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AD-773 597

INVESTIGATION OF ADVANCED STRUCTURALCONCEPTS FOR FUSELAGE

Robert J. Mayerjak, et al

Kaman Aerospace Corporation

Prepared for:

Army Air Mobility Research and DevelopmentLaboratory

October 1973

DISTRIBUTED BY:

National Technical Information ServiceU. S. DEPARTMENT OF COMMERCE5285 Port Royal Road, Springfield Va. 22151

DISCLAIMERS

The findings in this report are not to be construed as an officialDepartment of the Army position unless so designated by oLher authorizeddocuments.

When Government drawings, specifications, or other data are used for anypurpose other than in connection with a definitely related Governmentprocurement operation, the United States Government thereby incurs noresponsibility nor any obligation whatsoever; and the fact that theGovernment may have formulated, furnished, or in any way supplied thesaid drawings, specifications, or other data is not to be regarded byimplication or otherwise as in any manner licensing the holder or anyother person or corporation, or conveying any rights or permission, tomanufacture, use, or sell any patented invention that may in any way berelated thereto.

Trade names cited in this report do not constitute an official endorse-ment or approval of che use of such commercial hardware or software.

DISPOSITION INSTRUCTIONS

Destroy this report when no longer needed. Do not return it to theoriginator.

II

Unclassified

Kaman Aerospace Corporation Ucasfe

Bloomfield, Connecticut _______

INVESTIGATION OF ADVANCED STRUCTURAL CONCEPMS FOR FUSELAGE

4 DESCRIPTIVE NOTES (Ty'po Of report and Incusive daf..)

Final Report5 AU THORISJ (Fi.ars middle Inilial, lmst nam.)

Roburt J. MayerjakWilliam A. Smyth

0 RFPORT DATE 70. TOTAL NO OF PACES 7b, No, OF aces

October 1973 144 17 12$0. CONTRACT OR GRANT 1.0 90. ORIGINATOR's REPORT NUI(BERISI

DAAJO2-72-C-0057b. PROjCCT no USA4%11RDL Technical Report 73-72

Task IF162208AI7001_______________________C. 9b. OTHERRPRTNII(n LItnm..1.mab..fd

'UTIO STATMENTKaman Aerospace Report R-1164

ed for public release; distribution unlimited.

192 SPONSORING MILiTAR-' ACTIVETV

Eustis Directorate RDLbrtr

J Fort Eustis, Virginia231 AGSTRACT

Design concepts are presented and evaluated for the application of fiberreinforced composite materials to fuselage structures. The concepts are evalu-ated on the basis of cost-effectiveness and performance when specifically appliedJIto the aft fuselage of the Army AII-lG helicopter. Conparisons to the existingmetal structure show that several of the new concepts provide significantadvantages in initial acquisition cost, life-cycle cost, weight, safety, reli-ability, and maintainability.

DD oe"'"V,'?1473 OSOCT O ARM ~.JU59, JAUnclassifiei

UnclassifiedSecairsty Classification

t4 LINK A LINK 0 LINK CKIY WOROS ROLE WY ROLE WT RULE Wr

DesignComposite materials

StructuresFuselageAlI-IGHlel icopter

Costs

UOnclassi fied

Security Clasulifcation

DEPARTMENT CF THE ARMYU. S ARMY AIR MOBILITY PESEAR(H A OEVELOPMENT LABORATORY

EUSTIS DIRECTORATEFORT EUSTIS.VIRGINIA 23604

This report was prepared by the Kaman Aerospace Corporation underthe terms of Contract DAAJ02-72-C-0057.

Tihe information contained in this report is a result of research conductedto extend the basic understanding of adv-nced design concepts and compositematerials and their application to aircraft fuselage primary structures.

This effort is one of two parallel studies to investigate advancedstructural concepts for helicopter fuselages. The associated studyprogram, under the same title, was conducted by The Boeing Company.Vertol Division, under the terms of Contract DAAJ02-72-C-0056.

Numerous design approaches, material selections, and fabrication techniqueswere investigated for the All-IG Cobra helicopter tail section. A quanti-tative method (math model) for ranking the proposed design concepts wasdeveloped. Altlough each individual concept possesses specific areasof advantages and disadvantages, it was determined that a monocoque/sandwich structure was the preferred design concept.

The report has been reviewed by the Eustis Directorate, U.S. Army AirMobility Research and Development Laboratory and is considered to betechnically sound. It Is published for the exchange of informationand t ie stimulation of future research.

The technical monitor for this program was Mr. L. Thomas Mazza,Technology Applications Division.

C--i

$I

Task IF162208A17001Contract DAAJ02-72-C-0,)57

USAAMRDL Technical Repoxrt 73-72October 1973

INVESTIGATION OF ADVANCEDSTRUCTURAL CONCEPTS FOR FUSELAGE

Final Report

Kaman Aerospace Report R-1164

By

Robert J. MayerjakWilliam A. Smyth

Prepared by

Kaman Aerospace CorporationBloomfield, Connecticut

for

EUSTIS DIRECTORATEU.S. ARMY AIR MOBILITY RESERACH AND DEVELOPMENT LABORATORY

FORT EUSTIS, VIRGINIA

[Approved for public release; distribution unlimited.1

ii

SUMARY

The objective of this program is to prepare and evaluateseveral design concepts for the application of fiber-reinforced composite materials to fuselage structures. Theconcepts are evaluated on the basis of cost effectiveness andperformance when specifically applied to the aft fuselage ofthe Army AH-lG helicopter.

It is concluded from these studies that when compared to theexisting metal structure, the preferred new concepts providelife-cycle cost savings, weight savings, increased safety,reliability, and maintainability.

The preferred concepts are identified herein as boom-finconcept 3, elevator concept 4, and cover concept 3.

iii

FOREWORD

This design study of advanced concepts for fuselage struc-tures was performed under Contract DAAJ02-72-C-0057, TaskIF162208A17001 from the Eustis Directorate, U. S. Army AirMobility Research and Development Laboratory, Fort Eustis,Virginia, and was under the general technical cognizance ofMr. Thomas Mazza of the Technology Applications Division.

The authors acknowledge important contributions made to thisprogram by Messrs. Herbert Guay, Dr. John Hsu, Durwood Lathrop,Andrew Mank, Harold Showalter, Peter Sevenoff, Frank Starses,and William Whiteside, all of the Kaman Aerospace staff.

Preceding page blankv

FlTABLE OF CONTENTS

Page

SUMMARY ................................................. iii

FOREWORD ................................................ v

LIST OF ILLUSTRATIONS ................................... ix

LIST OF TABLES .......................................... xii

LIST OF SYMBOLS ......................................... xiv

MATERIAL SYMBOLS .................................... xiv

LAMINATE CODE SYMBOLS ............................... xiv

STRUCTURAL ANALYSIS SYMBOLS ............................ xv

COST MODEL SYMBOLS ..................................... xvi

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

DESCRIPTION ............................................. 3

EXISTING STRUCTURE ...................................... 3

BOOM-FIN CONCEPTS ....................................... 6

ELEVATOR CONCEPTS ....................................... 28

COVER CONCEPTS ...................................... 39

DESIGN CRITERIA ......................................... 44

SIGN CONVENTIONS .................................... 44

LIMIT LOADS ......................................... 44

STIFFNESS ........................................... 47

WEIGHT .............................................. 48

STRUCTURAL ANALYSES ..................................... 53

MATERIAL ............................................ 53

STRESS .............................................. c7

Preceding page blankvii

r -

Page

STIFFNESS ........................................... 68

WEIGHT .............................................. 76

RELIABILIMY ANALYSIS .................................... 78

BOOM-FIN FAILURE MODE AND EFFECTS ANALYSIS .......... 78

ELEVATOR FAILURE MODE AND EFFECTS ANALYSIS .......... 80

COVER FAILURE MODE AND EFFECTS ANALYSIS ............. 83

RELIABILITY ESTIMATES FOR ADVANCED DESIGN CONCEPTS.. 83

COST ................................................... 89

METHODOLOGY ......................................... 89

COST MODEL EQUATIONS .................................... 91

PRODUCTION COSTS .................................... 99

LIFE-CYCLE COSTS .................................... 101

DOLLAR VALUE FOR A POUND SAVED ........................ 106

CONCLUSIONS ............................................. 108

LITERATURE CITED ........................................ 110

APPENDIX. Failure Mode and Effects Analyses ........... 112

DISTRIBUTION ............................................ 126

viii

I

LIST OF ILLUSTRATIONS

Figure Page

1 Structural Skeleton of Existing Aft Fuselage.... 4

2 Contours at B.S. 194 ....... ............................... 6

3 Structural Configurations for Boom-Fin Concepts1 and 2 ......................................... 9

4 Bond Tooling for Boom-Fin ....................... 11

5 Materials for Boom in Boom-Fin Concept 1 ........ 13

6 Typical Damage From a Tumbled .30-06 RifleBullet .......................................... 14

7 Fitting at B.S. 41.30 ...... ..................... 1........5

8 Assembly at B.S. 41.30 .......................... 16

9 Fin Sections .................................... 18

10 Materials for Fin in Boom-Fin Concept 1..........19

11 Materials for Boom in Boom-Fin Concept 2.........20

12 Materials for Fin in Boom-Fin Concepts 2, 3,and 4 21

13 Structural Configurations for Boom-Fin Concepts3 and 4 ..................... ............................ 23

14 Access Panels for Boom-Fin Concepts 3 and 4 ..... 25



17 Structural Configuration for Elevator Concept 1. 29

18 Bond Fixtures for Elevator Concept 1 ............ 30




22 Structural Configuration fur Elevator Concept 4. 36

ix

Figure Page


24 Cover Concept 1 ................................. 40

25 Cover Concept 2 ................................. 41

26 Cover Concept 3 ................................. 42

27 Cover Concept 4 ................................. 43

28 Coordinates for Boom and Fin Loads .............. 44

29 Peak Turret Response Frequencies Versus FuselageLoading for Vertical Excitations ................ 49

30 Peak Turret Response Frequencies Versus Fuselage

Loading for Lateral Excitations ................. 49

31 Structural Mass Items for Existing Aft Fuselage. 50

32 Total Equipment and Hardware for Existing AftFuselage ........................................ 51

33 Typical Envelope of Allowable Stresses .......... 53

34 Typical Geometry for Boom Analyses .............. 59

35 Section Properties for Existing Boom and BoomConcept 1 ....................................... 69



38 Section Propertie.4 for Existing Boom and BoomConcept 4 ....................................... 72

39 Section Properties for Existing Fin and FinUsed in Boom-Fin Concept 1 ...................... 73

40 Section Properties for Existing Fin and FinUsed in Boom-Fin Concepts 2, 3, and 4 ........... 74

41 Cost Model ...................................... 90

x

Figure Page

42 Individual Cost Element Breakdown ............... 96 I43 Cost Diagnosis ................................. 97

44 Transporter Weight Versus Load .................. 107

45 Simplified Functional Block Diagram ............. 114

46 Simplified Reliability Bloc]: Diagram ............ 114

47 Aft-Fuselage Component Level Reliability BlockDiagram ......................................... 115

x

xi

LIST OF TABLES

Table Page

I Boom-Fin Concepts, Basic Characteristics ........ 7

Ii Elevator Concept-, Bas. Characteris ics ........ 28

III Cover Concepts, Basic Characteristics ........... 39

IV Limit Loads for Boom ............................ 45

V Limit Loads for Fin ............................. 46

VI Structural Mass Items for Existing Aft Fuselage. 52

VII Lamina Properties ............................... 54VIII Laminate Properties ............................. 55

IX Core Properties ................................. 56

X Load Intensities and Stresses for Uniform Shells 58

XI Typical Stresses for Boom Concept 1 ............. 60

XII Typical Stresses for Boom Concept 2 ............. 62

XIII Typical Stresses for Boom Concept 3 ............. 64

XIV Typical Stresses for Boom Concept 4 ............. 66

XV Section Properties for Uniform Shells ........... 68

XVI Weights for New Concept Assemblies .............. 77

XVII Reapportioned 3M Failure Data for AH-IG/HAircraft ........................................ 86

XVIII Predicted Failure Events for Fiberglass SkinAft-Fuselage Design Concepts .................... 87

XIX Predicted Failure Events for Advanced CompositeSkin Aft-Fuselage Design Concepts ............... 88

XX Summary of Production Costs and Weight .......... 100

XXI Total Cost for Each Unit of the Fleet ........... 102

XXII Relative Total Cost ............................. 103

xii

Table Page

XXIII Summary of Cost Model Input Parametersfor Best Estimates .... ............. ... 104

XXIV Detailed Costs in Best Estimate forExisting Structure and New Concepts ....... 105

XXV Failure Mode and Effects Analysis forBoom-rin Concepts ..... ............... 116

XXVI Failure Mode and Effects Analysis forElevator Concepts ..... .............. ... 121

XXVII Failure Mode and Effects Analysis forSecondary Structures ... ............ ... 124

xiii

LIST OF SYMBOLS

Symbols are listed in four categories. The categories areprez nted in the order of their appearance in the text.

MATERIAL SYMBOLS

E E Glass

D E. I. DuPont PRD-49

G graphile

PVC polyvinyl chloride

R300 rigid, closed-cell, expanded PVC foant at 3 lb/cu ft

R400 rigid, closed-cell, expanded PVC foam at 4 lb/cu ft(R300 and R400 are B. F. Goodrich designations)

S S Glass

T thickness, in.

U uniaxial

120 woven fabric, 120 Fyle

181 woven fabric, 181 style

913 woven fabric, 913 style(913 is a Ferro Corporation designation for a zerotwist fabric of a twi.ll weave with 53.7% of the fibersin the warp direction)

LAMINATE CODE SYMBOLS

In this report, a laminate code is used to describe a specificlaminate uniquely. The code is very similar to that describedin Reference 1 for description of orientation. Herein, thecode is extended to the description of material, style, angle,and thickness. The elements of the code used in this reportare listed below.A Each lamina is denoted by a symbol, A, B, C, etc.,

denoting its material, style, angle, or thickness.

A/B Individual adjacent laminae are separated by a slash,ma ~~ u o JC.1a- r o.%J£~,st~s n's j thicknesses are

different.

xiv

(A/B) The laminae are listed in sequence from one laminateface to the other, with parentheses indicating thebeginning and end of the code.

(AN) Adjacent laminae of the same material, style, angle,or thickness are denoted by a numerical subscript.

(A/B)s A subscript S indicates a symmetric laminate. Onlyone-half of the laminate is shown.

STRUCTURAL ANALYSIS SYMBOLS

11,22,12 The properties of the individual lamina aredescribed with reference to a local orthogonalaxis system identified by the numbers 1 and 2.In general, the 1-axis is chosen colinear withthe reinforcement. The subscript 11 denotes aproperty along the 1-axis. The subscript 22denotes a property along the 2-axis. The sub-script 12 denotes a property (such as shear orPoisson's ratio) along both the 1 and 2 axes.

xx, yy, xy The properties of the I.otal laminate are des-cribed with reference to a local orthogonalaxis system identified by the letters x and y.These axes lie in the plane of the laminate.In general, the x-axis is chosen to be nearlyFarallel to the central axis of the structure.The use of the subscripts is similar to thatdescribed for 11, 22, 12.

X,Y,Z The stiffnesses of the total cross-section andthe loads upon a cross-section are describedwith reference to a local orthogonal axis systemidentified by the letters X, Y, and Z. Ingeneral, the local X axis lies along the axisof the structure, be it boom or fin. The localY lies perpendicular to the midplane of the aix-craft and is positive to the left. When use&alone, X, Y, and Z denote coordinate positions.When used in a subscript, X, Y, or Z denotes aquantity along or about the axis.

