Application of composites to helicopter airframe and landing gear structures

CONTENTS Section Page SUMMARY . . . . . . . . . . . . . . . . . . . . . . 1 INTRODUCTION . . . . . . . . . . . . . . . . . . . 3 1.0 STRUCTURAL DESCRIPTION OF CH-539 AI3FRAME AND LANDING GEAR STRUCTURE . . . . . . . . . . . . 4 1.1 General Description . . . . . . . . . . . 4 1.2 Material Usage and Design Considerations . . . . . . . . . . . . . 8 2.0 DESIGN CRITERIA AYD DATA . . . . . . . . . . . . . 10 2.1 Material Candidates . . . . . . . . . . . 1 0 2.2 Design Criteria . . . . . . . . . . . . . 14 2.2.1 Shear Buckling of Skin Panels . . . . . . 14 2.2.2 Compression Buckling of Flange Elements . . . . . . . . . . . . . . . . 16 2.2.3 Combined Loading . . . . . . . . . . . . . 18 2.2.4 Laminate Failure Criteria . . . . . . . . 18 2.2.5 Impact Strength Criteria . . . . . . . . 2 0 2 . 2 . 6 Overall Vehicle Design Criteria . . . 20 2.3 Design Allowables . . . . . . . . . . . . 22 2.3.1 Material Allowables . . . . . . . . . . . 22 2.3.2 Bonded Joint Shear Allowable . . . . . . 22 3 . 0 DESIGN COXCEPTS . . . . . . . . . . . . . . . . . . 24 3.1 Airframe Structure . . . . . . . . . . . 24 3.1.1 Preliminary Considerations . . . . . . . 24 3.1.2 Stringer Ccnstruction . . . . . . . . . . 26 3.1.3 Skin Construction . . . . . . . . . . . . 2 8 V Page S e c t i o n S k i d s t r i n g e r Panel Construction . . . . 30 3.1.4 3 . 1 . 5 Frame Construction . . . . . . . . . . . 30 3.1.6 S h e l l Construction . . . . . . . . . . . 30 Floor Construction . . . . . . . . . . . 32 3.1.7 3.1.8 Connections . . . . . ; . . . . . . . . . 34 3.1.9 F a b r i c a t i o n Concept . . . . . . . . . . . 36 Landing Gear S t r u c t u r e . . . . . . . . . 48 3.2 Trunnion . . . . . . . . . . . . . . . . 48 3.2.1 A x i a l l y Loaded Menbers . . . . . . . . . 50 3.2.2 3.2.3 Torque A r m s . . . . . . . . . . . . . . . 50 3.2.4 F a b r i c a t i o n Concepts . . . . . . . . . . 50 4.0 COMPOSITE DESIGX APPLICATIOTJ . . . . . . . . . . . 52 4 . 1 Airfr.e . . . . . . . . . . . . . . . . 52 4.1.1 Cockpit S e c t i o n . . . . . . . . . . . . . 58

4.1.2 Cabin Sectioii . . e . . . . 58

4.1.3 Sponson S e c t i o n . . . . . . . . . . . . . 6 6

4.1.L) A f t S e c t i o n . . a e . e . . . . 68

4 . 1 . 5 Floor S e c t i o n . . . . . . . . . . . . . . 68 4.1.6 Main Rotor Pylon F a i r i n g . . . . . . . . 70 4.1.7 Tail Pylon . . . . . . . . . . . . . . . 70 .4. 1.8 H o r i z o n t a l S t a b i l i z e r . . . . . . . . . . 72 4 . 2 Landing Gear S t r u c t u r e . . . . . . . . . 74 4 . 2 . 1 Trunnion . . . . . . . . . . . . . . . . 74 4-2.2 Shock S t r u t . . . . . . . . . . . . . . . 74 4 . 2 . 3 Drag S t r u t Cylinder and P i s t o n . . . . . 76 vi

Section Page

5 . 0 COST E F F E C T I V E N E S S EVALUATION. . . . a . - . . . 7 8

5 . 1 Production V e h i c l e C o s t C o m p a r i s o n . . 7 8

5.2 Prototype Vehicle C o s t C o m p a r i s o n . . . . 80

5 . 3 T e n - Y e a r L i f e C y c l e C o s t f o r Production

Vehicle. . . . . . . . . . . . . . 8 0

6.0 RECOMMENDATIONS AND D I S C U S S I O N . . . . . . . . 8 2

6.1 R e c o m m e n d a t i o n s . . . . . . . . . . . . 8 2

6 . 2 D i s c u s s i o n . . . . . . . . . . . . - . 8 2

A P P E N D I X A . STRUCTURAL D Z T A I L S OF CURRENT CH-53D AIRFRAME AYD L U D I B J G

GEAR STRUCTURE . . . . . . . . . . . . . 8 5

A P P E N D I X B. VEIGI-IT TREND CURVES

FOR H E L I C O P T E R STRUCTURES. . . . . . . . 100

A P P E N D I X C. MATERIAL AND XMJUFACTURING

C O S T S . . . . . . . . . . . . . a e e 109

A P P E N D I X D. COST E F F E C T I V E N E S S A N A L Y S I S e e a o a 4 e e 112

R e f e r e n c e s a . # . + * a . 4 . e 1 2 3

vii FIGURES

1 Current CH-53D H e l i c o p t e r . . . . . . . . . . . . . 5

2 Current CH-53D A i r f r a m e and Landing Gear are

of Conventional Construction . . . . . . . . . . . 5

3 Current C H - 5 3 9 Nain Landing Gear i s of

Forged C o n s t r u c t i o n . . . . . . . . . . . . . . . . 7

4 Current CH-5 3 D i s Composed of Nine Major

Assemblies . . . . . . . . . . . . . . . . . . . . 7

Current CH-53D A i r f r a m e is P r i m a r i l y Aluminum Alloy with t h e Major Weight i n t h e Outer S h e l l

Construction . . . . . . . . . . . . . . . . . . . 9

Current CH-53D Landing Gear i s P r i m a r i l y

Aluminun And S t e e l Forgings. . . . . . . . . . . . 9

7 S p e c i f i c Tension Sti-ength and Yodulus f o r

Composite Materials. . . . . . . . . . . . . . . . 1 3

S p e c i f i c Compression S t r e n g t h and Modulus f o r

Composite Materials. . . . . . . . . . . . . . . . 13

9 Shear Buckling S t r e n g t h of Graphite/Zpoxy Skin Panels is Maintained o v e r a I.?ide Range of Ply

O r i e n t a t i o n s . . . . . . . . . . . . . . . . . . . 1 5

Shear Buckling S t r e n g t h of Boron/Cpoxy S k i n Panels i s 3 a i n t a i n c d over d Wide Range of Ply

O r i e n t a t i o n s . . . . . . . . . . . . . . . . . . . 1 5

C r i p p l i n g S t r e n g t h of U n i d i r e c t i o n a l G r a p h i t e l

Epoxy Flange Elements. . . . . . . . . . . . . . . 17

C r i p p l i n g S t r e n g t h of U n i d i r e c t i o n a l Boron/

Epoxy Flange Elements. . . . . . . . . . . . . . . 17

13 Graphite/Epoxy and Boron/Epoxy Skin Panels

have S i m i l a r Combined Load C a p a b i l i t i e s . . . . . . 19

14 O f f - A x i s Loading L i m i t s S t r e n g t h of Laminate . . . 19

15 PRD-49 can be used t o i n c r e a s e Impact S t r e n g t h . . 21

Composite S t r i n g e r Concepts. . . . . . . . . . . . 27

v i i i Page 17 Foan-Stabilized Composite Stringer has Lowest

Range. . . . . . . . . . 27

Weight over required Load 18 Graphite/Epoxy Skin has higher Shear Buckling

Efficiency than Boron/Epoxy or Aluminum Skin . . . 29

19 Graphite/Epoxy and Boron/Epoxy Skin/Stringer Panels have higher Combined Load Strength/

Weight Ratio than Aluminum Panels. . . . . e . . . 31

20 Stabilized Composite Frames are Lighter than

Current Construction . . . . . . . . . . . . e 31

21 Composite Shell Concept. . . . . . . . . . . . . . 31

22 Low Cost Fiberglass Hybrid F l o o r can Reduce

Weight by 2 8 Percent . . . . . . e . . . e . . . ., 33

23 Skin Splice Concepts . . . . . e . . . . . . . 35

24 Frame Splice Concepts. . . . . . . . . . e . 35

25 Stringer Splice Concepts . . e . e . . . e . . . 35

26 Flat Pattern of Layup for Cabin Skin Segment

and Frame Outer Caps . . . . e . . e e * 37

27 Composite Shell Construction Details Of

Components For First Cure Cycle. . . e * 39

28 Composite Shell Construction Details Of

Components For Second Cure Cycle b e . 41

29 Composite Shell Construction Splice Details. . 43

30 Composite Cabin Section. Assenbly Details . . . . 45

31 Composite Airframe Assembly. . . . . . . * . . 47

32 Composite Central Cylinder is most Practical

Concept for Landing Gear Trunnion. . . . . . . . . 49

33 Selective Replacement Concept is most

Practical for Axially Loaded Gear Elements . . . 51

Composite Landing Gear Torque A m Concepts . . . 51

35 Selective Replacement is most Pracitcal

Concept for Landing Gear . . . . . . . . . . . . e 51

ix Page 36 All-Composite Structural Assembly %eights are

Reduced, Compared with Current Structure . . . . 53

Composite Airframe. Major Assembly Connections. 55

Composite Cockpit Structure. Inboard Profile. . 59

39 Composite Cockpit Structure. Auxiliary Views. .

40 Composite Cabin Structure Details. . . . . . . . 61

Composite Sponson Structure Details. . . . . . 67

42 Composite Floor Structure. Typical Floor Panel.

43 Composite Main Rotor Pylon Fairing . . . . . . 71

44 Composite Tail Pylon Structure . . . . e . . . 71

4s Composite Tail Pylon Structure. Alternate

Manufacturing Concept. . . . . . . . . . . . . . 71

46 Composite Horizontal Stabilizer Structure. . . 73

47 Composite Landing Gear Trunnion. . . . . . . . . 75

48 Composite Shock Strut. . . . . . . .’ . . . . 75

49 Composite Landing Gear Drag Strut Details. . 77

50 Composite Production Vehicle Costs Three Percent

more than Current Production Vehicle . . 79

51 Composite Prototype Vehicle Costs Four P e r c e n t more than Prototype Vehicle of Conventional

Design . . . . . . . . . . . . . . . b , . . . . 81

B1 Weight Trend for Body Structure. . . . . . . . 101

B2 Ifeight Trend for Cockpit Structure . . . . . . . 102

33 Weight Trend for Cabin Structure . . . . . . . e 103

34 Weight Trend for Aft-Section Structure . . . . 104

35 \$eight Trend for Floor Structure . . e * . 105

B6 Weight Trend for Tail Pylon Structure, . . . , 106

X Page B7 I.Jeight Trend f o r H o r i z o n t a l S t a b i l i z e r

S t r u c t u r e . . . . . . . . . . . . . . . . . . . . 1 0 7

B 8 \ ? e i g h t Trend f o r Landing Gear S t r u c t u r e . . . . . 1 0 8

91 CE-53D Mission Environment . . . . . . . . . . . 118

xi TABLES Page

1 Comparison of Materials. . . . . . . . . . . . . 11

2 Material Properties Data . . . . . . . . . . . . . 2 3

3 Composite CH-53D. Type of Structure. Weight

. Summary.. . . . . . . . . . . . . . . . . . . . . 57

4 Composite CH-53D. Material Usage. . . . . . . . . 57

5 Cost and Weight Summary f o r Production

. . . 79

Vehicles . . . . . . . . . . . . . . . . . .

6 Cost Effectiveness Sunmary . . . . . . . . . . . . 81

7 . . . 8 3

Recommended Development Programs . . . . . .

. . . 8 9

A1 CH-53D Material Usage. . . . . . . . . . . .

A2 CH-53il Airframe, Summary of Structural Types

and Design Conditions. . . . . . . . . . . . 0 . . 9 0

CH-53D Cockpit Section, Structural Types and A3

. . . . . . . . . . . * . . 91

Design Conditions.

A4 C H - 5 3 3 Cabin Section, Structural Types and Design Conditions. . , . . * . . . .

A5 CH-532 S2onson Section, Structural Types and Design Conditions. . . . . . . . . . * . .

A6 CH-53D Aft Section, Structural Types and

Design Conditions, . . . . . . . . . . . a 0 . . 9 4

A7 CII-53D Floor Section, Structural Types and

. . . 9 5

Design Conditions. . . . . . . . . . . . . .

CII-539 Main Rotor Pylon Fairing, Structural A8

Types and Design Conditions. . . . . . . . . . . . ' 9 6

C ' r I - 5 3 D Tail Pylon Section , Structural Types

A9

. . . 97

and Design Conditions. . . . . . . . . .

A10 CH-531) fiorizontal Stabilizer, Structural

Types and Design Conditions. . . e e . . - . . . 9 6

x i i Page All C H - 5 3 9 Landing Gear. S t r u c t u r a l Types and

Design C o n d i t i o n s . . . . . . . . . . . . . . . . 99

D 1 CH-53D Weight and Cost Comparison of Composite

A i r f r a m e w i t h C u r r e n t Design . . . . . . . . . . 1 1 4

D2 CH-53D Cost E f f e c t i v e n e s s Summary . . . . . . . . 115

SYMBOLS A 1 Aluminux B/E Boron/Epoxy Butt Line BL Panel o r Flange Width b C S u f f i x f o r Compression d B e a m Depth E Elastic Modulus FS Fuselage S t a t i o n G Shear Modulus G / C Graphite/Epoxy G r a v i t a t i o n a l Acc'eleration g Compression i3uckling Constant " c Shear Buckling Constant KS M Bending >foment U l t i m a t e Load Factor Axial Load, and \?eight Parameter Shear Flow, and Design Dynamic Pressure Q Critical Buckling Shear Flow qcr Wetted Area S S t r u c t u r a l Design Gross Weight S G W STA S t a t i o n Titanium T i t Xaterial Thickness o r S u f f i x f o r Tension

v Shear Force

xiv Volume Fraction vf We 5 ght

w

'WL Water Line Ultimate Cabin Differential Pressure

%

E Strain Poisson's Xatio Density P Q Direct Stress Crippling Stress Qcc Critical Compression Buckling Stress Qcr Ultimate Allowable Compression Stress Qcu Ultimate Allowable Tension Stress Qtu Design Ultimate Direct Stress Qult ? Shear Stress Critical Shear Buckling Stress ?cr Design Ultimate Shear Stress fult xv APPLICATION OF COMPOSITES TO HELICOPTER AIRFRAME AND LANDING GEAR STRUCTURES M. J. Rich, G. F. Ridgley and D. W. Lowry Sikorsky Aircraft Division of United Aircraft Stratford, Connecticut SUMMARY A preliminary design study has indicated that advanced composite helicopter airframe structures can provide * significant system cost advantages in the 1980's. A seven percent increase in productivity and a five percent reduction in life cycle cost are projected. Due to their complexity, landing gear structures do not substantially benefit from the use of advanced composites.

Tne most successful concept was found to be all-molded composite modular panels, which provide integral skin/stringer and frame subassemblies. These subassemblies significantly reduce the number of parts relative to present construction.

The subassemblies are mechanically joined together for econom- ical, rapid final assembly and permit field replacement in the event of major damage. The use of 1872 lbm (849.1 kg) of graphite/epoxy and 466 lbm (211.4 kg) of PRD-43/epoxy is pro- jected to save 1118 lbm (507.1 kg), or 1.8.5 percent of the airframe weight of the CH-5 3 D helicopter. Graphite/epoxy was

selected for prima1.y s t r u c t u r e , and FRD-49/eyoxy for secondary

structure and reirifcrcement of primary structure f o r damage tolerance improvement.

The system cost effectiveness analysis showed that while the composite airframe increases unit helicopter flyaway cost by 3 percent, the increased productivity of seven percent reduces the fleet size required and provides an overall system cost reduction of five percent. Life cycle costs were based on production starting in 1978 and extending into the 1980's.

Based on present information, a prototype composite air- frame would cost approximately four percent more than a proto- type netal airframe. The difference is due primarily to the higher engineering design time, as the increased materials cost is largely offset by reduction of fabrication labor costs.

Application of advanced composites to helicopter airframes can be made cost effective f o r production in the 1 9 8 0 ' ~ ~ provided further development efforts are made. These efforts consist of further hardware design development, manu- facturing experience, and service experience to provide the necessary cost and technical data base.

INTRODUCTION The f e a s i b i l i t y of applying advanced composites t o aircraft s t r u c t u r e s h a s been amply demonstrated, and p r o j e c t e d weight savings f o r f l i g h t hardware have been shown t o be a c h i e v a b l e ( R e f . 1). Composite materials are now articles of commerce, and costs have d e c l i n e d t o a l e v e l t h a t provides some c o s t effec- t i v e a p p l i c a t i o n s . New usage i s a n t i c i p a t e d t o f u r t h e r reduce costs f o r t h e 1 9 8 0 t i m e frame ( R e f . 2 ) . The h e l i c o p t e r airframe, with i t s r e l a t i v e l y l i g h t l o a d i n g i n t e n s i t y , i s s i g n i f i c a n t l y d i f f e r e n t from t h a t of f i x e d wing aircraft. To e f f i c i e n t l y u s e advanced composites i n t h e h e l i c o p t e r airframe, very l i g h t gage composite s k i n s must be u t i l i z e d i n t h e post-buckled stress state. For t h a t reason t h e design concepts used f o r t h e h e l i - c o p t e r s t r u c t u r e s w i l l be a p p r e c i a b l y d i f f e r e n t from t h o s e c u r r e n t l y developed f o r f i x e d wing aircraft. 1 x 1 o r d e r t o e x p l o i t t h i s technology i n p r o d u c t i o n , an adequate c o s t and t e c h n i c a l base must be developed t o assess cost e f f e c t i v e n e s s .

Both s e r v i c e e x p e r i e n c e and manufacturing techniques must be developed i n a d d i t i o n t o d e t a i l e d design s t u d i e s .

The o b j e c t i v e of t h i s s t u d y w a s t o assess t h e a p p l i c a t i o n of advanced composite materials t o h e l i c o p t e r airframe and l a n d i n g g e a r s t r u c t u r e s and t o p r o j e c t q u a n t i t a t i v e improvements i n v e h i c l e c o s t s and performance. To a t t a i n t h i s o b j e c t i v e , t h e s t u d y u t i l i z e d t h e Sikorsky model CH-53D, a c u r r e n t production t r a n s p o r t h e l i c o p t e r , f o r comparison of composite w i t h c u r r e n t conventional c o n s t r u c t i o n . Composite materials p o t e n t i a l l y offer s u b s t a n t i a l weight s a v i n g , r e d u c t i o n i n number of p a r t s , and p o s s i b l e r e d u c t i o n i n manufacturing c o s t s for t h e h i g h e r s t r e s s e d s k i n / s t r i n g e r p a n e l s and forged frames and beams.

This i s p a r t i c u l a r l y t r u e of t h e s i z e , weight, and performance class of t h e C H - 5 3 D h e l i c o p t e r , which i n v o l v e s a g r e a t e r propor- t i o n of s t r u c t u r e than do smaller rotary w i n g a i r c r a f t .

This r e p o r t is d i v i d e d i n t o s i x s e c t i o n s . The first sec- t i o n reviews t h e c u r r e n t CH-53D s t r u c t u r e and determines where t h e emphasis should be placed f o r a p p l i c a t i o n of advanced compos- ites. The second s e c t i o n i s a compilation of s t r u c t u r a l d a t a and c r i t e r i a t o be uszd f o r composite design. The t h i r d s e c t i o n a p p l i e s the d e s i g n concepts t o t h e CH-53D and compiles t h e weight s a v i n g s t h a t can be achieved through t h e use of advanced com- p o s i t e s . The f i f t h s e c t i o n assesses t h e c o s t s of a composite production and p r o t o t y p e v e h i c l e and t h e l i f e c y c l e cost effec- t i v e n e s s f o r f l e e t o p e r a t i o n s . The s i x t h s e c t i o n reviews t h e f u r t h e r efforts r e q u i r e d t o achieve t h e cost e f f e c t i v e a p p l i c a - t i o n of advanced composites t o h e l i c o p t e r s t r u c t u r e s .

