Global and Local Stress Analyses of McDonnell Douglas Stitched … · 2013-08-30 · Global and...

NASA Technical Memorandum 110171

Global and Local Stress Analyses ofMcDonnell Douglas Stitched / RFIComposite Wing Stub Box

John T. Wang

Langley Research Center, Hampton, Virginia

March 1996

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-0001

https://ntrs.nasa.gov/search.jsp?R=19960020473 2020-04-12T00:34:51+00:00Z

Global and Local Stress Analyses of McDonnell DouglasStitched/RFI Composite Wing Stub Box

John T. WangNASA Langley Research Center

Hampton, Virginia

Abstract

This report presents the results of a pretest structural analysis of anall-composite stitched/RFI (resin film infusion) wing stub box which wasdesigned and fabricated by the McDonnell Douglas Aerospace Company.Geometrically nonlinear structural responses of the wing stub box werepredicted by using the finite element analyses and a global/local approachin which the global model contains the entire test article while the localmodel contains a large nonlinearly deformed region in the upper cover ofthe wing stub box. The wing box test article includes the all-compositewing stub box, a metallic load-transition box and a metallic wing-tipextension box. The two metallic boxes were connected to the inboard and

outboard ends of the composite wing stub box, respectively. The metallicload-transition box was attached to a steel and concrete vertical reaction

structure. In the global analysis, an upward load was applied at the tip ofthe extension box to induce bending of the wing stub. Global analysis found

that an upper cover region, which contains three stringer runouts, exhibitslarge nonlinear deformations. Hence, a local model refined in thenonlinearly deformed region was created to predict more accurate strainresults near stringer runouts. Numerous global and local analysis resultssuch as deformed shapes, displacements at selected locations, and strains atcritical locations are included in this report.

Introduction

The purpose of this report is to document the pretest structuralanalysis results for an advanced stitched/RFI (resin film infusion)graphite/epoxy wing stub box. This advanced stitched/RFI graphite/epoxywing stub box, representing the inboard portion of a civil-transport-aircraft high-aspect-ratio wing stub box, was designed and manufacturedby McDonnell Douglas Aerospace (MDA) Company under the support ofNASA's Advanced Composites Technology (ACT) Program. Theinnovative stitched/RFI process used to fabricate this all-composite wingstub box has the potential for reducing manufacturing costs while

producing damage tolerant composite aircraft primary structures.

A preliminary global analysis was performed by McDonnell Douglasengineers[l]. More refined nonlinear analyses of the global model and alocal model were performed by NASA Langley Research Center engineersto support the structural wing-stub-box tests at NASA Langley. The globalmodel contains the entire test article which includes an all-composite stubbox, an inboard metallic load transition box and an outboard metallic wing-tip extension box. The local model contains an upper cover regionsurrounding the location of three stringer runouts (where a stiffenerterminates at a rib). It was found that this region exhibits large nonlineardeformations. Hence, a more refined mesh was used in the local model toobtain more accurate stress analysis results. Displacements obtained by theglobal model analysis were applied to the boundaries of the local model.Deformed shapes, displacements at selected locations, and strains in criticalregions generated from the global and local analyses were used forinstrumenting the wing stub box. Only the analysis results are presented inthis report. These analytical results have been correlated with test resultsand the correlation will be presented in another paper [2].

Wing Stub Box Test Article

A photograph of the wing stub box test article is shown in Figure 1and the dimensions of the wing stub box are provided in Figure 2. Thecomposite wing stub box is about 12 ft long and 8 ft wide, and itsmaximum depth at the root is about 2.3 ft. The wing stub box test articleincludes the all-composite stub box, the inboard transition box (made ofsteel and aluminum), and the outboard extension box (made of steel). The

all-composite stub box weighs approximately 1200 lbs while the entirewing stub box test article weighs about 7600 lbs. In its test configuration,the transition box was attached to a steel and concrete strongback (test wall)at the NASA Langley Structural Mechanics Test Laboratory. The extensionbox is a load introduction structure and the ultimate design load, a verticalload at the tip of the extension-box front spar, is 166,000 lbs.

