NASA Technical Memorandum 110170

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Analysis of Selected CompressionSplice Joint Locations in a Graphite-Epoxy Transport Wing Stub Box

D. C. Jegley

Langley Research Center, Hampton, Virginia

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ANALYSIS OF SELECTED COMPRESSION SPLICE JOINT LOCATIONS IN

A GRAPHITE-EPOXY TRANSPORT WING STUB BOX

Abstract

Three critical compression splice joint locations in a stitched graphite-epoxy

transport wing stub box have been analyzed to determine their expected

structural performance. The wing box is representative of a section of a

commercial transport wing box and was designed and constructed by

McDonnell Douglas Aerospace Company as part of the NASA Advanced

Composites Technology (ACT) program. The results of the finite element

analyses of the splice joints are presented. The analysis results indicate that

failure will not occur in the splice joint regions for loads less than the Design

Ultimate Load of the wing box.

Introduction

To evaluate the potential of a stitched graphite-epoxy material form for use on

commercial transport aircraft wings, a short section of a wing box was

designed and fabricated by the McDonnell Douglas Aerospace Company under

the NASA Advanced Composites Technology Program (ACT) contract NAS1-

18862. This short wing-box section is referred to herein as the "wing stub

box." Current plans are to test the wing stub box at the NASA Langley

Research Center to evaluate its structural performance. Several structural

analyses were conducted by McDonnell Douglas Aerospace Company and by

NASA Langley Research Center in support of this effort. The present paper

describes the results of the analysis of three compression-loaded splice

locations in the wing-stub-box test specimen.

Wing-Stub-Box Test-Specimen Description

The wing-stub-box structure consists of a metallic load-transition structure at

the wing root, a composite wing stub box, and a metallic extension structure at

the wing tip as shown in figure 1. The load-transition structure and the wing-

tip extension structure are metallic end fixtures required for appropriate load

introduction into the composite wing stub box during the test. The load-

transition structure is located inboard of the composite wing stub box

(between the composite wing stub box and the vertical reaction structure at

the wing-stub-box root) and the wing-tip extension structure is located

outboard of the composite wing stub box. The load-transition structure is

mounted to a steel and concrete vertical reaction structure resulting in a near-

clamped end condition. The entire structure, including the composite wing

stub box and the metallic structures, is approximately 25 feet long. Details of

the geometry of the structure are presented in reference 1.

The composite wing stub box was fabricated from Hercules, Inc. AS4/3501-6

and IM7/3501-6 graphite-epoxy materials which are stitched together using E.

I. DuPont de Nemours Inc. Kevlar threads. IM7 graphite fibers are only used

for the 0 degree fibers in the lower skin. The composite skin and stiffeners

are composed of layers of the graphite material forms prekitted in nine-ply

stacks that have a [45/-45/02/90102/-45/45]T stacking sequence. Each nine-

ply stack is approximately .058 inches thick. Several nine-ply stacks of the

prekitted material are used to build up the desired thickness at each location.

The composite wing stub box was fabricated using a Resin Film Infusion (RFI)

process (see ref. 2).

Splice Description

Metallic splices are used to join the composite wing stub box to the metallic

wing-tip extension structure and to the metallic toad-transition structure (as

indicated in figure 1). These metallic splices are bolted to the ends of the

graphite-epoxy stiffeners and to the metallic extension and load-transition

structures. Three splice locations on the upper cover panel of the composite

wing stub box have been analyzed in detail and the results of the analysis of

the three splices are presented in the present paper. The locations of the

three splicesare shownin figure 2. Splice 1 is locatedat the outboardend of

the compositewing stub box. Splices2 and 3 are locatedat the inboardend ofthe compositewing stub box.

Each splice consistsof aluminum inner and outer splice plates and aluminumangle splices,as shownin figure 3. Bolts are usedto attachthe stiffener to the

angle splices and the skin to the inner and outer splice plates. All stiffeners

are approximately2.3 inches high and .46 inches thick with a 1.12-inch-wide

flange on either side of the blade. The distancebetweenthe end of the

composite upper-cover-panelskin and the end of the compositestringer is8.59 inches(as shownin figure 2) for all splicesconsidered. However, thedistancebetween the end of the compositestringer and the nearestrib is

different for the three splicesconsidered. The distancesfor splice locations 1,2, and 3 are shownin figure 2 and are 6.1 inches,6.4 inchesand 2.8 inches,

respectively.