S E C T I O N 1 . 0 STRUCTURAL D E S C R I P T I O N OF CH-53D AIRFTIIL’.IE AND LANDING GEAR STRUCTURE 0 %le Sikorsky model CH-53D i s a l a r g e , modern, h i g h speed t r a n s p o r t h e l i c o p t e r r e p r e s e n t a t i v e of conventional c o n s t r u c t i o n and material usage of the c u r r e n t g e n e r a t i o n of h e l i c o p t e r s .

1.1 GENERAL D E S C R I P T I O N The CH-53D u s e s high s t r e n g t h aluminum a l l o y as t h e primary s t r u c t u r a l material; 8 0 % of t h e airframe and l a n d i n g g e a r s t r u c t u r e weight i s of t h i s material, The C H - 5 3 3 h e l i c o p t e r , shown i n Figure 1, i s of s i n g l e main r o t o r c o n f i g u r a t i o n , powered by two t u r b i n e engines, and i n c o r - p o r a t i n g rear l o a d i n g f o r cargo and t r o o p s . This a i r c r a f t has sponsons t h a t provide water f l o t a t i o n , hydrodynamic and hydro- s t a t i c s t a b i l i t y , and support f o r the main l a n d i n g g e a r and f u e l t a n k s . The l a n d i n g g e a r i s of t r i c y c l e c o n f i g u r a t i o n , i n c o r - p o r a t i n g a i r - o i l o l e o s t r u t s and d u a l main and nose wheels.

There arc c u r r e n t l y 348 of these aircraft i n s e r v i c e , and it is a n t i c i p a t e d t h a t a t o t a l of 6 0 0 of t h i s model w i l l e v e n t u a l l y be b u i l t . The growth p o t e n t i a l of t h i s fast, l a r g e t r a n s p o r t makes it a l o g i c a l choice t o study t h e improvements p o s s i b l e through use of advanced composites.

The o v e r a l l v e h i c l e design criteria are given below:

Design Gross Weight ................. 33,500lbm

(15,196kg) Design L i m i t ( F l i g h t ) Load F a c t o r . , , 3 . 0

Design L i m i t (Landing) Sink Speed ... 8fps ( 2 . 4 4 m / s )

Design Limit Dive Speed.. .......... .195kts (100.31m/s)

Design Alternate Gross Weight.. .... .42,0001bm (19,051kg3

(Reduced F l i g h t Load F a c t o r ......... 2 . 3 9 )

The airframe s t r u c t u r e , shown i n Figure 2 , is of s e m i - nonocoque ( s k i n / s t r i n g e r / f r a m e ) c o n s t r u c t i o n , u s i n g m u i t i p l e s t r i n g e r s formed o v e r s t r u c t u r a l framing. Aluminum a l l o y s are used throughout for t h e primary s t r u c t u r e . F o r t h e l a n d i n g

gear , shown i n Figure 3 , forged aluminum a l l o y is t h e major

s t r u c t u r a l material, although high-strength s t e e l is used for many p a r t s .

The c u r r e n t CH-53D is composed of n i n e major subassemblies, as d e p i c t e d i n Figure 4. They are t h e c o c k p i t , c a b i n , sponsons, aft s e c t i o n , f l o o r , main r o t o r pylon f a i r i n g , t a i l pylon, hor- i z o n t a l s t a b i l i z e r , and l a n d i n g g e a r assemblies. A more d e t a i l e d d e s c r i p t i o n i s p r e s e n t e d i n Appendix A.

. .

"- I FIGURE 1 . CURRENT CH-53D HELICOPT^.

FIGURE 2 . cuRRE3"T C H - 5 3 AIRFRAME AND LANDING GEAR ARE OF CONVENTIONAL CONSTRUCTION.

ARMS FIGURE 3. CURRENT CH-’j3D MAIN LANDING GEAR I S O F FORGED CONSTRUCTION.

CURRENT CH-53D IS COMPOSED OF N I N E MAJOR ASSEMBLIES.

FIGURE 4.

1.2 MATERIAL USAGE AND DESIGN CONSIDERATIONS The major p o r t i o n of t h e p r e s e n t CH-53D weight is i n t h e o u t e r aluminum s k i n / s t r i n g e r / f r a m e s h e l l .

The d e s c r i p t i o n of s t r u c t u r a l weight i n t h e f u s e l a g e and l a n d i n g g e a r of t h e CH-533 by material, s t r u c t u r a l t y p e , design c o n s i d e r a t i o n , and subassembly is summarized i n F i g u r e s 5 and 6 . A more d e t a i l e d d e s c r i p t i o n i s p r e s e n t e d i n Appendix A.

From t h e w e i g h t breakdown f o r t h e c u r r e n t airframe s t r u c t u r e , shown i n Figure 5 , it can be seen t h a t a major p o r t i o n of t h e s t r u c t u r e i s i n the alusinum o u t e r s h e l l ( s k i n / s t r i n g e r / f r a m e s ) , which i s designed by s t r e n g t h consid- e r a t i o n s t h a t i n c l u d e c r i p p l i n g , buckling, and u l t i m a t e stress.

This r e g i o n of t h e airframe is, t h e r e f o r e , h i g h l y s u i t a b l e f o r a p p l i c a t i o n of t h e h i g h s p e c i f i c s t r e n g t h composite materials.

The minimum gage and n o n - s t r u c t u r a l r e g i o n s of t h e airframe appear s u i t a b l e f o r the a p p l i c a t i o n of lower s t r e n g t h b u t lower cost composite materials.

For the l a n d i n g g e a r , it can be seen from Figure 6 t h a t almost all t h e s t r u c t u r e weight i s designed by u l t i m a t e s t r e n g t h requirements. However, s i n c e t h e l a n d i n g gear s t r u c t u r e con- sists mainly of aluminum and steel forged parts, it is a n t i c i - pated t h a t t h e use of composite materials w i l l be l i m i t e d due .

t o t h e complexity of t h e r e q u i r e d shapes.

3000- TOTAL WT. = 6077 Ibm (2756.5 kel HORIZONTAL RlGlDlTY STABILIZER NON P Y L O N STRUCTURAL F A I R I N G MISC.

MIN. G A G E TAIL P Y L O N M A N U - L O N G E R O N S F L O O R FACTURING FAlRINGS FITTINGS COCKPIT 2000- B E A M S U J -1L S P O N S O N c I Lu Lu Iy A L U M I N U M ;KIN-STRING€, AFT SECTION U

t

c u) STRENGTH (CRI PPCl NG , 1000- BUCKLING.

6 ULTIMATE - STRESS1 C A B I N FRAMES MATERIAL STRUCTURE D E S I G N M A J O R TYPE C O N D I T I O N ASSEMBLIES FIGURE 5. CURREP?" CH-53D AIRFRAME I S PRIMARILY A L U M I m 3 ALLOY, 'WITH THE MAJOR WEIGIPT I N THE OUTER SHELL CONSTRUCTION.

300- * INCLUDES METALLIC HARDWARE NOT CONSIDERED FOR FABRICATION FROM COMPOSITE MATERIALS TOTAL WT. = 513 Ibm 1232.7 k@) * MISC.

NOSE GEAR MISC.

m I- I P

z

STEEL w m

z

U =I m I- v) BEAMS MAIN GEAR FlrYlNGS ALUMINUM STRUYS MATERtAi. STRUCTURE DESIGN MAJOR TYPE CONDITION ASSEMBLIES CURRENT C H - 5 3 LANDING GEAR IS PRIMAl3ILY ALUMINUM AND FIGURE 6 .

STEEL FORGINGS.

SECTIOX 2 . 0 DESIGN C R I T E R I A AXD DATA 0 S t r e n g t h and s t a b i l i t y d e s i g n c r i t e r i a f o r composites are s i g n i f i c a n t l y d i f f e r e n t f r o m t h o s e used f o r metals due to t h e a n i s o t r o p y of composites.

MATERIAL CAIIDI I),4TES 2.1 The h e l i c o p t e r airframe and l a n d i n g g e a r s t r u c t u r e s o p e r a t e i n a moderate temperature environment of -65OF ( 2 1 9 O K ) t o 16OoF (344OK) with the p r i n a r y d e s i g n c o n d i t i o n s being of s t a t i c n a t u r e . Candidate composite materials w e r e , t h e r e f o r e , e v a l u a t e d based on t h e room temperature s t a t i c s t r e n g t h of u n i d i r e c t i o n a l l a m i n a t e s .

O f t h e a v a i l a b l e n o n - n e t a l l i c and metallic natrices, o n l y epoxy w a s c o n s i d e r e d f o r t h i s s t u d y . Epoxy w a s s e l e c t e d because it has gained wide acceptance and t h e r e i s a c o n s i d e r a b l e body of e y p e r i e n c e e x i s t i n g f o r i t s use. Epoxy m a t r i x aaterials offer a good balance of p r o c e s s i n g , s t r e n g t h , and adhesion c h a r a c t e r i s t i c s . Metal matrices w e r e e l i m i n a t e d f r o m con- s i d e r a t i o n , because t h e y are p r e s e n t l y so much m o r e expensive and more d i f f i c u l t t o f a b r i c a t e t h a n r e s i n matrix materials.

The composite materials c o n s i d e r e d f o r t h i s s t u d y are summarized i n Table I , and t h e i r p r o p e r t i e s i l l u s t r a t e d i n F i g u r e s 7 and 8.

Both boron/epoxy and g r a p h i t e / e p o x y appear t o be t h e prime c a n d i d a t e materials f o r t h e major p o r t i o n of t h e primary s t r u c - t u r e . This i s due t o t h e i r h i g h specific s t r e n g t h and modulus i n both tensi.on and compression. PRD-49/epoxy is t h e prime c a n d i d a t e material f o r secondary s t r u c t u r e due t o i t s lower d e n s i t y and improvement i n modulus o v e r f i b e r g l a s s I n a d d i t i o n PRD-49/epoxy may be a candidate i n primary s t r u c t u r a l areas where i t s h i g h s p e c i f i c t e n s i o n strength can be u t i l i z e d .

P2D-49 combined w i t h g r a p h i t e may also be a c a n d i d a t e where moderate compression s t r e n g t h i s found adequate.

TABLE 1. COMPARISON OF MATERIALS ~ C O ' A B * 1973 1978 REASON FOR CONSIDERATION - .073 155 65 High t e n s i l e and exceptionally high compres- ?020.6) s i v e s t r e s s o f f e r a good choice f o r structures designed f o r reversal of s t r e s s e s . The 5 . 6 mil f i b e r is used i n cost estimates s i n c e it offers some savings over t h e 4 . 0 m i l s i z e .

,055 95 25 Graphite/Epoq Reason is similar t o t h a t for s e l e c t i n g boron/epoxy. I n addition, t h e smaller 1522.4) (High Strength, Mod I1 f i b e r s i z e allows g r e a t e r f l e x i b i l i t y i n or ms) producing conplex shapes.

.058 125 30 Graphite/Epoxy The moderately high strength coupled with a ~605.4) (High Modulus, ms) high modulus makes t h i s material a possible first choice f o r con?ression s t a b i l i t y l i m i t e d structures.

.055 75 Graphite/Epow (AS) Moderate strength e?d modulus. P r i m a r v 25 L522.4 advantage is low cost at current prices.

.065 2 . 5 0 2 . 5 0 A r e l a t i v e l y low cost composite material E-Glass/Epoxy 1799.2) with a high t e n s i l e strength properties.

While t h e lower conpression strength is a l i m i t i n g f a c t o r , t h e material i s a good choice f o r lii.Jitlv loaded structures.

6 6 .070 A choice t h n t ext.ends t h e raiipe where E-Glasslfpoxy would he used f o r increased 1937.6) i n t e n s i t y of l o a d 1 1 ~ s .

.UjO 22 High t e n s j l r s t . f e r e t h arid v+rv l o w density 13811 .o 1 o f f e r applicr.:?%a f secondxry stwwtures, or primary s t r w t u r e s deslmied f o r tension.

The increased modulus over t h e fiberglass/ epoxies extends t h e range of usefulness.

However, t h e material is severely 1in;ited i n e-ression.

1 1 ?O75-T6 A~u. Alloy (For R e f .

Based on 1973 dollars 5000 5000 rn X 4000 4000

w oPRD-4g-III/EPoxY

n \ I 3000 3000 O H S GRAPHITE/EPOXY O H S GRAPHITE/EPOXY - 0 BORON/EPOXY ( 4 MIL) 0 BORON/EPOXY ( 4 MIL) 0 E-GLASS /EPOXY O E-GLASS /EPOXY 2000 2000 0 HM GRAPHITE/EPOX 0 HM GRAPHITE/EPOX 0 BORON/ALUMINUM (REF ) 0 BORON/ALU~INUM (REF ) 3.000 3.000 O 7075-T6 ALUM ( R E F ) 0 7075-T6 ALUM ( R E F ) _ r L _ r L 0 100 200 300 400 500 0 100 200 300 400 500

SPECIFIC TENSION MODULUS - E/p, IN. X 10 5

r I I

0 5 10

SPECIFIC TENSION MODULUS - E / p , m x LO

FIGURE 7. SPECIFIC TENSION STRENGTH AND MODULUS FOR CO?@OSITE MATERIALS.

BORON/EPOXY (&MIL ) 0 m x * 4000

w

a Q \ I H S GkAPHITE/EPOXY 0 H rn 0 BORON/ALUMINUM ( R E F ) rn 2000 w

E

HM GRAJ?HITE/EPOXY 0 z . v 0 E-GLASS/EPOXY - 0 P R D - b 9 - I I I / E P O X Y 7 0 7 5 ~ 6 A LULI .

( REF )

S P E C I F I C COMPRESSION MODULUS - E / p , IN. x 10 6

I 1

0 ; 10

S P E C I F I C COMPRESSION MODULUS - E / p , m x 10 6

FIGURE 8. S P E C I F I C COMPRESSION STRENGTH AND MODULUS FOR COMPOSITE ?UiTERIALS.

2 . 2 DESIGN CRITERIA Some of t h e d e t a i l design criteria c u r r e n t l y used f o r m e t a l l i c a i r f r a m e s t r u c t u r a l elements cannot be used d i r e c t l y f o r elements c o n s t r u c t e d from composite materials. For example, s k i n panel s h e a r buckling and f l a n g e e l e m e n t compres- s i o n - buckling behavior cannot be p r e d i c t e d from known data u s i n g a simple elastic modulus r a t i o . These c r i t e r i a , t o g e t h e r w i t h criteria f o r laminate f a i l u r e and r e q u i r e d impact s t r e n g t h , used f o r design are p r e s e n t e d .

are h e r e reviewed and t h e c r i t e r i a 2 . 2 . 1 Shear Buckling of Skin Panels For metallic s k i n p a n e l s , t h e c r i t i c a l buckling s h e a r stress T C r i s given by t h e g e n e r a l r e l a t i o n where E = elastic modulus t = panel t h i c k n e s s b = panel width a n d Ks i s a f u n c t i o n of t h e p l a t e a s p e c t r a t i o and edge support c o n d i t i o n s .

. For a composite material t h e p r e d i c t i o n of t h e o n s e t of

shear buckling is n o t so simple. To e s t a b l i s h t h e c r i t i c a l b u c k l i n g shear stress, a computer program i s used ( R e f . 3 ) .

The program p r e d i c t s the buckling load f o r an a n i s o t r o p i c p l a t e by s e a r c h i n g f o r t h e e q u i l i b r i u m b i f u r c a t i o n p o i n t , using a n energy minimization t e c h n i q u e . Results from this program are shown i n Figures 9 and 1 0 , where the shear flow q is d e f i n e d t. A r e v e r s a l of shear. loading on t% panel as qcr = t CY c a u s e s t h e induced principal. stresses t o rotate 90". 'This is e q u i v a l e n t t o a d i f f e r e n t ply-layup sequence, and hence t h e p a n e l buckling load i s a f u n c t i o n of the d i r e c t i o n of t h e s h e a r l o a d i n g . For a l a r g e s k i n - s t r i n g e r panel of composite construc- t i o n , t h e sense of t h e s h e a r f l o w can vary i n i n d i v i d u a l sub- panels and can change f o r d i f f e r e n t airframe l o a d i n g c o n d i t i o n s .

For t h i s r e a s o n , t h e curves of Figures 9 and 1 0 are drawn t h r o u g h minimum v a l u e s of t h e buckling s h e a r flow. These mini- nium c u r v e s are a p p l i c a b l e f o r d e s i g n , with some conservatism, w i t h o u t r e g a r d t o ply-layup sequence o r d i r e c t i o n of s h e a r load- ing.

An a n a l y t i c a l procedure f o r p r e d i c t i o n of shear p a n e l ' f a i l u r e i s not c u r r e n t l y a v a i l a b l e , and test data are l a c k i n g .

The d a t a of R e f . 4 i n d i c a t e t h a t , f o r +45O ply, o r i e n t a t i o n of boron/epoxy l a m i n a t e s , t h e r a t i o of f a r l i n g load t o i n i t i a l 20c l o o * O 0 .02 .04 ' .06 .08 t S K I N THICKNESS, I N .

I I I 1 0 0 . 5 1 . 0 1.5 t S K I N THICKNESS, mm F I G U R E 9 . SHEAR BUCKLING STFENGTH OF GRAPHITE/EPOXY S K I N PANELS IS MAINTAINED OVER A WIDE RANGE OF P L Y ORIENTATION.

SO E \ I 5 : G I Er

ii

* 20

0 to2 .04 .08 t SKIN THICKNESS - IN.

I 1 I I 0 0.5 1.0 1 . 5

t SKIN THICKNESS - IIIZU

FIGURE 10. SHEAR BUCKLING STRENGTH OF BOROI?/EPOXY SKIN PANELS IS MAINTAINZD OVER A WIDE W G E OF PLY ORIENTATIOB.

buckling load i s approximately 8:l. One of t h e o v e r a l l v e h i c l e d e s i g n c r i t e r i a (see S e c t i o n 2.2.6) r e q u i r e s no shear buckling of s k i n p a n e l s at a 1.Og load l e v e l . This i m p l i e s s h e a r panel l o a d i n g of 3 X i n i t i a l buckling load a t t h e airframe design l i m i t load f a c t o r . O n t h i s basis, t h e f a i l u r e c r i t e r i o n f o r boron/epoxy and graphite/epoxy panels does not appear c r i t i c a l .

I n Figures 9 and 1 0 , t h e panel t h i c k n e s s i s considere6 made up as a symmetric, balanced laminate t o prevent out-of-plane warping.

2.2.2 Compression Buckling of Flange Elements As f o r s h e a r buckling, t h e c r i t i c a l compression buckling stress f o r a metallic f l a n g e e l e m e n t can be p r e d i c t e d by a simple r e l a t i o n s h i p . T h i s g e n e r a l r e l a t i o n is where E, = elastic compression modulus t = f l a n g e t h i c k n e s s b = f l a n g e width and Kc is a f u n c t i o n of t h e f l a n g e a s p e c t ratio and edge s u p p o r t c o n d i t i o n .

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

element of composite material and, as for s h e a r b u c k l i n g , t h e a l l o w a b l e l o a d i n g for a composite e l e m e n t i s obtained by u s e of a computer program. T h i s program, developed by F r a t t and Whitney A i r c r a f t , uses t h e s t i f f n e s s m a t r i x of t h e anisotropic f l a n g e element t o c a l c u l a t e the critical buckling stress levels.

The curves of Figures 11 and 1 2 present t h e allowable compres- sion stress for f l a n g e elements of u n i d i r e c t i o n a l boronlepoxy and graphite/epoxy d e r i v e d from this program.

1 . 5 cu cu

2 1 . 0 E

\ d k e P rl m tn w tn

w

crl PI 0.5 b U b 0 10 20 30 40 b/t RATIO CRIPPLING STRENGTH OF UNIDIRECTIONAL GRAPHITE/EPOxy FIGURE 11.

FLANGE ELEMENTS.

300,000 I I I : t I i i ! I i I !

