v i s i o n s f o r i n t e r f a c i n g systems and a t t a c h e d conponents (i.e., f l a p e r o n s , s p o i l e r s , e t c . ) and f u s e l a g e attachment from an s x i s t i n g tilt r o t o r a i r c r a f t .

D222-10060-2 The following <FE components from t h e e x i s t i n g tilt r o t o r a i r c r a f t , assumed t o be a v a i l a b l e i n t h e 1978 time frame, would be i n s t a l l e d i n t h e wlng;

. Nacelles

. S h a f t i n g

. Surface c o n t r o l s - L . E . umbrellas, f laperons , s p o i l e r s

. Controls

. Transmissions

. T i l t mechcnism

5 . 2 FABH ICAT ION AND ASSEMBLY The manufacturing e f f o r t i n c l u d e s t h e f a b r i c a t i o n of t e s t specimens, attachment f i t t i n g s , t o o l i n g , a t o o l proving wing box and a f l i g h t wing box.

The Government supplied hardware from an assumed e x i s t i n g tilt r o t o r a i r c r a f t would be i n s t a l l e d i n t h e wing and fix,al wing assembly would be accompliahed a t t h e c o n t r a c t o r ' s f z - i l i t y .

Following t h e completion of t h e f l i g h t wing assembly w i t h reseaxh instrumentat ion i n s t a l l e d ar'd c a l i b r a t e d , it w i ; l be shipped t o NASA f o r i n s t a l l a t i o n on the e x i s t i n g tilt r o t o r r e s e a r c h a i r c r a f t .

5.3 TESTS AND EVALUATION 5.3.1 Bench T e s t s The bench test e f f o r t i n c l u d e s t h e d e s i g n and f a b r i c a t i o c o f test f i x t u r e s , i n s t r u m e n t a t i c n and c a l i b r a t i o n o f t e s t s p e c i - mens, p r f o r r n a n c e o f tests and p r e p a r a t i o n o f test r e p o r t s .

~ h s type o f tests planned are:

. Coupons - c r a c k pi,opagation and Latigue

P a n e l s - compression and s h e a r

. J o i n t s - t e n s i o n and s h e z r

. Adhesive c o m p a t i b i l i t y

. Wing s e c t i o n - u l t i m a t e load

F u l l s c a l e wing - proof and dynamic shake

. Wing r o o t attachment - proof

. Tool ,-oving

. Environmental

5.3.2 Ground and F l i g h t T e s t s The ground and f l i g h t t e s t program included i n t h i s e s t i m a t e c o n s i s t s o f t h e following: Proof l o a d c o n t r o l s

. System f unct i o n a l s

. Dynamic shake

. S a f e t y of f l i g h t review

D222-10060-2

. P r e - f l i g h t checks

. ~ e l i c o p t e r mode

. T r a n s i t i o n and f i x e d winc mode

Tne time span considered from shipment of wing assembly through f l i g h t tests is approximately 6 months. ~ o e i n g V e r t o l ' s e f f o r t duri3g t h i s period i s i n support of NASA personnel who w i l l i n s t a l l t h e wing on t h e a i r c r a f t per Boeing furnished i n s t r u c t i o n s and per form t h e necessary groune and f l i g h t checks.

The planning c o s t s presented are bssed on projected L T 1977 planning d o l l a r s which is intended t o r e p r e s e n t an average f o r t h e p e r iod o f per f ormance .

Cost and schedule data comprising pages 5-5 to 5-8 has been removed from this volume since it is consjd~rcd proprietary information to the Boeing Vertol Company, D222-10060-2 6.0 REFERENCES 'V/STOL T i l t Rotor A i r c r a f t S t ~ d y , Volume I, Conceptual Design o f Useful M i l i t a r y and/or Commer,:jal A i r c r a f t " , NASA CR-114437, Boeing Company, V e r t o l D i v i s i o n , March 1972.

"I~/STOL T i l t Rotor A i r c r a f t Study, vslume 11, P r e l i m i ~ a r y Design o f Research A i r w a f t " , NASA C r - I.14438, Boeing Corn- pany, V e r t o l D i v i s i o n , March 1972 - "V/STOL T i l t Rotor A i r c r a f t Study, Volume 111, O v e r a l l Research A i r c r a f t P r o j e c t P l a n , Schedules, and Estimated Cost", NASA CR-114439, Boeing Company, V e r t o l D i v i s i o n , March 1972.

"V/STOL T i l t Rotor A i r c r a f t Study, Volune I V wind Tunnel I n v e s t i g a t i o n P l a n f o r a ~ u l l - ~ c a l e ~ i l t - ~ o t o r Research A i r c r a f t " , N A S A CR-114440, Boeing Company, V e r t o l D i v i s i o n , - .

March 1972.

"Model 222 S t r u c t u r a l Design C r i t e r i a " , Boeing Document D222-10029-1.

T a y l a r , R. J. : "Weight P r e d i c t ion Techniques and Trends f o r Composite Mater i a l s t r u c t u r e s , SAWE Paper NO. 887 , p r e s e n t e d a t t h e 3 0 t h annual conference o f t h e S o c i e t y of A e r o n a u t i c a l Weight Engineers, Inc. , May 1971.

6-1 D222-10060-2 7 . G a b r i e l , E., "Drag Eatirklation o f V/STOL A i r c r a f t " , Boeing Document D8-2194-1, Boeing Company, V e r t o l D i v i s i o n , May 1969.

8. Schoen, A. H., "User's Manual f o r VASCOMP 11, The V/STOL A i r c r a f t S i z i n g and Performance Computer Program", B o ~ ing Document D8-0375, Volume V I , March 1968 (Revised October 1971).

9. Wisniewski, John S., "Weight Trends Data f o r VASCOMP", Boeing Document D8-0375-2, 1967.