The upper-cover-panel construction of the all-composite wing stubbox shown in Figure 3 contains a skin panel, ten blade stiffeners, five ribs,two metal angles, two spar caps, and has a 22.5 in. by 12.5 in. oval accessdoor cutout. The upper cover skin, flanges, and blade stiffeners wereconstructed from AS4/3501-6 graphite/epoxy composite with repeatingsublaminates (stacks). The layup sequence of a stack is [45/-45/0/90/0/-45/45] and the nominal total thickness of a stack after curing is 0.058 in.The upper skin thickness varies from five stacks (0.29 in.) to ten stacks(0.58 in.) as shown in Figure 4. All the blade stiffeners of the upper cover

2

have the same thickness and each has eight stacks (0.464 in.). The upper-cover panel was fabricated by first assembling and stitching all the drypreform components together and then using a resin film infusion processto infuse resin into the preform. Each dry preform component such as theskin or the stiffener was stitched individually before assembling the upper-cover-panel. The stiffener flanges are stitched to the skin so that nomechanical fasteners are required. This reduces the manufacturing costs.However, at the stiffener runout locations, fasteners were installed after theRFI process to prevent skin stiffener debonding at these locations.Moreover, the access door cutout was reinforced by a composite land ringwhich was bolted to the panel skin using 0.3125 inch diameter bolts. Tofurther examine the region between Ribs 6 and 8, which contains threestringer runouts and two metal angles as shown in Figure 3, a local refinednonlinear analysis was carried out.

The interior of the wing stub box is shown in Figure 5 in which thefive ribs and the two spar webs are made of conventional AS4/3501-6prepreg material. The spar webs have a constant thickness of 0.31 in. andthe rib webs have a constant thickness of 0.15 in. Moreover, the ribs andspars are stiffened with composite stiffeners to prevent buckling. Thelower skin is made of stitched/RFI graphite/epoxy material with IM7 fibersin the 0-degree fiber direction and AS4 fibers in all the other fiberdirections. The thickness of the lower skin, (which is thicker than theupper skin), ranges from .33 in. to .82 in. The high moduli of IM7 fibersand thicker skin result in smaller strains for the lower skin; therefore nofailure is expected in the lower skin region and the results in this reportrelate mainly to the upper cover panel.

Material Properties and Allowables

The equivalent AS4/3501-6 material properties for the upper skinpanel laminates used in the analyses are

E x = 8.17 Msi

Ey = 4.46 Msi

G, = 2.35 Msi

v,_ = 0.459

and the equivalent AS4/IM7/3501-6 material properties for the lower skinpanel laminates used in the analyses are

Ex = 11.85 Msi

Ey = 4.55 Msi

G. = 2.57 Msi

v. = 0.409

Note that the x-direction, which is parallel to the rear spar, is coincident

with the 0-degree fiber direction and the y-direction is coincident with the

90-degree fiber direction. The ultimate compression strain allowable in

the x-direction of the undamaged upper skin laminate is 9330pe and the

0.3125 inch diameter filled hole B-base compression strain allowable is

8100/_¢ [1].

Global and Local Finite Element Models

The global finite element model shown in Figure 6 was used to

determine the global responses of the complete test article and this global

model contains 4408 quadrilateral elements (CQUAD4), 99 triangular

elements (CTRIA3), 1308 beam elements (CBEAM), 798 rod elements

(CONROD) and 742 rigid bar elements (RBAR). The finite element

analyses were performed using MSC/NASTRAN [3], V68. The upper skin,

lower skin, rib webs and spar webs were modeled as plate elements

(CQUAD4 and CTRIA3), and the stiffeners were modeled as beam

elements (CBEAM). A total of 5,266 grid points were used in the globalmodel.

A finer finite element mesh was used around the access door cutout

to better define the high stress concentration in that region. Grid points of

each composite stiffener were positioned along its centroidal axis.

Therefore, the stiffener elements, modeled as beam elements, were offset

from the skin. These stiffener beam elements were rigidly connected to the

skin by relatively stiff CBEAM elements. The finite element mesh of the

wing stub box interior region is shown in Figure 7 in which all the rib

webs including cutout holes were modeled. Both buckling analysis and

geometrically nonlinear analysis of the global model were performed by

using MSC/NASTRAN and the solution sequences used for buckling

analysis and nonlinear static analysis were SOL 105 and SOL 106,

respectively.

As mentioned earlier, the local finite element model contains an

upper cover region surrounding three stringer runouts as shown in Figure

8. The location of the local model on the stub box is clearly shown in

Figure 9 in which the local model is superposed on the global model.