# i I I i I I 200,000 (u \ % ft n m

w"

P; E-r m

w

GI pc ffi u 0 1oo,ooc t , C 10 20 30 b/t RATIO FIGURE 12. CRIPPLING STRENGTH OF UNIDIRECTIONAL BORON/EPOXY FLANGE ELEMENTS.

2.2.3 Combined Loading For p a n e l s c a r r y i n g both s h e a r and compre i o n l o a d i n g , i s an i n t e r a c t i o n curve of t h e c o m o n l y used design curve combined s h e a r and compression. To o b t a i n such a curve for t h i s s t u d y , t h e s h e a r buckling c h a r a c t e r i s t i c s of boron/epoxy and graphite/epoxy (Figures 9 and 10) are compared with t h e buckling c h a r a c t e r i s t i c s of an aluninum p a n e l of s i m i l a r a s p e c t r a t i o and panel t h i c k n e s s . From t h i s comparison, an effect- i v e elastic modulus is d e r i v e d for t h e composite materials.

This e f f e c t i v e modulus- i s t h e n used as i n p u t t o a computer program, t o produce t h e i n t e r a c t i o n curves of Figure 13. The p a n e l s i z e considered is t h e 6 . 0 i n . (152.4mm) x 2 0 . 0 i n .

(508.0nm) p a n e l , t y p i c a l f o r the c u r r e n t airframe. The maximum s t r i n g e r compression load c a p a b i l i t y shown i n Figure 1 3 is r e p r e s e n t a t i v e of t h e r e q u i r e d load c a p a b i l i t y f o r the panel s i z e considered.

2.2.4 Laminate F a i l u r e Criteria For an element of s t r u c t u r e b u i l t - u p f r o m p l i e s with d i f f e r i n g f i l a m e n t o r i e n t a t i o n , t h e c r i t e r i o n f o r element f a i l u r e i s more complex t h a n f o r a homogeneous material or a laminate w i t h u n i d i r e c t i o n a l f i l a m e n t s . Previous work i n t h i s area shows a d i v e r s i t y of opinion concerning t h e d e f i n - i t i o n of laminate f a i l u r e . Three varying summarized d e f i n - i t i o n s are given below, i n which t h e term laminate stress refers t o t h e nominal stress ( l o a d / g r o s s a r e a ) : Design u l t i m a t e laminate stress defined as t h e maxi- mum l a m i n a t e stress a t t a i n a b l e without r u p t u r e of any p l y ( R e f . 5).

0 Design l i m i t laminate s%r*css defined as the maximum l a m i n a t e stress a t t a i n a b l e without r u p t u r e of any p l y ( R e f . 6). T h i s d e f i n i t i o n is i l l u s t r a t e d i n Figure 14.

e Design u l t i m a t e laminate stress d e f i n e d as t h e maximum stress a t t a i n a b l e without r u p t u r e of more t h a n one p l y ( R e f . 7 ) .

The n a t u r e of the l o a d i n g a p p l i e d t o any s t r u c t u r a l element w i l l i n f l u e n c e t h e choice of f a i l u r e criteria used for d e s i g n . I n g e n e r a l , airframe s t r u c t u r a l elements are s u b j e c t e d t o r e p e a t e d l o a d i n g of varying i n t e n s i t y during t h e i r design l i v e s .

Any p l y f a i l u r e a t l i m i t load (which is p o s s i b l e by using t h e first and t h i r d d e f i n i t i o n s above) would impair t h e c a p a b i l i t y of such an element t o c a r r y f u r t h e r loading. For t h i s r e a s o n t h e second d e f i n i t i o n above, i l l u s t r a t e d by Figure 1 4 , is used f o r design.

1 8 P Y

8 50

0 1000 2000 M O O 4c STRINGER COMPRESSION WAD, 1bf I i I I 10 15 0 5 STRINGER COMPRESSION LOAD, kN GRAPHITE/EPOXY AND BORON/EPOXY SKIN PANELS HAVE FIGURE 13.

SIMILAR COMBINED LOAD CAPABILITIES.

FIGURE 14. OFF-AXIS LOADING LIMITS STRENGTH OF LAMIUATE.

2.2.5 I n p a c t S t r e n g t h Criteria Impact r e s i s t a n c e i s an important c o n s i d e r a t i o n f o r t h e l i g h t gage c o n s t r u c t i o n envisaged f o r e x t e r i o r s k i n p a n e l s i n t h e composite airframe. C u r r e n t d a t a ( R e f . 8) i n d i c a t e t h e following comparative Charpy impact c a p a b i l i t i e s f o r uni- d i r e c t i o n a l composite l a m i n a t e s :

PRI)-QS/Type 111 2 5 0 f t . lbf / i n 2 ( 5 6 5kmN/m2 1

PRD-49/Type I 1 5 0 f t . l b f / i n 2 ( 3 39kmN/m2 1

Boron/Epoxy ' 50 f t .lbf/in2(113kmN/m2)

Thornel-50(Graphite)/Epoxy 20 f t . l b f / i n 2 (45kmEJ/m2 1

25% ?R3-49/Type 111-75% Boron/

1 3 0 f t . l b f / i n 2 ( 2 94kmN/m2 1

EPOXY

Aluminum, 2024-T4 ( r e f . 1 2 2 0 f t . l b f / i n 2 ( 4 9 7kmN/m2 1

These v a l u e s are i l l u s t r a t e d i n Figure 1 5 .

There i s i n s u f f i c i e n t s e r v i c e experience t o i n d i c a t e t h e ' r e q u i r e d impact s t r e n g t h t o maintain impact damage a t an a c c e p t a b l e l e v e l . A minimum Charpy impact s t r e n g t h of 50 f t .

l b f / i n 2 ( 1 1 3 kml?/n21 f o r e x t e r n a l a i r f r a m e s u r f a c e s w a s a r b i t r a r i l y e s t a b l i s h e d as a c r i t e r i o n f o r t h i s study. With t h i s c r i t e r i o n , t h e use of graphite/epoxy would r e q u i r e an i n c r e a s e i n t h e material impact s t r e n g t h t o meet t h e minimum requirement. Such an improvement can be obtained by t h e use of PRD-49 material i n c o n j u n c t i o n w i t h t h e graphite/epoxy.

2.2.6 O v e r a l l Vehicle Design Criteria The f o l l o w i n g c r i t e r i a , which are independent of t h e material s e l e c t e d , are used f a r v e h i c l e d e s i g n : 0 U l t i m a t e factor of safety of 1 . 5 based on 1i.rni.t design l o a d s .

0 I?o buckling f o r 1.Og f l i g h t load c o n d i t i o n .

0 Material d e s i g n a l l o w a b l e s based on t y p i c a l room temperature v a l u e s with statistical r e d u c t i o n f o r d e s i g n r e l i a b i l i t y .

0 Material elastic c o n s t a n t s taken as t y p i c a l room temperature v a l u e s .

EL - 'jO/EPOXY

I

I

I

I

I BORON /EPOXY.

I

I I

I 125% PRD-49 TYPE I I I / E P O X Y 75% BORON

i

I

I

I PRD-49 T Y P E I

I

I

I PRD-49 T Y P E I11

I I t I I

I a I I I I 50 100 150 200 250 CHARPY IMPACT STRENGTH, FT lbf /IN.

I I I I I 0 100 20c 300 4 : O 500 CHARPY IMPACT STRENGTH, kmN/m2 FIGURE 15. PRD-49 CAN BE USED TO IIICREASE IMPACT STRENGTH.

2.3 DESIGN ALLOWABLES 2.3.1 Haterial Allowables For primary redundant a i r f r a m e s t r u c t u r e and a l l second- a r y s t r u c t u r e , it is customary t o use t h e '3' s t r e n g t h allow- a b l e , which i s d e f i n e d such t h a t t h e r e is a 90 p e r c e n t s u r v i v a b i l i t y with 95 p e r c e n t confidence, The ' B ' allowable s t r e n g t h i s used i n t h i s s t u d y , s i n c e t h e m a j o r i t y of t h e s t r u c t u r e falls w i t h i n t h e ' B f s t r u c t u r a l c a t e g o r y .

Material elastic c o n s t a n t s u s u a l l y e x h i b i t smaller v a r i a t i o n s t h a n t h e s t r e n g t h a l l o w a b l e s ; t h e r e f o r e , €or t h i s s t u d y , t y p i c a l v a l u e s are used.

Most of t h e f i b r o u s composites have no d e f i n e d y i e l d p i n t f o r u n i d i r e c t i o n a l f i l a m e n t s , i .e. t h e 0 . 2 p e r c e n t s t r a i n d e v i a t i o n f a l l s o u t s i d e t h e s t r e n g t h envelope.

For t h i s s t u d y , t h e r e f o r e , t h e u l t i m a t e a l l o w a b l e s are used.

The material p r o p e r t i e s used i n design are p r e s e n t e d i n Table 2, 2.3.2 Bonded J o i n t Shear Allowable For a given j o i n t , t h e a l l o w a b l e shear load i s a f u n c t i o n of s e v e r a l parameters, i n c l u d i n g adhesive t y p e and t h i c k n e s s , adherend materials and t h i c k n e s s e s , and j o i n t type and geometry.

I n view of t h e number and d i v e r s i t y of bonded j o i n t s c o n s i d e r e d i n t h e CH-53D composite airframe, a d e t a i l e d a n a l y s i s is n o t performed f o r each j o i n t . I n s t e a d an allowable average shear stress i s used far design purposes. U s e of t:liis d e s i g n allowable throughout t h e structurte r e s u l t s i n some j o i n t s that are n o t of optimum design locally. However, small changes in j o i n t design , although c i l i ti cal. for s t r e n g t h z e q u i r e r w n t s , do not s i g n i f i c a n t l y a f f e c t ttie s t r u c t u i ' e w e i g h t oz' manufacturing cost. The design average allowable shear stress is set a t 3000

p s i ( 0 . 0 2 1 G M / m L ) , which corresponds , for example, t o t h e s h e a r

allowable for t h e j o i n t sketched below ( d a t a f r o m R e f . 9 ) .

UNIDIRECTIONAL TITANIUM GWHITE/EPOXY D t t r(- m N

la H

n e o u d Y E : 0 ..I E R 8 8 8 8 , * * * 8 .

SECTION 3.0 DESIGX CONCEPTS The use of c o n p o s i t e materials p e r m i t s a r e d u c t i o n i n weight and number of p a r t s , t o g e t h e r with s i m p l i f i e d f a b r i c a t i o n and assembly p r o c e s s e s .

To u t i l i z e t h e design p o t e n t i a l of composite materials i n a c o s t e f f e c t i v e manner, e f f o r t is d i r e c t e d t o achieving t h e following f e a t u r e s i n t h e f i n a l d e s i g n : 0 Minimum number of s e p a r a t e d e t a i l p a r t s .

F a b r i c a t i o n of s e p a r a t e p a r t s u s i n g automated equip- nent such as i n f l a t p a t t e r n layup.

I n t e g r a l s h e l l c o n s t r u c t i o n t o reduce mechanical f a s t e n i n g t o t h a t r e q u i r e d f o r j o i n i n g of major assemblies.

0 Minimum number of s e p a r a t e c u r e c y c l e s .

0 Matching laminate layup s t r e n g t h p r o p e r t i e s t o d i r e c t i o n of loading.

The concepts for design are l i m i t e d by t h e following geometric c o n s t r a i n t s , which are considered unchanged from the c u r r e n t v e h i c l e : 0 E x t e r n a l airframe shape.

0 Location of major s t r u c t u r a l components, e , g . , l a n d i n g g e a r suppcirt frames, Geometry and location of doors, windows, access panels, etc.

0 Geometry and support l o c a t i o n of a l l n o n - s t r u c t u r a l systems AIRFRAME STRUCTURE 3.1 3.1.1 Preliminary Considerations are used throughout The following g e n e r a l i z e d concepts t h e composite s t r u c t u r e .

0 F i t t i n g s F i t t i n g s , both forged and machined, are used i n many r e g i o n s of t h e c u r r e n t airframe and l a n d i n g g e a r s t r u c t u r e .

They e x i s t , i n g e n e r a l , a t t h e i n t e r s e c t i o n of s t r u c t u r a l l o a d p a t h s , a t s p l i c e l o c a t i o n s , and a t p o i n t s of concen- t r a t e d l o a d i n g s e r v i n g t o d i f f u s e t h e l o a d i n g i n t o a d j a c e n t s t r u c t u r e . The f i t t i n g s are, t h e r e f o r e , of complex shape , c a r r y i n g high l e v e l s of combined l o a d i n g .

Consideration of p o s s i b l e cornposite materials f o r t h e of f i t t i n g s are l i m i t e d t o s h o r t (chopped) f a b r i c a t i o n f i b e r / e p o x y materials o r Soron/aluminum. Consideration of b u i l t - u p f i t t i n g s u s i n g metallic elements bonded t o laminated epoxy based composites i s n o t considered cost e f f e c t i v e .

For short f i b e r / e p o x y materials, t h e p o t e n t i a l weight r e d u c t i o n i s s m a l l . The d a t a of R e f . 1 6 s u g g e s t a s p e c i f i c ' t e n s i o n a l l o w a b l e of 7 0 0 , 0 0 0 i n . ( 1 7 , 8 0 O m ) , which is compar- able t o t h a t f o r aluminum. I n a d d i t i o n , material c o s t s are s u b s t a n t i a l l y h i g h e r t h a n alu.?inum, s o t h e concept i s n o t cost e f f e c t i v e .

U s e of boron/aluminum composite material has t h e poten- t i a l of b r a z i n g c a p a b i l i t y t o enable complex shapes t o be f a b r i c a t e d . IIowever, t h e p o t e n t i a l c o s t of f i t t i n g s made f r o m t h i s material is h i g h , an& t h e concept is n o t considered cost e f f e c t i v e .

F i t t i n g s r e p r e s e n t only 5.5 p e r c e n t of t h e airframe s t r u c t u r e weight and, based on t h e comments ahave, t h e air- frame f i t t i n g s are r e t a i n e d as m e t a l l i c elements. To i-ediice

t h e problems a s s o c i a t e d w i t h thema1 misma-tcli , t h e c u r r e n t

aluminum f i t t i n g s are rep1.ace.d by t i t a n i u m fittings of e q u i v a l e n t s t r e n g t h and s t i f f n e s s , a Graphite/Epoxy P l y Thickness For graphite/epoxy l a m i n a t e s , t h e p l y t h i c k n e s s a v a i l a b l e from s u p p l i e r s v a r i e s from 0 . 0 0 5 i n . (0,127mm) t o 0 . 0 2 0 in.

(0.508mm). Consideration of optimum p l y t h i c k n e s s f o r each s t r u c t u r a l element would r e s u l t i n t h e use of many d i f f e r e n t p l y t h i c k n e s s e s i n t n e composite airframe. T h i s s i t u a t i o n is not d e s i r a b l e f o r r e a s o n s of both material c o s t and q u a l i t y c o n t r o l d u r i n g manufacture. For these r e a s o n s , a s t a n d a r d p l y t h i c k n e s s of 0 . 0 1 i n . (9.254mm) is used i n t h e concepts f o r a l l graphite/epoxy l a m i n a t e s .

2 5 0 Laminate I.?arping To prevent out-of-plane warping, a l l laminates are con- s i d e r e d t o be b a l a n c e d , symmetric layups. Exceptions t o t h i s g e n e r a l r u l e o c c u r i n r e g i o n s where l o c a l s t r e n g t h e n i n g i s made by a d d i t i o n a l l a m i n a t e buildup and, f o r some sandwich panel face s h e e t s , where l o c a l bonding w i l l r e s t r a i n t h e warping.

0 Primary S t r u c t u r a l Naterial The r e l a t i v e s t r e n g t h s of boron/epoxy and graphite/epoxy laminates under combined l o a d i n g are shown by Figure 1 3 t o be of comparable magnitude. Since boron/epoxy has a p r o j e c t e d cost much h i g h e r t h a n t h a t for graphite/epoxy (see Table 11, t h e composite material considered f o r t h e primary airfraz!e and l a n d i n g gear s t r u c t u r e is graphite/epoxy, i n both HMS and MTS forms.

3.1.2 S t r i n g e r Construction The concepts considered f o r s t r i n g e r c o n s t r u c t i o n are shown i n Figure 1 6 . A comparison of t h e concepts on a weight b a s i s is p r e s e n t e d i n Figure 1 7 , which i n d i c a t e s t h e s u p e r i o r s t r e n g t h / w e i g h t c h a r a c t e r i s t i c s of t h e foam-stabilized con- posit; s t r i n g e r o v e r t h e a n t i c i p a t e d load range. I n a d d i t i o n , t h i s type of c o n s t r u c t i o n i s s u i t e d t o t h e concept of i n t e - g r a l s h e l l c o n s t r u c t i o n .

The aluminum s t r i n g e r s with composite reinforcement are c l o s e s t t o t h e f o a m - s t a b i l i z e d com?osite s t r i n g s r on a btisis of w e i g h t However, t h e C s e c t i o n aluminum stri.rigex'8 ~ A J C U ~ ~ i n t r o d u c e problems of thermal bowing due t o the bond c u r e temperature of 250° F (394OK) and would r e q u j x i e c l i p p i n g at fraine l o c a t i o n s t o g i v e torsi.ona1 s t a b i l i t y . The aluininum e x t r u s i o n i n f i l t r a t e d with composite would have similar pi-oblems and a p o t e n t i a l l y high c o s t .

For t h e s e r e a s o n s , t h e foam-stabilized graphite/epoxy is considered t h e m o s t cost e f f e c t i v e concept f o r s t r i n g e r s t r i n g e r c o n s t r u c t i o n .

U N I D I R E C T I O N i ,COMPOSITE (a) CURRENT ALUMINUM STRINGER (b) S E L E C T I V E REINFORCEMENT 1 UNIDIRECTIONAL UNID I R E CT IOM /coMposlTE LAMINATE ALUMINUM EXTRUSION ( e ) S E L E C T I V E REINFORCETvlENT 2 (d) ALL COMPOSITE STRINGER

U N I D I R E C T I O U A L -

LAMINATE (e) COMFJOSITE/FOAM STRINGEX (f) ALL COMPOSITE STRINGER FIGURE 16. COMPOSITE STRINGER CONCEPTS.

. O E

.ox

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n !3 x

E *oat

!xi

s!

z i H a I3 .001

TYPICAL LOAD

- RANGE FOR CII-53D .DO2 2000 4000 6000 8 1 CRITICAL COMPRESSION LOAD, lbf t I 1 i 0 10 20 30 CRITICAL COMPRESSION LOAD, kN FIGURE 1 7 . FOAM STABILIZED COMPOSITE STRINGER HAS LOVXST WEIGHT OVER REQUIRED LOAD RANGE.

3.1.3 Skin Construction The current CH-S3D outer skin in the cabin and aft sec- tion region is aluminum sheet, mainly of 0.025 in. (0.64mm) o r 0.032 in. (0.81mm) gage. The light gage composite laminate of equivalent shear strength would result in a low shear buckling allowable and.be damage prone. The use of sandwich panel construction for the outer skin would reduce both these problems. Xowever, a comparison of this type of construction with the current aluminum skin, on a weight basis, shows an adverse result. F o r composite sandwich panels, the sketch below indicates the minimum dimensions considered practicable from manufacturing considerations and normal service require- ments for minimum gages of airframe outer skin.

= 0.02 IN. (0.5l~1m)

1 to

= 0.25 IN. (6.35mm) ADHESIVE t core - 3 .

.06 lbn i/FT-/ layer

f = 0.01 IN. (0.25mm)

t i The following comparison is made: Sandwich panel W. lbm/ft2(kg/rn2 1 Graphite/epoxy skins ,255 (1.243) Core 4.01bm/ft3 (64.1 kg/m3) ,083 ( .405) Adhesive m120 ( .585) Total for panel cI__ ,458 (2,233) .O32lr aluminum skin (typ. f o r current airframe skins) .460 (2.243)

-

This comparison indicates essentially no weight saving potential for a sandwich panel replacing only the airframe skin.

