ADVANCED COMPOSITE ELEVATOR FOR BOEING 727 AIRCRAFT

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._. FIFTH QUARTERLY TECHNICAL PROGRESS REPORT "_ _ _' "_23 MAY 1978 THROUGH 22 AUGUST 1978 _ "t.,_ -,.j zj ,-y,,_.--

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'" NATIONAL AERONAUTICS AND SPACE ADMINISTRATION /_ '_._,. L_J_IGLEY RESEARCH CENTERHAMPTON. VIRGINIA 23666 j_% _*_- -_

IN RESPONSE1"0:

CONTRACT NASl-14962'P DRL LINE ITEM NUMBER 0186m

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A_rplane Company

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This Report is Submitted in Compliance

With DRL Line Item Number 018

- FIFTH QUARTERLY TECHNICAL PROGRESS REPORT

t 23 May through 22 August 1978

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Supervised by :O E Desper, Jr

i Program Administration and Data Manager

I Approved by: " _t_ , , iH. Syder

Engineering Deign Manager

9 ;.Ij/S ..... ,Approved by: <",_" ' _ " f."_,/. Cv-V-#_/_

I ._J.E. McCarty " - /

• Engineering Technology Manage_/V

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Approved by : _,_lll"V.S. Thomp

Operations Technology Manager

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W.D. Grant

Production Manager

Approved by : /

( Director of A )osites Programs

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1983002978-002

Boeing Commercial

Airplane CompanyContract NASI-14952

FOREWORD

This report was prepared by the goeing Commercial Airplar, e Company, .!

Renton, Washington, under Contract NASI-14952. It is the fifth quarterly

technical progress report covering work performed between 23 May 1978

and 22 August 1978. The program is sponsored by the National Aeronautics

and Space Administration, Langley Research Center (NASA-LRC). Dr. H. A. Leyboldp

is the Project Manager for NASA-LRC.

The following Boeing personnel are principal contributors to the program

_ during the reporting period: C. R. Zehnder, Design; R. D. Wilson, iStructural Analysis; M. Garvey, Manufacturing Specialist; D. Grant,

"_" Production Manager; L. D. Pritchett, Technical Operations Coordinator:

and D. B. Chovil, Business Support Manager.

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1983002978-003

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Boeing Connercial

- Airplane CompanyContract NAS1-14952

, SUMMARY

- Detail design activities are reported for a program to develop an advanced

composites elevator for the Boeing 727 commercial transport.

Design activities include discussion and results of the ancillary test

programs, sustaining efforts, weight status, manufacturing producibility

! studies, quality assurance development, and production status.

Engineering Design has completed the release of all scheduled production

drawings, and Operations has completed the required produclhility studies

for the elevator. The program for the fifth quarter is progressing as

scheduled.

_ v

1983002978-004

PRECEDINGPAG_BLANKNOTFILMED

Boeing Commercial

Airplane CompanyContract NAS1-14952

T_BLE OF CONTENTS

SECTION PAGE

FOREWORD ill

SUMMARY v

LIST OF FIGURES AND TABLES viii

1.0 INTRODUCTION i-i

2.0 DESIGN ANALYSIS 2-i

2.1 DESIGN LOADS CRITERIA AND ANALYSIS 2-1

2.1.1 Criteria and Analysis 2-1

2.2 ANCILLARY TESTING - COUPONS AND ELEMENTS 2-1

2.3 DESIGN 2-39

2.4 WEIGHT STATUS 2-40

2.4.1 Actual Weight ProBram 2-42

2.4.2 Static Balance Requirements Document 2-42

3.0 OPERATIONS DEVELOPMENT 3-1

3.1 PRODUCIBILITY STUDIES 3-1

3.2 ANCILLARY TEST COMPONENT FABRICAT__N 3-1

3.2.1 Allowables and Environmental 3-2

3.2.2 Concept Verification 3-2

3.3 QUALITY ASSURANCE DEVELOPMENT 3-2

3.3.1 Preliminary Standards 3-23.3.2 Production Standards 3-6

3.3.3 Fabrication Method 3-13

3.3.4 NDI Techniques 3-173.3.5 Test Results S-17

3.3.6 Conclusions 3-21

3.4 VERIFICATION HARDWARE 3-22

4.0 PRODUCTION 4-1

4.1 ASSEMBLY TOOLS 4-1

4.2 COMPONENT MANUFACTURE 4-3

5.0 REFERENCES 5-1

( vii

1983002978-005

Commercialuoelng _

Airplane CompanyContract NASI-14952 ---

LIST OF FIGURES AND TABLES

FIGURE PAGE

i-i Program Master Schedule 1-3

2-1 727 Advanced Composites Elevator Ancillary 2-2

Test Specimens


Test Specimens• r


Test Specimens

2-4 Cover Panel Padup at Ribs Static Test 2-20

2-5 Inplane Shear Test Front Spar Shear Web 2-20

2-6 Typical Moire Fringe Pattern Prior to Failure 2-25

2-7 Typical Moire Fringe Pattern After Failure 2-25

2-8 Honeycomb Panel Compression Test--Load = 13.3 kN 2-26

I-9 Honeycomb Panel Compression Test-toad = 39.9 kN 2-26

2-10 Honeycomb Panel Shear Test-_oad = 22 kN 2-27

2-11 Honeycomb Panel Shear Test-Load = 80 kN 2-27 "

2-12 Shear Panel Failure 2-28

2-13 Shear Panel Failure 2-28

2-14 Compression Panel Failure (Edge Band) 2-30

2-15 Compression Panel Failure (Normal View) 2-30

2-16 Failure in Panel Edge Band 2-31 ,

2-17 Front Spar/Actuator Fitting Splice Test 2-34 '

Specimen Failure !

2-18 Front Spar/Actuator Fitting Splice Test 2-34Specimen Failure

2-19 Front Spar/Actuator Fitting Splice Test 2-35

Specimen Failure _,

2-20 Panel Bending Test 2-36

2-21 Actuator Support Rib Test 2-36 :

viii i 'i|

1983002978-006

(Boeing Commercial


FIGURE PAGE

2-22 Actuator Support Rib Test 2-37

2-23 Actuator Support R_b Failure 2-37

2-24 Actuator Support Rib Failure 2-38

3-1 Assembly of Test No. 17 Outboard Test Box, 3-3

Showing Assembled Lower Skin, Ribs, and Rear Spar

3-2 Second Sonic Box, Test No. 18 Skin Pa ..i, Showing 3-3Outer Skin Plies in Place

3-3 Second Sonic Box, Test No. 18, Showing Honeycomb 3-4

Rib Ready for Cure%t

3-4 Second Sonic Box, Test No. 18, Showing Skin Panel 3-4in Autoclave

3-5 Second Sonic Box, Test No. 18, Showing Front View 3-5

of Completed Assembly

3-6 Second Sonic Box, Test No. 18, Showing Rear View 3-5

of Completed Assembly

3-7 NDI Reference Standard-Graphite/Epoxy Laminate 3-7

3-8 NDI Reference Standard-Graphlte/Epoxy Honeycomb 3-8

3-9 Laminate Step Standard 3-9

3-10 Front Spar Chord Standard 3-10

3-ii Fkin Panel Standard 3-11

3-12 Taper Edge Honeycomb Standard 3-12

3-13 727 Advanced Composites Elevator Skin Panel Standard 3-14

3-14 727 Advanced Composites Elevator Traillng-Edge 3-14

Assembly Standard

3-15 727 Advanced Composites Elevator Rear Spar Standard 3-15

3-16 727 Advanced Compositeq Elevator Front Spar Standard 3-15

3-17 727 Advanced Composites Elevator Rib Standard 3-16

3-18 Automated TTU Scanner 3-16

3-19 Graphlte/Epoxy Test Panel Standard, 3-18

Computerized C-Scan Recording

3-20 Semlportable TTU Scanner Without C-Scan Capability 3-19

3-21 Semiportable TTU Scanner, Showing Part Configuration 3-19Not Suitable for TTU Scan

iti ix

] 983002978-007

Boeing Commercial


FIGURE PAGE

3-22 Portable TTU Scanner, Showing Inspection of Flange 3-20Area by Hand-Held TTU Unit