AST area of stringer, in. 2

c compression, used in superscripts

E modulus of elasticity, ksi

F allowable limit stress, ksi

if stress, ksi

xv

G shear modulus, ksi

I bending stiffness/basic E, in.4

J torsional stiffness/basic G, in,4

N allowable limit load intensity, kips/in.

n load intensity, kips/in.

s shear, used in superscripts

T thickness, in.

t tension, used in superscripts

TF factor used to calculate contribution of cover in bend-ing analyses

principal Poisson's ratio, strain along 2-axis/strainalong 1-axis for load applied along 1-axis

COST MODEL SfMBOLS

CA attrition costs per aircraft life cycle

CAS cost of air shipping tail section from CONUS

CC cost of shipping container

CER cost of special ground support equipment required

CO acquisition cost for one tail section

CLC total tail section costs per aircraft life cycle

CM average cost of material per repair

CMD depot level labor rate

C MH organizational/intermediate-level labor rate

CMS shipping cost of material per repair as a fraction ofmaterial cost

CpS shipping preparation cost

CFD depot-level repair costs

CRO organizational/intermnediate-level repair costs

1 44 -- 1 C t v1

xvi

CSC container surface shipping cost

CSD depot-level scrap costs

Cso orgarnizational/intermediate-level scrap costs

CS tail section surface shipping cost to CONUSCT total operating costs per aircraft life cycle

CTO total initial costs per aircraft life cycle

Cw cost adjustment for weight

DRD average material and equipment costs per tail section

repaired at depot level

Dw dollar value per pound of structural weight

KRD fraction of damaged sections repaired at depot level

KRI fraction of damaged sections repaired at intermediatelevel

KRO fraction of damaged sections repaired at organizationallevel

KSD fraction of damaged sections sczapped at depot level

K so fraction of damaged sections scrapped at organizational/intermediate level

KTS total fraction of damaged sections scrapped

L aircraft life cycle, flight hours

M IMI to inspect and determine damage

M, same as M

M3 MMH to inspect and determine damage, remove and re-place tail section, requisition and obtain replacement

M4 MMH to receive and inspect at depot

M 5 MMH at depot level to receive, inspect, and dispose ofscrap

MMH maintenance man-hours

xvii

MTBD tail section mean time between damage, hr

NA number of tail sections lost to attrition per life cycle

N number of damaged tail sections per aircraft life cycleD

T mean field repair time, including time to remove andreplace 1.3% of damaged tail sections

Ws weight of existing tail section minus weight of newconcept

Y aircraft life cycle, yr

xviii

INTRODUCTION

Aluminum sheet is today the bargain among aircraft materials.It costs roughly 50 cents per pound. However, a conventionalaft fuselage made from this aluminum sheet costs about $35 per

pound, even for good designs in quantity production. Thematerial is inexpensive, but its fabrication and assembly aredear. The reason lies in the concept itself, namely, the handassembly of many individual piezes using fasteners. The con-cept also contributes to f lture maintenance costs because offatigue cracking at or near thousands of fasteners.

Fibrcas composite materials offer an opportunity to escapethese costs and in addition to get, pound-for-pound, a safer,more reliable structure. The opportunity will be wasted, how-ever, if old concepts are ietained. To take full advantage ofthe opportunity for low lUe-cycle cost, and the designer mustuse new design concepts that:

1. Reduce the number of pnrts2. Eliminate production operations3. Simplify assembly4. Improve field maintainability

These precepts were applied to develop new design concepts forthe aft fuselage of the AH-IG helicopter. The aft fuselageassembly is considered herein to consist of three components:the boom-fin, the elevators, and the covers over the driveshaft. Each of t:hese components has a distinctly differentfunction and presents a different problem for design. In thisreport, four new concepts are presented for each of the threecomponents for a total of 12 new design concepts.* The presen-tation includes estimates for production costs and weight foreach new concept. Productioni costs and weight are also shownfoc four aft fuselage assemblies. Each assembly consists of aboom-fin concept combined with the lowest cost elevators andlowest cost drive shaft covers if required.**

In addition, the total life-cycle costs are presented for eachof the four new concepts for aft fuselage assemblies and alsofor the existing metal aft fuselage. The life-cycle costs forthe existing fuselage provide the baseline for comparisons ofthe new concepts to conventional construction. The life-cyclecost model includes all the significant costs incurred in thesupport of a fLeet of 1000 helicopters over a 10-year lifecycle. The principal categories of cost in the model are ini-tial cost, operational cost, and attritional cost.**** Tables I, II, and III summarize the basic characteristics.** Table XX summarizes production costs and weiahts.*** Tables XXI and XXIII summarize life-cycle c6sts.

1

The concepts were developed within the following design con-straints:

1. Low life-cycle cost shall be a principal objective.This objective guarantees that consideration isgiven to low initial cost, high reliability, andgood repairability.

2. The new concepts shall provide improved safety.This requires high tolerance to ballistics damagein addition to adequate strength for design loads.

3. The weight of new concepts for the aft fuselageassembly shall not exceed that of the existingstructure. In this program, the emphasis is uponlife-cycle cost rather than weight reduction per se.The dollar value of weight reduction is accountedfor as one of many elements in a comprehensive modelfor life-cycle cost.

4. The new concepts shall be dynamically compatiblewith the aircraft. Adequate stiffness must beprovided to keep the natural frequencies wellpositioned in relationship to the principal ex-citing forces at main rotor harmonics.

5. The structure shall be constructed principallyfrom composite materials.

6. The new concepts shall be interchangeable withthe existing aft fuselage.

7. The existing tail rotor, gearboxes, and driveshafts shall be used.

2

DESCRIPTION

EXISTING STRUCTURE

Figure 1, from Reference 2, shows the structural skeletons forthe existing boom-fin and the elevator. In Reference 3, themanufacturer provides the following structural descziptions:

"The tail boom is of semimonocoque construction, com-posed of 75ST6 aluminum skins, longerons and stringers.The primary bending structure consists of four longerons,onc located in each of the four corners, and severalstringers spaced between longerons. Each longeron is a

built-up section consisting of one or more hats bondedand riveted together to form one part. A standard "J"section is used for the stringers on the sides andbottom of the boom. Two 900 angles, equally spaced fromthe vertical center line act as stringers on the upperskin surface. These angles also support th tail rotordrive shaft cover, which is a nonstructural fairingextending the length of the boom.

The longerons and stringers are supported by framesspaced at approximately 21.0 inches. Five of these bulk-head frames are considered basic structure as theyeither re-distribute or introduce load. These bulkheadsare located at boom stations* 41.32, 59.50, 143.28,206.0, 227.0.

The vertical fin is a sandwich construction camberedairfoil having an aluminum core and 7075-T6 aluminumskins. The internal structurF consists of a front spar,rear spar, trailing edge and Eeveral intermediate ribs.These components are made of 7075-T6 and 2024-T4 aluminumalloys. The leading edge of the vertical fin is a honey-comb fiberglas door which hinges off of the front sparto allow access to the tail rotor drive shaft. Itextends from the top of the boom to the 900 gear box.

The elevator is a swept airfoil of semimonocoque con-struction made of 2024-T3 aluminum. The elevator isactuated and supported by a 2024-T3 aluminum alloy spar.The leading edge is made from .050 2024-T4 aluminum inorder to resist the imr ct of the shell casings and/orammunition clips ejected from the GAU-2B/A Minigun andthe XM-129 40 mm grenade launcher."

* Abbreviated B.S. throughout this report

3

00

0*

IT

.03/0

Lef t;e' Elevator

C2

Spar Tube

F~guc 1.Structiira. Skeleton of Exi.3tinq Aft Fuselage.

4

The manufacturer also provides the following assemblydescriptions:

"The vertical fin is assembled (permanently) to the tailboom between B.S. 206 and B.S. 227. Support for thevertical fin is provided by the bulkheads at Boom stations206 and 227. At B.S. 206 the forward spar of the verticalfin assembles to the canted bulkhead. At B.S. 227 theaft spar and trailirT edge are assembled to the bulkheadthrough fittings. Additional continuity is provided byshear attachments between the fin panels and boomstructure.

The horizontal stabilizer is part of the cyclic controlsystem. As part of this system it helps to maintain theproper helicopter attitude. To install the stabilizerthe left and right hand spars are inserted into the209-001-908 Horn Assembly at Boom Sta. 140.35. This tiesthe elevator into the cyclic control system. The install-ation is supported by brackets and fittings extendingbetween the B.S. 143.28 bulkhead and a partial frame atB.S. 137.42.

The tail boom, vertical fin and horizontal stabilizer(elevator) together form one installation which is readilyassembled to the forward fuselage at F.S. 299.57 by fourbolts. These bolts are installed from the forward fuse-lage into barrel nut fittings located at the end of eachlongeron. Fuselage sta. 299.57 corresponds to boom sta.41.32.'

The drive train for the tail rotor is mounted on top of theboom a,.. in front of the forward spar of the fin. It includestwo intermediate supports at B.S. 80 and B.S. 143, a 420 gear-box at the junction of the boom and the fin, and a 900 gearboxat F.S. 0.

The forward end of the boom contains shelves of sandwich con-struction for the support of electronic equipment. The elec-tronic equipment can be reached from the forward fuselage andthrough an access panel on the left side at B.S. 90. Two otherlarge access panels are located on the bottcu of the fuselageat B.S. 133 and B.S. 174.

A landing skid is located at the aft end of the boom. It isattached to the bulkheads at B.S. 206 and B.S. 227. A pair ofsmall access panels is provided for the installation of theskid.

5

BOOM-FIN CONCEPTS

General

In this report, a design concept is defined as a singlecombination of:

1. structural configuration2. fabrication process

The boom-fin concepts presented herein differ in structural con-figuration rather than in fabrication process. The configura-tions for the boom include four external contours and two inter-nal arrangements. Each variation seeks to achieve greater struc-tural efficiency through improved geometrical effectiveness.For all contours, a side silhouette view of the boom-fin wouldbe unchanged. A plan view, however, would show less taper inthe boom for contours 2, 3, and 4. The greatest differencewould be seen at B.S. 194. Aft of this station all contoursfair in gradually to produce a conventional, streamlined boattail. Figure 2 shows cross sections at B.S. 194 for eachcontour.

Contour 3 and 4

Contour 2

Contouz I(Existing)

Figure 2. Contours at B.S. 194.

Only one fabrication process is used for all boom-fin conceptsbecause no other process was found to compete with it on acost-plus-quality basis. Essentially, it is a method for alarge single-phase bondment. Details of the fabrication planare described in the discussions for boom-fin concept 1.

6

... l -i ... . i i -I : i .. ., I -" ...¥ 'i --- ------" ' -• ° ... .. I

It is noted that the definition of a design concept excludesthe materials of construction. Thus, it would be possible toconsider a single design concept executed in several alterna-tive materials and to make comparative cost evaluations. Thescope of this study, however, did not permit such evaluationsto be made. Instead, only one combination oi materials isshown for each concept. These combinations do provide therequired strength and stiffness characteristics. In general,the prime material candidate for all concepts is glass-filamentreinforced epoxy. Glass deserves first consideration becauseof its low cost, high strength, ease of fabrication, and easeof lightning protection. However, graphite and PRD-49 areused for boom-fin concept 1 to achieve the required stiffnessand low weight using the existing contour. Details of thetype and placement of materials are presented in the dis-cussion which follows for each concept.

rable I summarizes the basic characteristics of the four newconcepts and the existing structure which is numbered 0.

TABLE 1. BOOM-FIN CONCEPTS, BASIC CHARACTERISTICS

Position of PrincipalNo. Drive Shaft Contour Type Material

0 External Existing Stiffened Sheet Aluminum

1 External Existing Sandwich PRD-49, Graphite

2 External New No. 2 Sandwich S Glass

3 Internal New No. 3 Sandwich S Glass

4 Internal New No. 4* Corrugated S Glass

*The outer envelope of the corrugations is coincidentwith Contour 3.

Boom-Fin 1

Figure 3 shows the structural configuration for boom-finconcepts 1 and 2. The principal element is a thin sandwichshell that forms the boom and gradually changes shape to be-come the fin. Several internal frames, bulkheads, and a ribdistribute concentrated loads to the shell and also contributeto its stability. When completed by the attachment of the fin

7

spars and the support panel for th' 420 gearbox, the shell it-self becomes a single cell tube chat supports all moments,shears, and torques without the aid of stringers or longerons.The tail rotor drive shaft and gearboxes, not shown in Figure3, are located on top of the shell and are enclosed by coversshown in phantom in Figure 3.

An intportant feature of the concept is its fabrication usinga method developed prior to the present study by Mr. FrankClark at Kaman Aerospace. The method produces with only onecure cycle a finished high-quality shell at low cost using low-cost tooling. Raw materials are placed in a female mold; afinished product emerges. The finished shell is a sandwichwith filament reinforcement and a foam core of PVC. The useof a heat-formable foam is a source of large savings of cost.The core can be purchased in flat sheets and then preformedto any shape after preheating it to 2200 F for 10 minutes.The core also accommodates itself to steps in the thickness offacings by yielding at the cure temperature. Upon cooling,the mechanical strength of the core returns. Through suchprocedures, the expensive steps of carving of cores and pre-curing of facings can be eliminated.

Tooling appropriate for the AH-lG boom-fin is shown in Figure 4.

The following bond sequence can be used:

1. Preform foam core segments.

2. Position preimpregnated composites for the outerface into the left-hand and right-hand female molds.

3. Position inserts such as attachment angles and localfacing reinforcements.

4. Position the four fittings at B.S. 41.3.

5. Position the PVC core and local core reinforcements.

6. Position preimpregnated composites for the innerface into the left-hand and right-hand female molds.

7. Position attachment angles for frames.

8. Place inner mandrel into right-hand mold.

9. Rotate right-hand mold and inner mandrel to placethem into left-hand mold.

10. Position splices and splice inserts.

8

Boom-Fin Primary Structure1 Shell

3 Frames2 Bulkheads1 Rib2 Spars1 Support Panel at 420 Gearbox5 Access Panels /15 Total Parts

Drive Shaft Covers -

42

Bottom AccessE rPanel

-'" " Elevator </ ,

Bottom Access oso Panel

°- -Frame 2/ Side Access Panel Electronics

Boom-Fin Shell Frame 1

Figure 3. Structural Configurations forBoom-Fin Concepts 1 and 2.

9

Fairing -- /"

Fin Spar, Upper 7Fin Rib

Fin Spar, Lower

Fin Access Panels /~/

I

420 Gearbox Support/ .-/ / // Tail Skid

Bulkhead 2

Bottom Access Bulkhead 1

Panel ./ "

- Frame 3

Frame 2

Electronics Shelf

for

SPLICE INSERT UPPER SPLICE LAYED UPAFTER L. H. & R. H. ARE JOINED

L. H. BOND TOOL R.H. BOND TOOL

L. H. PRE-BOND R. H. PRE-BOND

LOWER SPLICE FOAM ADHESIVELAYED UP INL. H. ASSEMBLY

BONDED ASSY

SEALER PEEL PLY

BLEEDER CLOTH

VACUUMPORT

BOND SURFACEBAG

BAG SUPPORT

M4ANDREL SUPPORT NC4

MANDREL END PLATE REMOVABLE

INSERT L. H. BONDTOOL USED DURINGLAY-UP OF SPLICE

Figure 4. Bond Tooling for Boom-Fin.

Preceding page blank 3.1

iI

BOND TC'OL SPLICE IN:

OPIICE LAYED UPH. & R. H. ARE JOINED

R. H. BOND TOOL

PRE-BOND

ESIVE

-I" B A G

oINNERCARRIAGE TO

TRANSPORT BONDASSEMBLY INTOAUTO CLAVE

REMOVABLEINSERT L. H. BOND ECCENTRIC SUPPORT FOR

TOOL USED DURING BOND TOOL, ALLOWS THE

LAY-UP OF SPLICE BOND TO ROTATE INTO VERTICALPOSITION & JOIN OPPOSITE HANDBOND TOOL OVER THE INNER MANDREL

!TOOL SPLICE INSERTS

BOND TOOLPLASTIC LAY-UP

FIN "NT L IS AFLEXIBLE BAG,

AIR INFLATABLE

MANDRELROTATES PLASTICSUPPORTLAY-UP BOND TOOL

TYPICAL FRAME

//.- SHEAR TIE MANDRELS

11. Seal and pressure test.

12. Cure in autoclave.

Figure 5 shows a cross section through the boom and indicatesconstruction typical throughout the length of the boom. Eachface of the sandwich is a balanced laminate. The inner andouter surfaces of each face are woven plies of PRD-49 stylecloth oriented at 450 to the boom axis. Between the plies ofPRD-49, uniaxial graphite is located at 00 to the axis of theboom. Typically, the core is PVC R-400 rigid foam .25 inchthick.

I Symbols are defined on page xiv.J Orientation is in degrees from

boom axis.

Face Region Material Style Angle (deg) T E T

Outer 1 D/G/D 181/U/181 45/0/45 .01/.0065/.0ll .02852 D/G/D 181/U/181 45/0/45 .01/.0130/.0lI .0350

Inner 1 D/G/D 181/U/181 45/0/45 .011/.0065/.0ll .02852 D/G/D 181/U/181 45/0/45 .011/.0130/.011 .0350

Figure 5. Materials for Boom in Boom-Fin Concept 1.

Preceding page blank 13

Such construction provides a structure with the followingcharacteristics-:

1. High survivability. The structure is highly redun-dant because its primary load paths are distributednearly uniformly around its perimeter. A single bul-let cannot cause disproportionate damage by strikinga concentrated longeron. Typical damage from a bal-listic strike will be quite local. Figure 6 shoqs atest panel that received a broadside hit from a .30-06 rifle bullet at approximately 20 feet. The bulletwas tumbled by six .25-inch plywood baffles. The testspecimen is a flat sandwich with two-ply, E glassstyle 181 facings on a PVC R-400 rigid foam core .25inch thick.