A more efficient use of sandwich panel construction would be as a replacement for skin and supporting structure. How- ever, for a large transport helicopter, such as the CH-53D, the internal concentrated loads require deep frames f o r struc- tural efficiency, Therefore, the sandwich construction would be limited to only replacing the skidstringer combination, For the CH-53D a typical aluminum skidstringer combination has a weight of . 6 5 1bm/ft2 (3.170 kg/m*). The required 20c E !

\ z 2 30 * c-7 R X m

g 20

0 ' 1oc 0 ,002 .006 .o t p S K I N WEIGET P W T E R , l b m / I N 2 c I I I 1 0 1.0 2 .o 3.0 4 .O t p SKIN WEIGHT PARAMETER, kg/m2 FIGURE 18. GRAPHITE/EPOXY S K I N HAS HIGHER SHEAR BUCKLING E F F I C I E N C Y THAN BOROK/EPOXY OR ALUMINUM S K I N . ' g r a p h i t e / e oxy sandwich c o n s t r u c t i o n i s e s t i m a t e d t o weigh

. 7 0 l b m / f t 3 ( 3 . 4 1 3 kg/m2). Therefore, it is concluded t h a t a

more e f f i c i e n t s t r u c t u r a l concept i s r e q u i r e d .

The a l t e r n a t i v e concept of a s i n g l e composite laminate as an o u t e r s k i n r e q u i r e s t h e comparison of composite s h e a r buck- l i n g c h a r a c t e r i s t i c s w i t h thos;! of an aluminkm p a n e l , s i n c e t h e c r i t i c a l design c r i t e r i o n f o r t h e s k i n i s t h a t no buckling ~ occur a t a 1.Og f l i g h t load c o n d i t i o n (see S e c t i o n 2 . 2 . 6 ) .

This comparison is i l l u s t r a t e d i n Figure 1 8 , which shows t h e s u p e r i o r shear buckling e f f i c i e n c y of t h e - + 4 5 O graphite/epoxy.

Based on t h i s comparison, t h e concept employed f o r t h e o u t e r

s k i n p a n e l s is a s i n g l e graphite/epoxy laminate o f balanced - +4S0

p l i e s .

3.1.4 S k i n / S t r i n g e r Panel Construction The concepts p r e s e n t e d f o r s t r i n g e r and s k i n c o n s t r u c t i o n are f u r t h e r v e r i f i e d by comparison of t h e i r combined l o a d i n g s t r e n g t h as a s k i n / s t r i n g e r panel combination w i t h t h e equiva- l e n t s k i d s t r i n g e r p a n e l of aluminum c o n s t r u c t i o n as i n t h e c u r r e n t airframe.

This comparison is p r e s e n t e d i n Figure 1 9 , which shows t h e h i g h e r combined s t r e n g t h c a p a b i l i t y of t h e composite s k i n p a n e l , p a r t i c u l a r l y i n t h e h i g h e r s h e a r flow r e g i o n . The com- p r e s s i o n a l l o w a b l e load f o r t h e aluminum c o n s t r u c t i o n i s t h a t for a s t a n d a r d aluminum C s e c t i o n s t r i n g e r + s k i n , as i l l u s t r a - t e d i n Figure 1 6 , and t h e composite s t i f f e n e r i s designed for t h i s l o a d .

3.1.5 Frame Construction Concepts considered f a r frame basic cross s e c t i o n are shown i n Figure 2 0 t o g e t h e r w i t h the w e i g h t / u n i t i e n g t h f o r each concept, based on a t y p i c a l l o a d i n g p o i n t and f r a m e depth r e q u i r e d i n t h e lower r e g i o n of t h e c a b i n frames. The 2 0 c , d , and e are seen t o have s i m i l a r concepts of Figures weight p r o p e r t i e s . For c o n s i d e r a t i o n of frames i n the c a b i n r e g i o n , the foam s t a b i l i z e d frame ( F i g u r e 2 0 d ) can be i n t e g r a t e d i n t o t h e concept of an i n t e g r a l s h e l l layup-and-cure c y c l e without p r i o r d e t a i l f a b r i c a t i o n of p a r t s , which is r e q u i r e d f o r t h e concepts of F i g u r e s 2 0 c , e.

For r e g i o n s of the s t r u c t u r e i n which d e t a i l p a r t f a b r i - c a t i o n i s considered t o precede the f i n a l assembly, t h e concept of Figure 2 0 c is considered an a c c e p t a b l e a l t e r n a t i v e .

WEB STABILIZING STlF FEN ERS (a) BUILT UP [b] SELECTIVE (c] COMPOSITE ALUMINUM REINFORCEMENT SANDWICH ALUMINUM NIDIRECTIONAL RECTIONAL EXTRUSION OSITE 90° FIBERGLASS COMPOSITE (dl COMPOSITE/ (el BUILT UP (f) COMPOSITE F O A M COMPOSITE INFILTRATED ALUMINUM

(J ~ ~ 1 0 0 , 0 0 0 I N . #

/ 1.4- # 10,000 2.0- 1.2-- STIFF.

T Y P I C A L D E S I G N LOADING E \ m a1.5- E 1.0-- STIFF. STIFF.

e- \ I €

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Lu + - 0.8--

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WEB 3 I U J

1.0- z

ADHfSlLE STIFF.

2 3 0.6-- WEB & U J FOAM WEB z CORE 0.4-- 0.5- CAPS WEB WEB CAPS CAPS CAPS CAPS

. " ' L

~ a b C d f a

WEIGH!" COMPARISON - COMPOSITE BEAM CONCEPTS

' (COMPOSITE MATERIAL CONSIDERED TO BE GWHITE/EPOXY) F I G U R E 20. S T A B I L I Z E D COMPOSITE FRAMES ME LIGHTER T M CuRR;ENT COIJSTRUCTION .

3.1.6 S h e l l Construction The concepts p r e s e n t e d f o r s t r i n g e r , s k i n , and frame con- s t r u c t i o n can be i n t e g r a t e d t o produce t h e concept f o r com- p o s i t e s h e l l c o n s t r u c t i o n . The concept is i l l u s t r a t e d i n Figure 21. The s t r i n g e r s are considered continuous, passing through a cut-out i n t h e frame web, and are jogged over t h e o u t e r frame c a p t o p r e s e r v e c o n t i n u i t y of frame bending cap- a b i l i t y .

3.1.7 . Floor Construction Cargo f l o o r design i s governed by t h e requirements of Reference 17 which s p e c i f i e s both d i s t r i b u t e d cargo loading and l o c a l wear and i n p a c t loading. The c u r r e n t f l o o r construc- t i o n is of aluminum sheet bonded t o C s e c t i o n aluminum e x t r u - s i o n s .

The concept considered f o r the composite airframe i s a hybrid sandwich p a n e l , shown i n Figure 2 2 , which a l s o i n d i c a t e s t h e v a r i a t i o n i n f l o o r weight f o r t h e composite materials considered. T h i s concept i s s i m i l a r t o a design c u r r e n t l y being evaluated by Sikorsky A i r c r a f t f o r f u t u r e h e l i c o p t e r a p p l i c a t i o n . An upper face sheet of t i t a n i u m , t o g e t h e r w i t h an upper l a y e r of dense aluminum core, i s considered i n order t o meet t h e design c o n d i t i o n s of s u r f a c e wear and local impact loading. The lower face sheet of composite material t o g e t h e r with t h e lower layer. of less dense aluminum honeycomb provide t h e necessary beam depth t o c a r r y t h e f l o o r bending moments and shears.

From Figure 2 2 , it can be seen t h a t little weight v a r i a t i o n occurs w i t h m a t e r i a l v a r i a t i o n for t h e composite el.eiiients of t h e f l o o r , u i t h the exception of PRD-4S/epaxy, due to its low

co~npr~essio~i str'ength ( s e e Table 2 ) . For tliis r e a s o n , fibex-glass

is chosen as t h e composite material for t h e floorb panels, due t o i t s much lower c o s t t h a n t h e o t h e r composites. On an over- a l l cost basis, a comparison with t h e c u r r e n t aluminum construc- t i o n shows t h e h y b r i d panel c o n s t r u c t i o n t o be less c o s t l y .

With t h e r e d u c t i o n i n f l o o r weight, shown by Figure 2 2 , t h e o v e r a l l concept i s judged cost e f f e c t i v e .

6.0 (TYP s: 4.0 hJ f3 Ert \ E ft n ~ 3 . 0

F

I I

Er; I %I 2.0 1 . 0

-t----

0.2 0.4 0.6 0.8 1.0 1.2 1.4 1 . 6 FLOOR T H I C K N E S S , I N .

I I I 1 0 10 20 30 FLOOR T H I C K N E S S , mm FIGURE 22. LOW COST F I B E R G L A S S HYBRID FLOOR CAN REDUCE WEIGHT BY 28 PERCENT.

.04

COMPRESSION LOADDANEX WT - lbf/lbm

I I I I I I I i 0 100 200 300 100 500 600 700 COMPRESSION LOAD/PANEL W l - N/kg G W H I T E / E P O X Y AND BORON/EPOXY SKIN/STRINGER PANELS FIGURE 19.

HAVE HIGHER COMBINED LOAD STRENGTH/WE3GHT RATIO THAN ALUMINUM PANELS.

WRTINU0V:I FEAW CAP (VIIIDIRECTIOIIAL) FIGURE 21. COMPOSITE SHELL CONCEPT.

3.1.8 Connections S t r u c t u r a l connections are considered t o be bonded where- e v e r p r a c t i c a b l e . For mechanical j o i n i n g , i n v o l v i n g f a s t e n e r h o l e s through composite l a m i n a t e s , t h e a s s o c i a t e d stress concen- t r a t i o n s r e q u i r e t h a t a d d i t i o n a l l o c a l material be added t o reduce the stresses t o an a c c e p t a b l e l e v e l o r t h a t some o t h e r means be used f o r j o i n t r e i n f o r c e m e n t , such as bearing s t r i p s bonded t o t h e composite material. Each of t h e s e e f f e c t s i n c u r s a weight and c o s t p e n a l t y .

However, t h e high s t r u c t u r a l e f f i c i e n c y of w e l l designed bonded j o i n t s , such as t h e stepped l a p j o i n t , i s o f f s e t , t o some e x t e n t , by t h e i n c r e a s e d manufacturing c o s t caused by t o l e r a n c e c o n t r o l and p o s s i b l e machining a t t h e connection i n t e r f a c e . For these r e a s o n s , emphasis is placed on bonded connections of simple c o n s t r u c t i o n .

Exceptions t o the concept of all-bonded connections are made a t t h e mating faces of the major s t r u c t u r a l assemblies and subassemblies. Consideration of mechanical connections a t t h e s e l o c a t i o n s w i l l assist i n assembly and f a c i l i t a t e d i s - assembly f o r p o s s i b l e i n - s e r v i c e replacement of s e v e r e l y dam- aged s e c t i o n s of s t r u c t u r e .

A t the mating faces of t h e c a b i n subassemblies, a s k i n s h e a r s p l i c e is r e q u i r e d . The concepts considered f o r t h i s s k i n s p l i c e are shown i n Figure 2 3 , which i n d i c a t e s the f i n a l design concept used. I n a d d i t i o n t o t h e skin shear s p l i c e , t h e cabin frames r e q u i r e a s h e a r a n d moment s p l i c e a t this location, and t h e concepts considered f o r this connection are stiown i n Figure 2 4 .

A t the m a t i n g faces of the c a b i n region with t h e cockpit and a f t fuselage s e c t i o n , the axial m z m b e m require a s p l i c e connection i n a d d i t i o n t o t h e s k i n shear s p l i c e . Concepts considered for t h i s connection are shown i n Figure 2 5 .

I

w U B \ a N \ 3.1.9 F a b r i c a t i o n Concept 3 . 1 . 2 through 3.1.8 are The concepts p r e s e n t e d i n s e c t i o n s i n t e g r a t e d i n t o a n o v e r a l l concept f o r airframe c o n s t r u c t i o n and assembly and are i l l u s t r a t e d , f o r t h e c a b i n s e c t i o n , i n F i g u r e s 2 6 through 2 9 .

The c a b i n s e c t i o n i s conceived as formed f r o m f o u r sub- assemblies, with each subassembly formed i n o n l y t w o c u r e c y c l e s . The first of t h e s e c y c l e s i n v o l v e s t h e o u t e r s k i n , o u t e r frane c a p s , s t r i n g e r s , and s p l i c e f i t t i n g s shown i n F i g u r e s 2 6 and 2 7 . The f l a t p a t t e r n u n i d i r e c t i o n a l l a m i n a t e f o r t h e s t r i n g e r i s first l a i d i n t o s t r i n g e r - s h a p e d pockets i n a s p l i t male mold f o r m . Molded-foam s t r i n g e r c o r e s are t h e n i n s e r t e d o v e r t h e s t r i n g e r l a m i n a t e s followed by t h e u n i d i r e c t i o n a l l a m i n a t e s t o form t h e o u t e r frame.caps. T o m i n i n i z e problems a t s t r i n g e r / f r a m e i n t e r s e c t i o n s , t h e o u t e r frame caps are t a p e r e d i n t h e width d i r e c t i o n . The o u t e r s k i n is t h e n l a i d up o v e r t h e mold s u r f a c e t o g e t h e r w i t h t h e t i t a n i u m s h e e t i n s e r t s (see Figure 2 6 ) and t h e s t r i n g e r end f i t t i n g s (see F i g u r e 2 7 ) . These elements are t h e n co-cured i n the first c u r e c y c l e .

The second c u r e c y c l e i n v o l v e s t h e frames, shown i n F i g u r e 28. The cured first s t a g e i s removed f r o m t h e s p l i t m a l e mold arid t r a n s f e r r e d t o a female mold. The molded foam frame c o r e is l a i d up o v e r t h e o u t e r f r a m e caps followed by t h e uni- d i r e c t i o n a l l a m i n a t e s f o r t h e frame i n n e r cap. The f l a t p a t t e r n layup of t h e +4S0 l a m i n a t e , and the frame s p l i c e f i t t i n g s (see F i g u r e 2-91 are t h e n added, and t h e second c u r e c y c l e i s performed t o c o m p l e t e t h e subassenttly. At t h i s s t a g e , d e t a i l trimming of s k i n c u t o u t s i s completed.

Vacuum l i f t i n g equipment i s considered f o r h a n d l i n g of parts i n t h e subassembly and final. assembly s t a g e .

The f i n a l assembly i s completed by mechanically connecting t h e f o u r subassemblies, as shown i n F i g u r e 30.

It is p o s s i b l e t o f a b r i c a t e c a b i n subassemblies u s i n g o n l y one c u r e c y c l e , t h u s r e d u c i n g t h e f a b r i c a t i o n costs. There i s however, t h e problem of t h e l a m i n a t e s t r e n g t h p r o p e r t y of t h e s t r i n g e r a t t h e s t r i n g e r / f r a m e i n t e r s e c t i o n . T h e r e f o r e , a t t h i s t i m e , t h e c o n s e r v a t i v e approach of two c u r e c y c l e s is used f o r cost e s t i m a t i o n purposes.

Other airframe a s s e m b l i e s are c o n s i d e r e d s u i t a b l e f o r f a b r i c a t i o n and assembly i n a manner s i m i l a r to t h a t ind'icated f o r t h e c a b i n . Detail drawings of t h e f i n a l airframe concepts are p r e s e n t e d i n S e c t i o n 4 . 0 , and a schematic i n d i c a t i o n of t h e complete airframe assembly i s shown i n F i g u r e 31.

E- m

F R E C E T ~ G PAGE BLANP; BOT m

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PRECEDING PAGE BLANK NOT FXLMED 3.2 L'liJDIiJrJ GEAR STRUCTURE The rolling gear is not included in the study of concepts for composite material application. This gear consists of wheels, tires, brakes, and miscellaneous hardware. With the exception of the wheels, all items of the rolling gear are not considered replaceable by equivalent parts of composite construc- tion. For the wheel, which is currently an aluminum forging, the only practical concept f o r composite material construction is considered to be a molded form of short fiber/epoxy material. Section 3.1.1 indicates that this material is not competitive on a cost effective basis compared with the .

aluminum forging.

The remainder of the landing gear consists of major steel and aluminum forgings together with miscellaneous hardware and nonmetallic elements. The structural elements considered f o r construction of composite materials are the oleo trunnion, shock strut, the drag strut cylinder and piston, and the torque arms. The current construction of the main landing gear is shown in Figure 3 , with the nose landing gear being of similar construction-.

The choice of elements of the landing gear considered unsuitable f o r replacement by equivalent parts of composite material is substantiated by other work in this field, such as Reference 18.

3.2.1 Trunnion Concepts considered for this p a r t are shown in F i g u r e 32.

F o r the all-composite construction, two mater*ials are potential candidates : short fiber/cpoxy m o l d i n g or boron/al umiiiurn w i . t f - i brazed connections, as outlined in section 3.1.1. The use of short fiber/epoxy material does not a p p e a r c o s t effective. Use of bcir*onialumirium composite ciarerial is 1 inited by t h e low tension strength allowable f o r O o , 9 0 ° laminates when based on the criteria illustrated by Figure 14. Using data from Reference 9 for unidirectional boron/aluminum laminates, the allowable tensile stress is reduced to 32.5 ksi (0.224G1J/m2) when the criteria of Figure 14 are applied. In addition, boron/aluminum construction is not considered comparable on a cost basis with the highly developed aluminum forging process.

F o r these reasons, the concept judged to have cost effective potential is that of selective replacement of the simple shaped elements of the trunnion with cylindrical shapes built up from graphite/epoxy laminates. The complex shape and loading of the connections require metallic fittings at these F o r a bonded joint connection of these fittings to the points.

composite cylinders, the weight of metal replaced in the inclined I (a) ALL COMPOS;TE CONSTRUCTION (b) SELECTIVE REPLACEMENT NOT COST EFFECTIVE METALLIC FITTINGS AT JOINTS, MANY PARTS, HIGH COST FITTINGS I ( c 1 COMPOSITE ~ENTRAL CYLINDER FIGURE 32. COMPOSITE CENTRAL CYLINDER IS MOST PRACTICAL CONCEPT FOR LANDING GEAR TRUNNION.

and h o r i z o n t a l a r m s of t h e t r u n n i o n would be very s m a l l . The concept considered t o have t h e g r e a t e s t p o t e n t i a l i s t h a t i n which the c e n t r a l c y l i n d r i c a l s e c t i o n a l o n e is r e p l a c e d by composite material (see Figure 3 2 c ) .

3.2.2 A x i a l l y Loaded Members The o l e o shock s t r u t and d r a g s t r u t p i s t o n and c y l i n d e r are similar i n t h a t t h e i r design l o a d i n g c o n d i t i o n s are a x i a l , although t h e o l e o shock s t r u t also carries s h e a r and bending moment due t o wheel d r a g and side l o a d s .

Each of these members i s e s s e n t i a l l y of c y l i n d r i c a l form, with complex machined d e t a i l s a t each end. Following t h e con- c e p t s o u t l i n e d for t h e t r u n n i o n i n s e c t i o n 3 . 2 . 1 , t h e concept considered for each of these members is a composite c y l i n d e r bonded at each end t o s t e e l f i t t i n g s t h a t c o n t a i n t h e n e c e s s a r y d e t a i l e d machined f e a t u r e s . This concept i s shown i n Figure 33.

3.2.3 Torque A r m s Concepts considered f o r t h e l a n d i n g g e a r t o r q u e a r m s are shown i n Figure 34. The all-composite concept is n o t consider- ed cost e f f e c t i v e , based on the r e a s o n i n g o u t l i n e d i n s e c t i o n 3.2.1 f o r t h e l a n d i n g g e a r t r u n n i o n . The a l t e r n a t i v e concepts shown i n Figure 34 are s i m i l a r i n t h a t t h e y both c o n t a i n many p a r t s , w i t h associated high manufacturing and assembly cost.

The t o r q u e a r m s r e p r e s e n t only 5.0 p e r c e n t of t h e landing g e a r s t r u c t u r e weight, and t h e i n h e r e n t complexity of the s t r u c t u r e for o t h e r t h a n a one-piece c o n s t r u c t i o n , s u g g e s t s a low cost e f f e c t i v e p o t e n t i a l . For t h e s e l’easoxis, tila t o r q u e arms a r e n o t considered changed from c u r r e n t c o n s t r u c t i o n in the composite l a n d i n g g e a r s t r u c t u r e .