3-23 Sondicator Inspection of Configuration not Suitable 3-2Gfor TTU Scan

3-24 Fokker Bond Testing of Configuration not Suitable 3-21for TTU Scan

3-25 Manufacturing Verification Hardware, Test No. 16 3-23

Skin Panel, Showing Doubler Being Laid Up

- 3-26 Manufacturing Verification Hardware, Test No. 16, 3-24

Showing Completed Skin Panel

3-27 Manufacturing Verification Hardware, Test No. 16, 3-24

Showing Skin Panel in TTU Inspection


Showing Front Spar Being Laid Up


Showing Completed Rear Spar Header


Showing Front Spar, Two Ribs, and Lower Skin BeingAssembled


Showing Aluminum Nose Ribs Being Attached to

Front Spar


Showing Front View of Completed Assembly


Showing Rear View of Completed Assembly

4-1 Front Spar with Nose Ribs Installed and Located 4-iin Tool

4-2 Final Assembly Jig for Left-Hand Rear Spar 4-2

4-3 Major Assembly Jig Being Set Up in Shop 4-2

4-4 Rear Spar with -5 Rear Spar Header 4-4

Installed, Showing Mismatch

4-5 Rear Spar Header, Showing Bow in Side 4-4

4-6 Detail Parts of 65C17529-950 Rib Assembly 4-6

4-7 Completed 65C17529-950 Rib Assembly 4-6

1983002978-008

_ Boeing CommercialAirplane CompanyContract NASI-14952

FIGURE PAGE

4-8 Hand Drill (right), Showing Standard 4-8Vacuum Unit and Modified Vacuum Unit (left)

4-9 Drilling Attach Angles oD Front Spar, Showing 4-8Interference Between Angle and Vacuum Collar

4-10 Nose Rib and Attach Angle, Showing Fiber Breakout 4-9of Fiberglass First Ply

TABLES

; _ 2-1 Test Result_ 2-5

2-2 Test Results 2-6










' 2-12 Test Results 2-17



2-15 Spar Web Shear T_st--Summary 2-21

2-16 Spar Web Shear Test (Test Panel No. I_ 2-22i

2-17 Spar Web Shear Test (Test Panel No. 2) 2-23

2-18 Spar Web Shear Test (Test Panel No. 3) 2-23

! 2-19 Honeycomb Panel Shear Test--Load- 80.1 kN 2-32(Back-to-Back Gage Average)

2-20 Honeycomb Panel Compression Test--Load = 41.6 kN 2-33

# (Back-to-Back Gage Average)

2-21 Predicted Weight Status 2-41

I

!

xi

1983002978-009

,- Boeing Commerciall Airplane Company

Contract NASI-14952

SECTION 1.0

INTRODUCTION

The escalation of Jet-fuel prices is causing a reassessment of technology

concepts and trades used in designing and building commercial airplanes.

The task is to incorporate fuel-savlng concepts into commercial aircraft

design.

f

The potential weight savings and fuel reduction resulting from the

extensive use of advanced composites in aircraft structure are significant.

However, the lack of technical confidence and cost data has delayed

their use in commercial aircraft.

Hardware programs considered in a production environment are required to

establish and demonstrate the safety, operating-life characteristics,

and manufacturing cost of advanced composite structures.

Boelng's approach to the problem is to obtain reliable production,

technical, and cost data bases by the integration of advanced composite

technology development under NASA contracts, which, when combined with

company effort, will accelerate the application of composites. This

approach addresses these data bases, developing realistic production

costs in a commercial transport manufacturing environment. Program

emphases are directed toward developing the information needed to obtain

an early production commitment decision by management, and will be

conducted in an environment consistent with production standards.

Preliminary development efforts, as covered in the first and second

quarterly reports, were devoted to conceiving, developing, and analyzing

alternate design concepts, and the preparation of a technical plan to

aid in selecting and evaluating material, identifying ancillary struc-

tural development test requirements, and defining full-scale ground-test

and fllght-test requirements necessary to ob_aln PAA certification.

IL ,

1983002978-010

Boeing CommercialAirplane CompanyContract NASI-14952

The program was built on precontract design activities as well as con-

tracted design activities that consider:

• Program management and plans development

• Establishing design criteria

• Conceptual and preliminary design

• Manufacturing process development

• Material evaluation and selection

• Verification test

• Detail design

• FAAapproval plan definition

This report describes work accomplished during the fifth 3-month period

of the contract. Design activities include the discussion of design

status, weight status, results of manufacturing producibillty studies,

ancillary test specimen configurations and results, and production

efforts. These activities are described under the headings: Design

Analysis, Ancillary Testing, Operations Development, and Production.

The overall program schedule status is stmmarized in Figure 1-1.

L I-2

1983002978-011

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1983002978-012

Boeing Commeccial


SECTION 2.0

DESIGN ANALYSES

2.1 DESIGN LOADS CRITERIA AND ANALYSIS

2.1.1 Criteria .and Anal>sis

As part of a regular meeting schedule with the FAA, a meeting was held in May.

A presentation concerning lightning strike threats, and required protection

systems for the elevator, was given by the Boeing El( odynamics staff.

Further discussions of test plans, and plans to _ce_ent test results in

the form of structural design values to the FAA, were included.

A presentation concerning moisture absorption in graphite/epoxy components

was given to the FAA at a scheduled June meeting. The presentation

discussed the mechanics of moisture absorption, and what can be expected

on actual in-service components.

Work was begun in June on preparing the elevator formal stress analysis

documentation.

2.2 ANCILLARY TESTINC,-COUPCNS AND ELEMENTS

Test No. 1-Coupons--Speclmen configurations and test cesults are shown in

Figures 2-1, 2-2, and 2-3, and Tables 2-1 through 2-10.

Test No. 5-Long-Term Environmental Assessment--SpeciNen configurations

and test results are shown in Figure 2-1, and Tables 2-7 and 2-8

(65C17706-7, -8, and -9).

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1983002978-026

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Boeing CommercialAirplane CompanyContract NAS1-14952

i-Test No. 4--Elements (Fastened Jolnts)--Specimen configurations and test

-- results are shown in Figure 2-1 and Tables 2-2, 2-3, 2-4, 2-11, 2-12,2-13, and 2-14. Repeated load test specimens are cycled to 500,000

- cycles at R= -I.0 at i0 Hz. The fully reversed test load used is approx-

Imately 25% of the static ultimate strength of the laminate. The 500,000

cycles are representative of service requirements for the 727 elevator.

!%

Test No. 8--Panel/Rib Pad--Figure 2-4 shows the test setup fox thi_ test.

Elevator skin vanel support at rib intersections was simulated for this

test. Specimens were tension tested to failure, using a hydraulic grip

machine setup. Test results are presented in Table 2-7.

Test No. 9-Spar/Shear Web-Figure 2-5 shows the setup for the first

spar/shear web picture frame test, conducted in Aprll 1978. Back-to-

back strain gage rosettes were installed on the six-ply (_ 45° orien-

| ration fabric) web between stiffeners.

A maximum principal shear strain of 0.0050, and a maximum normal strain

of 0.0041_ was obtained at the strain gage locations in the buckled web

prior to panel failure. The panel web failure was at a web-stlffener

attach fastener hole.

The shear flow in the web at tim, of failure was 164 N/m, which is

approximatel7 3-1/2 times the critical design shear load in the _ront

spar web.

A summary of the panel failure loads, stresses, and strains Is presented

in Table 2-15. Detailed strain gage result? for each panel are shown in

Tables 2-16, 2-17, and 2-18. As can be seen, the failure load values

i for three ranels are very close to one another.

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1983002978-032

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Table 2-15. Spar Web Shear Test-Summary _"

Maximum Maximum Max!mumTest Maximum shear normal shearpanelno. load, P stress strain strain

1 147 kN -143.4 0.0041 0.0050

2 154 kN -152.0 0.0044 0.0053

3 150 kN -144.2 0.0043 0.0050

Average -146.5 MN

m2

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1983002978-033

Boeing Comercial ORIGINALPAGE ISAirplane Company OF POOR QUALITYContract NAS1-14952

Table 2-16. Spar Web Shear Test (Test Pane/No. 1)

Principal strainsLoad

increment Load, kN Ox* Oy Txy E1 e2 712

1 17.8 2.3 2.7 -22.5 0.00042 -0.00035 0.00077

2 35.6 14.9 15.3 -40.7 0.00093 -0.00046 0.00139

3 53.4 28.1 26.3 -58.3 0.00142 -0.00058 0.00200

4 71.2 44.0 40.4 -75.5 0.00195 -0.00064 0,00259

5 80.1 52.1 47.8 -64.0 0.00222 -0.00067 0.00289

6 89.0 60.2 55.2 -92.7 0.00249 -0.00070 0.00319

7 97.9 68.3 62.1 -101.3 0.00275 -0.00074 0.00349

8 106.8 76.1 69.4 -109.8 0.00302 -0.00077 0.00379

9 115.7 83.8 76.3 -118.4 0.00328 -0.00081 0.00409

10 124.6 91.4 83.0 -126.2 0.00353 -0.00083 0.00436

11 133.5 1(30.1 90.8 -135.4 0.00382 -0.00086 0.00468

12 142.4 108.1 96.7 -143.4 0.00407 -0.00091 0.00498

*Stress in MN/m 2

Ex = 20 685 MN/m 2

Ey = 20 685 MN/m 2

Gxy = 0.293 04 MN/m 2Vxy = 0.68

PA

Failure location (t_

_--_- Stiffener

2-22

1983002978-034

(-" Boeing Commercial

Airplane Company ORtG|NAL p_G_ |SContract NASI-14952 OF pOOR QOALI'rf

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Table 2-17. Spar Web Shear Test (Test Panel No. 2)