Model Description

Three similar models were constructed to model accurately the geometry of the

three splice locations considered. In each case, the entire 8.59-inch-long

region of unstiffened skin and splice plates and 11 to 12 inches of the stringer

is included in the model. The stringer is centered laterally for each model.

The entire flange and 3 inches of skin is included on each side of the stringer,

resulting in a region to be analyzed for each splice location that is

approximately 20 inches long and 6 inches wide. The intercostal in the region

of each splice is also included in the model.

The skin thickness at splice location 1 is nine of the nine-ply stacks of

prekitted material. The skin thickness at splice locations 2 and 3 varies from

eight to ten of the nine-ply stacks. The variation in skin thickness is included

in the model. Quadrilateral and triangular finite elements are used to model

the skin, blade and splice parts. Beam elements are used to model the bolts

which connect the metallic parts to the composite parts. Each model has

approximately 2,200 elements and 17,000 degrees of freedom. The models for

the three splice locations are shown in figure 4.

The material properties for the stitched nine-ply stacks of the prekitted

materialusedfor the skin andbladesareEx=8.17Msi, Ey=4.46Msi, Gxy=2.35Msiand Vxy=.459 (ref. 3). The materialpropertiesused for the aluminumsplice

platesandanglesareE=10.6Msi, G=3.8Msi andVxy=.395.

Analysis

Boundary conditions for the local compression-splice analyses were obtained

from the global nonlinear finite element analysis results described in

reference 4. Nodal displacements for each skin node at the local-splice-model

edge were calculated from the global analysis displacement results. A

nonlinear analysis of the splice regions was conducted using the STAGS finite

element code (ref. 5). Maximum loads applied to the splice models correspond

to the Design Ultimate Load (DUL) condition for the wing stub box.

Results and Discussion

Calculated displacements in the splices at splice locations 1, 2 and 3 are shown

in figures 5, 6 and 7, respectively, for the DUL condition. Displacements shown

in figures 5a, 6a and 7a are not magnified but are to the same scale as the

model. Displacements in the axial (spanwise), lateral (chordwise), and out-of-

plane directions are also shown in figures 5-7.

The outboard end of the cover panel skin for splice location 1 deforms out-of-

plane 3.2 inches at DUL and the inboard end deforms out-of-plane 2.2 inches.

Most of this displacement is due to the global bending of the wing stub box

rather than local displacements in the joint region. The outboard end of the

upper-cover-panel skin in splice location 2 deforms out-of-plane .16 inches at

DUL while the inboard end deforms out-of-plane .11 inches. The outboard end

of the cover panel skin in splice location 3 deforms out-of-plane .18 inches at

DUL while the inboard end deforms out-of-plane .07 inches. The out-of-plane

displacement at splice location 1 is much larger than the out-of-plane

displacement at splice locations 2 and 3 since splice location 1 is farther away

from the root of the wing stub box. The inplane displacements are much

smaller than the out-of-plane displacements at splice location 1. The inplane

4

displacementsare of the sameorder of magnitudeas the out-of-planedisplacementsat splice locations2 and 3.

Calculatedstrains in the splicesat splice locations1, 2 and 3 are shown in

figures 8, 9 and 10, respectively,for the DUL condition.The axial strains in thesplices are shownin figures 8a, 9a and 10a, the lateral strains are shown in

figures 8b, 9b, and 10b, and the shearstrainsare shownin figures 8c, 9c, and10c. Away from the boundariesof the model, the maximumaxial surface

strains at DUL are approximately.008 in/in. The unnotchedallowable strain

for this material systemgiven in reference3 is .0093 in/in.

Away from the boundariesof the model, the maximumlateral strainsat DUL

are approximately.008 in/in. The maximum shear strains are .012, .004 and

.008 in/in, for splice locations 1, 2 and 3, respectively when the structure is

subjected to DUL.

Strain gages are to be attached to the wing stub box at the locations shown in

figure 11 for the three compression splices considered and the strain gages

are designated as strain gages A, B, C, D and E. Gages A and D are back-to-back

gages on the composite upper-cover-panel skin. Gage B is located on the

aluminum splice. Gage C is located on the aluminum inner splice plate and

gage E is located on the aluminum outer splice plate. Analytical predictions of

the strain at these strain gage locations are shown in figures 12, 13 and 14 for

splice locations 1, 2 and 3, respectively. The load for the ordinates in the

figures is the load applied at the outboard end of the wing-tip extension

structure. The applied load corresponding to Design Ultimate Load is 166,000

lb. The load-versus-strain curves shown in figures 12-14 indicate that the

maximum strain is predicted to occur at the location of strain gage A, and is

.0081, .0082, and .0071 in/in, at DUL for locations 1, 2, and 3, respectively. The

strains increase almost linearly with load, indicating that geometric effects do

not cause nonlinear response. Applied loads of approximately 195,000 lb,

185,000 Ib, and 220,000 lb would be required for the strain at strain gage A to

equal the unnotched allowable strain of .0093 in/in, at splice locations 1, 2, and

3, respectively. These load levels correspond to margins of safety of .17, .11

and .32, which indicate that failure should not occur at any of these regions

unless the load is 11 percent greater than Design Ultimate Load.