3.2.4 F a b r i c a t i o n Concepts The f i n a l d e s i g n concept for t h e l a n d i n g g e a r s t r u c t u r e is shown i n Figure 35. Details for i n d i v i d u a l components are p r e s e n t e d i n S e c t i o n 4 . 0 .

For p r o t o t y p e c o n s t r u c t i o n of t h e composite c y l i n d r i c a l elements of t h e l a n d i n g g e a r , a hand layup sequence would be used. F o r production a n automated f i l a m e n t winding or r o l l i n g p r o c e s s i s considered most e f f e c t i v e , with trimming and machin- ing r e q u i r e d p r i o r t o bonding ’to t h e steel e n d - f i t t i n g s .

5 0 N t m

SECTION 4.0 COMPOSITE DES1 GI4 APPLI CAT1 ON composites t o t h e CH-53D s i m p l i f i e s 9 A p p l i c a t i o n o f c o n s t r u c t i o n and reduces t h e s t r u c t u r e weight by 1 1 8 6 lbm (538 k g ) , or 18.0 p e r c e n t .

For purposes of t h i s s t u d y , it w a s assumed t h a t t h e i n t e r n a l l o a d i n g d i s t r i b u t i o n s remain e s s e n t i a l l y unchanged i n t h e composite s t r u c t u r e f r o m t h o s e i n t h e c u r r e n t s t r u c t u r e .

The l o a d s used for! design are, t h e r e f o r e , taken from t h e Sikorsky Load and S t r e s s Reports for t h e c u r r e n t CH-53D s t r u c t u r e .

Major f e a t u r e s of t h e f i n a l d e s i g n are p r e s e n t e d i n t h e f o l l o w i n g s e c t i o n s . A d e s c r i p t i o n i s p r e s e n t e d for each of t h e n i n e major a s s e m b l i e s i n d i c a t e d i n Figure 4 . A summary of comparative weights for each major assembly is p r e s e n t e d i n F i g u r e 36 and Table 3.

4.1 AIRFRAME D e t a i l e d a n a l y s i s f o r t h e composite airframe r e s u l t s i n a weight r e d u c t i o n of 1118 lbm (507.1 kg) (18.5 p e r c e n t ) compared w i t h t h e c u r r e n t s t r u c t u r e . Major usage of composite material i s i n graphite/epoxy E1872 lbm (849.1 k g ) l and PRD-49-III/epoxy E466 lbm (211.4)kgl. A breakdown of material usage i n t h e com- p o s i t e airframe i s p r e s e n t e d i n Table 4. An o v e r a l l view of t h e composite airframe s t r u c t u r e is shown i n Figure 37.

Appendix B c o n t a i n s weight t r e n d c u r v e s for t h e major airframe a s s e m b l i e s , w i t h t h e composite design p o i n t i n d i c a t e d on each curve.

MAIN ROTOR I I CURRENT STRUCTURE

PYLON F A I R I N G (13%) COMPOSITE S T R U C m E T A I L PYLON (19%) FLOOR (11%) FOR TOTAL AIRFRAME AM) LANDING GEAR, T H E WEIGHT I S I I REDUCED BY 18% LANDING GEAR (13%) (SEE TABLE 3 FOR NUMXRICAL DATA) C O C K P I T (16%) SPONSON ( 15% )

I CABIN (22%)

I I I I

500 1000 1500 2000

STRUCTURE WEIGHT - l b m

w ~ I I I - 1 I 400 600 800 1000

STRUCTURE WEIGHT - kg

F I G m 36. A L L - C O W O S I T E STRUCTURAL ASSEMBLY UXIGHTS ARE REDUCED, COMPARED WITH THE CURRENT STRUCTURE.

PRECEDING PAGE BUNK, NOT FlLMED I I I I I I I 4.1.1 Cockpit S e c t i o n The upper c o c k p i t s t r u c t u r e i s c u r r e n t l y of one-piece molded f i b e r g l a s s c o n s t r u c t i o n . Since it a l r e a d y embodies t h e manufacturing advantages of composite materials, no s u b s t a n t i a l change i s c o n s i d e r e d . The t y p i c a l s e c t i o n s shown i n Figure 38 i n d i c a t e t h e complex s t r u c t u r a l shapes t h a t are a t t a i n a b l e w h i l e m a i n t a i n i n g t h e one-piece c o n s t r u c t i o n concept. A r e d u c t i o n i n weight i s o b t a i n e d by c o n s i d e r a t i o n of PRD-49/ epoxy r e p l a c i n g t h e c u r r e n t f i b e r g l a s s material. Details of t h i s r e g i o n are shown i n F i g u r e s 38 and 39.

The lower c o c k p i t r e g i o n is c u r r e n t l y of b u i l t - u p con- s t r u c t i o n w i t h complex f i t t i n g s and m u l t i p l e c u t o u t s f o r t h e v a r i o u s systems l o c a t e d i n t h e r e g i o n . An e v a l u a t i o n of t n i s r e g i o n i n d i c a t e d l i t t l e p o t e n t i a l - f o r cost e f f e c t i v e a p p l i c a - t i o n of composite material, and t h e local s t r u c t u r e i s r e t a i n e d .

4.1.2 Cabin S e c t i o n The c u r r e n t CII-53D c a b i n s e c t i o n i s e s s e n t i a l l y a s t r e n g t h - designed aluminum s k i n / s t r i n g e r / f r a m e s h e l l and r e p r e s e n t s 4 0 p e r c e n t of t h e airframe s t r u c t u r e weight. The composite s h e l l c o n c e p t s developed i n S e c t i o n 3 are a p p l i e d t o t h i s s e c t i o n .

The l i g h t gage aluminum o u t e r s k i n is c o n s i d e r e d r e p l a c e d by a s i n g l e l a m i n a t e of +45O g r a p h i t e / e p o x y . Due t o t h e t h i n l a m i n a t e s t h a t r e s u l t from t h i s approach, PRD-49/epoxy is used f o improve t h e impact s t r e n g t h of t h e graphite/epoxy. A l a m i n a t e c o n t a i n i n g 2 0 p e r c e n t PRP-49 and 8 0 p e r c e n t g r a p h i t e s h o u l d raise t h e impact s t r e n g t h of t h e graphite/epoxy t o t h e minimum c o n s i d e r e d n e c e s s a r y for s e r v i c e use ( s e e s e c t i o n 2 . 2 . 5 ) .

Axial members (strlingerts arid longerons 1 are c o n s i d e r e d of f o a m - s t a b i l i z e d g r a p h i t e / e p o x y c o n s t r u c t i o n , as shown in F i g u r e 40. A s i m i l a r concept is used fori frane c o n s t r u c t i c n , as iridi- c a t e d i n F i g u r e 4 0 . For t h o s e framzs t h a t are c u r r e r i t l y of f o r g e d c o n s t r u c t i o n , t h e h e a v i l y loaded r e g i o n s w i t h complex f o r g e d d e t a i l s are c o n s i d e r e d r e p l a c e d by t i t a n i u m f i t t i n g s t h a t are bonded t o t h e composite frame. T y p i c a l frarnes are shown i n F i g u r e 4 0 .

For ease of f a b r i c a t i o n and assembly, t h e c a b i n s t r u c t u r e is c o n s t r u c t e d as f o u r subassembly u n i t s t h a t are mechanically connected t o f o r m t h e f i n a l assembly. The subassembly connec- t i o n d e t a i l s are shown i n F i g u r e 30.

5 8 F P

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> - a .

Typical design a n a l y s i s c a l c u l a t i o n s are p r e s e n t e d h e r e f o r some l o c a l d e t a i l s o f t h e c a b i n s t r u c t u r e .

Skin Construction The c u r r e n t aluminum s k i n is considered r e p l a c e d by a s i n g l e graphite/epoxy l a m i n a t e (see s e c t i o n 3.1.3).

Typical Analysis Lower c a b i n s k i n s between FS 2 2 2 and 302 are 0.025 i n .

(0.535mm) t h i c k and a r e designed f o r a naximum u l t i m a t e s h e a r flow of 1 5 7 l b f / i n . (27.4kN/m) w i t h a n u l t i m a t e v e r t i c a l l o a d factor of 4.5.

From t h e c r i t e r i a o f non-buckled s k i n f o r a l o a d f a c t o r of 1 . 0 (see s e c t i o n 2 . 2 . 6 1 , t h e c r i t i c a l buckling s h e a r f l o w q c r i s given by

q c r - - 4.5 = 34.9 l b f / i n . (6.lkI?/m)

From Figure 9 , t h e r e q u i r e d +-4S0 H?IS graphite/epoxy l a m i n a t e t h i c k n e s s f o r qcr = 34.8-lbf/in. ( 6 . 1 kX/m) i s 0.034 i n . (0.862mm). Following t h e concepts f o r laminate layup and p l y t h i c k n e s s o f s e c t i o n 3 . 1 . 1 , t h e d e s i g n l a m i n a t e i s a balanced four-ply - +4S0 laminate 0 . 0 4 i n . (1.02mm) t h i c k .

Aluminum s k i n weight = 0 . 3 6 0 l b m / f t * ( 1 . 7 6 kg/m2) Composite s k i n weight = 0 . 4 0 6 l b m / f t 2 (1.98 kg/m2) The h i g h e r weight f o r t h e graphite/epoxy s k i n , i n c l u d i n g 2 0 p e r c e n t PRD-49/epoxy3 i n d i c a t e s t h e d i f f i c u l t y of weight r e d u c t i o n f o r very t h i n gage aluminum s k i n . However, t h e f o u r - p l y graphite/epoxy laminate is also used t o r e p l a c e t h e 0 . 0 3 2 i n . (0.813mm) gage s k i n s t h a t form t h e major p o r t i o n of t h e c a b i n s k i n . For t h i s Fag?, t h e aluminum s k i n weight i s 0 . 4 6 1 l b i d f t 2 ( 2 . 2 5 kg/m2), i n d i c a t i n g a 1 2 p e r c e n t r e d u c t i o n for t h e composite c o n s t r u c t i o n . S i m i l a r weight r e d u c t i o n s are .

o b t a i n e d i n t h e h e a v i e r gage s k i n s around door and window c u t - outs.

6 2 S k i n - S t r i n g e r Panel The graphite/epoxy s k i n is considered s t i f f e n e d by foam- s t a b i l i z e d graphite/epoxy s t r i n g e r s (see s e c t i o n 3 . 1 . 4 ) .

Typical Analysis A local maximum u l t i m a t e s t r i n g e r compression load of 3030 l b f (13.4kX) with an a s s o c i a t e d u l t i m a t e p a n e l s h e a r f l o w of 38 l b f / i n . (6,6kM/m) o c c u r s i n t h e c u r r e n t s t r u c t u r e i n t h e lower cabin s i d e w a l l a t FS 3 2 2 . The c u r r e n t c o n s t r u c t i o n i s a 0.025 i n . (0.635mm) t h i c k aluminum s k i n r i v e t e d t o a 0.04 i n .

(1.02mm) t h i c k aluminum s t r i n g e r , g i v i n g a panel weight of 0 . 6 1 l b m / f t 2 ( 2 . 9 8 kg/m2). For t h e foam-stabilized graphite/epoxy c o n s t r u c t i o n , t h e s t r i n g e r laminate is 0 . 0 2 i n . (0.508mm) t h i c k and is combined with a 0 . 0 4 i n . (1.02mm) t h i c k +45O g r a p h i t e / epoxy 5 k i n , i n c l u d i n g PXD-49, t o g i v e a panel wcight of 0.54 , l b d f t ( 2 . 6 4 kg/m2).

Therefore, t h e composite s k i n / s t r i n g e r panel i s 1 2 p e r c e n t l i g h t e r t h a n t h e c u r r e n t aluminum panel. I n r e g i o n s of n i g h e r s h e a r l o a d i n g r e q u i r i n g 0 . 0 3 2 i n (0.813mm) aluminum s k i n , t h e same four-ply g r a p h i t e / e p o x y laminate i s used, g i v i n g a panel t h a t is 1 7 p e r c e n t l i g h t e r t h a n c u r r e n t c o n s t r u c t i o n .

Frame Construction A foam-stabilized graphite/epoxy c o n s t r u c t i o n i s considered f o r t h e c a b i n frames (see s e c t i o n 3.1 .SI.

Typical A n a l y s i s The frame a t FS 402 is t y p i c a l for frames a t FS 4 0 2 , 4 2 2 , 462, 482, and 5 0 2 . The frame, shown i n t h e following sketch, is designed by floor and hydrodynamic l o a d i n g , I

Section A-A (see Sketch) Axial load Pm = -489 l b f ( u 1 t ) (-2.19kN) Shear load V u = 690 l b f ( u 1 t ) (3.10kN) Bending moment X u = - 9 8 8 6 i n l b f ( u 1 t ) (-111.7kmN) Frame depth d = 2.0 i n . (S0.8mm) Outer cap ( t e n s i o n ) Pt = 4699 l b f (21.0kM) Q t u = 160,000 l b f / i n 2 ( 1 . 1 G N / r n 2 ) (from Table 2 f o r €ITS graphite/epoxy) A r e a r e q u i r e d = 4699 = 0.029 i n 2 ( 1 8 . 7 u m 2 ) 160,000 This is provided by 4 p l i e s of u n i d i r e c t i o n a l g r a p h i t e / epoxy 0 . 0 1 i n . ( 0 . 2 5 4 m ) t h i c k by 0 . 7 5 i n . (19.Omm) wide.

Inner cap (compression) Pc = -5187 l b f (-23.2kN)

Qcu - - 1 8 3 , 0 0 0 l b f / i n 2 (1.26G11/m2)

( f r o m Table 2 f o r HTS graphite/epoxy) Area r e q u i r e d = 5187 = 0 . 0 2 8 i n 2 ( 1 7 . 9 p m 2 ) 183,000 This i s provided by a laminate of t h e same dimensions as those f o r t h e o u t e r cap.

Shear web V = 690 l b f (3.10kX) = 6 6 , 0 0 0 l b f / i n 2 (0.454GY/rn2)

;%&n Table 2 for - +4S0 graphite/epoxy)

= 6 9 0 = 0.002 i n . ( 0 . 0 5 1 m ) Web t h i c k n e s s r e q u i r e d 2 ~ 2 ~ 6 6 , 0 0 0

This requirement i s covered by t h e minimm web laminate t h i c k n e s s of 0.02 i n . ( 0 . 5 0 8 mm) considered f o r use i n t h e frames.

S e c t i o n B3 (see s k e t c h ) is a d e e p e r , more h i g h l y loaded section.

Axial load PBB = 1 8 4 6 l b f ( u l t ) (8.25kfJ) Shear load V B B = 7708 l b f ( u l t ) (34.60kN) Bending moment M B B = -157,023 i n . l b f Cult) (-1,770 kmN) Frame depth d = 7.0 i n . (177 mm) Following t h e a n a l y s i s o u t l i n e d for s e c t i o n AA above and maintaining t h e frame width of 0.75 i n . ( 1 9 . 0 m m ) , t h e i n n e r cap r e q u i r e s 1 6 x .01 i n . ( 0 . 2 5 4 mm) p l i e s of u n i d i r e c t i o n a l ' graphite/epoxy and t h e o u t e r cap r e q u i r e s 2 1 p l i e s . The web requirement is maintained a t t h e mininum of 2 x .01 i n .

(0.254 mm) p l i e s of + 4 5 O graphite/epoxy. The deep cap l a m i - n a t e s r e q u i r e d a t s e c t i o n BB are t a p e r e d , as t h e l o a d i n g de- creases, t o t h e f o u r p l i e s r e q u i r e d a t s e c t i o n AA.

The composite frame a t FS 4 0 2 h a s a weight of 23.0 lbm (10.42 kg) compared w i t h 35.0 lbm (15.86 kg) f o r t h e c u r r e n t aluminum frame. This 34 p e r c e n t r e d u c t i o n i n weight is t y p i c a l f o r t h e l i g h t l y loaded c a b i n frames. S i m i l a r a n a l y s i s for t h e l a n d i n g g e a r support frame a t FS 4 4 2 (shown i n Figure 40) shows a composite frame weight of 9 6 . 0 lbm ( 4 3 . 5 kg) com- pared with t h e c u r r e n t weight of 1 3 6 . 0 lbm ( 6 1 . 6 kg). T h i s r e p r e s e n t s a 29 p e r c e n t weight r e d u c t i o n , which i s t y p i c a l for t h e more h e a v i l y loaded c a b i n frames.

F o r t h e c a b i n assembly, t h e f i n a l weight rleduction is 2 2 p e r c e n t , compared with c u r r e n t c o n s t r u c t i o n , as i n d i c a t e d i n F i g u r e 36.

6 5 _r 4.103 Sponson S e c t i o n The primary s t r u c t u r a l members i n t h e c u r r e n t s t r u c t u r e are t h e bulkheads connected t o t h e f u s e l a g e a t FS 3 0 2 , 3 8 2 , and 442. These c a r r y t h e v e r t i c a l l o a d i n g from t h e l a n d i n g g e a r and f u e l i n e r t i a l o a d i n g . I n a d d i t i o n main l a n d i n g g e a r d r a g l o a d s are c a r r i e d by major beams i n t h e a f t sponson r e g i o n .

These primary members d i c t a t e t h e i n t e r n a l load d i s - t r i b u t i o n i n t h e sponson and a d j a c e n t cabin s t r u c t u r e . They are, t h e r e f o r e , n o t considered removeable i n t h e composite d e s i g n . A s i m p l i f i c a t i o n of design is achieved f o r the bulk- heads by f a b r i c a t i n g them as sandwich panels with i n t e g r a l s t i f f e n e r s . The h i g h l e v e l of shear s t a b i l i t y , i n p l a n e s t r e n g t h and p l a t e bending s t r e n g t h of t h i s t y p e of c o n s t r u c t i o n make it s u i t a b l e f o r both t h e l a n d i n g g e a r bulkheads and f u e l c e l l bulkheads. Tne h i g h s p e c i f i c s t r e n g t h of t h e graphite/epoxy is f u l l y u t i l i z e d i n t h i s c o n s t r u c t i o n , due t o t h e i n h e r e n t s t a b i l i t y of t h e sandwich panel.

The remaining sponson s t r u c t u r e is t h e e x t e r n a l s k i n i n s t a l l a t i o n , t o g e t h e r with s u p p o r t i n g r i b s and i n t e r c o s t a l s .

Consideration of sandwich p a n e l s i n r e g i o n s o f l o w c u r v a t u r e e n a b l e s much of t h e s u p p o r t i n g s t r u c t u r e t o be renoved, which y i e l d s b e n e f i t s of reduced manufacturing t i m e and lower weight. I n r e g i o n s of high c u r v a t u r e , t h e i n h e r e n t s t a b i l i t y and load c a r r y i n g c a p a c i t y due t o shape enable s u p p o r t i n g s t r u c t u r e t o be removed without u s i n g sandwich c o n s t r u c t i o n .

PR3-49 i s used, w i t h t h e graphite/epoxy, t o improve t h e impact r e s i s t a n c e c h a r a c t e r i s t i c s of t h e o u t e r s k i n s .

Details of t h e f i n a l design are shown in Figure 4 1 .

c- W -j $ F ,

t

f

rl f , 4.1.4 A f t S e c t i o n The a f t s e c t i o n of t h e airframe is e s s e n t i a l l y an e x t e n s i o n of t h e c a b i n s e c t i o n , b u t i n c l u d e s t h e cargo loading ramp and overhead door, which produce, i n t h i s r e g i o n , a "C" s e c t i o n s h e l l s t r u c t u r e . Design d e t a i l s f o r t h i s s t r u c t u r e are s i m i l a r t o t h o s e p r e s e n t e d i n s e c t i o n 4 . 1 . 2 f o r t h e c a b i n .