/" _ Load Principalstratusincrement Load, kN Ox* Oy rxy e1 e2 1'12

j 1 27.5 11,7 10.2 -29.8 0.00068 -0.00034 0,001032 54.6 33.4 28.6 -55.3 0.00144 -0.00048 0,00193

3 84.6 57,4 49.6 -82.7 0.00227 -0.00062 0.00289

( 4 111.3 81.4 71.2 -109.7 0.00310 -0.00074 0.003835 133.5 100.8 88.8 - 131.5 0.00376 -0.00083 0.00459

6 154.4 119,8 106.0 - 152.0 0.00440 -0.00091 0.00531

*Stress in MN/m 2

Tab/e 2-18. Spar Web Shear Test (Test Panel No, 3)

t PrincipalstrainsLoad

increment Lo_d, kN ox • Oy _'xy el e2 3'12

i 1 27.5 10.9 11.2 -34.4 0.00076 -0.00042 0.00118

2 54.0 34.4 33.4 -59.2 0.00154 -0.00049 0.00202

3 83.4 59.5 55.9 -83.5 0.00233 -0.00054 0.00286

! 4 111,3 85.6 78.8 -108.8 0.00315 -0.00060 0.00375

5 ' 134.2 106.9 96.7 -129.5 0.00382 -0.00067 0.00449

6 150.3 122.1 110.2 -144.2 0.00430 -0.00071 0.00501

• Stressin MN/m 2

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1983002978-035

Boeing Commercial


The average test failure shear stress was 146.5 MN/m 2, as compared with

a maximum expected flight stress in the elevator front spar of 38 MN,'m2.

Typical moire fringe patterns, Just prior to failure and Just after

failure, are shown in Figures 2-6 and 2-7.

Test No. lO-Honeycomb Panel Stability-Shear and compression panels have

been tested. The panels measure approximately 56-cm square. The shear

panel was tested in a picture-frame fixture, as shown in Figures 2-10

and 2-11. The compression panel had the top and bottom edges potted to

prevent brooming at the load application point, and the vertical edges

were simply supported. The test setup is shown in Figures 2-8 and 2-9.

The shear panel was expected to fail in shear prior to elastic buckling,

and test results verified the failure mode. The first shear panel

failed parallel to a panel edge, in the transition area from the honey-

comb core edge to the laminate edge bond at a load of i00 kN/m, and the

failure ran out of a corner radius of the panel as shown in Figure

2-19. The failure load is approximately 2-1/2 times the airplane panel

design shear load. The next test specimen was modified to eliminate the

corner radius. The failure is shown in Figures 2-12 and 2-13. The

failure load was approximately the same as the original configuration

load, but the failure mode was different, as shown. Further testing

will be accomplished with the modified configuration.

Moire fringe evaluations were done on both the shear and compression

tests, as shown in Figures 2-8 through 2-11.

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2-24

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1983002978-036

_FBoeing Commercial

AirplaneCompany ORIGINAL PAGE IS

F Contract NASI-14952 OF POOR OUAT,ITY

F

•r: %: Figure 2-6. Typical Moire Fringe Pattern Prior to Failure

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%t |Figure 2-7. Typical Moire Fringe Pattern After Failure ,,

2-25

1983002978-037

Boeing Commercial ORIGINAL PAGE ISAirplane Company OF POOR QUALrrYContract NASI-L4952

r

Figure 2-8. Honeycomb Panel Compression Test-Load = 13.3 kN

- .,:..-_",_,_ _.

•Figure 2-9. Honeycomb Pane/Compression Test-Load = 39.9 kN

2-26

1983002978-038

Boeing Commercial

Airplane Company ORIGINALPAGE IS

. Contract NASI-14952 OF POOR QUAI,[I_

r i ._/4 I , i _ g 0, Ij I j #, t 4/ q

/ " I o

r '-

F ., ,* %_. _#

r % ,,

%, , //

r J t ,w'

Figure 2-10. Honeycomb Pane/Shear Test-Load = 22 kN

r,

"3

j •

O

o"

i

Figure 2 13. Shear Pane/Failure

tF

1983002978-0

#

- Boeing CommercialAirplane CompanyContract NA_I-14952

!°Typical compression panel failures occurred in the edge band, as shown

in Figures 2-14, 2-15, and 2-16. The load at failure was approximately

700 N/cm (400 ib/in), or approximately twice the airplene panel design

" end load.

- Strain gage results for the first shear and compression panel tests are

. shown in Tables 2-19 and 2-20.

The compression panel will be sectioned through the failure location,

and examined in detail for failure mode.

" Test No. ll-_par/Aluainua Splice-_he first of three front spar/actuator

fitting splice test spec/mena was statically tested to failure at room

• temperature and dry (not preconditioned h_ moisture). The failure

occurred in the tension flange at the end fastener common to the aluminum

actuator fitting, as shown in Figures 2-17, 2-18, and 2-19.

The test s_ _claen failed at approximately 80% of the predicted failure

load. Post teat analysis, using strain sage data Just inboard of the

failure, indicated that local splice eccentricity effects were greater

• than anticipated for the test speclaen configuratlon. Although local

lateral straps were used to stabilise the spar chords, the continuous

skin support that will exist on the airplane elevator component was not

duplicated in this test. The eccentricity effects will be greatly

reduced on the full-scale hardware.w

At the point of failure, _ne test load was equivalent to approximately

" 1.9 times the critical design ultimate load for a flight condition. In

the test specimen spar constant section, the load at failure was equiva-

}. lent to approximately 1.4 times the critical design ultimate load for a

flight condition.

i 2-29

1983002978-041

Boeing Commercial ORIGINALPAGEAirplaneCompany BLACKAND WHITE PHOTOGRAPH'Contract NASI-_4952

./="- Edge band __

,_ *: ._,__,,t failure !_

_| ,.:"

Figure 2-14. Compression Panel Failure (Edge Band)

edge" :" "band fadure '_;'_-'_"

• appearance _ • ".'

"1

Figure 2-15. Compression Panel Failure (Normal View)

2-30

1983002978-042

_" Boeing Commercial

Airplane Company

Contract NASI-14952 ORIGINAL PAGE IS

_ OF POOR QUALITYP

' t- eliminated ;k

t

Failure / _-- Failure (original configuration)(modified corn

P i H°neyc°mb shearpanel !

PP

, ,.,,are/j _1"_-"i'ur°,.n.,appearance---I _ A-A °_'°o_r_F_e,,oo,,r,p;__3 _ ,. ,,,,

L ,111

Finger.tight j,/="a

l fasteners-- 8-BP

i ! Honeycomb compressionpanel I

Figure 2-16. Failure in Panel Edge Band

7

2-31

1983002978-043

Boeing Commercial ORi IAL PAGE ISAirplane Company OF POOR QUALF_Contract NAS1-14952

Table 2-19. Honeycomb Pane/Shear Test, 80. I kN Load (Back-to-Back Gage Average)

Principal strains

Gage ox * Oy Txy e1 _2 3'12

1 - 13.6 -28.0 199.9 0.00383 -0.00463 0.00846

2 30.2 20.3 192,8 0.00440 -0.00370 0.00810

3 74.3 53,1 161.8 0.00431 -O.00251 0.00681

4 9.3 5.7 157.1 0.00340 -0.00320 0.00660

_ 5 J 53.5 30.6 164.2 0.00400 -0.00290 0.00690I

• Stressin kN/m 2

Ex = 64.5 103 kN/m 2

Ey = 26.3 103 kN/m 2 _°_"

Gxy = 28.3 103 kN/m 2 2Vxy = 0.68

Tface = 0.267 mm

Tcore = 14 mm 5

F

+ \//'xy

2-32

1983002978-044

rBoeing Commercial

F Airplane CompanyContract NASI-14952

ITable 2-20. Honeycomb Pane/Compression Test, 4 1.6 kN Load (Back-to-Back Gage Average)

Principal strains

Gage a x * ay rxy e1 e2 _12

1 0.0 -65.5 0.0 0.00000 -0.00250 0.00000

2 0.0 -68.1 0.0 0.00000 -0.00260 0.00000

"_" 3 0.0 -70.7 0.0 0.00000 -0.00270 0.00000

4 0.0 -60.3 0.0 0.00000 -0.00230 0.00000

5 1.2 -70.3 0.0 0.00000 -0.00270 0.00360

"" _ 6 0.0 -70.7 0.0 0.00000 -0.00270 0.00000

"Stressin MN/m 2

t

IINy

Ex = 54.5 103 kN/m 2

* Ey = 26.2 103kN/m 2 J Jl J2 13_ , Gxy = 23.8 103 kN/m 2 x

Vxy = 0.68 J

e ,, Tface = 0.267 mm 56.6(typical)cmI 4 j/5 j 6 Ji Tcore=14mm = w

!