Structural details near a typical stiffener runout region are shown in Figure

4

10. Note that the tapering of the blade stiffener flange and the blade webin the runout region and a small gap existing between the stiffener flangeand rib flange were also modeled in the local model. The blade stiffenerweb is attached to the rib web, but it is not connected to the rib flange. Thelocal finite element model shown in Figure 8 has 2642 quadrilateral shellelements, 8 triangular elements and 2738 nodes. Displacement boundaryconditions of the local model were obtained through spline interpolation ofthe global displacement results. NASA Langley's COMET (ComputationalMechanics Testbed) finite element code [4] was used for the nonlinearanalysis of this local model.

The global model was originally created by engineers at theMcDonnell Douglas Aerospace Company and some modifications wereperformed by engineers at NASA Langley. PATRAN [5] was used tomodify the global model and was also used to create the local model.Results from both the global and local analyses were postprocessed usingPATRAN.

Analysis Results

Global analysis results

A buckling analysis of the wing-stub-box global model was

performed first and the critical buckling load was found to be 11.2 %

higher than the ultimate design load of 166,000 lbs with the buckling mode

shape shown in Figure 11. The second analysis performed was a

geometrically nonlinear NASTRAN analysis and the predicted deformed

shape of the whole test article at ultimate load is shown in Figure 12. The

jack loading displacements predicted at ultimate load were u= -0.2169 in,

v=0.1629 in, and w=13.4887 in. The predicted relationship between the

load and the vertical displacement at the loading point is approximately

linear as shown in Figure 13. The close-up view of the deformed shape of

the composite wing stub box is shown in Figure 14. Note that the wavy

deformation of the upper skin is likely caused by a lack of longitudinal

stiffener support near the access door cutout and its two adjacent outboard

bays as shown in Figure 3. Figure 15 illustrates the out-of-plane (z-

direction) deflections in the deformed region along two lines, labeled A-A

and B-B, parallel to the major axis of the elliptic cutout. Line B-B, located

at 15.875 in. from Line A-A, is in location which sees less wavy

deformation and the difference of the out-of-plane displacements between

these two lines provides a good measurement of the nonlinearity

developing in the two bays outboard of the access door.

5

The predicted out-of-plane displacements at six LVDT (LinearVoltage Displacement Transducer) locations on the lower skin are plottedin Figure 16. Locations 1 to 3 are on the rear spar while Locations 4 to 6are on the front spar (see insert in Figure 16 and the Douglas AircraftCompany Tasks Assignment Drawings TAD Z7944503). All the out-of-plane displacements are seen to be linearly related to the applied load.

All the strain results presented for the global model are based on theNASTRAN element coordinate system. The element x-axis is coincidentwith the global x-axis except in the access door cutout region where theelement x-axis is tangent to the edge of the cutout (in the circumferentialdirection). Strain contour plots (exx) of the upper composite skin panel

are plotted in Figures 17-19. High strains occur at the edge of the accessdoor cutout. Close-up strain contour plots for the skin side, mid-surface

and stiffener side of the upper cover in the cutout region are in Figures 20to 22. High strains are predicted in the circumferential direction of thecutout: 9100/.re at the skin side, 6650/.te at the stiffener side, and 6770/.teat the midsurface.

Predicted strains at all strain gage locations for 100% and 50% ofthe ultimate load are listed in Tables I to IV in which the strain gagelocations and the completed term of every abbreviation can be found inDouglas Aircraft Company Tasks Assignment Drawings TAD Z7944503.

Note that the strain gage patterns for the upper and lower skins are shownin Figures 23 and 24.

Strain versus load plots are generated for some strain gages locatedin high strain regions including these gages near the access door cutout(gages 78, 79, and 612) and in the highly nonlinear deformed regions(gages 63, 64, 67, and 68). Locations of these gages are shown in Figure25 and predicted strain versus load plots at these gage locations can befound in Figures 26 to 30.

Local analysis results

A geometrically nonlinear analysis was performed with the localmodel using the COMET code and the displacement boundary conditionsobtained from the nonlinear global analyses. These displacement boundaryconditions were applied on the four edges of the local model and the rootof the middle rib. The predicted deformed shape of the local model at theultimate loading condition is shown in Figure 31. Note that the skin isdeformed to the stiffener side which may be due to the lack of longitudinal

6

stiffener support in one of the skin bays. The skin strain contour plots atthe ultimate load, as derived from the nonlinear analysis, are plotted inFigures 32 to 34 at the skin side, stiffener side, and mid-surface. Thestrains in the skin side are much higher than in the stiffener side due to thesignificant skin bending deformation shown in Figure 31.