The cargo l o a d i n g ramp becomes, i n t h e open p o s i t i o n , an e x t e n s i o n of tne cargo f l o o r and i s s u b j e c t e d t o f l o o r design l o a d i n g . The c u r r e n t i n n e r s k i n of supported aluminum s h e e t i s r e p l a c e d by t h e h y b r i d sandwich panel used f o r t h e cargo f l o o r (see s e c t i o n 4 . 1 . 5 ) . I n t h e open p o s i t i o n , t h e o u t e r s u r - face of t h e cargo ramp rests on t h e l o c a l ground s u r f a c e t o p r o v i d e , with t h e h i n g e s , t h e ranp s u p p o r t . Due t o t h e uncer- t a i n n a t u r e of l o c a l ground s u r f a c e i n s e r v i c e use, t h e o u t e r s k i n of t h e ramp is maintained as an aluminum s h e e t . This is due t o c o n s i d e r a t i o n of t h e lower impact s t r e n g t h of g r a p h i t e / epoxy laminates (see Figure 1 5 1 , which may n o t be adequate f o r this l o c a l area.

The i n t e r n a l support beams of t h e ramp, which are c u r r e n t l y of b u i l t - u p aluminum c o n s t r u c t i o n , are r e p l a c e d by composite sandwich beams of t h e t y p e shown i n Figure 2 0 c .

4.1.5 F l o o r S e c t i o n The concept f o r f l o o r c o n s t r u c t i o n p r e s e n t e d i n s e c t i o n 3.1.7 and shown i n Figure 2 2 i s analyzed f o r t h e l o a d i n g requirements of Reference 1 7 . The f i n a l design s e c t i o n i s shown i n Figure 4 2 .

A t t h e f l o o r / c a b i n f r a m e i n t e r s e c t i o n s t a t i o n s , cargo t i e - down f i t t i n g s are l o c a t e d i n c i r c u l a r cups t h a t are i n t e g r a l w i t h t h e f l o o r . Local d e t a i l s of t h e s e connections arc shown i n F i g u r e 4 0 , 6 8 4 . 1 . 6 Main Rotor Pylon F a i r i n g The c u r r e n t c o n s t r u c t i o n c o n s i s t s of an o u t e r s k i n , p r i m a r i l y o f f i b e r g l a s s , supported by aluminum frames of minimum gage c o n s t r u c t i o n . E x i s t i n g geometry and l o c a t i o n w i t h i n t h e f a i r i n g of access p a n e l s , vent screens, etc. and geonetry and support l o c a t i o n f o r t h e v a r i o u s systems housed w i t h i n t h e f a i r i n g are n o t changed i n t h e composite s t r u c t u r e .

This maintenance of e x i s t i n g geometry combined with e x i s t - i n g l i g h t gage c o n s t r u c t i o n reduces t h e p o s s i b i l i t y of s i g n i f - i c a n t c o s t and weight r e d u c t i o n f o r t h i s assembly. A r e d u c t i o n i n weight i s obtained by t h e use of PRD-49-III/epoxy i n p l a c e of t h e f i b e r g l a s s used i n t h e e x i s t i n g f a i r i n g . No change i s envisaged f o r t h e l i g h t gage aluminum s t r u c t u r e w i t h i n t h e f a i r i n g .

An o v e r a l l view of t h e f a i r i n g i s shown i n Figure 43.

4.1.7 T a i l Pylon The major p o r t i o n of t h e curpent s t r u c t u r e i s of box beam c o n s t r u c t i o n , with t h e s t r u c t u r a l -leading edge combining t o form a two-cell t o r q u e box. For composite c o n s t r u c t i o n , shown i n Figure 4 4 , t h e beam webs are considered of sandwich c o n s t r u c t i o n w i t h cap material bonded between t h e face s h e e t s . The i n h e r e n t l o c a l s t a b i l i t y of t h e caps e n a b l e s f u l l use t o be made of a composite with- high s p e c i f i c a x i a l s t r e n g t h . Graphite/epoxy i n t h e HTS f o r m i s used f o r t h i s a p p l i c a t i o n and for t h e beam face s h e e t s .

Within t h e box s e c t i o n , t h e f o u r formed s t i f f e n e r s on t h e o u t e r s k i n of t h e c u r r e n t c o n s t r u c t i o n are r e p l a c e d by two composite s t i f f e n e r s s i m i l a r t o t h o s e considered for t h e cabin s t r u c t u r e (see Figure 4 0 ) . Although t h i s produces a l a r g e r s k i n panel width, with a s s o c i a t e d r e d u c t i o n in shear buckling s t r e n g t h , t h e r e s u l t i n g s t r u c t u r e has fewer p a r t s and lower weight. The o u t e r graphite/epoxy s k i n s are considered t o i n c l u d e PRD-49/epoxy, as described for t h e c a b i n s e c t i o n , t o improve t h e s k i n impact r e s i s t a n c e .

The t r a i l i n g edge s e c t i o n c o n s i s t s of minimum. gage detach- a b l e f i b e r g l a s s f a i r i n g s i n t h e c u r r e n t s t r u c t u r e . Replacement of t h e f i b e r g l a s s by PR3-49-III/epoxy y i e l d s a weight r e d u c t i o n without change i n manufacturing t i m e .

Two concepts are considered f o r assembly of t h e composite s t r u c t u r e . The conventional approach, shown i n Figure 4 4 , h a s t h e m e r i t of simple subassembly u n i t s being assembled'in one bonding f i x t u r e . An a l t e r n a t i v e concept i s shown i n Figure 4 5 , 7 0 m f f t which i s an a t t e m p t t o s i m p l i f y t h e f i n a l assembly bonding by using fewer, more complex subassemblies. However, t o l e r a n c e problems are envisaged a r i s i n g from t h e simultaneous bonding a t t h r e e p o i n t s a l o n g t h e s t r u c t u r e c e n t e r l i n e , and t h e concept of Figure 44 i s considered p r e f e r a b l e .

4.1.8 H o r i z o n t a l S t a b i l i z e r The basic c o n f i g u r a t i o n of t h e c u r r e n t three-beam, t h r e e - c e l l box s t r u c t u r e i s n o t considered changed for t h e composite s t r u c t u r e (see Figure 4 6 ) . The two major beams, which provide t h e connection t o t h e t a i l pylon, are of sandwich c o n s t r u c t i o n , enabling the high s p e c i f i c s t r e n g t h s of graphite/epoxy uni- d i r e c t i o n a l and c r o s s p l y l a m i n a t e s t o be u t i l i z e d due t o t h e high l e v e l of local s t a b i l i t y f o r both web and caps.

The t i t a n i u m f i t t i n g connecting t h e s t a b i l i z e r t o t h e t a i l pylon is s p l i c e d t o t h e major beans u s i n g t i t a n i u m a n g l e s .

This simple connection is considered p r e f e r a b l e , from a manufacturing s t a n d p o i n t , t o t h e a l t e r n a t i v e of a c a p / f i t t i n g machined connection.

Current aluminum cover s k i n s are r e p l a c e d by c r o s s p l y graphite/epoxy i n HMS form, using PRD-49-III/epoxy t o improve t h e s k i n impact r e s i s t a n c e , as d e s c r i b e d i n s e c t i o n 4 . 1 . 2 .

Close t o t h e s t a b i l i z e r ! r o o t , t h e s k i n p a n e l s are designed by combined shear and a x i a l loading. For t h i s r e a s c n , r i b and s t i f f e n e r spacing i s n o t changed i n t h i s area t o maintain local buckling s t r e n g t h .

Details of c o n s t r u c t i o n f o r t h e composite h o r i z o n t a l s t a b i l i z e r are shown i n Figure 4 6 . , , 4.2 L A N D I N G GEAX STRUCTURE The cost e f f e c t i v e a p p l i c a t i o n of composites t o t h e l a n d i n g g e a r s t r u c t u r e i s s e v e r e l y r e s t r i c t e d by t h e complexity of the s t r u c t u r e . S e l e c t i v e replacement of simple s t r u c t u r a l shapes by e q u i v a l e n t s t r u c t u r e o f composite material reduces t h e weight by 68 l b m ( 3 0 . 8 kg) ( 1 3 . 2 p e r c e n t r e d u c t i o n ) , using 37 lbm (16.8 kg) graphite/epoxy. A breakdown of material usage i n t h e composite l a n d i n g g e a r i s p r e s e n t e d i n Table 4 .

Details o f t h e f i n a l composite design are d i s c u s s e d below.

However, due t o i t s l o w cost e f f e c t i v e p o t e n t i a l , t h e composite l a n d i n g g e a r i s n o t considered i n t h e f i n a l e v a l u a t i o n o f t h e composite CH-533.

4.2.1 Trunnion The f i n a l design f o r t h e main l a n d i n g g e a r t r u n n i o n , based on t h e concepts of s e c t i o n 3.2.1, i s shown i n Figure 4 7 .

, Graphite/epoxy i n HTS forrn is chosen, s i n c e the design r e q u i r e s h i g h s t r e n g t h and local s t a b i l i t y f o r t h e c y l i n d e r w a l l . The b a s i c laminate i s of O o , f4So, 90° c o n s t r u c t i o n with a d d i t i o n a l Oo and 90° p l i e s i n t h e r e g i o n of the c e n t r a l f i t t i n g t o assist in forming t h e bonded s h e a r connection and t o c a r r y hoop com- p r e s s i o n l o a d i n g f o r g e a r side-loading c o n d i t i o n s . A metallic l i n e r i s considered necessary f o r t h e c y l i n d r i c a l c o n t a c t s u r f a c e t o provide a smooth, w e a r r e s i s t a n t s u r f a c e . S u i t a b l e wear tests would be n e c e s s a r y t o s u b s t a n t i a t e t h i s requirement.

Tne central metallic f i t t i n g , which connects t h e drag s t r u t and side brace a r m s of t h e t r u n n i o n t o t h e c e n t r a l c y l i n d e r , i s e s s e n t i a l l y a t h i c k c y l i n d e r u n d e r t h e a c t i o n of a x i a l and r a d i a l l o a d i n g . Comparison of aluminum, t i t a n i u m , and s t e e l f o r t h i s f i t t i n g shows lowest weight f o r aluminum. I n a d d i t i o n aluminum h a s t h e lowest material and fabrication c o s t s , For t h e s e r e a s o n s , aluminum is considered preferable, althougli no d e t a i l e d a n a l y s i s i s performed for the l o c a l problems, due t o thermal m i s m a t c h between t h e aluminum f i t t i n g and graphite/epoxy c y l i n d e r .

4.2.2 Shock S t r u t Following t h e concepts o u t l i n e d i n s e c t i o n 3.2.2, t h e shock s t r u t i s designed w i t h t h e c u r r e n t c e n t r a l c y l i n d r i c a l s e c t i o n

r e p l a c e d by a-graphitet'epoxy c y l i n d e r , using- a O o , +45O

l a m i n a t e , w i t h bonded connections t o steel end f i t t r n g s . The f i n a l design i s shown i n Figure 4 8 .

. ..

4) -f .

A t the l o w e r bonded connection, t h e basic c y l i n d e r w a l l t h i c k n e s s is i n c r e a s e d i n order t o reduce t h e induced bending stresses t o a l e v e l t h a t t h e bond can c a r r y . T h i s i n t r o d u c e s an adverse f e a t u r e of t h e d e s i g n , s i n c e t h i s i n c r e a s e d w a l l - - t h i c k n e s s is r e q u i r e d t o be maintained over an a p p r e c i a b l e l e n g t h of t h e c y l i n d e r w a l l , due t o t h e d e s i g n requirement f o r c o n s t a n t r a d i u s c o n c e n t r i c c o n t a c t s u r f a c e s .

4.2.3 Drag S t r u t Cylinder and P i s t o n As o u t l i n e d i n s e c t i o n 3 . 2 . 2 , these components, designed p r i m a r i l y by a x i a l l o a d i n g , are considered as composite c y l i n d e r s bonded t o steel end f i t t i n g s . The f i n a l design is shown i n Figure 4 9 .

Consideration o f w a l l t h i c k n e s s f o r s t r e n g t h requirements and local buckling shows both components c r i t i c a l f o r local buckling. For t h i s design c o n d i t i o n , t h e IIMS form of g r a p h i t e / epoxy is used. The laminate c o n s i s t s p r i m a r i l y of Oo ( u n i d i r e c t i o n a l ) p l i e s , although some 90° p l i e s , approximately $0 p e r c e n t of t o t a l , are considered necessary t o prevent s p l i t t i n g of t h e u n i d i r e c t i o n a l p l i e s . The 90° p l i e s are combined with t h e O o p l i e s i n a symmetric layup t o prevent l a m i n a t e warping.

Metallic l i n e r s are considered f o r r e g i o n s s u b j e c t e d t o w e a r caused by s l i d i n g c o n t a c t s u r f a c e s .

7 6 N I C K E L P L A T E ON OUTER SURFACE STEEL END F I T T I N G S P I S T O N BONDED TO CYLINDERS T Y P I C A L CONNECTION DETA 3.25 I N . 1 . D .

19.1 I N . ( 4 8 5 . b ) CKEL PLATE ON C Y L I r n E R F I G U R E 49. COMPOSITE LANDING GEAR DRAG STXUT D E T A I L S .

SECTION 5 . 1 3 COST EFFECTIVENESS EVALUATION 8 A composite CH-53D a i r f r a m e reduces weight by 1 1 1 8 lbm (507,l kg) ( 1 8 . 5 p e r c e n t of airframe w e i g h t ) , i n c r e a s e s a c q u i s i t i o n cost by 3 p e r c e n t , and r e s u l t s i n a 5 p e r c e n t reduc- t i o n i n t h e 10-year l i f e c y c l e c o s t .

5.1 PRODUCTION VEHICLE COST COMPARISON T h e c o s t comparison for c u r r e n t and composite production v e h i c l e s is p r e s e n t e d i n Figure 5 0 , i n which the composite pro- d u c t i o n v e h i c l e c o s t s are based on t o t a l production of 6 0 0 v e h i c l e s i n a 1980 t i m e frame, s t a r t i n g i n 1 9 7 8 . The details of t h i s a n a l y s i s are contained i n Appendix C .

The moderate i n c r e a s e i n v e h i c l e fly-away c o s t (under 3 p e r c e n t ) i s shown by t h e c o s t breakdown of Figure 50 t o be l a r g e l y a t t r i b u t a b l e t o the i n c r e a s e d c o s t of t h e composite m a t e r i a l s , compared with t h e metals used i n t h e c u r r e n t v e h i c l e .

' T h i s i n c r e a s e i s o f f s e t , t o some degree, by t h e r e d u c t i o n i n l a b o r c o s t for f a b r i c a t i o n using composite materials.

Table 5 shows t h a t t h e r e d u c t i o n of airframe weight by 1 , 1 1 8 l b m ( 5 0 7 . 1 kg) is achieved a t a c o s t of $ 8 4 . 2 / l b m ($186/kg) r e d u c t i o n .

F 5.2 PROTOTYPE VEEICLE COST CO?IPARISON The c o s t comparison between prototype composite and con- v e n t i o n a l material v e h i c l e s i s presented i n Figure 5 1 , with both v e h i c l e costs r e f l e c t i n g a 1 9 8 0 t i m e frame.

From t h e c o s t breakdown shown i n Figure 51, it i s a p p a r e n t t h a t a major source of t h e 3 . 6 p e r c e n t h i g h e r c o s t for t h e com- p o s i t e v e h i c l e is t h e h i g h e r e n g i n e e r i n g c o s t . This i s i n t r o - duced by t h e a d d i t i o n a l d e s i g n and a n a l y s i s e f f o r t , which would be r e q u i r e d for p r o t o t y p e d e s i g n using composite materials.

This c o s t i n c r e a s e , t o g e t h e r w i t h t h e h i g h e r composite material costs, is o f f s e t t o sone e x t e n t by t h e r e d u c t i o n i n f a b r i c a - t i o n l a b o r c o s t s u s i n g composite materials. However, t h e t o o l - i n g i s an unknown f a c t o r . F u r t h e r work i s r e q u i r e d i n t h i s area. Vhile it i s b e l i e v e d t h a t t h e t o o l i n g f o r a composite airframe should c o s t less t h a n t h a t for t h e conventional m e t a l d e s i g n , t h e c o n s e r v a t i v e estimate is made t h a t t o o l i n g c o s t s are t h e same.

5.3 TEN-YEAR LIFE CYCLE COST FOX PRODUCTION VEHICLE A c o s t e f f e c t i v e n e s s a n a l y s i s w a s conducted f o r t h e CH-53D on t h e b a s i s of a 600-vehicle f l e e t with a n average u t i l i z a t i o n of 500 hours p e r v e h i c l e p e r y e a r . The f l e e t o p e r a t i o n c o n s i s t s of t h e prirnary t r a n s p o r t mission r o l e , with 30 p e r c e n t t r o o p and 70 percent car20 usage. For t h i s r o l e , t h e average gross weight is 4 0 , 7 7 0 lbn ( 1 8 , 5 2 0 kg). The r e s u l t s of a 10-year l i f e c y c l e c o s t of o p e r a t i o n a n a l y s i s are summarized i n Table 6 .

F u r t h e r d e t a i l s of t h i s a n a l y s i s are p r e s e n t e d i n Appendix D . - From Table 6 it can be seen t h a t , d e s p i t e t h e 3 p e r c e n t increase i n v e h i c l e a c q u i s i t i o n cost, t h e composite v e h i c l e h a s a 7 percent g r e a t e r p r o d u c t i v i t y t h a n t h e c u r r e n t v e h i c l e , r e s u l t i n g i n a 5 p e r c e n t r e d u c t i o n i n fleet l i f e c y c l e cost.

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SECTION 6 . 0 RECOMMENDATIONS A?ID DISCUSSION

8 Cost e f f e c t i v e a p p l i c a t i o n of advanced composite

materials can be achieved by t h e use of all-molded composite s h e l l s f o r airframe c o n s t r u c t i o n ; however, f u r t h e r development is r e q u i r e d t o provide a complete c o s t and data base.

6 . 1 RECOMMENDATIONS Development programs are r e q u i r e d , aimed a t an e x t e n s i o n of manufacturing t e c h n i q u e s l e a d i n g t o f a b r i c a t i o n of composite hardware and e v e n t u a l s e r v i c e use. Such programs should i n c l u d e airframe components designed s p e c i f i c a l l y f o r h e l i - c o p t e r a p p l i c a t i o n s .

The manufacturing program should i n c l u d e such d e t a i l s as i n s p e c t i o n techniques and bonding p r o c e s s e s , i n a d d i t i o n t o a c h i e v i n g experience i n t h e f a b r i c a t i o n of l a r g e molded assemblies. A t e c h n i c a l data b a s e , f u r n i s h e d by t e s t i n g of b o t h d e t a i l components and major a s s e m b l i e s , would be r e q u i r e d t o supplement the composite hardware f a b r i c a t i o n program.

areas of damage t o l e r a n c e A d d i t i o n a l q u a n t i t a t i v e d a t a i n t h e and r e p a i r a b i l i t y should t h e n be o b t a i n e d by s e r v i c e use.

Some recormended programs are summarized i n Table 7 .

6.2 DISCUSSION H e l i c o p t e r airframe s t r u c t u r e s are l i g h t l y loaded compared w i t h f i x e d wing aircraft. The a p p l i c a t i o n of composite materials t o h e l i c o p t e r airframe s t r u c t u r e s r e q u i r e s , there- fore, a d i f f e r e n t t y p e of s t r u c t u r a l design t o provide t h e m o s t e f f e c t i v e use of t h e materials.

Due t o t h e r e l a t i v e l y low airframe l o a d i n g s , t h e composite h e l i c o p t e r s t r u c t u r e r e q u i r e s t h e use of t h i n composite. l a m i -

n a t e s , both f o r s k i n and support s t r u c t u r e , in o r d e r t o

a c h i e v e a s t r u c t u r a l weight lower t h a n t h e c u r r e n t aluminum s t r u c t u r e . T h i s requirement l e d t o t h e development, i n t h i s s t u d y , of t h e all-molded composite s h e l l concept. The a p p l i - c a t i o n of t h i s concept i n v o l v e s t h e use of i n i t i a l buckling c a p a b i l i t y and post-buckled s t r e n g t h f o r t h e l i g h t gage s k i n s .