Y

I

p

_ I/I//_'/II/I/,'/IIIIIIIIII/IIIIII///IIIIIIIIIIIIIIIIII_U/1/////,

_"_ 2-33

1983002978-045

Boeing Commercial

Airplane Company ORIGINAL PAGEContract NASI-14952 BLA("K AND WHITE PHOTOGRAPF'f

i

Figure 2-18. Front Spar/Act,:ator Fitting Splice Test Specimen Failure

2-34

1983002978-046

r Boeing Commercial

Airplane Company ORIGINAL PAGE k

r Contract NASI-149_ OF POOR QUALITY :

' ,- __ " " " Aluminum

.- _ ,- ;_- _ ,_,;;,,, , actuator, , " support

/

P" - J I1_i

* Figure 2-19. Front Spar/Actuator Fitting So/ice Test Specimen Failure

f; Test No. 12--Panel Edge Shear and Bending-The setup for this test is

• shown in Figure 2-20. Test specimens were loaded to failure by apply-

_ ing either an upload as shown in F_gure 2-20 or a download (not

shown). The test failures to date typically occur away from the edge

1"_ band in the honeycomb section. Preliminary results indicate the detail

sustained approximately twice the airplane panel edge design load.

I;Test No. 14--Actuator Support Rib Verification--The actuator support rib

I; test was completed in June. The test setup is shown in Figures 2-21and 2-22, and the failed part in Figures 2-23, and 2-24.

• I: The test specimen failed at approximately 1.06 times the predicted

• failure load. The test failure load was equivalent to approximately 3.5

}:_/,_, _ 2-35/,

"%, i

1983002978-047

Boeing Commercial ORIGINALPAGE

Airplane Company BL_C,_ AND WHITE PHOTOGRAPHContract NASI-14952

/

, L

Figure 2-20. Pane/Bendmq Test

2-36

o.

198300P.q7R_OaR

11

t Boeing Commercial

Airplane Company ORIGINAL PA_b i_"-'_ OF POOR QUALITY

!- _ ContracL NASI-14952

Cp

I

_ rib

Aluminum :;.1_

support fitting

Boeing Commercial OEIGINALPAGEAirplane Company BI._CK AND WHITE PHOTOGRAPI_Contract NASI-14952

"* .# ''_ G

. ©• G

/ ," ".............. A

Figure 2-24. Actuator Support Rib Failure

times the actuator capacity on the airplane component. The graphite/epoxy

rib was sized to achieve approximately the same stiffness and strength

as the current metal design, which was designed to approximately 3.0

times the airplane actuator capacity.

Test No. 15--Flrst Sonlc Box--Testing of the first sonic box has been

completed. In su_mar>, the box was tested for 15 hr at 155 dB (overall)

and 4 hr at 158 dB (overall). This testing represented two lifetimes of

in-service accumulated damage. The box was then damaged by impacting

the honeycomb panel and edge bands at 12 locations. The damaged areas

were evaluated, using a hand-held ultrasonic inspection instrument. The

box was then tested for an additional in-servlce lifetime (4 hr at

158 dB overall). Post-test inspection revealed no apl _rent propagation

of the damaged areas.

A detailed presentation of data will be given when the laboratory test

report is complete.

2-38

1983002978-050

f

Boeing CommercialAirplane CompanyContract NASI-14952

fTest No. 17-Outboard Test Box (lO-ft box)--Testing was begun on this

test article in mid-July. Torsional and bending stiffnesses are being

evaluated. Testing will include subjecting the test article to the

"_ design ultimate load, and is planned for completion by the end of August.

_-- A finite element model of the test article, as it is supported and

loaded during test, is planned. The model will be extracted from the

overall elevator finite element model. Strain gage and deflection

• _ results will be compared to calculated values for finite eleanentmodel

analysis verification.

tTest No. 18--rearSpar Sonic Box-Testing of the rear spar sonic box has

.... been completed. This box was the same overall size as the Test No. 15

box, but incorporated a rear spar instead of a tapered, closed trailing

edge. The box was tested for 8 hr at 158 dB (overall), and then damaged

by impacting the panel and _ear spar at a total of 18 locations. An

additional 4 hr of testing at 158 dB (overall) did not cause any apparent

propagation of the inflicted damage.

2.3 DESIGN

_. Project is continuing to perform design-sustaining activity. The major-

ity of the work is not involved in changing part-configuratlon, but it

is providing drawing interpretation and clarification, correcting drawin8

and inconsistancies, and incorporating manufacturing toerrors requests

simplify fabrication of parts. However, two design changes were made.

The first change was made on the transverse horn rib _65C17536), where

• " the rib flanges were found to be inadequate when additional detail

stress analyses were conducted. Excessive flange bending was encountered

in reacting the horn weight design loads into the web of the transverse

rib. This bending was corrected by nestin8 0.18-cm (O.071-in) thickd

_ aluminum angle reinforcing clips against the rib inside mold line at

T each end of the rib at both the upper and lower flanges.

) _. 2-39

"1983002978-05"1

Boeing Commercial

Airplane CompanyContract NAS1-14952

The second design change involved the ply buildup of the rear spar

detail (65C17546-3). The first production rear spar detail had an

unacceptable spanwise warpage of 0.91 cm _0.36 in). This warpage was

diagnosed as resulting from differential thermal contraction between the

web and the flanges of the rear spar. The aft portion of the flanges

contained a stiffener insert (65C17546-11), which was constructed of

100Z spanwise (0°) fiber. Also, the outside ply in the aft portion of

the flanges was 100Z spanwise fiber. The remainder of the spar was

essentially of a quasl-isotropic layup (25Z 0°; 50Z 450; 25Z 900). The

thermal contraction of the 100Z 0° fibers is much less than that of the

quasi-isotropic construction. Thus, when cool-down from the 350°F

curing temperature takes place, warpage occurs. The fix was to build

the stiffener insert with a quasl-lsotropic fabric layup, and to make

the outside ply, which had previously been spanwlse tape, one ply of 0°

or 90° fabric. A spar of the new layup configuration has been fabricated

and the warpage has been eliminated.

2.4 WEIGHT STATUS

The current weight status of the advanced composites elevator system is

shown in Table 2-21. These weight data contain the following changes:

(1) Corrosion protection, which has previously been identified as a

line itemp has been partially distributed to the following itelns:

Front and rear spars 1.3 k8 C2.9 lb)

Ribs 0.1 ks C0.3 lb)

Skin panels 0.7 kg _1.6 ib)

Control tab 0.1 kg C0.2 lb)

The corrosion protection llne item has been reduced by a llke emount.

2-40

1983002978-052

Boeing Commercial


(2) The following items have been updated to reflect production drawing

predictions :

Fxont and zear spars +0.3 ks (+0.6 Jb)

Ribs -0.2 kg (-0.4 ib)

Control tab +0.4 kg (+0.9 ib)

Balance panet hinges +2.9 kg (+6.3 ib)

Elevator adjust weights +0.5 kg (+i.0 Ib)

Nose ribs and skins +2.0 kg (+4.5 ib)

Balance panel structure +0°2 kg (+0.5 ib)

Total weight of the advanced composites elevator system is now 191.3 kg

(421.8 lb), and the overall weight reduction is 26%.

2.4.1 Actual Weight.Program

Actual weight data of graphite/epoxy production detail parts are being

analyzed, and will be reported when completed.

2.4.2 Static Balance Requirements Document

The Static Balance Requirements Document, D6-46023, has been compiled

and is being reviewed, prior to preliminary issue.

The dccument details static balance requirements for the total elevator

system, the weight requirement of each balance panel, the required

moment of the elevator surface, and weight and balance requirements of

the control tab.

2-42

1983002978-054

£

_' Boeing CommercialAlrplane CompanyContract NASI-14952

%

SECTION 3.0,w

OPERATIONS DEVELOPMENT

IThis section dlscusses results of the manufacturin 8 produclbility

studies, ancillary test hardware fabrication, quality assurance develop-

ment, and verification hardware manufacture.

$ 3.1 PRODUCIBILITY STUDIES

Manufacturing requested a producibility study of the rear spar C65C17546-3),i

" because the detail as fabricated warped 0.91 cm (0.36 iz,)in the spa_wlse

direction. A review of the problem indicated the precured graphite/epoxy

, filler of 100% fibers in the spanwise direction was causing most of the

problem. Engine=ring changed the graphite/epoxy filler to a quasi-

isotropic layup of fabric. At the same time, the exterior graphlte/epoxy

tape ply was also changed to fabric.