LL

I 4I''! I I liltIPP j ~

Entrance Exit

Figure 6. pical Dmaqe From a Tumbled .30-06 Rifle Bullet.

2. Good repairability. Again as a consequence of thenearly uniform distribution of load paths, the loadintensity pcr running inch of perimeter is very low.For example, at limit load, at the most criticallocation, the highest direct load, nx, in each faceis less than 600 pounds per inch. Repair in such a1,% l i g fowA .:=k _%r~ ai- mm ~lap joints.

14

3. Good low-velocity impact resistance. The poor low-velocity impact toughness of graphite is overcome bythe extensive mix with PRD-49, which has very goodimpact properties. The faces are by volume morethan two-thirds PRD-49.

4. Low weight. Stiffness requirements are a very signif-icant factor in the design of the boom. The use ofgraphite provides the required stiffness with lowweight.

The attachment of the boom to the forward fuselage is a criti-cal area for strength, alignment, and maintainability. Thisattachment is accomplished with four aluminum fittings thatare bonded into the shell and actually become a portion of theshell that is partially wrapped around the line of action ofthe bolt. Figure 7 shows an isolated view of a typical fit-ting. Figure 8 shows the final relationships between the boom,fitting, and frame at B.S. 41.3. Forged aluminum fittings areused to provide adequate strength. The fittings are subjectto multiaxial stresses and require the interlaminate shearcapabilities of a metal for efficient design. Precise align-ment and mating to the forward fuselage is assured by finalmachining and drilling of the contact faces af'-er final cureand assembly of the interior structure. Good ainitainabilityis assured by the fact that the bolt nuts are d4rectly access-ible and visible from the exterior of the boom.

r--Flange for Attachmentto Closing Frame Composite

L-Aluminum Fitting(Forging)

* r . L k M 1. A 1 'J.

cn Cl)4J (a

C))

co

4-P4

0 01- "-

tpP

4J4

;:z4 4~4+Q)rI -

16

The fitting concept shown in Figure 7 provides several advan-tages. These include:

1. The elimination of mechanical fasteners. Thealuminum fitting is bonded to the compositematerial to make a preassembly for insertion intothe female mold for the boom fin.

2. The opportunity to use higher pressure for thebond between the fitting and the composite.

3. The opportunity to inspect critically the fittingassembly before its use in the boom-fin bondment.

4. The opportunity to retain bolt preload with a highdegree of reliability. Bolt loosening is unlikelybecause little wear is possible in the stackupof clamped parts due to broad areas of contact andbecause a conventional nut with a large nylon self-locking ring is specified for one-time usage.

5. The opportunity for easy ii. ';ction in the field.A torque stripe painted bc en the nut and thefitting allows easy detection of rotation of thenut and possible loosening of the connection.

6. Low weight, low cost, and simplicity.

Figure 8 also shows the attachment of the frame at B.S. 41.30.The frame is riveted directly to the flange on the fitting andindirectly to the shell through four angles. The use of me-chanical fasteners in this area is very desirable because theyare highly loaded and a high-quality secondary bond would bedifficult to achieve. The frame also serves as a terminalboard for electrical connectors and a support for the forwardend of the electronics shelf. The electronics shelf itself isa simple sandwich supported by and attached to the frames byonly four bolts. The shelf is restrained in the fore-and-aftdirection by two angle struts that connect it to the shell atB.S. 41.3.

Figure 9 shows a cross section through the fin and details forthe attachment of the front cover, the spar and access panels.When the front cover is opened, the drive shaft and the controlcables are exposed. Two large, removable panels in the sparprovide access to the interior of the fin. These panelsfacilitate repair of extensive ballistics damage to the finand also permit periodic inspections of the interior facings.Figure 10 indicates the materials of construction for the fin.

17

44

0

U)I44

r I0

O 0

04,0

180

I~,> I/2-,-

L .........

18

Region it at Fin Station 5.lb at Fin Station 50.

PVC, R400, T=.375 in.

/PVC, R300

Symbols are defined on page xiv.orientation is in degrees from boom axis.

Face Re-gion Material Style Angle (deg) T F

outer3 It (DID-) S (18l/181)s (0/45)s (.0ll/.0ll) S .044

lb (DAb/G) S (181A18/tJ) (0/45/0)S (.0ll/.0ll/.026)S .096

2 DA) 181/181 0/45 .011/.011 .022

3 E/E 181/181 0/45 .009/.009 .018

4 (E/E) S (181/181)~ (0/45)S (. 009/. 009)s .036

5 AlumTinum .050 .050 ~

Inner it (DA))S (181/181)s (0/45)s (.011/.O011) S .044

lb (D/D/G) s (181/181/U) 5 (0/45/0)S G.011/.OUl/. 013)S .070

2 DAD 181/181 0/45 .011/.011 .022

Figure 10. Materials for Fin in Boom-Fin Concept 1.

191

Boom-Fin 2

Boom-Fin 2 differs from boom-fin 1 only in contour and ma-terials. The new contour has already been described and shownin Figure 2. It is considerably more efficient. For example,at B.S. 194, an equal weight of material uniformly distrinutedon contour 2 produces a 38% increase in torsional stiffnessand a 76% increase in lateral bending stiffness over that forcontour 1. The improvement is less at all other stations, andzero at B.S. 41.30 where the contours 1 and 2 are identical.The stability is also improved by the increased curvature ofthe shell. The improved efficiency of contoar 2 permits anall-glass boom-fin having the stiffness of existing metalstructure with less weight. Figure 11 shows the constructionthat is typical throughout tie length of the boom. Figure 12shows the materials for the fin.

Symbols are defined on page xiv.I Orientation is in degrees from

boom axisCore is PVC R400, T=.25 inch.

Face Region Material S Angle (deg) T E T

Outer 1 S/S/S 913/U/913 45/0/45 .010/.030/.010 .050

Inner 1 S/S/S 913/U/913 45/0/45 .010/.010/. 010 .030


20

Region It at Fin Station 5.

lb at Fin Station 50.

~ R400, T=3/8 in.

Symbols are defined on page xiv.

Orientation is in degrees from boom axis

Face Region tlateria1 Style Angl e (deg) T E

Outer it (S/S)S (913/913)S (0/45) S00/.10 .040

lb (S/ S/S) S (913/913/J) S (0/45/0)S (.010/.010/.040)S .120

3E/E 181/181 0/45 .009/.009 .018

4 (E/E) S (l8l/181)S (0/45)S (.009/.009)S .036

5 Aluiznni

Inner it (s/S) s (913/913)s (0/45)S ( 0101.*010) S .040

lb (S/S/S)~ (9l3/9l3/U)s (0/45/0)S (.010/.010/.020)S .080

2 s/S 913/913 0/45 .010/.010 .020

Figure 12. Materials for Fin in Boom-Fin Concepts 2, 3 and 4.

21

Boom-Fin 3

Figure 13 shows the structural configuration for boom-finconcepts 3 and 4. In these concepts, the fin is identical tothat for concept 2. The boom, however, is changed appreciably.The boom is fabricated as an open shell with a 12-inch-wide gapat its top, running along its entire length. This gap is clos-ed by a system of removable structural panels. The drive shaft

is located within the contour of the boom. The bearings forthe drive shaft are supported directly by the two intermediateframes. Hinged doors are provided at these locations foraccess to the drive shaft bearings and couplings for daily in-spections. The control system can also be inspected through

these doors. Figure 14 shows additional details for the covers.

Favorable characteristics of concept 3 include:

1. Low weight. Concept 3 uses the entire availableenvelope with corresponding increase in structuralefficiency. At B.S. 194, an equal weight of material,uniformly distributed on contour 3, increases torsion-al, vertical, and lateral stiffnesses by 68%, 50%, and71%, respectively,over that for contour 1.

2. Excellent access to the interior of the boom.

3. Elimination of secondary covers for drive shaft.

4. Elimination of two bottom access panels.

The principal disadvantage is the requirement for removal ofstructural panels to replace drive train components. Figure15 shows typical construction throughout the length of theboom.

Boorn-Fin 4

Boom-Fin 4 has the same structural layout as boom-fin 3.However, it uses single-walled construction over major portionsof the boom. Stability is provided by longitudinal corruga-tions without the use of stringers or longerons. The corruga-tions are tapered bc-th in depth and in width over the length ofthe boom. At the forward end, B.S. 41.30, the convolutionsprovide a natural transition to the attachment fittings.Sandwich construction is used at selected locations, such asthe top covers, the edge of the opening for the covers, theareas around the side access panel, and tie areas around theattachment of the elevators. Figure 16 shows constructionat a typical section through the boom. The fin is identicalto that described for boom-fin 2.

22

Fairing

Boom-Fin Primary Structure

1 Shell Drive Shaft3 Frames Cover2 Bulkheads

1 Rib2 Spars01 Support Panel at 42 Gearbox

7 Access Panels17 Total Parts

To,. Access Panel Typ.

424(Nonstructural) 1

Drive ShaftSupport Bearings

0 Frame 2

SideAccss PnelElectronics 9

Boom-Fin Concepts 3 and 4.

23

Fairing

Drive Shaft Fin Spar, Upper -'

Cover >

Fin Rib

Fin Spar, Lower

Fin Access Panels//420 Gearbox Support

.1-Tail Skid

Drive ShaftSupport Bearings

Bulkhead 2

S- °Bulkhead 1

Frame 3

Framce 2

Electronics Shelf

for

4 IIX4 E4Z

Co 4 111

E- P U) W

Z X >H H d

rI4 t P4 U H T

00 ~0WW0HE4 Wf H 0 0 U P 0

H E-4 0 z z 44A U) H P4 w A00 Ca Z 00U

Oa00W0 H 0 4

:0 1 FE- )w q 0>4~ z O>;44H -

Ozo, 7l 0

zz

H E-1 CI

W IZ w CAH

04A

toU

9 0

iWHA

-0-

I/

Preceing pge blank25

ISymbols are defined on page xivOrientation is in degrees from

Iboom axis.Core is PV" R400, T=.25 inch.

Face Region Material Style Angle (deg) T ET

Outer 1 E/S/S 120/U/913 45/0/45 .004/.010/.010 .024

2 E/S/S 120/U/913 45/0/45 .004/.030/.010 .044

3 E/S/S 120/'U/913 45/0/45 .004/.020/.010 .034

Inner 1 E/S/E 120,/U/120 45/0/45 .004/.010/.004 .018

2 E/S/E 120//120 45/0/45 .004/.030/.004 .038

3 E/S/E 120/U/120 45/0/45 .004/.020/.004 .028


26

2

Symbols are defined on page xiv.Orientation is in degrees fromboom axis.Core is PVC R400, T=.25 incb.

Face Region _______a Style Angle (deg) T ___

Outer 1 E/S/S 120/UJ/913 45/0/45 .004/.010/.010 .024

2 S/S/S 913/U/913 45/0/45 .010/.040/..010 .060

3 S/5/5 913/U/913 45/0/45 .010/.050/.010 .070

4 (S/S/s), (913/913/U)S (45/0/0)S (.010/.010/.005) .05011 ~S*

Inner 1 E/S/E 12OAtJ/120 45/0/45 .004/.010/.004 .018

2 S/S/S 913AJ/913 45/0/45 .010/.010/.010 .030


27

ELEVATOR CONCEPTS

General

The four elevator concepts presented herein differ both instructural configuration and fabrication process. The sequen-tial numbering of the concepts is also the historical order oftheir development, and it will be found that each conceptevolves from its predecessor. The last is the simplest, and

cost analysis also shows it to be cleirly the best. Table IIsummarizes the characteristics of each concept.

TABLE II. ELEVATOR CONCEPTS, BASIC CHARACTERISTICS

Principal MaterialsNo. Bending Element Facing Type

Spar Facings

0 Internal tubular spar Stiffened sheet Aluminum Aluminum

1 Internal tubular spar Shell & sandwich Aluminum E-Glass

2 Internal tubular spar Shell & sandwich Aluninum E-Glass

3 External faces Shell & sandwich E-Glass E-Glass

4 External faces Sandwich E-Glass E-Glass

Elevator 1

Concept 1, shown in Figure 17, contains an aluminum tubularspar and attachment fittings similar to .hose in the existingstructure. This spar tube is the main s;ructural element thatsupports the bending moments and transfers them by socketaction to a cross tube located in the boom. Two aluminum ribs,bonded to the tube, support the outer envelope, which consistsof a leading-edge "C" and a trailing-edge arrow, both madefrom glass fabric. The leading edge is a symmetric laminatewith a total thickness of six plies. The trailing-edge arrowis a sandwich with two-ply facings over a core of PVC R-300rigid foam. The extreme trailing edge of the arrow also con-tains an aluminum strip 2 in. x .032 in. to provide resistanceto handling impacts.

28

Plan View /Elevator Primary Structure_ 1 Spar Tube

1 Attachment Fitting, Bonded to- - Tube

2 Ribs1 Leading-Edge "C"1 Trailing-Edge Arrow1 Tip Cap2 Clip Angles9 Total Parts

Leading

EdgeSpar Tube

Clip Angle

Tip Cap

Channel Insert Trailing-Edge Arrow

Figure 17. Structural Configurationfor Elevator Conc:ept 1.

The following lay-ups are used:

Item Material Style Angle (deg) ET

L. E. "C" (S/E/E) S (913/181/181)S (0/0/45)S .06

T. E. Arrow E/E 181/181 0/45 .02

29

INTENSIFIER

LEADING-EDGE BOND FIXTURE

BOND TOOL

FIBERGLASS FACE SHEETAL SHEET METALSPAR RIGID FOAM

Figure 18. Bond Fixtures for Elevator Concept 1.

30

Elevator 2

Concept 2 is similar to concept 1 in layout and materials. Itdiffers principally in that the outer envelope is made in a one-phase bondment as shown in Figures 19 and 20. This method pro-duces a D-shaped leading edge of all-fiberglass construction.

Plan View Elevator Primary Structure

1 Spar TubeJ , -_ _ -1. - 1 Attachment Fitting, Bonded to

L ..l- -I! " f Tube2 Ribs1 Envelope1 Tip Cap2 Clip Angles8 Total Parts

,-Spar Tube

Envelope- Clip Angle

Figure 19. Structural Configuration for Elevator Concept 2.

31

Elevator 3

In concept 3, the aluminum tubular spar is eliminated. In itsplace a fiberglass spar is used. For most of its length, thespar has a channel shape. At its root end, however, it changesto become circular. A hard-coated, thin-walled, .06-inchaluminum tube is bonded over the spar to provide wear-resistantbearing surfaces for socket action with the cross-tube withinthe boom.

The outer envelope itself is the load path for bending moments.Figure 21 shows that the outer envelope and the tip cap aremade as a single unit. The outer envelope is both bonded andriveted to the spar. The outer envelope can be divided intothree general regions: the leading edge forward of the spar,the middle region aft of the spar, and a trailing-edge arrow.Lay-ups for these regions are listed below.

Region Material Style Angle (deg) ZT

Fwd. of Spar (E4)S (1814)S (0/0/45/45)S .08

Aft of Spar (E4)s (1812)S (0/45)S .04

T. E. Arrow E/E 181/181 0/45 .02

The envelope also contains plies of uniaxial E-Glass orientedsparwise at the location of the spar. On the average theuniaxial cap is .04" x 4.0". The trailing edge is a sandwichsimilar to that used in concepts 1 and 2, including thealuminum strip for handling impacts.

32

PRESSURE BAG

BOND TOOL -BOND TOOL (HINGED) "

REMOVABLE NOSE HALF

TO ALLOW LAY-UPOF SPLICE RIGID FOAM

ZBOND TOOL FIXED

NOSE SPLICE

• BAG

REMOVABLEBOND TOOL SEGMENT


33

PRESSURE BAG

BOND TOL

BOND TOOL (HINGED)

REMOVABLE NOSE HALFTO ALLOW LAY-UPOF SPLICE RIGID FOAM

BOND TOOL FIXED

-k'NOSE SPLICE P S

BAG

B00


33

TRIM AFTER BOND

Lz ___________SURPLUS PREPREG CLOTH WILL~~'L~-GATHER IN THIS AREA WHEN BOND

TOOL CLOSES, ALLOWING THE

GID O:71LAY-UP TO OCCUR~ WHILE TOOL IS OPEN

PRESSURE

COSUR OR

ENT

BOND TOOL

Concept 2. r ±ALU

I

Elevator Primary Structure

1 SparPlan View 1 Sleeve

1 Attachment Fitting2 Rib Segments1 Envelope6 Total Parts

Rib SegmentAttachment

Envelope


Ratuinug jftv- -Bank

351

Elevator 4

Figures 22 and 23 show the structural configuration and thebond fixture for concept 4. In this concept the entire outer

surface, tip cap, end rib, interior spars and interior rigidfoam are laid up and cured in one step. A single-piecealuminum forging is then fastened to the elevator to completethe assembly. This concept has several advantages, which in-clude: low cost, low weight, extremely low parts count, andexcellent strength, reliability, aerodynamics, salvageabilityand repairability.