The design a n a l y s i s is based on a n a l y t i c a l methods, with some c o r r e l a t i o n f r o m diagonal t e n s i o n ( s h e a r ) t e s t i n g . A n a l y t i c a l methods are also used t o p r e d i c t t h e combined load c a p a b i l i t y ( s h e a r and compression) f o r t h e s k i n - s t r i n g e r combination. Tine all-molded c o n s t r u c t i o n f o m s a monolithic s t i f f e n e d s h e l l and t h u s t e n d s t o i n c r e a s e the s t r e n g t h c a p a b i l i t y . S t r e n g t h t e s t - i n g of such a s t r u c t u r e would be necessary t o v a l i d a t e t h e i g n a n a l y s i s . Since t h e major weight of t h e airframe i s i n t h e o u t e r s h e l l , such t e s t i n g could r e s u l t i n a f u r t h e r weight TABLE 7. RECOMMENDED DEVELOPMENT PROGRAMS STRUCTUBE MAVJFACTURIXG SERVICE i'ECH."IICALDATA Foam-Stabilized Develop rnolding process Install foam stabiliz- Test foam stabil- ized stringers for Composite for foam cores. fabrica- ed stringer on heli- cor?.pression Structures tion and inspection copter to determine effect of load E. strength. Develop techniques for built-up crippling data.

foam stabilized composite environment stringers and frames. Test foam stabil- Fabricate sample struc- ized frames in tures. bending and shear.

Light-Gage Fabricate large composite Install composite skin Conduct shear tests Composite skin panels, both flat and section on helicopter on various ply Skins to determine ability orientations for curved. Develop rapid of panel to withstand flat panels. Ex- methods for including cutouts (for windows, environmental damage. tend tests to access, etc.). Investi- curved panels.

gate possibility of Determine initial single cure construction buckling E strength for skin/stringer/frame characteristics as shell. dependent on ply orientation and effect of panel curvature. Con- duct impact tests on G/E and P.W-49 coabination of plies for damage tolerance.

Install simple joints Conduct experi- Composite-to- Fabricate joints using for secondary struc- mental tests using Metal Joints composites and titanium Tures to evaluate narmai ( 250° to and aluminum fittings.

load and environmental 350° F cure) and Develop inspection room temperature techniques for bond effects.

integrity. cures. Evaluate different bonding techniques witn combination of composites and metals.

Conduct combined Light-Gage Fabricate skin/stringer load tests (shear panels. Evaluate pro- Composite Skins and compression, duction methods for With Foam- and shear and ten- build-up of structures.

Stabilized 'sion). Both flat Stringers and curved panels to be evaluated.

Install se,pent of Conduct tests on All-Molded Develop build-up of shell construction typical shell con- Shell of Skin/ fabrication. Evaluate and evaluate effect struction for Stringer and possibility of single- cure layup. of service usage. typical loadings Fra-e Construc- on shell. Evalu- tion ate effects of all molded constrw- tion on strength characteristics.

8 3

reduction and increased cost effectiveness for the all-molded

construction. The all-bonded concept raises the problem of

thermal mismatch in the region of metallic fittings. Titanium

fittings, replacing the current aluminum fittings, alleviate

this problem to some extent. Ilowever, the development of a

lower temperature adhesive bond, preferably a room temperature

cure, may allow the use of aluminum fittings, which would further

increase the cost effectiveness of the composite shell construc-

tion.

The manufacture of the shell conceived in this study

requires experience in the handling, laying-up, and curing of

large, thin, composite panels. In addition, the use of such

thin laminates for external airframe skins requires knowledge

of the damage tolerance capability required for service use.

In this study, PRD-49 is used with the graphite/epoxy, to

increase the impact resistance to the level judged necessary

for damage tolerance.

Service experience is necessary to confirm the damage

tolerance assumptions and to provide confidence in the type of

construction and fabrication methods used. Such experience may

well provide data to further increase the cost effectiveness of

the composite application.

All the factors mentioned above indicate that further

efforts are required to provide the confidence and necessary

data for cost effective application of composites to helicopter

airframe structures.

APPENDIX A STRUCTURAL DET-ULS OF CUflXEIIT CH-5 3D AIRFRAME AX3 M N D I I I G GEAR STWCTURE 1 . 0 of t h e r e p o r t p r e s e n t s a b r i e f d e s c r i p t i o n S e c t i o n of t h e c u r r e n t CH-53D s t r u c t u r e . F u r t h e r d e t a i l s are p r e s e n t e d here, i n c l u d i n g t h e material usage, p r e s e n t e d i n Table Al, and t h e s t r u c t u r e weight breakdown, p r e s e n t e d i n Table A 2 . The s t r u c t u r e i s c o n s i d e r e d broken i n t o n i n e major a s s e m b l i e s , as i n d i c a t e d i n F i g u r e 4 , and a weight breakdown is p r e s e n t e d for each assembly (see Tables A 3 through A l l ) .

AIRFWIE STXJCTURE e The airframe i s c o n s t r u c t e d mainly of aluminum a l l o y , w i t h t h e major weight being i n t h e o u t e r s h e l l . A b r i e f d e s c r i p t i o n of material usage, t y p e o f s t r u c t u r e , and governing d e s i g n c o n d i t i o n s i s p r e s e n t e d f o r each major s t r u c t u r a l assembly shown i n F i g u r e 4.

Cockpit S e c t i o n The c o c k p i t s e c t i o n extends from f u s e l a g e s t a t i o n (FS) 84 t o 1 6 2 and r e p r e s e n t s 9 . 5 % of t h e airframe s t r u c t u r e w e i g h t . .

- The upper canopy; which forms t h e p i l o t ' s e n c l o s u r e , i s an i n t e g r a l l y c o n s t r u c t e d f i b e r g l a s s s h e l l ( s k i n / f r a m e ) . The lower c o c k p i t s t r u c t u r e , which p r o v i d e s s u p p o r t f o r nose land- i n g g e a r lEads t o g e t h e r - w i t h eqiipment and- p e r s o n n e l l o a d s , i s of aluminum c o n s t r u c t i o n w i t h t h e o u t e r s h e l l s u r f a c e being - aluminum honeycomb sandwich p a n e l s .

The primary materials used i n t h i s s e c t i o n are aluminum ( 5 9 % ) and f i b e r g l a s s ( 2 2 % ) . M a t e r i a l usage f o r t h i s s e c t i c n is p r e s e n t e d i n Table Al, and t h e weight breakdown by t y p e of s t r u c t u r e and d e s i g n c o n d i t i o n is p r e s e n t e d i n Table A3.

Cabin S e c t i o n T h i s s e c t i o n e x t e n d s from FS 162 t o 5 2 2 and c o n t a i n s 4 0 . 5 % of t h e airframe s t r u c t u r e weight. The s t r u c t u r e i s g e n e r a l l y of semi-monocoque c o n s t r u c t i o n , employing s h e e t m e t a l f o r n e d

s t r i n g e r s and b u i l t - u p frames. Major , h e a v i l y loaded frames

are brought o u t t o the s k i n line, i n t e r r u p t i n g t h e l o n g i t u d i n a l members ( s t r i n g e r s and l o n g e r o n s ) . Minor frames are a t t a c h e d t o t h e s t r i n g e r s and n o t brought o u t t o the s k i n l i n e (float- ing frame c o n s t r u c t i o n ) .

8 5 Major frames are l o c a t e d a t t h e c o c k p i t / c a b i n and a f t s e c t i o n / c a b i n i n t e r f a c e s as r e d i s t r i b u t i o n bulkheads. Other major frames c a r r y loading; from t h e engines, main r o t o r , and main l a n d i n g g e a r .

The s t r i n g e r s , longerons, and s k i n act t o c a r r y t h e design bending moment and shear l o a d s . S t r i n g e r spacing is g e n e r a l l y 6.0 i n . ( 1 5 2 . 4 mm) around t h e cabin p e r i p h e r y , and frames are spaced a t 2 0 . 0 i n . ( 5 0 8 . 0 m m ) , g i v i n g a t y p i c a l s k i n p a n e l s i z e of 6.0 i n . ( 1 5 2 . 4 mm) x 2 0 . 0 i n . ( 5 0 8 . 0 m m ) .

The major material used for t h i s s e c t i o n i s aluminum ( 9 5 % ) . Material usage f o r t h i s s e c t i o n i s p r e s e n t e d i n Table Al, and t h e s t r u c t u r e weight breakdown i s p r e s e n t e d i n Table A 4 .

Sponson S e c t i o n The sponson i s an a i r f o i l - s h a p e d s t r u c t u r e a t t a c h e d t o t h e c a b i n between FS 2 7 2 and 4 9 4 . The forward s e c t i o n houses t h e f u e l c e l l s and i s designed by f u e l l o a d i n g c o n d i t i o n s and walk- i n g l o a d s on t h e upper s u r f a c e . Construction i s p r i m a r i l y of aluminum s h e e t supported by s t i f f e n e r s f o r i n t e r n a l bulkheads and supported by i n t e r c o s t a l s f o r o u t e r s k i n .

The a f t sponson s e c t i o n houses t h e n a i n l a n d i n g g e a r (MLG) , which i n t r o d u c e s t h e p r i n c i p a l design l o a d i n g c o n d i t i o n .

i n t o t h e s e c t i o n . Walking and hydrodynamic l o a d s aFe o t h e r design c o n d i t i o n s i n t h i s r e g i o n . MLG l o a d i n g is carried by bulkheads and beams of b u i l t - u p c o n s t r u c t i o n .

The o u t e r s k i n i s supported by i n t e r c o s t a l s t o c a r r y t h e normal loading.

Two back-to-back balkheads a t t h e forward and a f t s e c t i o n i n t e r f a c e permit s t r u c t u r a l i s o l a t i o n of t h e t w o s e c t i o n s .

Aluminum a l l o y is t h e major m a t e r i a l used for t h i s s e c t i o n ( 8 8 % ) . Material usage for t h i s s e c t i o n i s p r e s e n t e d i n Table A l , and t h e s t r u c t u r e weight breakdown i s p r e s e n t e d i n Table A 5 .

A f t S e c t i o n This s e c t i o n l i e s behind the c a b i n s e c t i o n and extends f r o m FS 5 2 2 t o 749. Primary design l o a d s are t h e t o r s i o n , s h e a r , and bending noment from t h e t a i l pylon, which is connected t o the rear of t h i s s e c t i o n . The presence of t h e cargo ramp and overhead door i n t h e forward r e g i o n produces an open channel s e c t i o n . A f t of t h i s r e g i o n , t h e s t r u c t u r e is a f u l l y c l o s e d s e c t i o n .

The s t r u c t u r e is of b u i l t - u p s h e l l c o n s t r u c t i o n ( s k i n / s t r i n g e r / f r a m e ) , with forged aluminum f i t t i n g s l o c a t e d a t t h e i n t e r f a c e with t h e t a i l pylon.

8 6 Aluminum a l l o y i s t h e major material used f t h i s s e c t i o n ( 9 2 % ) . Material usage i s presented i n Table Al, and t h e s t r u c t u r e weight breakdown i s presented i n Table A 7 .

Floor S e c t i o n This s e c t i o n i s designed by c a r g o , v e h i c l e , and personnel l o a d i n g and c o n s i s t s of an aluninun s h e e t s t i f f e n e d by aluminum e x t r u s i o n s . The f l o o r s u p p o r t s , a t t h e cabin s e c t i o n frames, are designed t o i s o l a t e t h e f l o o r from primary airframe load- ing.

The major material used f o r t h i s s e c t i o n i s aluminum ( 9 9 % ) . Material usage f o r t h e s e c t i o n is p r e s e n t e d i n Table A l , and t h e s t r u c t u r e weight breakdown i s p r e s e n t e d i n Table A 7 .

Main Rotor Pylon F a i r i n g This s t r u c t u r e provides an aerodynamic f a i r i n g around t h e main r o t o r s h a f t and i t s associated s y s t e m above t h e c a b i n .

It c o n s i s t s of a f i b e r g l a s s s k i n supported by formed aluminum framing.

Major materials used are aluminum ( 5 9 % ) and f i b e r g l a s s ( 3 8 % ) . Details of material usage are presented i n Table Al, and t h e s t r u c t u r e weight breakdown is p r e s e n t e d i n Table A B .

T a i l Pylon S e c t i o n The t a i l pylon extends from FS 749 t o 8 9 0 . Design l o a d s are t h e bending moments, s h e a r s , and t o r s i o n s introduced by t h e t a i l r o t o r , t a i l r o t o r geap box, and h o r i z o n t a l s t a b i l i z e r , which are a l l supported by t h e t a i l pylon.

The s t r u c t u r e is of c l o s e d s e c t i o n , with two i n t e r n a l beams,' b u i l t up from aluminum s h e e t and e x t r u s i o n s with forged f i t t i n g s a t t h e connections t o t h e h o r i z o n t a l s t a b i l i z e r and aft s e c t i o n . A f i b e r g l a s s f a i r i n g houses t h e t a i l r o t o r d r i v e s h a f t , which r u n s along the a f t face of t h e pylon.

Aluminum is t h e major material used i n t h e t a i l pylon (84%). Details of material usage are presented i n Table Al, and the s t r u c t u r e weight breakdown is p r e s e n t e d i n Table A 9 , Horizontal Stabilizer The horizontal stabilizer is an asymmetric structure mounted on the upper right hand side of the tail pylon. Design loading comes from stabilizer air loads and is carried to the tail pylon by a three-beam, three-cell box structure. Beams and outer skin are built up from aluminum sheet with aluminum stiff- ening and support members. The leading edge, carrying local airload only, is a fiberglass skin supported by formed alumimun ribs.

Aluminum is the major material used for the stabilizer ( 8 9 % ) . Details of material usage are presented in Table Al, and the structure weight breakdown is presented in Table A10.

LANDING GEAR STRUCTURE

The landing gear uses aluminum alloy as the major structural material ( 4 2 % ) , but with a high proportion ( 3 3 % ) of high- strength steel parts.

The landing gear is of tricycle configuration, incorporating air-oil oleo struts and dual main and nose wheels.

The main landing gear is housed in the sponson structure and retracts forward. The nose landing gear is housed in the , cockpit section and retracts aft.

Both gears are of similar construction. The main landing gear is shown in Figure 3 . Material usage for the landing gear is presented in Table Al, and the structure weight break- down is presented in Table A l l .

8 8 TABU A 1 C H - 5 3 MATERIAL USAGE MATERIAL WEIGHT l b m (kg) MISC. NON FIBER MET AND MAJOR ASSEMBLY STEEL GLASS HARDWARE SUBTOTAL Cockpit Section 333 3 564 ( 1 . 4 ) (255.8) (151) Cabin Section 4 2424 (1049.9) (1.8) (1099.31 Sponson Section

-

685 767 ( 310.7 (347.9) A f t Section 1066 1 (483.4) (525.65 (045) Floor & Supports 445 449 Section (201.8 (1.8) (203.6) Main Rotor Pylon 1 176 29 7 (134.65 F a i r i n g (79.8) ( . 4 5 ) 264 1 T a i l Pylon ( 142.85 (119.7 ) (0451 Section Hori zcnt a1 2 102 S t a b i l i z e r (41.3j (.9> (46.3) 248 Total Airframe 16 5 375 (112.5 (2438.1) (7.25) (185.5)

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66 Main Landing 146 131 (29.9) Gear (66.2) (59.4)

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Nose Landing 59 73 38 Gear (26.8) (33.1) (17.2) 219 169 5 13 T o t a l Landing (232.7) (99.31 (76.6) G e a r Total Airframe & Landing 6590 lbm Gear Structural. Weight (2988.65 kg) n h h h n n h ln t-• F U J o n r l m K I K Wt- cu W tn H B

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v W n n M m V ) .

V ) . I I 1 I l l I I mu3 MV) I-l rl u .

U v I I I I I 1 1 1 I I I I I I I 1 1 1 I I I I I 1 I W

! D

k m c a, +J 4 w k

c 2 . d

0 4J ld 4 tn i k

APPENDIX B

WEIGHT TREND CURVES FOR HELICOPTER STRUCTURES

The curves of Figures B1 through B8 indicate the weight

trends for the major assemblies of helicopter structures.

Points corresponding to the current and composite CH-53D

structure are given in all tables.

10,000 PB, BODY WEIGHT PARPLlETER I - W f = 1071.8 PB, l b m (4125 PB,sI kg) NZ = ULT LOAD FACTOR SGW = STRUCTURAL DESIGN GROSS UT. lbm (kg) Sf = FUSELAGE VETTED AREA, FT2 (m"\ A P = ULT CABIN D I F F E R E $ T l A L 2 PRESSURE, lbf/IN. <N/m * LESS SPONSONS, RAE.IpS, T A I L PYLON, AND PYLON FOLD q = D E S l G N g Y I a V I $ PRESSURE, l b f / I N . (X/m )

I I 1 I 1 1 1 1

100 1

.1 1 10 FIGURE B1. WEXGHT TREND FOR BODY STRUCTLTRE.

10 1 Po Y , - E 2 Ei R

g 1000

V m

E

V

& 500

XV 1oc

““t

I I I I I I I I 100 1000 10 ,a Sf, COCKPIT W E I T E D AREA, FT2

(sf,sI = .0929 Sf’ ”)

FIGURE B2. WEIGHT TREND FOR COCKPIT STRUCTURE.

10 2 PCAB, C A B I N WEIGHT PARANETEX 0.3056 0.875 ,0.133 10,000 1000 -0983 Pcm PCAB, 2000 M Y Q

*-

E

k ! 500

R

m

s

.

m NZ = ULT MAD FACTOR S G W = STRUCTURAL GROSS W, lbm (kg) 8CAb = CAB114 k%lTF3 AREA, F I ' * (m2) 1oc AP = ULT C A B I N D I F F E R E g T I A L PRESSURE. l b f / I N . (N/m ) I I I I I 1 1 1 FIGURE B3. WEIGHT TREND FOR CABIN STRUCTUF.E.

10 3

I I I I l i l t 10 100 1 ( Sf. WETTED ARZA, FT2

(.,,,, = .0929 Sf . ' )

WEIGHT TREND FOR AFT S E C T I O N STRUCTURE.

F I G U R E B4.

10 4

PERSONNEL FLOOR I I I I I I I I I 10

sm, F'LOOR AREA, FT

= .0929 SFL .'>

FIGURE B5. WIGHT TREXD FOR FLOOR STRUCTURE.

10 5

El f:

* *

g 100 5; PI

d

d 50

* LESS PYLON

FOLD F I T T I N G 2 0 ~ 10 I I I I I I I I . I I I I I I l l 10 100 1000 STp, TAIL PYLON PLKTFORM AREA, FT2 WEIGHT TREND FOR TAIL PYLON STRUCTURE; FIGURE B6.

SHs, HORIZOIiTAL STABILIZER PLANFORM AREA, i”p (‘HS,SI ‘HS, FIGURE B7. WEIGHT TREND FOR HORIZONTAL STABILIZER STRUCTURE.

* 2000

* INCLUDES GEAR SUPPORT STRUCTURE

M Y DOES NOT INCLUDE RUNNING GEAR, U Z S , ETC.

t

* *

- 9

II

P

20c €IO-3s-1

I I I 1 1 1 1 1

1 . 0 1 3 1

SGW/lOOO , lbm

SGWsI = .454 SGW,kg) WIGHT TREND FOR LANDING GFXS STRUCTURE.

FIGURE ~ 8 ,

no8

APPENDIX C MATERIAL AND MANUFACTURIXG COSTS Material Costs For t h e p r o t o t y p e and production composite v e h i c l e s , t h e material costs used are v a l u e s f o r 1 9 7 8 p r o j e c t e d from c u r r e n t c o s t t r e n d s . P o s s i b l e cost i n c r e a s e s due t o i n f l a t i o n are excluded.

costs used are given below: The p r o j e c t e d T i PRD-491 MATERIAL G/E FORGING ALUM FOAM EPOXY ADHESIVE U N I T * 2 5 . 0 0 7.00 1 . 0 0 3.00 2.50 1 . 9 6 COST (55.00) (15.45) ( 2 . 2 0 ) (6.60) ( 2 . 7 4 ) (21.50) * For a d h e s i v e , t h e u n i t cost is f o r one l a y e r .

N f t 2 For PRD-49/epoxy, t h e u n i t cost is $/yd ( $ / m ) f o r 0 . 1 i n .

( 2 . 5 4 mm) f i n i s h e d f a b r i c 3 8 i n . ( 0 . 9 2 6 m ) wide.