• A second lear spar was fabricated and tzL_med using the new construction.

i The second rear spar had a warpage of 0.13 cm C0.050 in) in the spanwise• direction as compared to 0.86 cm (0.36 in) for the first spar. The

0.13-cm (0.050-in) spanwise warpage is acceptable, and can be removed

with slight hand pressure,

3.2 ANCILLARY TEST COMPONENT FABRICATION

" The ancillary test program includes allowables and envlro_mental, concept- , verification, and repair. The following describes the fabrication and

assembly status.

1983002978-055

Boeing Commercial


3.2.1 Allowables and Environmental

This part of the ancillary test program includes coupons (Test No. i),

joint/elements (Test No. 4), and environmental specimens (Test No. 3).

All allowable and environemntal specimens have been fabricated and

assembled.

3.2.2 Concept Verification

This part of the ancillary test program includes panel/rih pad (Test

No. 8), spar/shear web (Test No. 9), honeycomb panel stability (Test

No. i0), spar/alumlnum splice _Test No. ii), panel edge spar (Test

No. 12), rib design (Test No. 14), sonic test box (Test No. 15), 10-it

outboard test box (Test No. 17), second sonic test box (Test No. 18).

All concept verification hardware has been fabricated and assemhled for

Lasting.

Figure 3-1 shows the 10-it outboard test box [Test No. 17) during

assembly. Figures 3-2 through 3-6 show the second sonic test bo_ (Test

No. 18) during fabrication and assembly.

3.3 QUALITY ASSURANCE DEVELOPMENT

This section discusses the Preliminary Standards, Production Standards,

Fabrication Method, NDI techniques, Test Results, and Conclusions.!

3.3.1 Preliminary Standards

Preliminary standards were designed to represent the anticipated production

design, and were fabricated per BAC 5562. The following preliminary !

standards were built:

3-2

, Figure 3-2. SecondSonic Box, Test No. 18 Skin Panel, Showing Outer Skin P/ies in Place

a_ 3-3

1983002978-057

Boeing Commercial ORIGINAl PAGEAirplaneCompany BLACK AND WHITE pHoTOGRAPHContract NASI-14952

¢=

_k._' "_

Figure 33 Second Sonic Box Test No 18 Showing Honeycomb Rib Ready for Cure

J!

,:__: .....

\ J

" I. t

: • • • o o_ _,

i

Figure 3-4. Second Sonic Box, ;rest No. 18, Showing Skin Panel in Autoclave

, #

3-4

1983002978-058

t"Boeing Comercial

Airplane Company ORIGIIq#xL pAGE IS

Contract NASI-14952 OF POOR QUALITY

_t

1" Figure 3-5. Second Sonic Box, Test No. 18, Showing Front View of Completed Assembly

t

\4

'll r I" v !

4 • ......

Figure3_. SecondSonicBox, TestNo. 18,ShowingRearViewof Comp/etedAssemblyr

3-5

........ lirl = III I '_ ................ Illl ................ I ..... _i

1983002978-059

Boeln8 Commercial

Airplane CompanyContract NASI-14952 4,

• Laminate £tandard--This standard was built with 0.64- x 0.64-=m

(0.25- x 0.25-in) defects. Iu was required to qualify NDI techniques

per Boeing Advanced Composites Process Specification (J_C 5562).

See Figure 3-7 for details.

• Honeycomb Standard-Thls standard was built with 0.64- x 0.64-ci

(0.25- x 0.25-in) defects. It was required to qualify NDI techniques

per BAC 5562. See Figure 3-8 for details. °

• Laminate Step Standard-This standard consists of eight laminate

steps with 0.64- x 0.64-cm (0.25- x 0.25-in) and 1.27- x 1.27-cm

(0.5- x 0.5-in) defects in each step. The thickness of steps are

2P, 4P, 6P, 8P, lOP, 12P, 14P, and 15P. See Figure 3-9 for details.

• Front Spar Chord Standard--This standard has 1.27- x 1.27-cm (0.5-

x 0.5-in) defects in varlous thickness levels. The details are

shown on Figure 3-10 .

• Skin Panel Standard-This standard has 0.64- x 0.64-cm (0.25- x

0.25-in), 1.27- x 1.27-cm (0.5- x 0.5-in), and 2.54- x 2.54-cm

(1- x 1-in) defects. The standard was designed to duplicate upper

and lower skin panels. The details are shown on Figure 3-11.

• Taper Edge Standard--This standard was designed to investigate

defects in the taper area of the skin panels. The defect sizes

were 1.27 x 1.27 cm (0.5 x 0.5 in) and 2.54 x 2.54 cm (I in x

1 in). The details are shown on Figure 3-12.

3.3.2 Production Standards

Production NDI standards with built-ln defects were designed to duplicate

exactly the various sections of the production parts, and were fabricated

in accordance with BAC 5562. Following is a list of the production

standards:

3-6

1983002978-060

t

Boein8 Commercial O_ poO_ Q_LI_


(9in)

I _ 7.6 cm I 7.6 cm _I_ 7.6 cm _1k r'--- (3 in) _-i _ (3 in) -I - (3 in)

"' T "-_ Ply No. Orientation

t I 0°7.6 cm 2 +45°(3 in) 3 0°

*" / 15.25 cm 4 -45 °

f (6 in) 5 0°

" ['--] r-] 6 0°7 -45 °

; _ 7.6 cm 9 +45 °(3 in) 10 0°

P,,.,o,*_" --_ 10 Plies

// " BMS 8-212Type II

'= = Class2" '= Style 3K-70-P

t.

[_> O.64-cm2 (0.25-in2) disband

q ' _ 0.64<m 2 (0.25-in2) fabr;ccutout

_ _ Two 0.64-cm 2 (0.25-in2) Teflon shims-2 mil

Fabricate per BAC 5562

|-Iv

t _" Figure 3-7. NO/Reference Standard-Graphite/Epoxy Laminate

t b.P

,lb,

?

i _, 3-7

.. "_,

.,,mmmw - ..=-mmBF ...... v ,mw--_. _ _ ' _-_'- ' mr ,

1983002978-061

Boeing Commercial

Zirplane Company _I_,AL pAGE_ rTY|_ "_'

Contract NASI-14952 ,_ pOO_ _Ij_.,. -

I_ 23 cm :-I(9 in)

7-6cm I 7.6cm Ii 7.6cm"=--- (3 in) _ - (3 in) --- -- (3 in)

' (37.6Tin)Cm "_ Ply21No. Orientation0°0°

.__ 15.25 cm

[---] [_] (6in)/

7.6 cm r

1l

Nomex core 2 PliesBMS B-212

1.8 cm (0.50 in) Type IIthick Class20.32 cm (0.125 in) Style 3K-70-Pcell

0.64 cm2 (0.25 in2) disbond

0.64 cm2 (0.25 in2) adhesivecut out

Two 0.64 cm2 (0.25 in2) Teflon shims-2 mil

Fabricate per BAC 5562

Figure 3-8. "NDI Reference Standard-Graphite�Epoxy Honeycomb

3-8

1983002978-062

rBoeing CommercialAirplane Company

Contract NASI-14952

oRIGiNALPAGEIS

OF POORQUALITY

F 7.62 cm

I_-(3 in)_l_ _- --I: :_= -L _,_ _,__ :ITypical 8 places

i

I ! i I lli..,r •.15 -13 .11 "9 • 7 • 5 • 3 7.62cm

_" _ .16 i ,14 1.12 .101. 8 I. 6 i. 4 (3in)

._'_ _ 61 cm _1; (24 in) - I

_._. r---? (1-15)

I

_ -- i, /-. -- __ It/ i_

_ --:.- _ -- J

Defect No. 1 and 2 BetweenP1 and P2Defect No. 3 and 4 BetweenP3 and P4 Material:

_ Defect No. 5 and 6 BetweenP4 and P5 BMS 8-212• Defect No. 7 and 8 BetweenP5 and P6 Type II, Class2

Defect No. 9 and 10 BetweerlP6 and P7 Style 3K-70-PDefect No. 11 and 12 BetweenP7 and P8

Defect No. 13 and 14 Betweer,P8 and P9Defect No. 15 and 16 BetweenPy and P10

_:::" 0.64 x 0.64-cm (0.25 x 0.25-in) defect

-_1.27-cm x 1.27-cm (0.5 x 0.5-in) defect

[_> Two (0.002-in) shimsof specifieddefect size0.O05-cm Teflon

Ply No. Oriontation Ply No. Orientation

1 0° 8 +45°J 2 +45 ° 9 0°

3 90° 10 90°4 0 ° 11 -45°5 -45 ° 12 0°

! 6 90° 13 90°7 0° 14 +45 °

15 0°

i L

{ Figure 3-9. Laminate Step Standard

3-9

1983002978-063

Boeing Commercial -,

Airplane Company ORIGINAL-p/',G__n'_I'3Contract NASI-14952OF pOOR _uAL,. -.