Elevator Primary Structure

1 End Fitting1 Envelope

2 Total Parts

Envelope En

Fitting


36

TRAILING-EDGE ALUMINUM STIFFENERBOND TOOL TRIM

RIGID FOAM

WIND BOND TOOL

JIA M

SPLITMANDRELS I

AIR BAG FIBERGLASS

FIBERGLASS SPARS IFIBERGLASS END CLOSURE

FIBERGLASS DOUBLERS& TWO STAINLESS STEEDOUBLERS

AIR BAG

WIND JAMMER

SPLITMANDREL 1

- RIGID FOAM /MANDRELS

Figure 23. Bond Fixtures for Elevator Concept 4 FIBERGLASSSPARS

37-

I

TRAILING-EDGE ALUMINUM STIFFENER

BOND TOOL TRIM AFTER BOND

AM ' -GATHERING AREA FOR

SURPLUS PREPREG CLOTH

BOND TOOL

BOND TOOLUPPER HALF

ALUMINUM STRIP

i IBERGLASS AIR BAGFACE SHEET

CAUL PLATE

FIBERGLASS END CLOSURE

FIBERGLASS DOUBLERSTWO STAINLESS STEEL '

DOUBLERS /

AIR BAG .

SPLITMANDREL/i

RIGID FOAM FIBERGLASS- D LAY-UPSPLIT RIGID CORE

MANDRELS CAUL PLATE

F TBE(T -ASSConcept 4. SPARS

Structurally, the principal element is a box beam formed by thetwo channel-shaped spars and the outer envelope. Stainlesssteel doubler plates are interleaved into the lay-up to assistin the transfer of the loads at the attachment to the end fit-ting. The end fitting itself is a single-piece aluminum forg-ing that interfaces properly with the existing cross tube inthe boom.

Typical lay-uos for the envelope are shown below:

Region Material Style Angle (deg) IT

Fwd. of Box E4 1814 (0/45)S .04

Over Box * E8 (1812/U2)S (0/ 4 5 /0/ 0 )S .08

Aft. of Box E/E 181/181 0/45 .02

* Average values, more uniaxial at root, less at tip

COVER CONCEPTS

Table III summarizes the principal characteristics of fourconcepts presented for covers to enclose the drive shaft overthe boom. Concept 3 is the best of the group, with an out-standingly low life-cycle cost and good durability.

TABLE III. COVER CONCEPTS, BASIC CHARACTERISTICS

Principal MaterialsHinge Cover Type

Hinge . Cover

0 1 Side, Continuous Chem-Milled Sheet Aluminum Aluminum

1 1 Side, Discrete Sandwich Aluminum E-Glass

2 1 Side, Continuous Reinforced Sheet Polyalcmer*

3 1 Side, Continuous Formed Sheet " Polyalcmer*

4 2 Sides, Continuous Formed Sheet

Propylene-ethylene type crystalline copolymer

Preceding page blank39

Cover 1

Figure 24 shows cover concept 1. Like the boom, it is a sand-wich shell made in a single-phase bondment. Typical lay-upsare:

Item Material Style Angle (deg) XT

Outer face E/E 181/120 0/45 .014

Inner face E 181 0 .010

The core is R400 with a thickness of .25 inch. The coreextends over the central regions. At the ends, single-wallconstruction is used to simplify the mechanical attachment ofthe hinges and the quarter turn fasteners. The lower long-itudinal edges are also single wall to facilitate fabricationand trimming. The perimeter is locally reinforced with addi-tional plies of cloth and a small aluminum sheet at thefastener holes.

E GLASS FACES,PVC CORE SANDWICH

EXTRUDED ALUMINUM QUARTER TURN FASTENER

HINGE

Figure 24. Cover Concept 1.

One feature of this design is the attachment of each cover tothe hinge by only two fasteners per cover. This provideseconomy in assembly and replacement, and also allows con-vienent removal of the cover for obstruction-free maintenanceof the drive train. Another fetu r s th sort lei-t_ ~ ~~~ gtesor e1h of the

hinges. Each hinge is attached with only two fasteners to the

40

shell of the boom. The hinges can also be very durable with

permanently entrapped hinge pins.

Cover 2

Concept 2, shown in Figure 25, is conventional. It is verysimilar in concept to the existing metal cover. However, itshould be superior to the existing cover because of betterimpact resistance. Its only new feature is the use of a plastichinge; polypropylene and polyalomer are leading candidatesfor the hinge.

STIFFENING ARCHE GLASS COMPOSITE, T = .080 IN.

i TEFLON ANTI-FRET PA

PLASTIC HINGE E GLASS COMPOSITE

RIP STOPPER T = .030 IN.

STAINLESS STEEL, T = .010 IN.


Cover 3

Figure 26 shows cover concept 3. This cover is vacuum moldedin one piece from a .10-inch-thick sheet of polyalomer orpolypropylene. It features extremely low cost and excellenttolerance to impact and mishandling. One concern is failureof the hinge, loss of the cover, and damage to the tail rotor.Such behavior is precluded by the presence of a small steelstrap, called a rip stopper. It is a standby load path thatcomes into pidy only if the hinge begins to jail. The rip

41

stopper restrains the cover and prevents the progressivefailure of the hinge.

STIFFENING ARCH

INTEGRAL HINGEj

RIP STOPPER POLYAL014ER ORSTAINLESS STEEL, .010 IN. POLYPROPYLENE


Cover 4

Concept 4, shown in Figure 27, uses the same fabrication pro-cess as concept 3. It differs only in configuration. Hingesare provided on both sides, permitting service from either sideof the aircraft. This is a m..xed blessing, however. Itincreases serviceability, but it also increases the likelihoodof forgetting to secure both sides of the cover. In addition,the presence of the hinge precludes carrying the stiffeningarch to the base, and thus concept 4 has less lateral resis-tance than concept 3. On the whole, concept 3 is the betterdesign.

42

STIFFENING ARCH

--INTEGRAL HINGETYPICAL BOTH SIDES

RIP STOPPER POLYALOMER ORSTAINLESS STEEL, .010 IN. POLYPROPYLENETYPICAL BOTH SIDES


43

DESIGN CRITERIA

SIGN CONVENTIONS

In this report, two local coordinate systems, shown in Figure28, are used to define the loadings on the boom-fin. In eachsystem, a consistent left-hand sign convention is used for co-ordinates, forces, and moments. The local X-directions lie inthe midplane of the aircraft and along the axis of the boomand along the forward spar of the fin. The positive sense forthe X-directions has an aft component. The Y-direction liesperpendicular to the midplane of the aircraft and is positiveto the left; the Z-direction is orthogonal to X and Y withleft-hand sense. Forces and moments are positve when theirvectors lie in the direction of the positive sense of thecoordinate axes. In all tables of internal loads, the statedloads are those acting upon the forward segment, as shown inFigure 28.

z x

XX

y

y

FUSELAGE STATION

Figure 28. Coordinates for Boom and Fin Loads.

LIMIT LOADS

According to Reference 3, the critical destgn conditions forthe boom are: condition Vb, yaw + 150 recover; condition Vb,yaw - 150 recover; and condition XIV, tail-down landing. Inthe present study, these conditions are identified as cases.12, .13, and .15, respectively. The critical design cond-itions for the fin are cases .12 and .13 only. Tables IV andV summarize the limit loads for the boom and fin.

44

TABLE IV. LIMIT LOADS FOR BOOM

BOOM Fx Fy Fz Mx My Mz

Case Sta. (kips) (kips) (kips) (in.-kips) (in.-kips) (in.-kips)

.12 41.32 -0.09 -2.66 -1.45 75.91 -166.46 -597.7859.50 -0.08 -2.73 -1.35 80.12 -140.15 -549.7380.44 -0.08 -2.73 -1.35 83.64 -11.89 -492.86

101.38 -0.07 -2.96 -1.13 85.12 -83.62 -435.86122.33 -0.07 -2.96 -1.13 81.60 -61.52 -492.73143.28 -0.03 -3.03 -0.47 92.16 -37.80 -312.78164.23 -0.03 -3.03 -0.47 96.07 -27.82 -249.59185.18 -0.02 -2.92 -0.28 102.15 -17.95 -186.54194.30 -0.02 -2.92 -0.28 103.84 -15.28 -159.19

.13 41.32 -0.09 1.67 -1.43 -57.98 -165.30 406.2659.50 -0.08 1.74 -1.33 -61.10 -189.32 376.1480.44 -0.08 1.74 -1.33 -63.35 -111.47 339.75

101.38 -0.07 2.00 -1.11 -63.99 -83.61 293.24122.33 -0.07 2.00 -1.11 -65.98 -60.11 261.11143.28 -0.03 2.09 -0.49 -67.81 -38.63 220.05164.23 -0.03 2.09 -0.49 -70.52 -28.44 176.39185.18 -0.02 2.00 -0.29 -75.51 -18.25 132.82194.30 -0.02 2.00 -0.29 -76.67 -15.52 114.01

.15 41.32 0.11 0.0 1.85 0.0 367.04 0.059.50 0.11 0.0 1.85 0.0 333.*39 0.080.44 0.11 0.0 1.85 0.0 294.63 0.0

101.38 0.11 0.0 1.85 0.0 255.87 0.0122.33 0.11 0.0 1.85 0.0 217.09 0.0143.28 0.11 0.0 1.85 0.0 178.31 0.0164.23 0.11 0.0 1.85 0.0 139.53 0.0185.18 0.11 0.0 1.85 0.0 100.75 0.0194.30 0.11 0.0 1.85 0.0 83.87 0.0

Notes:

1. For cases .12 and .13, the points of application ofthe loads lie at aircraft B.L. = 0 and W.L. = 63.09.

2. For case .15, the points of application of the loadslie at aircraft B.L. = 0 and W.L. = 76.81.

3. The above loads were calculated from load data pre-sented in Reference 3, pg. 4.003 by applying a rota-tion of -3.550 about the Y-axis and appropriate signh~ng_ t-- tansform te- loads into the luc i (true)axes for the boom.

45

TABLE V. LIMIT LOADS FOR FIN

SPAR FLIN SPAN MX M y Mz

Case Sta. Sta. Sta. (kips) [in.-kips) (in.-kips (in.-kips

.12 39.55 70.25 0 -3.21 12.26 -18.95 -164.260.00 49.80 14.22 -2.69 11.65 -18.95 -105.980.00 29.80 28.13 -2.18 10.91 -18.95 - 57.03

100.00 9.80 42.03 -1.69 10.42 -18.95 - 18.401112.61 3.00 50.80 - .27 2.40 0 .1iI

.13 39.55 70.25 0 2.36 3.51 -18.95 129.'60.00 49.80 14.22 2.06 2.96 -18.95 86.j.80.00 29.80 28.13 1.75 3.39 -18.95 47.37

100.00 9.80 42.03 1.43 3.93 -18.95 14.85112.61 3.00 50.80 .16 -.02 0 .02

Notes:

1. Spar stations have the same sense as the local Xdirection.

2. Fin stations run down the spar. Their origin is atthe location of the tail rotor; see Figure 1.Fin station = 109.8 - spar station.

3. Span stations are used in Reference 2. They arerelated to aircraft waterline positions as follows:Span station = W.L. - 67.5.

4. Fx and Fz are zero.

5. Points of application of load lie along the centerline of the forward spar.

6. The above loads were calculated from load data pre-sented in Reference 3, pages 5.01 and 5.02 by applyinga rotation of -44.050 about the Y-axis and appropriatesign changes to transform the loads into the local(true) axes for the fin.

46

STIFFNESS

Stiffness criteria are required to achieve two objectives:

1. To avoid resonances with main rotor harmonics,particularly 1/rev and 2/rev.

2. To assure that changes in the angle of attack ofthe fin as a result of lateral loads are similarto those that occur in the existing structure.

The first objective avoids excessive vibration levels. Thesecond objective avoids changes in the lateral control of theaircraft. Such control changes may not be harmful, perhapseven desirable, but their investigation would be beyond thescope of this study and thus, to be safe, are avoided.

One approach to defining stiffness criteria would be to requirethat the bending and torsional stiffness match the correspondingunbuckled stiffnesses* of the existing structure at each sta-tion along the boom and fin. This approach will lead to satis-factory new designs if the total mass and its distribution aresimilar to those of the existing structure. Such is the generalcase because approximately one half of the total mass of theassembly is composed of nonstructural items common to all de-signs. The match-stiffness-everywhere approach is undesirableas a rigid requirement, however, because it restricts designfreedom unnecessarily and, in some concepts, requires that thestructure be heavier and more costly than necessary.

An alternative, less restrictive, and still valid, approach canachieve the stated objectives. The stiffness criteria can bedefined by deflection requirements. Consider the aft fuselageto be cantilevered in normal position from a rigid wall. Thenthe following three deflections must be satisfied:

1. The vertical deflection from a vertical load of 1kip applied at the tail rotor 900 gearbox shall bebetween 100% and 120% of the corresponding deflec-tion for the existing aft fuselage, calculated tobe 1.136 inches.

*Unbuckled stiffnesses for the existing structure were cal-culated. They are presented and compared to stiffness ofthe new concepts. See page 68 and Figures 35 and 39.

47

2. The lateral deflection from a lateral load of 1 kipapplied at the tail rotor 900 gearbox shall be be-tween 90% and 100% of the corresponding deflectionfor the existing aft fuselage, calculated to be2.81 inches.

3. The change in angle of attack of the fin at the loca-tion of the tail rotor 900 gearbox from a 1-kip loadapplied laterally at the 900 gearbox shall be between80% and 120% of the corresponding change in angle ofattack for the existing aft fuselage, calculated tobe 1.062 degrees.

The above deflection requirements are based upon the observa-tion that the tail rotor, the q0 qearbox, and the fin to-gether constitute 28% of the tot.%! mass of the aft fuselage.This mass dominates the behavior because it is in the antinodalposition for the first two natural modes of vibration in boththe lateral and vertical directions. Thus, when the stiffnessfor this important mass is maintained similar to that of theexisting structure, the frequencies cannot deviate signifi-cantly.

Figures 29 and 30, from Reference 4, show the positions of thenatural frequencies of the existing aircraft in relationshipto the main rotor harmonics. From Figure 29, it is seen thatthe first vertical mode lies between 1/rev and 2/rev, beingmuch closer to 2/rev, and that the second mode lies between2/rev and 4/rev, being closer to 4/rev. Thus, it is appareni-that some reduction of stiffness in the vertical direction isdesirable and should be allowed. Accordingly the criteriastate that the allowable vertical deflections are 100% to120% of those in the existing structure. Figure 30 shows thatthe first lateral mode lies almost midway between 1/rev and2/rev. The second mode lies, unfortunately, in resonance with4/rev. It is believed that the good position of the firstmode should not be disturbed, even if it allows the secondmode to be in resonance. Accordingly, the criteria statethat the allowable lateral deflections are 90% to 110% ofthose in the existing structure.

WEIGHT

It is required that the new concepts weigh no more than theexisting aft fuselage. Figures 31 and 32 provide concise sum-maries of the weight and balance of the existing aft fuselage.Table VI further identifies the items of weight in the exist-ing structure. These figures and tables were prepared fromdata presented in Reference 5.

48

25 Main Rotor

HarmonicsN2 2nd Mode 4/Rev

oA15

71st Mode XXX 2/e10 1/Rev

rrxx: l /Rev5 $4 •

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0 z z f a

Fuselage Loading

Figure 29. Peak Turret Response FrequenciesVersus Fuselage Loading forVertical Excitations.

25 Main Rotor2nd Mode Harmonics

\- 4/RevN20

>1or 15

!0 1st Mode A\XXkXX 2/Rev

4J

14 H 0 1/Rev5 $4 •

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0 1214Fuselage Loading

Figure 30. Peak Turret Response Frequencies

Versus Fuselage Loading forLateral Exritai-ine

49

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51

TABLE VI. STRUCTURAL MASS ITEMS FOR EXISTING AFT FUSELAGE

Item Boom Fin Elevators

Bulkheads, frames, ribs 13.7 2.0 2.2

Cover 56.8 19.9 10.4

Fittings 2.6 1.0 0.8

Formers, stiffeners, angles 2.6 2.8

Longerons 22.7 - -

Stringers 13.3 - -

Spars - 15.4 10.2

Doors and frames, access 6.0 4.0 -

Fairings and fillets 2.3 3.9 0.4

Exterior finish 2.5 0.7 -

Hardware, miscellaneous 3.1 1.2 1.0

Bumper, aux. tail gear 4.5 - -

Trailing-edge strips - - 0.4

Supports - - 2.9

TOTAL WEIGHT, POUNDS 130.1 50.9 28.3

52

STRUCTURAL ANALYSIS

MATERIAL

Material analyses were made to determine elastic constants andlimit strengths for the specific laminates used in the designconcepts. Limit loads and limit strengths are used throughoutthis report.