Figures 35 to 37 display the variation of the strain in the x-directionwith load for the three gaps between the stiffener flange and rib flange,(see Figure 10), located at the stiffener runouts, (labeled #1 to #3 in Figure8), for the skin side, mid-surface, and stiffener side locations. Predictedlateral strains (in y-direction) in the rib webs are plotted in Figures 38-40.Significant rib-web bending was predicted at the runouts where thestiffener webs attach to the ribs.

Predicted strains at all strain gage locations in the local model regionfor 100% and 50% of the ultimate load are listed in Tables V and VI. Notethat the strain gage locations can be found from Figure 23.

The existence of the stiffener runouts and the large skin bendingdeformations may induce the stiffener-skin separation loads, defined as thevertical or normal stress resultants between the blade and the skin. Theblade stiffeners in the local model are numbered from 1 to 4 as shown inFigure 8. The separation load of the front spar may not be an issuebecause the front spar cap is mechanically fastened to the skin, thus onlythe separation load results for blade stiffeners are presented and plotted inFigures 41 to 44. The horizontal axes of these plots are the distancemeasured from an edge of the local model in the positive x-direction (seeinsert of each plot). These results can be compared with blade stiffenerseparation allowables [6]. The ultimate failure load of blade stiffenerseparation tests performed in Reference 6 is about 2500 lbs/in which ismuch higher than the predicted separation loads here. Therefore, nostiffener-skin separation failure is expected.

Concluding Remarks

This report documents the global and local geometrically nonlinearfinite element analysis results, generated by engineers at NASA Langley, ofthe McDonnell Douglas Stitched/RFI wing stub box. These analyses wereperformed by the Computational Structures Branch of NASA LangleyResearch Center. Results have been used for strength prediction of the stubbox and determination of the instrumentation and gage locations of the stubbox test article. When the actual structural tests are performed on the stubbox, analysis results will be correlated with the test data.

7

Acknowledgment

The author wishes to acknowledge the analytical support provided by

Thiagaraja Krishnamurthy, Brian Mason and Tina Lotts of AnalyticalServices and Materials Inc.

References

1. Hinrichs, S. C., "ICAPS Stub Box Structural Analysis," Vols. I to VII.,MDC94K9101, Preliminary, Feb. 6, 1995

2. Wang, J. T., Jegley, D. C., Bush, H. G., and Hinrichs, S. C.,"Correlation of Structural Analysis and Test Results for the McDonnellDouglas Stitched/RFI All Composite Wing Stub Box," to be presented inthe 1 lth DoD/NASA/FAA Conference on Fibrous Composites inStructural Design, Fort Worth, TX, Aug. 26-29, 1996.

3. Anon., MSC/NASTRAN Reference Manual, Version 68, The MacNeal-

Schwendler Corporation, April. 1994.

4. "The Computational Structural Mechanics Testbed User's Manual,"NASA TM-100644, Oct. 1989.

5. Anon., PATRAN Plus User Manual, Release 2.4, PDA Engineering,Publication Number 2191023, Sept. 1989.

6. Hinrichs, S. C., "Stitched Stringer Tension Pull-Off Subcomponent TestReport," MDC94K9103, Jan. 5, 1995.

Table I. Strain prediction from global model at ultimate load of 166,000 lbs(including all gages on external upper skin).

Strain

Gage No

2

4

5

Angle

0.0

6 45.0

7 90.O

11

12

15

16

0.0

Surface

extext

ext

ext

ext

ext0.0 ext

0.0

0.0

17 0.0

18 0.0

0.0

0.0

0.0

0.0

20

22

25

26

27 45.0

28 90.0

32 0.037 0.0

39 0.0

46 0.0

46 0.0

48 -13.0

extext

ext

ext

ext

ext

ext

ext

ext

ext

ext

extext

ext

ext

ext

Location

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

MicroStrain

-3465.4

-3758.9-5032.7

-1761.4

1821.7

-4978.8

-4978.8-4556.0

-4556.0

-4224.8

-4842.3

-4953.4

-5194.9

-4210.0

-4433.7

-805.2

1395.3

-4085.5

-4750.7-4609.1

-4924.3

-4924.3-4522.1

NodeNumberin FEMModel

196

194

65

65

65

1359

1359

614614

71

414

437

371

101

403

403

403

401

474523

885

885

47

9

Table I (Continued)

Strain prediction from global model at ultimate load of 166,000 lbs.