l- 0.6cm- (16 in)

7" 7.62 cm(3 in)Typical 5 places

1 • o5 =9 _-t

2 • • 6 • 10 .--]q-

, 38 cm 1(15 in) J3 • e7 ol1

T

4 • e8 e12 -_

l-- J

7.62 cm I

Typical 5 places

P (1-17) _7 P (1,4, 7, 11, 14, 17) 6 plies (fabric

/ Material:

BMS 8-212

Type II, Class2Style 3K-70-P

11 plies(tape)

All defects1.27 x 1.27-cm (0.5 x 0.5-in) 2-m;I Teflon shims Material:

BMS 8-212Defect No. 1 BetweenP17 and P16 Defect No. 7 Be_een P1 _ and P10 Type II, Class1Defect No. 2 BetweenP16 and P15 Defect No. 8 BetweenP10 andP9 Grade 145Defect No. 3 BetweenP15 and P14 Defect No. 9 BetweenP17 and P14Defect No. 4 BetweenP14 and P13 Defect No. 10 Be_een P14 and P11Defect No. 5 BatweenP13 and P12 Defect No. 11BetweenP11 and P7Defect No. 6 BetweenP12 and Pll Defect No. 12 BetweenP7 and P4

Figure 3-10. Front Spar Chord Standard

3-10

1983002978-064

!

Boe£ng Commerc£al ORIGINAL ['AGE !_A£rplane Company OF POOR QUALITYContract NASI-14952 1

- 1= (9 in)

" T7.62 cm

(3 in)• Typical 3 places

2,54 cm2 1.27 cm2(1 in2) • • (0.6 in2) _ Ply No. Orientation

23 cm _ 1 0°" (9 in) 2 00

" 1.27 cm2 0.64 cm2(0.5 in2) • • (0.25 in2) _' _

J

i I'--'0cm_I_ _I_• (3 in)' _11Typical 3 places

Teflon shims

i AF- 143 adhesive

BMS 8-212•_ Type II, Class2

•I. I¢ _

3 PCF Nomex "] " " " Style3K-70-P --" 0,32-cm (0.125-in)

cellsqt

" P 11,2)_/AF-143 adhesive

" Figure3-11. Skin Pane/Standard ,

_ t'r4

_ _V 3-II '.

_L_-_': ...................@-

1983002978-065

Boeing Commercial

Airplane Company tO_'IG|_:L Q_ALfI'YContract NASI-14952 OF pOOR

I= 51 cm _11(20in)

P (1, 2) ---_ Ply No. Orientation• 1 0°

2 0°

AFo143adhesiveX//_I --

--t

_X'X.__ p (1,21Detail1

Material:

BMS8-212"{ypeII, Class2Style3K-70-P

Core- Nomex, 0.32-cm(0.125-in)cells

1.27x 1.27-cm{0.5x 0,5-in)defect2.54 x 2.54-cm(1 x 1-in)defect

Figure 3-12. Taper Edge Standard

3-12

1A., •

T i i

, =i==,-

1983002978-066

6Boeing Commercial


-- • Upper and lower skin panels were fabricated per drawing 65C-17707-902

and 65C-17707-903 from station 136.50 to 173.21. The built-in

defects, ranging from 0.64 x 0.64 cm (0,25 x 0.25 in) to 2.54 x2.54 cm (I x i in), were introduced thloughout the panels, as shown

- in Figure 3-13.

_ • The two skin panels were sealed at the tapered end to make a

trailing-edge assembly. The defect size ranged from 1.27 x

1.27 cm (0.5 x 0.5 in) to 5.08 x 5.08 cm (2 x 2 in). The details

are shown in Figure 3-14.

• The rear spar standard, Figure 3-15, was built per drawing

65C-1707-7E, from station 117.0 to 99.79. The defects ranged from

0.64 x 0.64 cm (0.25 x 0.25 in) to 2.54 x 2.54 cm (i x 1 in).

• The front spar standard, Figure 3-16, was built per drawing

_. 65C-17707-9006E, from station 149.34 to 129.19, with defects ranging

from 0.64 x 0.64 cm (0.25 x 0.25 in) to 2.54 x 2.54 cm (i x I in).

• The rib standard, Figure 3-17, was built per drawing 65C-17707-4E,

full size. The defects ranged from 0.6& x 0.64 cm (0.25 x 0.25 in)

to 2.54 x 2.54 cm (i x 1 in).

| 3.3.3 Fabrication Method

The NDI standards were fabricated per BAC 5562. The defects were

introduced in the laminate panels, as outlined in FiguTe 3-7. In the

1 production standards, however, the cutting out of fabric or tape wasdiscontinued, and 2 shims of 2-mil Teflon were introduced in between the

fabric or tape layers. In the honeycomb assemblies, _he defects were

( introduced as outlined in Figure 3-18.

(3-13

i

1983002978-067

Boeing Commercial

Airplane Company ORIGINALPAGEContract NAS1-14952 BLACK AND WHITE PHOTOGRAPH

a mLPjl

• , L, ,',. , ,,;, • . • .

P _

- • . "r "co""; ", :" _"_ " "''"

, '' . _ ' "*¢ ,._t fr'lT_ P _=," "_ ' _'a& ,t_P_<irll'4_# 'lt_llM' "

titI

Figure 3-13. 727 Advanced Composites E_evator Skin Panel Standard

l*

/

..................................... __, _"_ _ _ _

II IIII II . ,.... _

Figure 3-14. 727 Advanced Composites Elevator Trailing-Edge Assembly Standard i

i

3-14

, . - _.

1983002978-068

Figure 3-15. 722 Advanced CompositesElevator RearSpar Standard

_- i¢ .• 1 =

Io

¢

• c, •

i Figure 3.16. 727 Advanced CompositesElevator _-ront SparStandard

1983002978-069

Boeing Con_ereial

Airplmne Company

-I Contract NASI-14952

,: ,:_

,e

• .., f,:.:.'."

Fiqure 3-17. 727 Advanced Composite: Elevator Rib Standard

Figure 3-18. Automated TTU Scanner

o_ _oo_ _,A_t,--_ 3-16

1983002978-070

6

C Boeing CommercialAirplane CompanyContract NASI-14952

3.3.4 NDI TechniRue 9

f NDI techniques evaluated during this investigation were the following:

" ultrasonic (TTU)with automated andThrough-transmlsslon scanning

computerized C-Scan recording

f-$ Through-transmisslon ultrasonic, with manual scanning and a visual

_ signal display having C-Scan recording capability

$ The Sondicator, Mode' $-I or S-2B

• The Fokker Bond Tester

(• Low kV X-ray (15-40 kV)

('" 3.3.5 Test Results

" All NDI standards were evaluated by the above-mentloned techniques.

Through-transmlssion ultrasonic was the most sensitive method for

" production application. When a permanent record is not required, the

portable instruments can be used. The automated through-transmis31on

" ultrasonic unit with computerized C-Scan recording capability is shownon Figure 3-18. Figure 3-19 shows a sample of computerized C-Scan

recording. Voids, delaminations, or porosity indications are shown by

(_ darker recordin?_ (and higher ntunbers). A _emiportable model of through-

transmission ultrasonic without C-Scan capability is shown in Figure

1_ 3-20. Most production parts can inspected by through-of the be

transmission ultrasonic, except some configuration, (Figure 3-21), are

I_ not suitabl for this inspection. Other examples of parts uninspectableby through-transmlssion ultrasonic are small corner radii, flanges of

i_ spars and ribs, and sealed area of trailing-edge assembly. These areascan be inspected by other techniques, as shown in Figures 3-22, 3-23, and

3-24.

t;'V +

3-17 'j

12 .. _ i i _mm,_ : -" C_•--.... ..