Composite theory as described in Reference 6 was used tocalculate the laminate properties from the properties of theindividual lanina. Table VII presents the input propertiesfor the individual lamina and the sources for the data. TableVIII presents the calculated properties for the laminates.

It is believed that tests would show that the calculatedstrengths are underestimated because the following conservativeassumptions are made in the failure theory.

1. Limit strength for each lamina is taken as .5xstrength. In effect, this provides an extra 33%margin in ultimate strength over that normallyrequired by aircraft specifications.

2. Failure of the total laminate is assumed when anyindividual lamina reaches its limiting strain. Thelimiting strains can be found from the data in TableVII by dividing the allowable limit stresses by thecorresponding elastic constant.

In general, no weight penalty is incurred by the use of theseconservative strengths because stiffness requirements, ratherthan strength, control the selection of thicknesses.

Figure 33 shows a typical envelope of allowable stresses. Forthe laminates used in this study, the allowable design space istypically a rectangle or a near rectangle.

xy for f = 0

'L:'. ...fc Fc:Y ***.F t

Fxy -Allowable Design Space

Flaure 33. Typical Envelope of A11owbl Stresses

53

TABLE VII. LAMINA PROPERTIES

Allowable Limit Stresses Cksi) Elastic Constants (ksi)Materia,IStyle F 1StFle 1 F52 F 2 F12 Ell E22 G

E,120 -28.3 29.1 -24.9 28.90 +6.95 3410. 3210. 570. .078

D,181 -16.0 35.0 -14.0 34.80 +4.50 4700. 4200. 550. .10*

G, U -70.0 92.5 -15.0* 3.50 +5.00* 21000. 1200 700.* .30*

S, U -63.0 109.5 -12.5 2.38 +4.53 7100. 1740. 516. .25

S,913 -39.6 56.8 -39.6 56.80 +7.00* 4420. 4420. 700.* .10*

Notes

1. Allowable limit stress = .5 x ultimate strength.

2. Properties for E-glass style 120 are based upon thosein Reference 7 for style 7781 at 260°F cure. Ell andE2j are the average of the initial values for tensionan compression. G is the secant modulus to .25 xultimate strength.

3. For PRD-49 style 181, see Reference 8.

4. For Graphite uniaxial style, see Reference 9, SP-286T2.

5. For S-Glass uniaxial style, see Reference 10, 1002S.

6. For S-Glass 913 style, 3ee Reference 11.

*Data reported in the literature is often incomplete. Insuch cases, estimates have been made for the missingproperties. These estimates are identified by anasterisk.

54

TABLE VIII. LAMINATE PROPERTIES

Allowable Limit Allowable Limit Elastic Constants

Ication Loads (kips/in.) Stresses (ksi) (Curvature=O) (ksi)

Concept Region c tC

Boom 1 1 -1.178 1.558 .630 -20.67 27.33 11.05 6204. 2254. 1725.

2 -2.092 2.766 .692 -29.89 39.51 9.89 8970. 2152. 1535.

Boan 2 1 -1.152 2.596 .888 -14.40 32.45 11.10 4693. 2228. 1263.

Boa 3 1 -.592 1.303 .429 -14.10 31.02 10.21 4438. 2071. 1164.! -1.824 -3.484 .610 -22.24 42.49 9.84 5753. 1938 848.3 -1.179 2.382 .520 -19.02 38.42 8.38 5310. 1991. 955.

Bom 4 1 -.592 1.303 .429 -14.10 31.02 10.21 4438. 2971. 1164.2 -1.421 3.126 .932 -15.79 34.73 10.36 4967. 2190. 1180.3 -1.456 2.918 .579 -20.80 41.69 8.27 5743. 2053. 943.4 -1.012 1.640 .520 -20.24 32.80 10.40 4214. 3192. 1187.

Fin 1 it -1.048 1.908 .736 -11.91 21.68 8.36 3498. 3270. 1290.lb -6.488 8.572 1.104 -39.08 51.64 6.65 11722. 2434. 1013.2 -.524 .954 .3681 -11.91 21.68 8.36 3498. 3270. 1290.3 -.403 .415 .2311 -22.39 23.06 12.83 2701. 2607. 1054.4 -.806 .830 .4621 -22.39 23.06 12.83 2701. 2607. 1054.

Fin 2 it -2.168 2.168 1.0841 -27.10 27.10 13.55i 3492. 2392. 1355.lb -5.668 8.880 1.496 -28.34 44.40 7.48! 5658. 2499. 851.2 -1.084 1.084 .542 -27.10 27.10 13.55 3492. 3492. 1355.3 -.403 .415 .231 -22.39 23.06 12.83, 2701. 2607. 1054.4 -.806 .830 .462 -22.39 23.06 12.831 2701. 2607. 1054.

Notes:

1. See Figures 5, 10, 11, 12, 15, and 16 for furtherdescriptions.

2. The properties for each region include both faces.

3. For boom concept 4, the properties shown abovecorrespond to the descriptions in Figure 16. For thestress and stiffness analysis, the thicknesses shownin Figure 16 for regions 3 and 4 are multiplied by1.228 to account for the developed length of thecorrugations.

55

Table IX presents allowable limit strengths and stiffnesses for

core materials. The values shown are derived from the averagetest values in the following manner:

1. The allowable limit strength = 4/9 x the averageultimate strength at 700 F. The factor 4/9 is theproduct of a 2/3 x 2/3. The first 2/3 accounts for a1.5 factor between limit and ultimate load; the second2/3 accounts for a reduction in properties for useat 1600 F.

2. The elastic constant = 2/3 x the average elasticconstant at 700F. The factor 2/3 again accounts fora reduction in properties for use at 160 0F.

It is believed that the average values are attainable in theapplications described herein because the modest thicknessesinvolved permit inspection and subsequent rejection or repairof any area with a significant flaw.

TABLE IX. CORE PROPERTIES

Density Allowable Limit Stress (ksi) Elastic Constants (ksi)

Mtra b/uft) ECFs GR200 2.A .020 .018 1.0 1.0

R300 3.0 .035 .031 1.7 1.7

R400 4.0 .062 .049 3.0 2.3

R500 5.0 .093 .062 4.0 3.0

1600 5.7 .110 .075 4.7 3.7

Notes:

1. Allowable limit stress = 4/9 x average ultimatestrength at 700 F.

2. Elastic Constants = 2/3 x average elastic constantsat 700F.

3. Average data obtained by B. F. Goodrich Tests,Reference 12.

56

-ii

STRESS

The boom-fin is a slender beam-like structure suited toconventional analysis over most of its length. Elementarybeam and torsion theories are used to determine sectionproperties and stresses.

Table X provides a concise summary of a preliminary studyof the load intensities and stresses for each contour. Thedata in Table X correspond to a shell of uniform thicknessand moduli subjected to the limit loading conditions. Thefollowing conclusions are apparent:

1. Case .12, yaw +150 recover, produces the higheststresses.

2. The highest stresses occur in the vicinity of B.S.122.

3. Even the highest stresses are relatively low.

4. On an equal area basis, stresses for contours 2 and3 are approximately 15% lower than those for theexisting contour at B.S. 194.3.

More detailed stress analyses were also performed using theappropriate material properties for the three loading con-ditions at several stations on the boom and the fin for eachdesign concept. In general, it was found that the materialsdescribed for each concept are quite adequate. Tables XIthrough XIV present typical results from such analyses for thelimit loads from Case .12 at B.S. 122.33 for each designconcept. The following notes pertain to Tables XI throughXIV:

1. The idealized structure for these analyses is shown inFigure 34.

2. The element line refers to either a bending element ora shear element.

3. Tension areas are used for the bending elements. Eachbending area is the sum of the transformed areas ofthe concentrated stringer, AST, and TF x the areas ofthe shear elements adjacent to the bending element.

4. The margins of safety, M. S., apply only to thebending element. M. S. = F/f - 1.0.

57

TABLE X. LOAD INTENSITIES AND STRESSES FOR UNIFORM SHELLS

i ldmit LWad Intensities (kips/in.) Relative stress

Boan Perimeter Case .12 Case .13 Case .15 for Equal Areas

Sta, 0contour (in) nXx nxy! nxx nXy nxx nxy nxx nxy

80.4 1 95.3 .686 .115 .521 .096 .376 .060 1.00 1.002 96.8 .662 .119 .492 .096 .371 .058 .98 1.033 102.3 .626 .107 .467 .085 .345 .050 .98 .97

122.4 1 79.5 1.007 .149 .553 .126 .400 .068 1.00 1.002 82.8 .914 .147 .497 .120 .385 .006 .95 1.033 885. .835 .128 .454 .106 .340 .056 .92 .96

164.2 1 63.6 7.94 .243 .567 .176 .402 .079 1.00 1.00

2 69.4 .648 .214 .463 .154 .371 .077 .89 .963 74.9 .581 .185 .412 .134 .306 .062 .86 .90

194.3 1 52.2 .749 .356 .541 .257 .361 .089 1.00 1.002 60.2 .537 .278 .389 .201 .314 .085 .83 .903 65.3 .478 .236 .344 .170 .245 .067 .80 .83

.otes:

1. nxx is the maximum direct stress per inch of perimeterIt is a close approximation to consider that both apositive and a negative nxx occur. In reality, thepositive and negative values are slightly differentbecause a small axial load exists along the boom. Theaxial load is a component from the resolution of thevertical shear into the boom axes which are inclined-3.550 to the vertical.

2. nxy is the shear flow.

58

View Looking Forward

Typical bending24 1 2 element, stringer

24 1 2

22 4 Typical shearelement, skin

20 z /etc62

19 yJ X I 7

18i B.S. 194.3

16 i 10

14 13 12

ContourB.S. 41.0

Figure 34. Typical Geometry for Boom Analyses.

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67i

STIFFNESS

The three boom contours differ appreciably in their efficiencyfor stiffness. Table XV summarizes the results of a prelim-inary study to assess the relative efficiencies, whereinsection properties were determined for a shell of uniformmoduli and uniform thickness of .01 inch. The columns ofrelative properties for equal areas in Table XV may be inter-preted as factors for an increase in effective modulus due togeometry. The improvements are significant, particularly intorsion and lateral bending.

TABLE XV. SECTION PROPERTIES FOR UNIFORM SHELLS

Relative Properties

Boom Area Area Properties (in. 4 ) for Equal Areas

Sta. Contour (in. 2 ) J IV 17 J IV I

80.4 1 .953 198.8 .23.9 90.9 1.0 1.0 1.02 .968 213.6 126.0 100.2 1.06 1.0 1.083 1.023 245.5 161.3 304.9 1.15 1.21 1.07

122.4 1 .795 115.2 71.8 52.7 1.0 1.0 1.02 .828 137.0 72.4 67.3 1.14 .97 1.233 .885 163.2 102.4 71.1 1.27 1.31 1.21

164.2 1 .635 59.1 36.8 27.0 1.0 1.0 1.02 .694 81.2 41.1 43.1 1.26 1.02 1.463 .749 101.7 60.1 45.8 1.46 1.39 1.44

194.3 1 .522 32.7 20.4 15.0 1.0 1.0 1.02 .602 52.1 24.6 30.4 1.38 1.04 1.763 .653 68.7 38.4 32.1 1.68 1.50 1.71

Detailed analyses using the appropriate materials for eachconcept were also made. The results are shown and comparedto the unbuckled section properties for tLe existing structurein Figures 35 through 40. The unbuckled properties werecalculated from data presented in reference 3. Calculationsare required because the section propertieis presented inreference 3 are the buckLed properties co::responding to limitload conditions. To obtain the unbuckled properties, newcalculations were made using the tension areas for each elementas reported in reference .1.

68

9

8

Z upEIy existing

7 unbuckled

Left

54

GJ

43

2

1

40 t0 80 -100 120 -140 160 -180 200

Boom Station, inchez

Figure 35. Section Properties for Existing Boomi and Boom

Concept 1.

69

r ......

9

8

7 un\uckld zu

0

6 y

* Left

.4

0

I-f

II8

z upEly existing

7 ~unb''-:kled

6 y

- Left

4GJN

4.44.4-.H 3

2

3-3

0 40 60 80 100 120 140 160 180 200

Boor# Station, inches

Figure 37. Section~ Properties for Existing Boom and Boom

Concept 3.

71

9

8

My~ existing Z up

7 unbuckled

% sr Left.A

44GJ

4.4-43

*rI 3

2

14

0 40 60 80 100 120 140 160 180 200

Boom Station, inches

Figure 38. Section Properties for Existing Boom and Boom

Conlcept 4.

72

35

Z Forward

30Y Note:30 Left

EI y-n lOxEI z

25I(N

~20

.1 x EI~ existing

~~.15

00

~44'44

4.3

10

0 10 20 30 40 50 60

Fin Station, inches

Figure 39. Section Properties for Existing Fin and

Fin Used in Boom-Fin Concept 1.

73

30,0

EEX

0

C 10102

30 40 50

30 60Figure 40* inchesFin used Prp.O

c p ~ 2 ne~ts2, 3and q

74

It is apparent from Figures 35, 36, and 39 that concepts 1and 2 approximately match the stiffness of the existing un-buckled structure at every station. Concepts 3 and 4, however,deviate significantly from the existing structure. In general,the torsional stiffness is less than that of the existing boomand the lateral stiffness is slightly higher than that of theexisting boom. Additional analysis shows that these deviationsare self compensating, and that the entire boom-fin assembly hasessentially the same stiffness as the existing structure. Forexample, if boom-fin concept 3 and the existing assembly wereeach cantilevered from a rigid wall and loaded with a 1-kipload at the 900 gearbox, the following deflections and rota-tions at the 900 gearbox would occur:

1. vertical deflection from a vertical load1.136 inches for the existing design1.262 inches for concept 3

2. lateral deflection from a lateral load2.81 inches for the existing design2.65 inches for concept 3

3. angular rotation (change in angle of attack) of thefin from a lateral load1.062 degrees for the existing design1.064 degrees for concept 3

Additional calculations were made to provide some insight intothe relative importance of the stiffnesses of the boom and thefin. Of particular interest is the question: what are theorigins of lateral deflection at the 900 gearbox in responseto lateral load at the 900 gearbox? In the existing structurethe following are the origins:

44.3% from lateral bending of the boom19.6% from torsional twisting of the boom33.7% from lateral bending of the fin2.4% from torsional twisting of the fin

These percentages help in visualizing the consequences ofchanges in a particular component of stiffness. In onerespect, however, the percentages do not reflect the fullsignificance of the lateral stiffness of the fin itself.Lateral stiffness of the fin is strongly coupled to changes inangle of attack because of the sweepback of the fin. Lateralbending of the fin in response to a control input produces achange in angle of attack of the fin that would in forwardflight produce air loads that reduce the effectiveness of thecontrol input. The s-gnificant fact is that the control"ee f th airraft can be affected by the lateral bending

75

stiffness of the fin.

One additional feature of the new design is noteworthy. Inthe new fin concepts described herein, the elastic axis ofthe fin moves forward to approximately the location of theforward spar. This is advantageous because it is more nearlyin line with the line of action of the lateral loads on thefin, resulting in less torsional deflection.

WEIGHT

Table XVX lists the calculated weights for each of the newdesign concepts. It can be seen that the designs include asignificant allowance of 15 pounds in concepts 1, 2, and 3,and 21 pounds in concept 4 under the category of "Other" toaccount for miscellaneous reinforcement, fasteners, brackets,etc. The allowance for concept 4 is larger to account forsandwich reinforcement around the top opening.

76

TABLE XVI. WEIGHTS FOR NEW CONCEPT ASSEMBLIES (LB)

Concept

Shell, BS 41 to BS 194 51.15 80.43 72.04 71.62

BS 194 to FS 50 14.12 20.51 17.96 17.84

FS 50 to FS -22.5 17.36 19.98 19.80 19.78

Covers, fin and 420 gearbox 5.15 5.15 5.15 5.15

Electronics shelf 3.66 3.66 3.66 3.66

Elevators, concept 4 23.00 23.00 23.00 23.00

and frames 8.90 8.90 9.50 9.50

Bumper, aux. tail gear 4.50 4.50 4.50 4.50

Fittings, BS 41 4.40 4.40 4.40 4.40

Exterior finish 3.20 3.20 3.20 3.20

.*alties 3.75 3.75 2.50 2.50

Covers, boom, concept 3 8.20 8.20 0.00 0.00

Fasteners, boom opening .00 .00 2.41 2.41

Other 15.00 15.00 15.00 21.00

Lightning protection 10.00 .00 .00 .00

Total Weight 1".39 200.68 183.12 188.56

77

RELIABILITY ANALYSIS

EjThe reliability analysis consists primarily of performing fail-ure mode and effects analyses of those boom-fin, elevator andsecondary structure design concepts selected for detailed eval-uation in order to estimate qua-itatively the reliability ofadvanced aft-fuselage designs compared to the existing AH-IGstructure. The results are then applied to available histor-ical failure data, and quantitative reliability estimates aregenerated for input to the life-cycle cost evaluation model.