Strain

Gage NoAngle Surface

50 0.0 ext51 0.0 ext

52 45.0 ext

53 90.0 ext

58 0.0 ext

62 0.0 ext

63 0.0 ext

67 0.0 ext

65 0.0 ext

67 0.0 ext

69 0.0 int

70 45 int

71 90 int

72 0.0 ext

74 - 13.0 ext75 0.0 ext

-13.0

0.0

77

78extext

79 0.0 ext

81 0.0 ext

84 0.0 ext

92 0.0 ext

93 0.0 ext

94 0.0 ext

95 0.0 ext

96 0.0 ext

97 0.0 ext

Location MicroStrain

Upr.Skin -4465.8

Upr.Skin -4430.3Upr.Skin -543.4

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinInt.Coast

1679.8-3050.7

-3581.4

-4809.6

-1910.8

-4257.0

-1910.8

1201.7

Int.Coast 535.5

Int.Coast -681.0

Upr.Skin -3473.1

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

-2434.0-6015.1

-1324.6-8430.2

-7868.0

-6647.8


518

877

877

877125

872

1328121

864

121

Q3098/99

Q3098/99

Q3098/9968

107117

221

4325

62

-5017.2 63-5082.1 1234

-5485.3 1292

-2647.8 614

-2647.8 614

-4978.8

-4978.8

1359

1359

10

Table I (Continued)


Strain

Gage NoAngle Surface Location Micro

StrainNode

Numberin FEM

Model

102 0.0 int Lwr.Skin 2231.2 1467123 0.0 int Lwr.Skin 1957.8 2244

125 0.0 int Lwr.Skin 2475.8 2260128 0.0 int Lwr.Skin 2616.2 2276

136 0.0 int Lwr.Skin 2493.4 1371

140 0.0 int Lwr.Skin 3138.0 1452

144 0.0 ext Lwr.Skin 3184.9 1605

-45.0201 aft

205 -45.0 aft

209 -45.0 aft

302 0.0 aft

303 45.0 aft

304 90.0 aft306 0.0 aft

307 45.0 aft

308 90.0 aft

310 0.0 aft311 45.0 aft

312 90.0 aft

444 45.0

90.0

0.0

45.0

90.0-45.0

445

448

inb

inb

inb

inbinb

inb

454

Fwd.spar

Fwd.spar

Fwd.spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.SparRIB .4

RIB .4

RIB .4

RIB .5

RIB .5Rib.6

455

-280.4 3007

-139.7 2955

-664.6 2891

-1102.3 3324

27.0 3324

352.9 3324

-779.5 3307

136.2 3307

348.4 3307

-194.1 3292

1044.3 3292-383.5 3292

-380.3

-114.1

986.5

530.0

56.2

464 168.7

465 90.0 inb Rib.6 50.6

3703

3651

3674

3918

3823

39603961

11

Table I (Continued)


Strain

Gage NoAngle Surface Location Micro

StrainNode

Numberin FEMModel

468 0.0 inb Rib.6 -318.3 4167

474 -45.0 inb Rib.7 - 173.0 4348

475 90.0 inb Rib.7 92.6 4173

478 0.0 inb Rib.7 - 195.7 4250

484 -45.0 inb Rib.8 - 180.8 4475

485 90.0 inb Rib.8 82.1 4382

488 0.0 inb Rib.8 549.8 4507501 0.0 fwd Ext.Box 138.6 5104

502 -45.0 fwd Ext.Box -2483.7 5104

503 90.0 fwd Ext.Box -316.5 5104

0.00.0

0.0

0.0

507 int

ext

ext

ext

511

513

Ext.Box

Tr.Str.Up

Tr.Str.Up

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.skin

Upr.skin

Upr.skin

Upr.skin

602

282

55.147.3

-4670.7

Q41604852

4865

387-4924.9 421

-5083.4 453

-1618.8 221-3885.6 49

604 0.0 ext606 0.0 ext

610 -13.0 ext611 0.0 ext

612 0.0 ext

613 0.0 ext

614 0.0 ext

615 0.0 ext

616 0.0 ext617 0.0 ext

-8596.O

-2281.1

39

120

-2489.0 69

-4350.3 444

-4551.4 449-4924.3 885

12

Table II. Strain prediction from global model at ultimate load of166,000 lbs (including all gages on internal upper skin).