Since the margins of safety in the regionsof the upper cover panel away fromthesesplice locations,such as near the accessdoor, are less than .05 (ref. 6),

failure is not expectedto occur in thesesplice regions

Concluding Remarks

Three compression splice joint regions of the upper cover panel of a graphite-

epoxy wing stub box were analyzed using a finite element analysis. Loads up

to Design Ultimate Load (DUL) were considered. The predicted maximum strain

in each compression splice joint region at DUL is approximately .008 in/in.,

which is less than the allowable strain of .0093 in/in, for the material. The

wing stub box is not expected to fail at these compression splice joint regions

for applied loads up to 11 percent greater than DUL.

References

1. Markus, Alan M.; and Grossheim, Brian G.: Manufacturing Development for

Stitched/RFI Transport Wing Stub Box Structures. Proceedings of the Fifth

NASA/DoD Advanced Composites Technology Conference, NASA CP 3294, pp.

773-786, May 1995.

2. Markus, Alan; Thrash, Patrick; and Grossheim, Brian: Manufacturing

Development and Requirements for Stitched/RTM Wing Structure.

Proceedings of the Fourth NASA/DoD Advanced Composites Technology

Conference, NASA CP 3229, pp. 503-523, 1993.

3. Hinrichs, S. C.: ICAPS Stub Box Structural Analysis Volume 1: Material

Properties. McDonnell Douglas Aerospace Transport Aircraft-West, Report

Number MDC94K9101, March, 1995.

4. Wang, John: Global and Local Stress Analyses of McDonnell Douglas

Stitched/RFI Composite Wing Stub Box. NASA TM 110171, October 1995.

5. Brogan, F. A.; Rankin, C. C.; and Cabiness, H. D.: STAGS User Manual.

Lockheed Palo Alto Research Laboratory, LMSC Report P032594, 1994.

6

6. Hinrichs, S. C.: ICAPS Stub Box Structural Analysis Volume 4: Upper Skin

Strength Analysis. McDonnell Douglas Aerospace Transport Aircraft-West

Report Number MDC94K9101, March, 1995.

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Form ApprovedREPORT DOCUMENTATION PAGE OMBNo. 0704-0188

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1. AGEI_CY USE ONLY (Lmave blank) 2. REPORT DATE

October 19954. TITLE ANDSUBTITLE

Analysis of Selected Compression Splice Joint Locations in aGraphite-Epoxy Transport Wing Stub Box

6. AUTH(3'RiB) ....

Dawn C. Jegley

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Langley Research CenterHampton, VA 23681-0001

9. SPONSORING/ MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

3. REPORTTYPEANDOA1"ESCOVERED

Technical Memorandum

5. FUNDING NUMBERS

WU 510-02-12-01

8. PERFORMING ORGANIZATIONREPORT NUMBER

10. SPONSORING / MONITORINGAGENCY REPORT NUMBER

NASA TM 110170

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION / AVAILABILITY STATEMENT

Unclassified.Unlimited

Subject Category 39

12b. DISTRIBUTION CODE

13. ABSTRACT (Maxlmum 200 words)

Three critical compression splice joint locations in a stitched graphite-epoxy transport wing stub box have beenanalyzed to determined their expected structural performance. The wing box is represenlative of a section of acommercial transport wing box and was designed and constructed by McDonnell Douglas Aerospace Companyas part of the NASA Advanced Composites Technology (ACT) program. The results of the finite elementanalyses of the splice joints are presented. The analysis results indicate that failure will not occur in the splicejoint regions for loads less than the Design Ultimate Load of the wing box.

_. SUBJECTTERMS

composite materialstitching

finite element analysis

17.SECURITYCLASSIFICATION

bolted joint

18. SECURITY CLASSIFICATION 19. SE_iJI_ITY' CLASSIFICATIONOFREPORT OFTHISPAGE OFABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500

15. NUMBER OF P_,GES

33

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Pte_:libed by ANSi Bid. Z39-182go- 10_