. _ i¢_ _ _, _ __ _- v

1983002978-071

Boeing Commercial _G%:'?"---4"[_I.[ _

Airplane Company O1: F" '"

Contract NASI-14952

GRAPHITE TEST PANELPART NUMBER _ 11423A SERIAL NUMBER • 0@8DATE - 3/ 9/78 CALIBRATION PROCEDURE o L4SCANNER INDEX - G@ THOUSANDTHS SCAN PATTERN - AX SCALE o 1 Y SCALE - 1 SCAN DIRECTION - LEFTPLOTTING SEQUENCE - 1,0,_,0,0,@, NUMBER OF OPERATIONAL CHANNELS - 1

SIGNAL IDENTIFICATION LEVELS PLOTTED SYMBOLSLEVEL MAX YOLTS_I@@

O I@O@ IOO@ .GT.BLANK .GE, 70_1 ?00 70@ .GT. : .GE. 335

335 335 ,GT. 1 .GE, 28S3 28S 2_S .GT. _ .GE. 2SS4 255 2S5 .GT. 3 .GE. 22SS 22S 225 .GT. 4 .GE. t9_6 lge lge .GT. S .GE. 1SS7 1SS 1ES .GT. 6 .GE. 11S8 115 IlS .GT. 7 .GE. ?S9 7S 7S .GT. 8 .GE, $0

I@ SO S@ .GT g .GE. 3811 30 3@ .GT. I@ .GE. IS12 IS IS .GT. II .GE. 0

TOTAL NUMBER OF BLOCKS " 21@ ................................ z,,:,ll, • . , ,.,H,3 ........ - ,i , h , I I ............. I .,,ll,.,I .... ih,I ........ IIh

.................................... '" '" " ..... ' ..... ..:,'"..... " '" " ." .-..:"_'.'_:".:'" ':'. ' ."'"'"'. ....... " .... ' '" " _Y_:_L`_:`:_`_`:_._'_::::_._'!__.__'. . . '' . . '.: ". . . . ' ,,,,,, ,i I........... ...............II''HI' .... ,,.,,lll.,. , ,, , ',,,,,,,,,,,'.... H,, ......... ,,,.,.,

• . ', . ,, . . : '" ,',":,X,;:':,::',,;:::::;::":'.',.:.• I _ ..... illawl_mll*l .... IIIIqlallallll ..... a

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.. .:.. : .'. .." ..... .,, ,, . , • .' : ..... ". .: :

........ , .... , ..... ,.,, ................. ,, ......... , ..... Vl ....... ,ll.,.,.l,,,,l ..... I,,.: I*,,,IO ....... o.,,,,, .... , ...... 1,1,,,i., 4, ..... , .......... ,...,,,,..,,,," i,

............................................................... ,, ................................................. ,,.,..'., ,,.,,, ................ .................................. ,, ...... , ..... , ........... , ...... ,.. ,............................... ,...., ....... ;; ............... .,_............................m _

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., . , ;: ! • . • : • . .... • . . , ., ..... :• . . ' ' '".... " '.. : . .':: . ............." ...... ,. ' .'".:,, 'i;• i • • I ,II, ii I II, I, I • i ......

I , , • •

if ii I ........ 1 .I I Iii ,1 i ,i i Illll Ill " J,l,'ll II ....... IIIII l Dill II ....... Illlll I I I I lr I IIi1.1 ,, ..... ,I I1,1, I, '1 1,1,11, Ill'llll' IIIllh I ,,I,_1' IIII I,., IIIII

.......... I ...... I, I II I'll II II .......... I ....... I ..... III II1.11111 IIIIIII IIIIIII I I IIIIIIIII I III I I I# Illi% I1 1111111 I IIIIIII I1[111 11 IIIII Illl*llll IIIIIIIIIII IIIII II IIIIIIIIIIII

_: I1,1 I I I .... II111 I III I I1,111111 I11 III IIIII1'11111 .... III III11111 IIIII 1111 l I.IIIIIIIIIIIII IIIIIIIIIIIII r ii i i111111 Ill i11 ii iiii1,1 I iiiii iii i II iiiii i i I iii ii II i111111tl

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Figure 3-19. Graphite�Epoxy Test Panel Standard, Computerized C-Scan Recording

3-18

L

.... -,,-,,ira-

1 983002978-072

Boeing Commer¢ial

Airplane Company ORIGINAL PAGE ISOF POORQUALITY

Contract NASI-14952

_hi,-

• I

Figure 3-20. Semiportable TTU Scanner Without C-Scan Capability

, _,__--- _ ....

Figure 3-21. Semiportable TTU Scanner, Showing Part Configuration Not Suitable for

TTU Scan

_ 3-19

,t .a. _ #' .a .=.=,_.=..,,_amlw.,,_m.=.=_ '" ''" - ....

- '- .......... -................... 1_)83-0"-0"297807:

Boeing Commercial

Airplane Company ORIGINAL PAGEContract NAS1-14952 _ DID WHITE PHOTOGRAPH

,,. _'_==.,,,,_..._•

o'. , _

Figure 3-22. Portable TTU Scanner, Showing Inspection of Flange Area by Hand-HeldTTU Unit

=.L _ . r

'--,,qll ....

,, /i" ,_1_,,, _J

Figure 3-23. Sondicator Inspection of Configuration not Suitable for TTU Scan

3-20

i

........_._,.,.._,_.m,_J,_,,_ == ....... , , i,,

1983002978-074

Boeing Commercial

Airplane Company ORIGINAL PAGE I_

t Contract NASI-14952 OF POOR QUALITY

CIr. ' metolml

oelt_l "Jlct;_tle #'Joee_P'eel,*.... :.......l';",, .:l.........

.........;"'-, ,;::

q.

Figure 3-24. Fokker Bond Testingof Configuration not Suitable for TTU Scan

3.3.6 Conclusions

• All defects in the NDI standards are detectable by one or more of

the NDI techniques

• " • Production parts can be insp cted by using through-transmission

ultrasonic to detect defects, 0.64 x 0.64 cm (0.25 x 0.25-in) or- larger.

. • Some part configurations, such as rear spar, flange areas of spars

and ribs, small corner radii, and sealed area of traillng-edge

assembly, cannot be inspected by through-transmlsslonultrasonic

technique

{

1983002978-075

Boeing Commercial


• Rear spar, and flanges of spars and ribs, can be inspected by using

hand-held through-transmission ultrasonic, the Sondicator, or the

Fokker Bond Tester. Because of usual limitation of the manual

: scanning, the realistic defect detection capability is 1.27 x

1.27 cm. (0.5 x 0.5 in).

• The sealed area of the traillng-edge assembly can be inspected by

the Sondicator or X-ray. The detection capability of the Sondlcator

is 1.27 x 1.27 cm (0.5 x 0.5 in), while X-ray technique can detect

minute voids in the sealant.

• In-service and maintenance inspection can be accomplished by the

Sondicator, the Fokker Bond Tester, and/or X-ray (when practical).

Since these inspection conditions are not as favorable as the

laboratory or production environment, the realistic defect detection

capability for the Sondicator and the Fokker Bond Tester is expected

to be 2.54 x 2.54 cm (i x i in) or larger.

3.4 VERIFICATION HARDWARE

Details and assembly of the 203-cm (80-in) long, full-scale section of

the verification hardware have been completed (Flgures 3-32 and 3-33).

The structure verified the production detail tools, and assured Ma_Lufac-

turing that no ma_or assembly problems will be encountered during produc-

tion of the 5-I/2 shipsets.

The verification hardware was also used as a training aid for production

assembly personnel. Two significant facts were learned from the verifica-

tion hardware:

I. It was concluded that using only a section of the structure to

verify the fabrication and assembly processes will not always

identify all of the problems associated with the full-size atructure.

3-22

i

q

1983002978-076

df Boeing CommercialAirplane Company

Contract NASI-14952 ORIGINAL PAGE 18

I OP POOR QUALITYf

The 203-cm (80-in) front spar and short rear spar used for the

!_ verification hardware showed only slight warpage after fabrication.

However, the full-slze production spars showed excessive warpage,

I . causing rejection of the front spar and scrapping of the rear spar.

Based on this experience, full-slze detail parts should be fabricated.

2. During fabrication of the concept verification hardware (test

( hardware), dimensional problems occurred with the ribs. It was

J concluded that these problems occurred because all ribs were

fabricated on male tools that were made by scaling the photo

," contact master (PCM) drawing. Variations in the PCM drawing and

the scaling process accounted for the dimensional dlscrepencies.

(The verification ribs were fabricated on production male tools, which

. were made using master dimensioning index (MDI) data, and numerical#

control machining. No dimensional descrepencles were found in ribs

I fabricated from these tools.

Figures 3-25 through 3-33 show the verification hardware (Test No. 16),

I I_" during fabrication and assembly.

,j _-

% - .

- ? •

Figure 3-25. Manufacturing Verification Hardware, Test No. 16 Skin Pane/,Show;ng i

i Doubler Being L_id Upt

liililalmalm,im_ ..... '........... _ ........... iwl II I I I ...... | II -- _..... =_ .................. _ .... i lall ..... _ ............... _" ....... _'

"iw_-- _-- -- ..... , m_ u.-- _ - --'ilF" v- _ W _"IF'_r .......

1983002978-077

Boeing Commercial

Airplane Company ORIGINALPAafContract NASI-14952 Bt:ACKAND WHITE PHOTOGRAPH' 'qP

Figure 3-26. Manufacturing Verification Hardware, Test No. 16, Showing CompletedSkin Panel

!