Detailed tabulations of the results of the failure mode andeffects analyses are presented in Tables XXV through XXVII ofAppendix II. A discussion of these results for the boom-fin,elevator and secondary structure design concepts considered ispresented in the following paragraphs.

BOOM-FIN FAILURE MODE AND EFFECTS ANALYSIS

Failure mode and effects analyses (FMEA) have been performedfor each of the four boom-fin concepts selected for detailedevaluation. In performing these analyses, it was evident thatthe structural design features which significantly affectreliability are quite similar for the four concepts. Theconceptual design level FMEA applied during this program doesnot involve sufficient quantitative analysis to permit easyrecognition of slight differences in failure modes and effectsamong the competing design concepts. For this reason, and toavoid repetitive listing of identical failure modes, a single

boom-fin FMEA is presented in Table XXV. This FMEA is appli-cable to the four design concepts under evaluation and makesliberal use of notes throughout the failure mode tabulation toindicate those differences among the concepts which should behighlighted.

Table XXV shows that the failure probability codes assigned tothe assumed failure modes are all quite low, either class C orclass D. This is so primarily because the possibility of thesefailures occurring was considered during the design process,and specific design features were incorporated in the boom-finconcepts to eliminate, or at least to minimize, the probabilityof such failures. For example, metal-to-metal contacts areused at all attachment and fastening points, wherever possible,to minimize the probability of loosening and wear at theselocations, even when this requires the insertion of the metalinto the outer skin during the bonding process. Similarly,the outer skins of the advanced concepts have been designedwith sufficient strength to tolerate delamination, bond separ-ation, punctures and tears over bmall areas with m.0 Zif aeffect on the load-carrying capability of the skin.

778

The large majority of assumed failures listed in Table XXVhave little effect on safety of flight and have been assignedcriticality codes in the negligible and/or marginal categories.The only failures which might be serious enough to be categor-ized as critical or catastrophic involve fracture and/orbuckling of the outer skin or internal frames/bulkheads, orthe loss of frame to skin attachments; however, these failuresare all assigned to class D, indicating that they are notexpected to occur during the service life of the aircraft.The design strength of the advanced boom-fin concepts makesfailures of this type extremely unlikely.

As mentioned previously, there appear to be only slight dif-ferences in the predicted reliability of the advanced boom-fin concepts evaluated. The thickness of the fiberglasssandwich outer skin of concepts 2 and 3 is dictated by stiff-ness requirements, and as a result, this skin is actuallystronger than the advanced composite sandwich skin of concept1. Thus concepts 2 and 3 exhibit greater resistance to cracks,fractures, buckling, dents and punctures. The corrugatedfiberglass laminated skin of concept 4 is judged to be superiorin resistance to delamination and is also highly dent resist-ant. On the other hand, concepts 3 and 4 have a largerenvelope because of the location of the drive train within theshell. These concepts also require an increased number ofmechanical fasteners to secure the panels. Concepts 1 and 2,however, mount the drive train on top of this primary structureshell and require secondary structure covers to enclose it.in general, it is felt that the concepts utilizing fiberglassskin construction are more desirable from a reliability stand-point than concept 1, which uses an advanced composite sandwichskin. However, at this stage of the design process there islittle basis for establishing differences in the reliabilityof the three boom-fin concepts that utilize fiberglass skins.

Any one of the four advanced boom-fin concepts should producea structure with reliability characteristics significantlysuperior to the existing AH-lG metal boom-fin. This increasedreliability results primarily from a reduction in the totalnumber of parts, a reduction in the number of mechanicalfasteners used, the elimination of production operations, andsimplification of assembly. Each of the advanced conceptsyields a clean structure with high fatigue strength andexcellent resistance to impact damage. Structural efficiencyand long fatigue life are enhanced by the absence of a mechan-ical joint between the boom and the fin and by the improveddesign of removable panels and covers.

79

ELEVATOR FAILURE MODE AND EFFECTS ANALYSIS

The FMEA for the four elevator advanced design concepts select-ed for detailed evaluation is presented in TableXXVI. For thesame reasons cited in discussing boom-fin failure modes andeffects, a single FMEA applicable to all four elevator conceptshas been tabulated.

As a result of design efforts to maximize the resistance ofelevator concepts to both inherent and external failure causes,all assumed failures listed in TableXXVI have been categorizedas either class C or class D failures. This indicates thattheir probabilities of occurrence are relatively low. Inaddition, it is noted that almost all of the assumed failureshave been assigned to either the negligible or the marginalcriticality categories, the one exception being inherent crackor fracture failures of the elevator spar resulting from over-load or fatigue. Since this type of failure is difficult todetect in its early stages, it is conceivable that a seriousfracture failure might eventually occur sequ,'ntially from aninitial minor crack. In the most extreme ,;ase imaginable,separation of the elevator from the aircraft could result,which warrants categorization of spar fr7-tures as possiblecritical failures. All elevator design concepts, including theexisting AH-lG elevator design, are susceptible to this failuremode. Again, it is pointed out here that spar crack/fracturefailures have been recoqnized as a potentially critical failuremode, and every effort has been made to preclude the occurrenceof such failures in evolving the elevator design concepts.Thus, it is unlikely that serious spar fracture failures willbe experienced during the service life of the aircraft.

Improved reliability has been a factor of prime concern indeveloping each of the elevator advanced design concepts;because of this, the differences in predicted reliabilityamong the four concepts evaluated are expected to be small.An aluminum tube and two ribs are the main load-bearing mem-bers in concept 1. The leading- and trailing-edge outer skinof this design is well supported and lightly loaded. As a re-sult, concept 1 is not likely to experience cracks, fracturesor buckling of the outer skin from inherent causes. Further-more, the increased thickness of the leading edge increases itsresistance to damage from foreign object impacts. Concept 2is similar to concept .; nowever, the one-piece, D-shaped spareliminates the leading edge to spar bond found in concept 1.Concept 3 utilizes a fiberglass spar bonded to a one-piece,load-carrying skin. This design eliminates the buried outerrib and is more resistant to impact damage because ot theincreased strength of the outer skin. Concept 4 is unique in

80

that it uses only two pieces: the elevator itself and the endfitting. This is a highly reliable design since it is complete-ly integral and involves only primary bonds. A redundantfeature is provided by the two spars of the box beam. Stainlesssteel bearing plates, interleaved in the caps of the box beam,aid in the transfer of load to the fitting at the root. Theend fitting is a forging with integral lugs and attachmentfittings, thus eliminating metal-to-metal bonding operationsfound in the other elevator concepts.

Although these advanced elevator design concepts exhibit wear,fretting and corrosion failures similar to those experiencedby the existing-elevator and will utilize essentially the sameone-bolt attachment to the control shaft, it is evident fromthe preceding discussion that each of the advanced conceptsincorporates features which result in an improvement in reli-ability compared to the existing AH-lG elevator design, throughthe principles of parts reduction, elimination of fabricationoperations, and simplified assembly.

COVER FAILURE MODE AND EFFECTS ANALYSIS

Table XXVII presents the FME. zor four secondary structuredrive-shaft cover design concepts that have been selected forevaluation. As was done previously for the boom-fin and ele-vator, a single FMEA is applied to the four drive-shaft coverconcepts, with appropriate notes being used to highlight differ-ences among the concepts. The data of Table XXVII show thatnone of the failure modes assumed for the drive-shaft coverswill seriously affect aircraft performance; therefore, all havebeen assigned to the negligible criticality category. Thesuccess of attempts made during the design effort to minimizefailure occurrences is evidenced by the fact that most of thedrive-shaft cover failures listed in the FMEA are not expectedto occur in service and havt been assigned to probabilityclass D. The remainder of the failures ire considered to berandom occurrences, and thus they are ranked in probabilityclass C. These include fastener wear and fretting and battledamage for all four concepts evaluated, torn hinges for thoseconcepts using polypropylene hinges, and foreign object damagefor the two fiberglass design concepts.

There is considerable similarity in the reliability aspects ofthe four drive-shaft cover design concepts, but there are alsodifferences which uniquely affect reliability. All use thesame type quick-release fasteners and overlap adjacent covers,so each is mechanically entrapped in the event of a fastenerfailure. Overlapping the covers also precludes foreign matterbeing carried by the airstream from blowing in under thecovers.

81

Concept 1 is a fiberglass sandwich composite structure boltedat the corners on one side to a short-length aluminum hingewhich in turn is bolted to the fuselage at two points. Adja-cent covers are attached to the same hinge. Since the coverscan be removed without extracting the hinge pin, the pin canbe locked in place. Although this hinge is of heavy gaugealuminum and uses a cadmium-plated steel hinge pin, it mightstill experience wear, fretting and corrosion failures to somedegree.

Concepts 2, 3 and 4 all use a full-length one-piece polypropy-lene molded hinge. This type of hinge is not susceptible towear, fret and corrosion failures; however, possible tearingof the hinge is of some concern, particularly at very lowtemperatures. As a precautionary measure, a thin-gauge stain-less steel rip-stopper is used at the leading fastener hole ofeach polypropylene hinge to limit any tearing that might occur.In concept 2 the hinge is riveted to the laminated fiberglasscover. Concepts 3 and 4 are molded in one piece so that thehinge is an integral part of the cover, along one edge in con-cept 3 and along both edges in concept 4. This design elimi-nates hinge-to-cover attachment problems.

Concept 1 is susceptible to bond separation failures, whileboth concepts 1 and 2 might experience delamination failuresat the edges of the cover. Because concepts 3 and 4 aie moldedpolypropylene, neither is susceptible to either of these fail-ure modes. In addition, the resiliency of polypropylene makesboth concepts 3 and 4 relatively immune to puncture/tear andforeign object damage. Since concept 4 does not have a stifflateral load path, as provided in concept 3 by the molded-inarch-shaped stiffeners extending all the way to the fastenerholes at the free edge, it will be subject to more hinge flex-ing and will probably be fabricated in a heavier gauge. Thiswill further enhance the resistance of concept 4 to externaldamage events.

Based on these considerations, it seems possible that concepts3 and 4 might prove to be somewhat more desirable from a reli-ability standpoint than either concept 1 or 2. However, allfour of the advanced concepts evaluated show promise ofincreased reliability compared to the existing AH-lG drive-shaftcover designs, due primarily to reduction in the number of partsand fasteners, choice of material, and simplified assembly.

These advanced design concepts are also under consideration forsecondary structure fairing and access door applications.Thus, fairings and access doors are expected to exhibit essen-tially the same failure modes and effects as the drive-shaftcovers. No separate FMEA is presented here for the fairings

82

and access doors, since the FMEA for the drive-shaft coversshown in Table XXVII is also applicable to other secondarystructure components using the same design concepts.

The FMEA for the tail skid secondary structure is summarizedin Table XXVII. The simple steel tube structure of the tailskid is susceptible to very few failure modes, none of whichis considered serious enough to warrant a criticality code

classification beyond marginal. Of the three assumed failuremodes tabulated, corrosion and battle damage are expected tobe highly unlikely occurrences for the tail skid and have beengrouped in probability class L. Should a projectile hit resultin a hole in the tail skid tube, the tube should be replaced;thus, battle damage is assigned a possible marginal criticalitycode as indicated in Table XXVII. A bent tail skid tube causedby a hard landing is considered to be a random occurrence andis assigned to probability class C. If the tube is bentseverely, it should be replaced; therefore, a marginal criti-cality code is appropriate for this failure mode.

The proposed tail skid design concept does not differ radicallyfrom the existing AH-lG tail skid; however, a bending failureof the new design is less likely to result in the fuselagestructure's contacting the ground. This in itself is a note-worthy reliability improvement.

RELIABILITY ESTIMATES FOR ADVANCED DESIGN CONCEPTS

In discussing the failure mode and effects analyses for theboom-fin, elevator and secondary structure advanced designconcepts, it was noted that all the concepts under considera-tion are expected to result in a significant reliability im-provement compared to the existing AH-lG aft-fuselage struc-ture. Possible areas of improved reliability have been men-tioned based on a qualitative assessment of the proposed de-sign concepts. However, in order to evaluate the readuction inprogram costs that can be achieved by using an edvanced aft-fuselage design for the AH-lG, it is necessary to derive aquantitative reliability comparison. Rellabiliy estimates forthe existing AH-lG aft fuselage and for the promising advanceddesign concepts can be made based on available 3M data for theAH-lG/J helicopter.

The most recent available 3M data summary of the detailedmaintenance actions list for Marine Corps AH-lG/J aircrafttabulates all recorded malfunctions/failures for the two-yearperiod from June 1969 through July 1971. Since the aft-fuselagesection is listed as a segregated work unit code in this sum-mary, all the failure modes of interest are conveniently qroupedtogether. Under the aft-fuselage work unit code, the recordedmalfunction/failure events are classified among approximately

83

twenty-five malfunction codes. For the purpose of this analy-sis, these events have been reapportioned among nine generalcategories which correspond more closely to the assumed failuremodes of the boom-fin, elevator and secordary structure FMEA'sdiscussed previously. TableXVII shows the 3M failure/damagedata grouped under these nine categories, as well as the per-centage of total events falling within each category. The two-year period covered by the 3M data involved a total of 23,948flight hours. Thus, the 225 failures recorded duringthis period yield a mean time between failures of 106.4 hours.The data given in Table XVII provide a basis for estimating thenumber of malfunctions/failures and the coincident MTBFthat might be expected for the advanced aft-fuselage designconcepts, operating under the same conditions, through theapplication of engineering judgment and a detailed knowledgeof the promising concepts.

As noted previously, the advanced aft-fuselage design conceptsselected for evaluation are all expected to exhibit improvedreliability characteristics compared to the existing AH-lGaft fuselage, resulting primarily from their common advantageousfeatures of a reduction in total number of parts and totalnumber of mechanical fasteners used, elimination of productionoperations, and simplification of assembly. In general, dif-ferences expected in reliability among the various aft-fuselagedesign concepts themselves are difficult to evaluate at theconceptual design level of analysis. One significant differenceis noted, however, in that the outer skin of the fiberglassboom-fin concepts is stronger than the outer skin of the ad-vanced composite boom-fin concept. Because of this, thefiberglass design3 of boom-fin concepts 2, 3 and 4 offer greaterresistance to cracks, fractures, buckling, dents and puncturesthan the advanced composite design of boom-fin concept 1. Inaddition, the technology involved in fabrication and productionof fiberglass structures is further developed than the corres-ponding technology for advanced composite materials. For thesereasons it is possible to distinguish between the fiberglassand advanced composite skin structures in estimating the ex-pected number of malfunction/failure events from the 3M dataof Table XVII.

Table XVIII presents the predicted number of malfunction/failureevents for advanced aft-fuselage designs u~ing the fiberglassouter skin of boom-fin concepts 2, 3 or 4, based on estimatedpercentage reductions in the number of events recorded in the3M data as indicated for each of the nine failure event cate-gories. A reduction of 70 percent is estimated for broken/loose/missing hardware failures because of the reduction intotal number of parts and particularly in the number of mechan-ica! fastcncrs used in Lhe advanced desigr concepts. In the

84

crack/fracture/luckling category, the estimated reduction is60 percent, due primarily to the increased skin strength of thefiberglass designs. Note that these two categories, where thelargest percentage reductions in the number of malfunction/failure events are expected, account for more than half thetotal events recorded. Reductions are estimated to be 50 per-cent in the wear category due to design features for minimizingwear, and 50 percent in the foreign object damage and puncture/tear categories due to inherent skin strength; while 30 percentreductions are estimated for the corrosion and improper ad-justment/alignment categories. The noncorrosive propertiesof f:.berglass are responsible for the estimated redaction incorrosion failures, and fewer number of mechanical fastenersand simplified assembly account for the estimated reduction inimproper adjustment/alignment failures. All predicted reduc-tions ii, the number of malfunction/failure events are consid-ered to be conservative, but reasonable. No reductions areestimated for the number of events in the battle damage andno defect categories in the interest of maintaining thisconservative approach. The resulting number of predicted mal-function/failure events for the fibeiglass skin aft-fuselagedesigns is 116.9, which yields a mean time between failures of204.9 hours.