A l l o t h e r u n i t costs are $/lbm ($/kg).

Manufacturing Costs The cost e v a l u a t i o n assumes t h e accomplishment of a manu- f a c t u r i n g r i s k r e d u c t i o n program for f a b r i c a t i o n techniques prior t o composite p r o t o t y p e f a b r i c a t i o n . The cost of t h i s program has been i n c l u d e d i n t h e t o t a l c o s t f o r t h e composite p r o t o t y p e .

For both p r o t o t y p e and production composite v e h i c l e manu- f a c t u r e , t h e r e n t - f r e e use and a v a i l a b i l i t y of e x i s t i n g facil- i t i e s and equipment are assumed.

The effects of i n f l a t i o n on manufacturing l a b o r costs are not c o n s i d e r e d , and c u r r e n t ( 1 9 7 3 ) l a b o r rates are used.

Engineering and Tooling Current ( 1 9 7 3 ) rates are used for t h e assessment of engineering d e s i g n and a n a l y s i s costs f o r t h e composite proto- t y p e v e h i c l e , and f o r t o o l design and s u s t a i n i n g e n g i n e e r i n g effort f o r both p r o t o t y p e and production composite v e h i c l e s .

1 0 9 Cost Estimation Based on the assumptions outlined above an estimation of the cost f o r both prototype and production composite vehicles is made. The breakdown of these cost estimates is given in tabular form below.

Production composite vehicle cost. Cost/unit f o r total production of 600 vehicles.

(Tabulated values are in $1000 units) CONVENTIONAL Total

I

681.19 Airframe

I I

I Non-Airframe I I I

Acquisition 2,701.81.

2,701.81

I I I

F I I i

Total 3,383.00 3,290.00 Vehicle I 1 Prototype Composite Vehicle (Tabulated values are in $1000 units)

1 COMPOSITE I CONVENTIONAL

VEHICLE ' VEHICLE (REF)

Engr . 1 6,555.00 1

Material Total

I 16,893.60 I 16,213.00

Airframe Non-Airframe Acquisition 2,701.81 2,701.81

T o t a l I 19,595.41 I 18,914.81

Vehicle I I A P P E N D I X D COST EFFECTIVENESS ANALYSIS In comparing current and composite designs, the two are evaluated for mission performance and fleet cost. Performance is measured on the basis of aircraft mission productivity expressed in ton-knots (kg.m/s). Cost is measured on the basis of fleet cost to maintain a constant fleet effectiveness.

To determine mission productivity representative of the conditions in which the CII-53D operates in its primary transport mission, a probabilistic mission environment was established, and 1000 simulated missions were flown. The mission environment used in the simulation was defined by the cumulative probability distributions of the following parameters: 1. sea level temperature (standard altitude lapse rate was assumed) 2. take-off pressure altitude 3 . sortie radius 4. cruise altitude elevation above take-off 5. percentage of outbound payload carried inbound 6 . required payload 7 . down time per sortie 8 . hover time per sortie 9 . take-off hover power margin (fraction of hover out- of-ground-effect power actually required) Cumulative probability distributions are shown in Figure D1. Take-off pressure altitude was based on 20% of time take- off is at sea level, 5 0 % of time take-off is at 5 0 0 feet (153 m) or less, 904 of time at 2 5 0 0 feet (762 m) or less, and 100% of time at 10,000 feet (3048 m) or less. This is representative of land area from shore to 50 n. mi. (92.6 km) inland typical of CH-531) operations. Sea-level temperature distribution is based on an average world-wide temperature distribution of 76OF (298OK) f o r potential areas of engagement on or near coastlines; 85% of time temperature is 83OF (302OK) or below, and 100% of time temperature is at 12OOF (322OK) or below. This distribu- tion also approximates the sea-level temperature distribution Cruise pressure altitude for regions of anticipated operation.

above take-off is e s t i m a t e d t o average 2 0 0 0 feet (610 m > w i t h 90% of t i m e f l y i n g 2500 feet (762 m) or less, and 1 0 0 % f l y i n g 6000 feet (1830 m ) o r less above take-off p r e s s u r e a l t i t u d e .

Required payload, a demand f u n c t i o n independent of c a p a b i l - i t y , is based on c a r r y i n g t r o o p s 3 0 % of t h e t i m e and cargo 70% of t h e t i m e . Cargo d i s t r i b u t i o n i s based on redeployment of a i r - c a r g o f r o m C-130 and C-141 f o r 30% of cargo l o a d i n g s , and redeployment of 2 1/2-ton (2268 kg) and 5-ton (4536 kg) t r u c k c a r g o for 70% of cargo l o a d i n g s . Required payload-out averages 16.8 t o n s (15,240 kg). T h i s requirement exceeds t h e CH-533 payload c a p a b i l i t y . Therefore, any i n c r e a s e i n payload c a p a b i l - i t y w i l l produce an i n c r e a s e i n p r o d u c t i v i t y .

Inbound payload averages 12% of outbound payload w i t h 50% of f l i g h t s r e t u r n i n g empty. S o r t i e r a d i u s average i s 25 n. m i .

(46.3 kn), w i t h 50% of missions being 1 5 n . m i . (27.8 km) or less and 90% of missions being 50 n. m i . (92.6 krn) o r less.

,Take-off power margin range i s .60 t o 1 . 0 , with an average v a l u e of .75. Hover t i m e p e r s o r t i e averages 1/3 minute, w i t h t w o minutes maximum hover tine p e r s o r t i e . Down t i m e p e r s o r t i e a v e r a g e s t w o minutes, w i t h 90% of t i m e less t h a n seven minutes and 100% of t i m e less t h a n 30 minutes.

Other i n p u t s t o t h e mission a n a l y s i s i n c l u d e CH-53D r o t o r p a r a m e t e r s , engine performance, b a s i c o p e r a t i n g weight, and c o n s t r a i n t s imposed by maximum g r o s s weight, d r i v e system r a t i n g , speed l i m i t , and f u e l c a p a c i t y . The CH-53D parameters are : Rotor D i a m e t e r

7 2 ' - 2 . 7 " (22.0 m)

T o t a l Blade A r e a 469 sq. f t . (43.5 m*) Basic Operating Weight 2 4 2 1 0 lbm (11.,000 kg) ( i n c l u d e s 725 l b s . (329 kg) of f i x e d u s e f u l l o a d ) Engine T64-GE-41 Drive System Maximum Power 7 , 0 0 0 HP (0.523 MU) Maximum Gross Veight 4 2 , 0 0 0 lbm ( 1 9 , 0 6 0 kg) Red Line Speed L i m i t 1 7 0 K t s . (87.3 m / s ) 630 g a l s . (2.39 m3) Fuel Capacity S i m u l a t i o n of t h e c u r r e n ? CtI53D i n t h e e s t a b l i s h e d m i s s i o n environment y i e l d e d t h e f o l l o w i n g r e s u l t s : Average Take-Off Gross Weight 4 0 , 7 7 0 lbm (18,520 kg) Average Outbound Payload 1 4 , 6 2 0 lbm (6,630 kg) 2,285 lbm/hr (1,517 kg/ Average F u e l Flow h r ) Average S o r t i e T i n e 0.362 h r Average Mission P r o d u c t i v i t y 372.8 Ton-Kts. (173,500 kgm/s) Comparison of t h e e x i s t i n g and composite airframe d e s i g n s on t h e b a s i s of weight and c o s t f o r a s i n g l e p r o t o t y p e f l i g h t v e h i c l e and a f l e e t of 6 9 0 aircraft are g i v e n i n t h e f o l l o w i n g T a b l e 31. F u r t h e r d e t a i l s of v e n i c l e c o s t s are g i v e n i n Appendix C .

TABLE 01.

C H - 5 3 WEIG,W AND C O S T COMPARISON OF C O M P O S I T E A I R F W I E W I T H CURRENT D E S I G N .

Prototype C o s t ($1 16,213,000 Current Prcduction Cost ($1 C H - 5 3

I

Airframe Weight l b m 6,077 (2756.5 (kg)

-

Prototype Cost ($1 16,093,600 Composite Production Cost ($1 681,190 CH-53D A i r f r a m e 4,959

(kg) I (2249.4)

For a single aircraft, the increased cost is $680,600 to achieve 1,118 lbm (507.1 kg) of weight savings by use of compos- ites. For a 600-aircraft fleet, the cost is reduced to $83.2 per pound of weight saved ($184/kg). In order to relate the impact of the candidate design characteristics on aircraft performance and cost, it is necessary to evaluate the operation- al changes in aircraft productivity and mission effectiveness achieved by the use of composites. Change in fleet cost of the composite design to maintain the same fleet effectiveness as the current design is used to evaluate tne impact of cost and technical factors.

Table 9 2 compares the two designs, considering (a) perfor- mance in the CH-53D primary transport mission role and (b) the total system cost over the expected 10-year service life to maintain a constant fleet effectiveness of 600 baseline aircraft flying an average 500 hours per aicraft annually.

TABLE D2. CH-53D COST EFFECTIVENESS SUMMARY.

COMPOSITE CHANGE CH-53D Acquisition Cost (million $1 3.290 3.384 +. 094 per Aircraft Weight Empty lbm 23,485 22,367 -1,118 (kg) (10,670 (10,143) (507 1 Miss ion Availability 90'3 .go9 Mission R e l i a b i l i t y 992 0992 Average A i r c r a f t Productivity ton-knots 372.8 397.4 +24.6 (kg .m/s ( 173 , 500 (184,950 1 (11,450: Operating Cost per F l i g h t Hour ($1 715 715 F l e e t Size 562.86 600 -37.14 F l e e t Life Cycle Cost $1 (million 4 119 -202.1 3916.9 A c q u i s i t i o n c o s t , e s t i m a t e d a t $ 3 , 2 9 0 , 0 0 0 , i n c l u d cost, i n i t i a l s p a r s , ground support equipment (GSE), a i n g c o s t s . I n i t i a l s p a r e s and GSE c o s t are assumed un from t h e c u r r e n t CH-53D. The use of PRD-49 i n of t h e airframe i s considered t o provide a dam l e v e l s i m i l a r t o t h a o f t h e c u r r e n t aluminum s t r u c t u r e . F o r t h i s r e a s o n , no change is considered i n MElH/FH, s o t h e i n c r e - mental c o s t t o t r a i n maintenance personnel i s zero. Changes i n flyaway c o s t are o b t a i n e d from Table D1 f o r t h e production air- craft, a d j u s t e d f o r a m o r t i z a t i o n of composite a i r c r a f t non- r e c u r r i n g c o s t , and are added t o t h e c u r r e n t CH-53D a c q u i s i t i o n cost t o o b t a i n the composite aircraft a c q u i s i t i o n c o s t .

Weight empty of t h e c u r r e n t CH-53D i s based on s p e c i f i c a - t i o n S D 5 5 2 - 1 - 3 . The composite CH-53D empty weight is o b t a i n e d by adding t o t h e c u r r e n t aircraft t h e incremental change due t o t h e composite d e s i g n o b t a i n e d from Table D1.

Mission a v a i l a b i l i t y is based on a down-hour r a t e of 1 . 6 p e r f l i g h t hour f o r t h e c u r r e n t CH-53D. The use of composite materials may reduce t h i s rate through r e d u c t i o n of c o r r o s i o n and r e l a t e d i n s p e c t i o n . However, i n t h e absence of s e r v i c e experience i n t h i s area, t h e rate is considered unchanged, g i v i n g t h e same a v a i l a b i l i t y f o r t h e composite CH-53D as f o r t h e c u r r e n t v e h i c l e . Mission r e l i a b i l i t y i s based on an a b o r t rate of 2 3 p e r thousand f l i g h t hours and an average mission t i m e of . 3 6 2 hour. For t h e composite v e h i c l e mission, r e l i a b i l i t y is considered unchanged, due t o l a c k of s e r v i c e information.

Average mission p r o d u c t i v i t y of t h e c u r r e n t CH-53D i s o b t a i n e d from t h e mission s i m u l a t i o n p r e v i o u s l y d i s c u s s e d .

The mission c a p a b i l i t y of t h e composite CH-53L)- is obtained by adding t o t h e c u r r e n t aircraft v a l u e t h e incremental change i n p r o d u c t i v i t y due t o t h e incremental change i n weight. The p a r t i a l d e r i v a t i v e o f mission p r o d u c t i v i t y with r e s p e c t t o weight is -.022 t o n - k t s . p e r pound ( - 2 2 . 6 kg.m/s/kg).

Operation cost of t h e c u r r e n t CH-53D, e s t i m a t e d a t $357,500 a n n u a l l y , i n c l u d e s naintenance, replacement s p a r e s , replacement GSE, replacement t r a i n i n g , f u e l , and crew costs.

Replacement s p a r e s , replacement GSE, and crew costs are assumed n o t t o change. Change from t h e c u r r e n t aircraft naintenance cost and replacement t r a i n i n g c o s t of maintenance personnel i s a f u n c t i o n of t h e incremental I.ZEIH/FH change f o r t h e composite aircraft. A s mentioned p r e v i o u s l y , t h e MMH/FH are considered unchanged for t h e composite v e h i c l e compared with t h e c u r r e n t CH-53D. Therefore, maintenance and replacement t r a i n i n g costs do n o t change. The effect of composite design on annual f u e l cost is o b t a i n e d from t h e incremental f u e l flow due t o change i n aircraft weight empty c1.8 lbm f u e l / h o u r ( 0 . 8 1 k g / h r ) l t i m e s t h e c o s t of f u e l 11.85 cents/pound of f u e l , ( 4 . 0 8 c e n t s / k g ) l t i m e s t h e annual 500 f l i g h t h o u r s , i . e . , $ 1 6 . 7 a n n u a l l y , c o n s i d e r e d n e g l i g i b l e . T h e r e f o r e , t h e o p e r a t i n g cost o f t h e composite CH-53D i s c o n s i d e r e d t o be t h e same as t h e c u r r e n t CH-5 3D * F l e e t s i z e of t h e composite G I - 5 3 3 i s based on t n e nuqber of a i r c r a f t r e q u i r e d t o m a i n t a i n t h e f l e e t m i s s i o n e f f e c t i v e n e s s of t h e c u r r e n t CH-53D, where mission e f f e c t i v e n e s s i s d e f i n e d as t h e product o f p r o d u c t i v i t y , a v a i l a b i l i t y , and r e l i a b i l i t y .

F l e e t l i f e c y c l e c o s t i s t h e summation of a c q u i s i t i o n c o s t , a s s u n i n g t h a t t h e b a s i c aircraft development c o s t has been a m o r t i z e d , p l u s o p e r a t i n g c o s t for t h e r e q u i r e d f l e e t s i z e f l y i n g a n a v e r a g e o f 5 0 0 h o u r s a y e a r p e r aircraft o v e r a 10-year s e r v i c e l i f e .

The i n c r e a s e d p r o d u c t i v i t y of t h e composi’te aircraft off- sets i t s i n c r e a s e d a c q u i s i t i o n cost, r e s u l t i n g i n a reduced f l e e t c o s t o f $ 2 0 2 , 1 0 0 , 0 0 0 o v e r t h e 10-year s e r v i c e l i f e .

!O OF I I 1 280 300 320 OK 1 . 0 .a . 6 .4 .2 I

I 1 I -

I 0 2 4 6 8 1 FT x 103 1 I I i 0 1 . 0 2 . c 3.0 FIGURE D1. C H - 5 3 M I S S I O N ENVIROHME2iT.

!O N. M I .

$.I r 1 I

E 0 Go 80 G o 160 200

km (a) CRUISE ALTITUDE ELEVATIOP INCREMENT ABOVE T@-OFF b I I # 0 0.5 1.0 1.5 k m F I G W D1. CH-53D M I S S I O N ENVIRONMEDT. (CONTINUED) PERCENT T O N S $00 1 0 . 0 0 0 15:OOO 20;OOO 25,000 kt3 FIGURE D1. CK-53D M I S S I O N ENVIRONMENT. ( C O N T I N U E D ) MINUTES 0 1 2 M I N U T E S F I G U R E 'D1.

CH-53D M I S S I O N ENVIRONMENT (CONTINUED) FRACTION OF HOVER OGE POWER FIGURE D1. CH-53D MISSION ENVIRONMENT (CONCLUDED) RCFE XEITCE S 1. M . J. S a l k i n d and G . S . Molister, ed. A p p l i c a t i o n of Composite !.laterials, S p e c i a l T e c h n i c a l P u b l i c a t i o n 524, ASTM, P h i l a d e l p h i a , P a . , J a n u a r y , 1973.

2 . Composites Recast, Panel Reports of Composite Recast, AF/!\JASA Long Range Planning Study, February, 1 9 7 2 .

3. J. E . Ashton, A n i s o t r o p i c Plate A n a l y s i s , Report FZ1.I-4899 AFML, Advanced Filament and Composite D i v i s i o n , Wright P a t t e r s o n AFB, Ohio, October , 1 9 6 7 .

4. B. E . Kaminski and J . E . Ashton, Diagonal Tension Behavior of Boron-Epoxy Shear P a n e l s , J o u r n a l of Composite Materials, Vol. 5 , October, 1 9 7 1 .

5 . F-15 Composite Wing F l i g h t T e s t , T h i r d Q u a r t e r l y T e c h n i c a l Report MDC A 1546, AFML, Wright P a t t e r s o n AFB, Ohio, February, 1 9 7 2 .

6. Same as R e f . 9 , F i r s t E d i t i o n , August, 1 9 6 9 .

7 . Filament Composite Landing Gear Program, T e c h n i c a l Report AFFDL-TR-72-78, AF F l i g h t Dynamics L a b o r a t o r y , AF Systems Command, Wright P a t t e r s o n AF3, Ohio, J a n u a r y , 1 9 7 3 .

8 . J. W . Moore, PRD-49 A IJew Organic High Modulus R e i n f o r c i n g F i b e r , E. I. Du Pont de IJemours Co., I n c . , T e x t i l e F i b e r s D i v i s i o n , Wilmington , Delaware.

9 . Advanced Composites Design Guide, AFML, Wright P a t t e r s o n AFB, Ohio, T h i r d E d i t i o n , November, 1 9 7 1 .

1 0 . P l a s t i c s for F l i g h t V e h i c l e s , MIL-HDBK-17.

11. J. !?. Davis, N . R . Zurkowski, Put S t r e n g t h and S t i f f n e s s

Where You Heed I t , T e c h n i c a l R e p o r t , 311 C o . , Reinforced

P l a s t i c s D i v i s i o n , S t . P a u l , Plinnesota.

1 2 . T e c h n i c a l Data S h e e t for ''Scotchply" Type 1002, 3 M C o . , St. P a u l , Minnesota.

13. K. 11. Boller, Effect of Hotches on F a t i g u e S t r e n g t h of Conposite Materials , T e c h n i c a l Report ATML-TR-69-6, AFML, Wright P a t t e r s o n AF3, Ohio, J u l y , 1969.

14. Advanced Composite Wing S t r u c t u r e s , , T e c h n i c a l Report AFML-TR-70-231, Vols. I , 11, AFML, Wright P a t t e r s o n AFB, Ohio, December, 1970.

1 2 3 1 5 . E. F. O l s t e n , U. S. Naval Systems Command Report, Contract N00019-71, A p r i l , 1971 t o A p r i l , 1972.

1 6 . F l i g h t w o r t h y G r a p h i t e Reinforced Aircraft Primary S t r u c t u r a l Assemblies, T e c h n i c a l Report AFNL-TR-70-207, Vol. I1 Tasks 3 t h r o u g h 1 2 , AFML, Wright P a t t e r s o n AFB, Ohio , October , 1970, 1 7 . General S p e c i f i c a t i o n f o r Design and C o n s t r u c t i o n of Aircraft Weapon Systems, SD 24 J , Vol. 11, Dept. of t h e Navy, Bur. of Naval Weapons, Washington, D . C.

18. S. Yurenka, I n v e s t i g a t i o n of Advanced Filament Wound Aircraft Landing Gear S t r u c t u r e s , T e c h n i c a l Report AFML-TR-69-229 , AFML, Wright P a t t e r s o n AFB , Ohio, J a n u a r y , 1 9 7 0 .