Figure 3-27. Manufacturing Verification Hardware, Test No. 16, Showing Skin Panel inTTU Inspection

3-24

1983002978-078

,qmm=

* Boeing Commercial

Airplane Company ORIGINAL PAGE ISf Contract NASI-14952 ()FPOOR QUALITY• I

i

_J

lI

D

, "_ ; : _, ,,. . ;. .,_.._-_ •

F/gum 3-28. Manufacturing Verification Hardware, TestNo. 16,Showing Front SparBeing Laid Up

/

/

. _ ;....,_f,@,,,,===.=.=m=....m=--==mmmm ==mmmm

F/gum 3.29. Manufacturing Verification Hardware, TestNo. 16,Showing Completed.... RearSpar Header

1

1983002978-079

BoeLng Ct_mmerci,l]

Ai rplane Compr.,ny ORIGINALPAGI_Contract NAS1-14952 |LACK AND WHITEPHOTOGRAPH

/

Figure 3-30. Manufacturing Verification Hardware, Test No. 16, Showing Front Spar,Two Ribs, and Lower Skin Being Assembled

m i

"'g '7 °

Figure 3-31. Manufacturing Verification Hardware, Test No. 16, Showing Aluminum Nose

Ribs Being Attached to Front Spar

3-26

1983002978-080

i

TT Boeing Commerc ial OR£_/NAL PAGI_ I8

Airplane Company OP POOR QUALI'I_Contract NASI-14952

l_ 2;:a. =- _.-_,-,\;Ifr"

T i '_ - _F_\\< > '.

Figure 3-32• Manufacturing Verification Hardware, Test No. 16, Showing Front View ofComple ted Assembly

,,r _ u.

"" - " " - .... II ........ _ _',_"

2 ._' t -i

I_igure 3-33. Manufacturing Verification Hardware, Test No. 16, Showing Rear View ofCompleted Assembly

1

1983002978-081

_J

qt

4- Boeing CommercialAirplan _. Company

" Contract NASI-14952

. ORIGINAL PAGE Ib

SECTION 4.0 OF POOR QUALITY

PRODUCTION

This section discusses the production status of the program.

4.1 ASSEMBLY TOOLS

+

Fabrication of left-hand assembly tooling is proceeding per schedule.

; The front spar final assembly Jig is in the final steps of tool tryout

(Figure 4-1). The rear spar final assembly Jig was completed dlring the

report period, and is presently set up in the assembly shop, where tool

tryout is being accomplished (Figure 4-2). The major assembly final

assembly Jig is also cemplete, and being set up in the assembly shop

(Figure 4-3). Assembly work and tool tryout will activate in the _mJor

position during the next reporting period. Remaining activity under

this woz_ breakdown structure involves completion of rewor_ to the

existing elevator tab tool to provide dual usage, and fabrication of the

transportatlon/storage dolly for the left-hand unit.

i _ _ . - ._ _,._ .... • _i _ .... ,-....

ram"--L ' ' ..... ! 'q , i ....

4 --l_--_.-_ I _. " . _- .

"7

°, 1°. - __.

( Figure 4.1, Front Spar with Nose Ribs Insta//ed and Located in Too/

_ ,-.. 4-1

Ig83002978-082

_'_ ORIGINAL PAGE IS

OF POOR oUArxPv

Boeing Cmmercial


4.2 COMPONENT MANUFACTURE

Fabrication division has completed production of all advanced composites

i components for the left-hand test unit, except for the skin panels and

tabs. These components were scheduled for completion during August.

Significant problem areas encountered during the fabrication of advanced

c _pcsites production parts, and the changes instituted to correct them,

are as follows:

: Problem_o. i: Difficulty in maintaining proper fitup of

65C17546-5 and -13 rear spar headers to the

65C17546-3 spar, due to corner thlcknessea.

Fix: Layup of the rear spa_ headers was ahanged so

that only the inner and outer plies of fabric

will be overlapped in the four corners. The

middle five plies wii_ be butt-spliced at the

corners. This change will eliminate the thick

corners that caused the improper fltup conditions

(Figures 4-4 and 4-5).

Problem No. 2: Fabrication of skin panels encountered difficulties

with engineering requirements for overlapping

splice doublers and filler plies.

Fix: Fabrication division was given authority to

overlap the sp%ice doublers only, and to butt-

splice all filler plies per hAC 5562. This

method adapts t., a production en¢Ironment,

produces higher quality skins, and has a higher

material utilization.

Activities associated with assembly of the left-hand test elevator,

which began July i0, 1978, are as follows: '

i

__1_ 4-3 i

1983002978-084

Boeing Commercial


Figure 44. Rear Spar with .5 Rear Spar Header Installed, Showing Mismatch

_,,e • , .e_

Figure 4-5. Rear Spar Header, Showing Bow in Side

OF vOO_4-4

k

1983001297_

Boeing Commercial

_ Airplane CompanyContract NAS1-14952

i. 65C17529-950 Rib Assembly

This is the backup rib that attaches to the main actuator fitting.

All c( )onents are of advanced composites material. It is now

complet_ and in storage. No difficultles were experienced in fltup

of parts oz dri111ng operations (Figures 4-6 and 4-7).

2. 65C17546-950 Rear Spar Assembly

" No assembly work has been accomplished as yet. Assembly parts are

being utilized for tool tryout. Relative to the close-down of the

channel-shaped rear spar in its cured condition, there appears to

be no problem with springing the channel sides open far enough to

install the headers and tab hinge fittings.

3. 65C17545-950 Front SparlLeadin_-Edge Assemb1_-

Assembly work is proceeding, along with tool tryout work. Nose rib

attach angles have been installed, and the spar halves have been

attached to the main actuator fittlnE. Nose rib assemblies have

been prefitted to the attach angles , and fastener locations have

been drilled.

Rel_tive to the warp in the spars that occurs durlnE the cure

cycle, i_ is noted that installation of the nose rib attach angles

corrects the condition.

Utilizing the special drills and 18,000 r/min motors for drilling of

fastener holes through tie advanced composites part& did not

present any problems where toollnE provisions included a posltve

guide. Drilllng free-hand caused some concern, due to the abillty

of the special drills to act like a router and cause oversize

holes. Free-hand drilling techniques were covered during tralnin5

class to help control this problem. A number of holes must be

• free-hand drilled.

,i

4-5 _

1983002978-086

Boeing Comercial


• _'-

Boeing Commercial

" Airplane CompanyContract NASI-14952

Modification to vacuum dust collection features developed for the

drill motors has been required to correct interference problems (Figures

4-8 and 4-£). MR&D is working on further refinements to improve efficiency.

Breakout of graphite/epoxy fiber has not been experienced. There is

breakout in the fiberglass first ply, but this is not a problem. The

fiberglass ply is a corrosion barrier only (Figure 4-10).

In most cases, the grip length callout for fasteners is too short,

because graphite/epoxy parts are on the high side of the thickness

tolerance range. All drawings have now been reexamined by Engineer-

ing, and revised EAMRs calling out longer grip lengths, where

required, lmve been issued. The 707/727/737 Materiel department,

working from advanced information and the revised EAMRs, is in the

process of determining supply sources and expediting shipments of

new grip length fasteners. It is noted tt_t the majority of fasteners

are unique to the elevator program, are not shelf-stock items at

Boeing or the suppliers, and require extended lead times for procure-

ment. The Manufacturing department is working with Engineering to

determine where a_ailable substitute fasteners can be used. Produc-

tion schedule impact will he assessed, when all aspects of fastener

availability have been examined.

Assembly progress has been paced by tool tryout. The major assembly load

date of August 17, 1978, will be met. Fastener availability would be the

only constraint.

4-7

1983002978-088

Boeing Commercial ORIGINALPAGEAirplane Company BLACKAND WHITE P_OTOP,RAPHContract NASI-14952

h_m_m_mmm

Figure4-8. HandDrill (right),ShowingStandardVacuumUnit andModified VacuumUnit (left)

.,,.,---/ - (",, ,

• ,_r

Figure4.9, DrillingAttach Angleson FrontSpar,ShowingInterferenceBetweenAngleand', VacuumCollar

4-8Jk

198300297R_nRo

It

I¢I¢

,, #

¢

t +,f4-9

1983002978-090

C"Boeing Commercial

Airplane Company-- Contract NASI-14952

%

SECTION 5.0

f

REFERENCES

i. "727 Structural Design Loads," Boeing Document D6-5863.

2. "727 Horizontal Tail Stress Analysis," Boeing Document D6-5873.

" 3. "Boeing Design Standards," _oeing Document D-5000.

4. "Advance_ Composites Design Guide," Volumes I through IV, ThirdEdition.

5. "Requirements for Epoxy Resin Impregnated Graphite Fibers," BoeingDocument BMS 8-212.

6. "M_ufacture of Autoclave Cured Graphite/Epoxy Structural Parts,"Boeing Document BAC 5562.

5-I i

1983002978-091

ADVANCED COMPOSITE ELEVATOR FOR BOEING 727 AIRCRAFT

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Transcript of ADVANCED COMPOSITE ELEVATOR FOR BOEING 727 AIRCRAFT