The aft-fuselage design using the advanced composite skin ofboom-fin concept 1 is considered in Table XIV. Here it isnoted that the estimated reduction in the number of malfunc-tion/failure events for the advanced composite skin design isthe same as that for the fiberglass skin designs for allfailure categories except those where failure occurrencesdepend on the strength of the fuselage outer skin; namely,crack/fracture/buckling, foreign object damage, and puncture/tear. Because the advanced composite outer skin of boom-finconcept 1 is not as strong as the fiberglass skin of boom-finconcepts 2, 3 and 4, a reduction of only 30 percent is esti-mated in the number of crack/fracture/buckling failures. Forsimilar reasons, and in keeping with the conservative approachof the analysis, reductions of only 25 percent are assumed inthe number of failures occurring in the foreign object damageand puncture/tear categories. With these differences accountedfor, the number of malfunction/failure events predicted forthe advanced composite aft-fuselage design is 138.2, yieldinga mean time between ilures of 173.3 hours.

In summary, available 3M data records for the AH-lG/J aircrafthave been used to predict the expected mean time between fail-ures for promis. ig aft-fuselage design concepts. The analysisindicates that design concepts using fiberglass fuselage skinsare expected to show an increase in MTBF of approximately 93percent. For the aft-fuselaqe concept using advanced composite

85

skin, the expected increase in MTBF is about 63 percent.Because the analysis has been performed at the conceptualdesign level, no attempt has been made to evaluate slight dif-ferences in the overall reliability of the promising aft fuse-lage designs that might result from possible differencus inthe reliability of the elevator and secondary structure designconcepts described previously.

The expected reliability of advanced aft-fuselage design con-cepts has been compared to the reliability of the existingAH-lG aft fuselage in terms of the mean time between failures.This particular reliability parameter has been chosen becauseit is a primary input to the cost model, which is the principaltool of the methodology for evaluating the cost effectivenessof promising design concepts.

TABLE XVII. REAPPORTIONED 3M FAILURE DATA FORAH-lG/H AIRCRAFT

Number of PercentageFailure Event Category Recorded of Total

Events Events

Wear/Fretting 17 7.6

Corrosion 5 2.2

Broken/Loose/Missing 67 29.8Hardware

Crack/Fracture/Buckle/ 50 22.2Bond Failure

Foreign Object Damage 7 3.1

Puncture/Tear 18 8.0

Battle Damage 12 5.3

Improper Adjustment/ 29 12.9Alignment

No Defect 20 8.9Totals 225 100.0

Total Flight Hours = 23,948MTBF = 106.4 Hours

86

_ - ...... . ..- - .. . " I i i i~ ' i . o o " -....... ...... j...

TABLE XVIII. PREDICTED FAILURE EVENTS FOR FIBERGLASSSKIN AFT-FUSELAGE DESIGN CONCEPTS

Expected %Failure Event Category Reduction Estimated % of

in No. of No. of TotalEvents Event s Events

Wear/Fretting 50 8.5 7.3

Corrosion 30 3.5 3.0

Broken/Loose/Missing 70 20.1 17.2Hardware

Crack/Fracture/Buckle/ 60 20.0 17.1Bond Failure

Foreign Object Damage 50 3.5 3.0

Puncture/Tear 50 9.0 7.7

Battle Damage -0- 12.0 10.2

Improper Adjustment/ 30 20.3 17.4Alignment

No Defect -0- 20.0 17.1

Totals 116.9 100.0


87

TABLE xix. PREDICTED FAILURE EVENTS FOR ADVANCEDCOMPOSITE SKIN AFT-FUSELAGE DESIGN CONCEPTS

Expected %Failure Event Category Reduction Estimated % of

in No. of No. of TotalEvents Events Events

Wear/Fretting 50 8.5 6.2

Corrosion 30 3.5 2.5

Broken/Loose/Missing 70 20.1 14.5Hardware

Crack/Fracture/Buckle/ 30 35.0 25.3Bond Failure

Foreign Object Damage 25 5.3 3.8

Puncture/Tear 25 13.5 9.8

Battle Damage -0- 12.0 8.7

Improper Adjustment/ 30 20.3 14.7Alignment

No Defect -0- 20.0 14.5

Totals 138.2 100.0


88

COST

METHODOLOGY

The objective of the cost analysis is to determine and comparethe total cost per aircraft life cycle for the existing struc-ture and four new concepts for the aft fuselage. The life-cyclecost per aircraft is defined herein as the total cost for eachaircraft in a fleet maintained at a constant size for a spe-cified time. For this study, a fleet size of 1000 and a lifetime of 10 years are used.

The principal tool for cost analysis is the cost model shownschematically in Figure 41. Input data for the model can-siders the structural design concept itself and a scenario oflife-cycle events which is applied to each design concept.The scenario has been developed by Kaman from pa-. field ex-perience of the AII-IG/J helicoper as reported ir marine Corps3M data covering the two-year period from June 1969 throuhJuly 1971 involving a total of 23,948 flight hours. Thi:baseline data defining structure damage causes, types anddistribution is further developed and refined by reliabilityand maintainability analysis and then applied to each designconcept. Both external damage causes, such as normal incidents,hangar rash and ballistic impacts, and inherent damage events,such as fatigue, bond failures and loose/broken fittings, areconsidered. With the composite tail assembly repair/scrapcriteria established, cost elements either supplied by theGovernment or determined by Kaman are applied to compute thetotal cost per aircraft for the tail assembly during the lifecycle of the AH-IG helicopter.

Tail assembly life-cycle costs are generated in the followingbasic categories:

1. Acquisition2. Operation3. Attrition

Acquisition costs are incurred in producing the particulartail assembly design and providing pipeline spares and mate-rials. They include, but are not limited to, the followingelements:

1. Engineering design, research, development, test, andevaluation costs

2. Unit cost of producing the tail assembly as a functionof quantity

3. Nonrecurring and rate tooling costs4. Profit

89

Sceari ofctLifDesign Concepts Cycle Events

Reliability&

MaintainabilityAnalyses

I------------ ----------- --- -Damage/Failure Rates, Repair/Scrap Criteria,jRepair/Scrap Fraction vs. Maintenance Level,Time and Material Repair Requirements I

I ACLife Cycle Unit- Cost of New F4-- Tail Section

GSE Cost, Repair Mater _ial Cost & Shipping | alSipn ot

Maintenance Labor Cost Attrition Rate

70[Value of Weight Saved

I INITIAL COSTS OPERATING COSTS ATTRITION COSTS

Tail Unit Cost Damage Repair CostPipeline Spares Damage Replacement Replacement Cost

Weight/Cost Saved Cost

Total Costper AircraftlLife Cycle

Figure 41. Cost Model.

90

Operating costs are determined from the results of the reli-ability and maintainability analysis for the scenario of life-cycle events applied to the tail assembly design concept. Thefollowing cost elements and maintenance parameters are con-sidered in this category:

1. Mean time between damage (MTBD) and mean time to re-pair (MTTR) criteria

2. Military and civilian labor including inspection,disposition, requisition and repair requirements

3. Packaging and shipping costs4. Repair or replacement material costs, other GSE and

support equipment osts5. Repair level of maintenance (organizational, inter-

mediate, depot)

Attrition costs consist of replacement costs for damage due toothar than tail assembly damage-related causes and are in-cluded in the model to provide a realistic appraisal of totalunit quantities required during the aircraft life cycle.

The cost model generates a summation of the tail assembly costper aircraft life cycle in the three basic categories dis-cussed above. This cost per aircraft multiplied by the desiredfleet size yields the total program tail assembly life-cyclecost.

COST MODEL EQUATIONS

Number of Damaged Tail Sections per Aircraft Life Cycle

ND-LMTBD

where N = number of damaged tail sections perD aircraft life cycle

L = aircraft life cycle, flight hours

MTBD = tail section mean time 1---ween damage

Initial Costs

The initial costs consist of tail section unit cost, pipelinespares costs, and weight cost.

a. Unit Cost

CO= acquisition cost for one tail section

91

b. Weight Cost

Cost comparisons of different designs should be based uponequal effectiveness in terms of mission performance. Thetail sections in this study have essentially equal missionperformance with only one significant exception: weight.In order to account for this difference, a dollar value isdetermined for the weight difference and is applied toeach design as an adjustment in the initial cost. Theexisting metal tail section is the baseline; its weightcost is zero. Typically the new concepts are lighterthan the existing structure, and thus the weight costadjustment is a lowering of the cost.

CW= DwW

where C= cost adjustment for weight

DW= dollar value per pound of structuralWweight

WS= weight of existing tail section minusweight of new concept

c. Pipeline Spares

The initial spares requirement assumes that all tail sec-tions scrapped and all tail sections repaired at the depotlevel, as well as all repair meterials, over a 6-monthperiod must be procured to establish the pipeline.

CS= ND (0.5/Y)((KTs+KRD)(Co+Cc+CAs)+(KRo+KRI)

(CM+CMC.MS))+CER(O.5/Y)

where CS= initial spares cost per aircraft life cycle

KTS= total fraction of damaged sections scrapped

KRD = fraction of damaged sections repaired atdepot

KRO= fraction of damaged sections repaired atorganizational level

KRI= fraction of damaged sections repaired atintermediate level

CM= average cost of meterial per repir

92

CMS= shipping cost of material per repair as afraction of material cost

CEP,= cost of special ground support equipment

required

CC= cost of shipping container

CAS= cost of air shipping tail section fromCONUS

Y= aircraft life cycle, years

d. Total Initial Cost

CTO= Co+Cs+Cw

where CTO= total initial costs per aircraft lifecycle

Operating Costs

The operating costs consist of tail section repair and scrapcosts summed for organizational/intermediate and depot levelmaintenance.

a. Organizational/Intermcdiate Level Repair CostsCRO= ND (CMH (Ml+TRo) KRI+CMH (Ml+TRo) KRO

(KRO+KRI) (CM+CMCMS))+CER

where CRO= organizational/intermediate level repair costs

CMH= organizational/intermediate level laborrate

M = MMH to inspect and determine damage

T = mean field repair time, including time toRO remove and replace 1.3% of damaged tailsections

b. Organizational/Intermediate Level Scrap Costs

Cso= NDKSO(CMHM3+CO+CAS+Csc)

where CSO= organizational/intermediate level scrapcosts

93

KSO= fraction of damaged sections scrapped atorganizational/intermediate level

M3= MMH to inspect and determine damage,remove and replace tail section, requisi-tion and obtain replacement

CSC= container surface shipping cost

c. Depot Level Repair Costs

CRD= NDKRD(CMHM3+Css+ 2Cps+CAs+CMM 4+DRD)

where CRD= depot level repair costs

DRD= average depot repair cost per tail section

CSS= tail section surface shipping cost toCONUS

Cps= shipping preparation cost

CMD= depot level labor rate

M4= MMH to receive and inspect at depot

d. Depot Level Scrap Costs

CSD= NDKSD-(CMHM3+Cps+Css+CMDM5+Co+CAs)

where CSD= depot level scrap costs

M5= MMH at depot level to receive, inspect,and dispose of scrap

K = fraction of damaged tail sections scrappedat depot level

e. Total Operating Costs

CTM= CRO+CSO+CRD+CSD

where CTM= total operating costs per aircraft lifecycle

Attrition Costs

The attrition costs consist of the cost of new tail sectionsand the associated- transportation rchages to -pic tail

94

sections lost to causes other than damage to the tail section.

CA NA(CO+CAS)

where CA= attrition costs per aircraft life cycle

NA= number of tail sections lost to attrition

Total Tail Section Ccss

The total .,ail section costs are the sum of the initial, op-erating and attrition costs.

CLC= CTO+CTM+CA

where CLC= total tail section costs per aircraft lifecycle

Lists and Diagrams

Figures 42 and 43 list and diagram the cost elements.

I

95

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96

TAIL DAMAGEEVENTS PER

A/C LIFE CYCLE

1P N COST OFTAIL SECTION VALUE OF I SPARES INSPECTINGPROCUREMENT WEIGHT SAVED

-PROCUREMENT DMG

TAIL PRODUC- VALUE PER 1 TAIL DAMAGE REPAIR NO COST OF RE-

TION COST POUND OF I EVENTS PER TAIL ON MOVING DAM-

PER UNIT WEIGHT ] A/C LIFE CYCLE A/C AGED TAIL

Y IlI YESNRE AND WEIGHT SAVED REPAIR NO COST OF SPARE COST OF TAIL 'AIL

RDTE COSTS ICOST REDUCTION TAIL SECTION REPAIR AT BEYOND

PER UNIT FIE__ ______ [ORIG/INT REPAIR

XEE YES

PROFIT COST OF REPAIR COST OF COST OF GSE COST OF NOIPER UHIT MATERIAL CONTAINER FOR REPAIR TAIL SECTI0-

COST OF GSE I COST OF AIR ISTLLINFOR REPAIR SHIPPING INSTALLING

SPARE TAIL

INITIAL COSTS 7 PRTOA OT T0

Figure 43. Cost Diagnosis.

97

TAIL DAMAGE

EVENTS PER

A/C LIFE CYCLE

PIPELINE COST OFSPARES INSPECTINGSUREMENT DAMAGE

ELPA COST OF RE- COST OF RE-DAMAE NOCEIVING ANINTS PERTAIL MOVING DAM- INSPECTINGLIFE CYCLE AGED TAIL ISETN

LIFE CYCLE AT DEPOT

COST OF SPARE COST OF TAIL TNO INSTALL SPARE TAILTAIL IN REPAIR AT BEYOND & SHIP DAMAGED' I

FIE ORIG/INT TAIL TO DEPOT I

I"YEIE YES

T OF REPAIR COST OF COST OF GSE COST OF NEW COST OF AIRH SHIPPINGCOTFMATERIAL CONTAINER FOR REPAIR TAIL SECTION DEPOT REPAI 2

COST OF COST OF SHIP- COST OF PREEST OF GSE COST OF AIR

INSTALLING PING CONTAIN- ING REPAIREEREPAIR SHIPPING S IdTI O H

SPARE TAIL ER TO DEPOT TAIL FOR SHIMENT TO FIEI

S JK OPERATIONAL COSTS AT ORG/INT LEVEL OPRIONA

Cost Diagnosis.

D7

I

COST OFINSTALLINGSPAR TAILIN FIELD

COST OF PRE-PARING DAM-

AGED TAIL FORSHIPMT TO DEPOT

COST .OF RE- COST OFCEIVING AND SHIPPING

INSPECTING _ _AT DEPOT T DEPO

SINSTIL SPARE TAIL NO COST OF -ATTRITED TAIL& SHIP DAMAGED REPAIRABLE SCRAP SECTIONS PERTAIL O DEPOT I DISPOSAL A LIFE CYCLE

YES

COST OF AIR FSHIPPING COST OF COST OF NE COST OF NEWSIING DEPOT REPAIR TAIL SECTION TAIL SECTIONTO FIELD _ _]

COST OF sHIP- COST OF PREPAR COST OF AIR I COST OF AIRPING CONTAIN- ING REPAIRED SHIPPING

TAIL FOR SHIP- IIER TDPO T ENTO D FIELD FIELD TO FIELD

J.EVEL 7OPERATIONAL COSTS AT DEPOT LEVEL ATRTINL CO0STSI

!I

r-j

I

PRODUCTION COSTS

The cost estimating department at Kaman Aerospace Corporationprepared estimates for the production costs for each of thenew design concepts. Table XX summarizes these estimates andalso lists the weight for each concept. The following notesdescribe the basis for the estimate.

1. All costs are the fully burdened cost to the customer.The costs include all overheads and a 12% profit.

2. The total production run is 2000 aft-fuselage assemblies.This quantity is the approximate number required to main-tain a constant fleet size of 1000 for a 10-year periodusing the historical attrition rate of 10% per year.

3. Constant 1974 dollars are used.

4. Tooling costs include both materials and labor.Tooling costs are distributed equally to each unit.

5. Material costs for composite materials and core materialsinclude a 15% allowance for losses from cutting & trimming.

6. Material costs for other materials include a 1.9% allow-ance for losses.

7. Production labor costs are computed using $17/hour.Average hours for the total production run were used withan 87% learning curve.

8. Prices used for the most significant raw materials arelisted below. All composite materials listed are pre-impregnated with a low-temperature epoxy resin.

Material Price, dollars/sq yd

Fabric 913 Style, S-Glass 7.25Fabric 120 Style, E-Glass 3.75Fabric 181 Style, E-Glass 5.50Fabric 181 Style, PRD 49 21.00Unidirectional E Glass, T=.009 in. 2.40Unidirectional S-1014 Glass, T=.010 in. 8.10Unidirectional Graphite SP-288-T2 37.00PVC core R-400, T=-.25 in. 4.95

These prices were received from the following suppliers:

B. F. Goodrich CorporationMinnesota Mining and Manufacturing CorporationUnitLe6States Polymeric corporation

99

TABLE XX. SUMMARY Of" PRODUCTION COSTS AND WEIGiT

Average Cost (dollars)Weight

Concepc Tool.ng Material Labor Total (ib)

Elevators 1 38 186 368 592 26.4(set of 2) 2 57 198 351 606 26.0

3 76 187 302 565 25.24 76 179 266 521 23.0

Covers 1 it 83 281 388 5.7(set of 4 2 32 65 387 484 6.4over boom) 3 10 23 28 61 8.2

4 10 23 28 61 8.2

Boom-Fin 1 122 5200 3230 8552 141.12 122 1939 3335 5396 169.33 153 1843 3296 5292 160.04 169 1899 3173 5241 165.6

Aft Fuse!age 1 208 5402 3524 9134 172.3Asse !ly 2 208 2141 3626 5975 200.5

3 229 2022 3562 5813 183.04 245 2078 3439 5762 188.6

Notes:

1. Average costs per unit are calculated assuming an 871learning curve and a total purchase of 2000 aftfuselage assemblies.

2. The boom-fin includes the primary structure for eachconcept and the leading edge cover on the fin.For concepts 1 and 2, it does not include secondarycovers for the drive shaft over the boom. For con-cepts 3 and 4, no such covers are required.

. Concept 1 requires ai. unknown additional cost forlightning protection The weight for concept 1 in-cludes a 10-lb allowance for lightning protection.

4. The aft fuselage assembly includes:

numbered boom-fin+ elevator concept 4+ covers concept 3, for boom-fin concepts 1 and 2

100

LIFE-CYCLE COSTS

Tables XXI and XXII present the final results from the analyrsis of life-cycle costs using the cost model. Table XXIshows the cost in dollars for each unit of the fleet, andTable XXII shows the cost relative to the cost of the exist-ing aft fuselage, concept 0, taken as 100. In order toassess the sensitivity of the costs to changes in value ofinput parameters, several variations are shown. These

r include:

best estimate

+15% changes in purchase price

+50% changes in mean time between damage

+50% changes in cost of a typical repair

+50% changes in the scrap rate

+100% change in the value of a pound saved

In all cases, these changes wiare found to have a relativelyminor effect upon the costs and the relative rankings of theconcepts. The most significant savings (30 to 40%) arefound in concepts 2, 3, and 4, which use glass reinforcement.

Table XXIII summarizes the input data used for the best-es-timate cost analysis. The origins of these data have alreadybeen presented in the sections describing the cost model andthe reliability analysis. Table XIV shows in more detailthe distribution of costs corresponding to the best estimateof cost for each concept.

101

-ON

TABLE XXI. TOTAL COST (DOLLARS) FOR EACH UNIT OF THE FLEET

Concept

Circumstances 0 1* 2 3 4

Best Estimate 21251 20922 14644 13684 13768

.85 x Unit purchase price 18517 17933 12714 11804 11907

1.15 x Unit purchase price 23986 23911 16674 15558 15629

.5 x MTBD 25775 23913 16689 15703 15780

1.5 x MTBD 19744 19928 13961 13008 13097

.5 x Cost of typ. repair 20390 20394 14197 13235 13321

1.5 x Cost of typ. repair 22112 21451 15091 14130 14215

.5 x Scrap rate 19903 19986 14095 13146 13235

1.5 x Scrap rate 22599 21859 15193 14219 14301

0 x Value of lb saved 21251 22217 14952 14603 14493


* Concept 1 requires an unknown additional cost for lightning

protection.

102

TABLE XXII. RELATIVE TOTAL COST (SEE NOTE 1)

-'-"Concept

Circumstances 0 1 2 3 4

Best Estimate 100 98 69 64 65

.85 x Unit purchase price 87 84 60 56 56

1.15 x Unit purchase price 113 113 78 73 74

.5 x MTBD 121 113 79 74 74

1.5 x MTBD 93 94 66 61 62

.5 x Cost of typ. repair 96 96 67 62 63

1.5 x Cost of typ. repair 104 101 71 66 67

5. x Scrap rate 94 94 66 62 62

1.5 x Scrap rate 106 103 71 67 67



Note:

1. Best estimate for concept 0 taken as 100.

2. Concept 1 requires an unknown additional cost for

lightning protection.

103

iA

TABLE XVIII. SUMMARY OF COST MODEL INPUT PARAMETERS FORBEST ESTIMATES

Concept

Symbol Units 0 1 2 3 4

CAS dollars 850.0 same, unless noted

CC dollars 200.0

CER dollars 0.0 100.0 100.0 100.0 100.0

CM dollars 25.0

CMH dollars/hour 4.0

CMS fraction 0.10

CO dollars 7939.0 9134.0 5975.0 5813.0 5762.0

CSC dollars 125.0

Cw dollars 0.0 -1295.0 -308.0 -920.5 -725.0

KRI fraction 0.994

KRO fraction 0.0

KSO fraction 0.007

KTS fraction 0.006

L flight hours 5000.0

Ml hours 0.5

M3 hours 59.7

MTBD flight hours 106.4 173.3 204.9 204.9 204.9

NA fraction 1.0

TRO hours 2.0

Y years 10.0

104

TABLE XXIV. DETAILED COSTS IN BEST ESTIMATE FOR EXISTINGSTRUCTURE AND NEW CONCEPTS*

Concept

Symbol Definition 0 1 2 3 4

ND number of dmagedtail sections 47.0 28.9 24.4 24.4 24.4

CS initial spares cost 191 133 90 89 88

CID total initial cost 8130 7972 5757 4981 5125

CRO repair cost, org/intlevel 1752 1175 1010 1010 1010

CSO scrap cost, org/intlevel 2581 1791 1053 1029 1021

CRD repair cost, depotlevel 0 0 0 0 0

CSD scrap cost, depotlevel 0 0 0 0 0

C M total operatingcost 4332 2967 2062 2038 2031

CA attritional cost 8789 9984 6825 6663 6612

CIC total life-cycle cost 21251 20922 14644 13682 13768

* All costs in dollars per aircraft life cycle

105

DOLLAR VALUE FOR A POUND SAVED

One parameter, CW , the adjustment for weight, deserves addi-tional discussion. C is the product of a difference in weightand the dollar value #er pound of structural weight. Thelatter is potentially controversial because in engineeringpractice it can take on several different values during thedesign cycle. First, there is a true economic value which is

believed to be $35 per pound. Rationale for this number ispresented in this section. Second, there is a legal orcontractual value which can vary from near nil to severalhundred dollars, depending on the penalty terms of a contractand the relationship of the aircraft to the contractual spec-ifications.

The economic value can be derived by considering the aircraftto consist of a transporter and a load to be transported.The transporter supports the load and must grow in weight inresponse to an increase in the load if performance (speedand maneuverability) is kept constant. For the AH-lG heli-copter, the transporter consists of the following groups:

Group Weight, lb (Ref. 5)

Rotor 929.1Tail 59.2Total Basic Structure 801.3Landing 108.3Flight Control 382.2Propulsion 1404.4Hydraulic/Pneumatic 91.0

3775.5

The load consists of all other items, such as crew, instruments,armament, fuel, etc. For its basic mission, the gross weightof the AH-IG is 8621.0 lb (Ref. 5). Thus

load = gross - transporterload = 8621.0 - 3775.5 = 4845.5 lb

On the average, 3775.5/4845.5 = .779 lb of transporter is re-quired to support each pound of load.

Now, consider that two hypothetical tail sections exist thatare functionally identical but differ in weight by 1 pound.If the lighter tail section were used, one could reason thatthe transporter could be designed .779 lb lighter. At cur-rent prices for the AH-lG, the transporter costs $60 per

106

pound. The value of a pound of weight saved in this case isthen .779 x 60.00 = $46.70 in initial cost. Fuel costs overthe lifetime of the aircraft would add another $1.40, bringingthe total to $48.10. This estimate, however, is too high fortwo reasons:

1. The slope of the tangent to the curve of transporterweight versus load is less than that of the secant,as shown in Figure 44.

7 - Estimated curve for helicopter class with AH-lG-q speed and maneuverabilityA! 6-

4J 5- Known pointon curve

Tangent

$34J

P2-1) Secant, slope = .779

(0

0 1 2 3 4 5 6 7 8Load, kips

Figure 44. Transporter Weiqht Versus Load.

2. The law of marginal utility. The last increment ofload is less useful than the first.

For these reasons, the true economic value of a pound savedis estimated to be only $35 in the design stage. This valuewas used in the calculation of the best estimates of life-cycle

cost. Life-cycle costs with the value of a pound saved at$0 and $70 are also shown in the sensitivity analysis, Tables

XXI and XXII.

107

I -- o

CONCLUSIONS

.. The new concepts for the aft fuselage provide significantsavings in life-cycle costs over the existing metal aftfuselage. The preferred concepts save 36% in life-cyclecost.

2. The new concepts have a lower unit purchase price. Theunit purchase price of the preferred concepts is 26% less.The new one-phase fabrication methods for the boom,elevators, and covers are very significant factors inachieving the low unit purchase prices.

3. The new concepts provide increased safety, reliability,and maintainability.

4. The new concepts save weight. The preferred concept (No. 3)is 12% lighter.

5. Fuselage concepts 2, 3, and 4 (the glass compositeconcepts) are significantly more cost effective thanconcept 1. The principal reason for their advantage isthe lower cost of the material.

6. The differences in cost between fuselage concepts 2, 3, and4 are not great. The respective life-cycle cost savings forfuselage concepts 1, 2, 3, and 4 are 1%, 31%, 36%, and 35%.

7. Fuselage concept 3 is selected as the best concept becauseit offers the greatest cost saving and significant weightsavings. It also provides excellent access to the interiorof the boom via removable panels over its entire length.

8. Fuselage concept 2 is a very close second to concept 3.Concept 2 provides less, but still good, cost savings; lessweiqht savings; and less access to the interior. However,concept 2 has some compensating virtues. These include:(1) increased structural integrity that accrues from acompletely monolithic boom without joints and an increasein thickness for the outer facings, and (2) easier accessto the drive shaft over the boom.

9. Elevator concept 4 is an easy winner. It enjoys the lowestcost, lightest weight, greatest reliability, and bestsalvageability.

10. Cover concept 3 is also an easy winner. It has very lowcost and is the most rugged.

108

11. In general, the low life-cycle costs reported herein areattributed to the fac* that the design concepts:

a. Reduce the number of parts.

b. Eliminate production operations.

c. Simplify assembly.

d. Improve field maintainability.

109

LITERATURE CITED

1. Arvin, G. H., et al, STRUCTURAL DESIGN GUIDE FOR ADVANCEDCOMPOSITE APPLICATIONS, North American Rockwell Corpora-tion; Advanced Composite Division, Air Force MaterialsLaboratory, Air Force Systems Command, wright-PattersonAir Force Base, Ohio, January 1971.

2. Anon, DS, GS, aND DEPOT MAINTENANCE MANUAL ARMY MODELAH-lG HELICOPTER, TM55-1520-221-35, Headquarters, Depart-ment of the Army, April 1969.

3. Ambrose, E., and Clarke, D., TAILBOOM, VERTICAL FIN ANDHORIZONTAL STABILIZER STRUCTURAL ANALYSIS, Bell HelicopterCompany Report 209-099-056, Fort Worth, Texas, January1967.

4. Montgomery, B. L., and Balke, R. W., RESULTS OF GROUNDVIBRATION SURVEY, VOLUME I OF II, Bell Helicopter CompanyReport 209-099-017, Fort Worth, Texas, February 1969.

5. Noyes, J. K., ACTUAL WEIGHTS AND BALANCE A-IlG, BellHelicopter Company Report 209-099-264, Fort Worth, Texas,June 1971.

6. Ashton, J. E., Halpin, J. C., and Petit, P. H., PRIMER ONCOMPOSITE MATERIALS: ANALYSIS. Technomic PublishingCompany, Stamford, Connecticut, i969.

7. Anon, PLASTICS FOR AEROSPACE VEHICLES, PART I.REINFORCED PLASTICS, MIL-HDBK-17A, Department of Defense,Washington, D. C., January 1971.

8. Anon, PRD-49, DU PONT'S NEW HIGH MODULUS ORGANIC FIBERFOR PLASTICS REINFORCEMENT, BALLISTIC ARMOR, AND TENSIONCABLE APPLICATIONS, PRELIMINARY DATA, E. I. DuPont deNemours and Company, Wilmington, Delaware, September 1971.

9. Anon, "SCOTCHPLY" SP-286 PREPREG GRAPHITE TAPE, MinnesotaMining and Manufacturing Company, Saint Paul, Minnesota,January 1972.

10. Durchlaub, E., Sacco, P., FRACTURE MECHANICS AND DAMAGETOLERANCE OF S GLASS COMPOSITES, AFFDL-TR-72-82, TheBoeing Company, Vertol Division; Air Force FlightDynamics Laboratory, Air Force Systems Command, Wright-Patterson Air Force Base, Ohio, August 1972.

110

I :l -i 1 - 1 i :f -- -I I ' f j - I I -i" . . . ... . . .... ......... .. .

11. Anon, FERRO ZERO TWIST FABRICS, S GLASS, Ferro Corpora-tion, Nashville, Tennessee, undated.

12. Anon, RIGICELL, RIGID VINYL FOAM DATA SHEET,B. F. Goodrich Industrial Products Company, Akron, Ohio,undated.

II

111

I

APPENDIX

FAILURE MODE AND EFFECTS ANALYSES

The failure mode and effects analysis (FMEA) is a basic analyt-ical tool for design evaluation and reliability improvement.This technique involves a systematic assessment of everypossible potential failure mode within a hardware design basedon an empirical knowledge of the design and hirtorical statis-tical data when available. By following this approach, theweakness or limitations of a design can be pinpointed andappropriate engineering attention directed toward improvingcritical reliability areas of the design.

The simplified functional flow block diagram of Figure 45 showsthe basic functions which the aft fuselage must perform as anassembly of the helicopter system. For the purposes of theFMEA's, it is sufficient to summarize the functional requirementsas follows:

Transfer Loads - The aft-fuselage structure must havesufficient strength to receive input loads from the tailrotor, elevator, and tail skid, and to transfer these tothe main fuselage.

Meet Stiffness Requirements - The advance design conceptsmust provide adequate stiffness to assure that naturalfrequencies are not adversely affected.

Provide Access - The advanced concepts must have enoughaccess panels, covers and/or doors to permit drive trainmaintenance, flight control maintenance, structuralrepair and inspection.

Provide Environmental Durability - The advanced conceptsmust be at least comparable to the existing aft fuselagein resistance to erosion, corrosion, foreign object andballistic damage, and hangar rash incidents.

The functional flow block diagram of Figure 46 and preliminarydesign drawings of the boom-fin, elevator and secondary struc-ture advanced design concepts serve as a basis for preparingthe simplified aft-fuselage reliability block diagram and theaft-fuselage component level reliability block diagram presentedin Figures 46 and 47 respectively. The latter diagram isolatesand itemizes those components whose successful performance isessential to proper performance of the overall aft fuselagestructure. The FMEA methodically evaluates each of theseessential components in turn to establish possible failure

112

modes, causes of failure, effects of failure on aft fuselagestructural performance, methods of detection, qualitativeestimates of failure criticality to safety and mission success,and subjective classifications of expected failure probabil-ities of occurrence. To facilitate the analysis procedure,all potential inherent and external modes of failure aresorted into the following basic failure groups:

Inherent - Wear/FrettingCorrosionLoose/Broken HardwareDelaminationBond SeparationCrack/Fracture/Buckling

External - DentPuncture/TearBattle Damage

Failure criticality codes are defined as follows in accordancewith MIL-STD-882:

Category I - Negligible - Any nuisance failure notserious enough to be classified in ahigher category, but requiring correctiveaction during routine preventivemaintenance.

Category II - Marginal - Any failure which is expectedto degrade performance, which can betolerated throughout a mission, but whichshould be corrected immediately uponcompletion of the mission.

Category III- Critical - Mandatory Abort - Any failureexpected to cause personnel injury/hazardor major aircraft system damage, orrequiring immediate corrective action forsurvival. A time-dependent repair actionwhich can cause a catastrophic failure ifuncompensated.

Category IV - Catastrophic - Any failure expected tocause death/serious injury, or loss ofthe aircraft.

113

To provide a qualitative estimate of the probability of failureoccurrence, the following subjective classifications are used:

Class A - Probability of failure is not remote.Class B - Probability of failure is remote.Class C - Subject to rare, random failures.Class D - Not expected to fail in s.cvice.

Tables XXV through XXVII present the detailed results of thefailure mode and effects analyses.

TRANSFER PROVIDE E ROVMELASSTIFFNESS ACESENVIRONMENTALLOADS ACS

REQUIREMENTS DURABILITY

Figure 45. Simplified Functional Block Diagram.

~SECONDARYBOOM-FIN - ELEVATOR - STRUCTURES

Figure 46. Simplified Reliability Block Diagram.

114

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