StrainGage No

Angle

0.0

Surface

int

3 0.0 int9

10

13

14

30

31

33

34

35

36

38

40

4142

43

44

47

64

6880

87

91

101

103

0.0 int

0.0 int

0.0 int

0.0 int90.0 int

0.0 int

0.0 int

45.0 int

90.0 int

0.0 int

0.0 int

0.0 int

45.0 int

90.0 int

0.0 int

0.0 int

-103.0 int

0.0 int

0.0 int

0.0 mt

0.0 int0.0

0.0

0.0

45.0104

105 90.0

int

ext

ext

extext

Location

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.SkinLwr.Skin

Lwr.S_nLwr.S_n

Lwr.SMn

MicroStrain

-4015.2

-4488.6

-4780.7


196

194

1359-4780.7 1359

-4389.4 614

-4389.4 6141607.8 399

-4032.7 401

400.9 4336

520.7 4336

708.9 4336

-4242.4 474

-4326.3 523

500.4 4388

279.8 4388

595.0 4388

-4986.1 885

-2526.6 885

1627.0 493

-2639.8 1328

-3500.6 121

-5589.5 25

-5071.2 1292-5176.1 1234

2582.9

2803.8

496.4

-1096.2

1496

1550

1550

1550

13

Table II (Continued)


Strain

Gage No

107

109

Angle

0.0

0.0

Surface

ext

ext111 0.0 ext

112 45.0 ext

113 90.0 ext

115 0.0 ext

116 0.0 ext

118 0.0 ext

119 45.0 ext

120 90.0 ext

13.0

13.0

122

124ext

ext126 0.0 ext

0.0129

Location

!

Lwr.SkinLwr.SNn

MicroStrain

2640.0

3342.7


1667

1558Lwr.Skin 2898.7 1574

Lwr.Skin 975.5 1574

Lwr.Skin -1076.0 1574

Lwr.Skin 3001.5 1446

Lwr.Skin 2535.1 1691

Lwr.Skin 2296.7 1784

Lwr.Skin 570.7 1784

Lwr.Skin -706.9 1784

794.6

2099.3

Lwr.Skin

Lwr.Skin

1385

1374

Lwr.Skin 2354.4 2273Lwr.Skin 2627.3 2287ext

131 0.0 ext Lwr.Skin 2620.1 1808132 45.0 ext Lwr.Skin 405.3 1808

133 90.0 ext Lwr. Skin -843.0 1808

137 13.0 ext Lwr.Skin 3169.2 1371

138 0.0 ext Lwr.Skin 2936.5 2276

139 0.0 ext Lwr.Skin 3050.8 2291

0.0

-45.0

90.0

202

203204

fwd

fwdFwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

fwd

fwdfwd

206207

0.0

-673.2 3007

-370.2 3007

-574.1 2007

-541.3 2955

-107.8 2955

98.9 2955

-340.4 2891

6.1 2891

43.0 2891

-3958.6 3172

-45.0

208 90.0 fwd

210 0.0 fwd

211 -45.0 fwd212 90.0 fwd

213 0.0 fwd

0.0 fwd214 -4064.6 2671

14


Strain

Gage No

215

216

217

218

301

Angle

0.0

0.0

0.0

0.0

Surface

fwd

fwdfwd

fwd

45.0 fwd

305 45.0 fwd

309 45.0 fwd

441 0.0

442 45.0

443 9O.O

449 0.0

451 0.0

452 45.0

90.00.0

-45.0

90.0

453

461

462

463

out

Location

Fwd.sparFwd.spar

Fwd.sparFwd.spar

Rear.Spar

Rear.SparRear.Spar

RIB .4

RIB.4out

out RIB .4

out RIB .4

out

out

outout

out

out469 0.0 out

471 0.0 out

472 -45.0 out

RIB.5

RIB .5

RIB .5

Rib.6

Rib.6

Rib.6

Rib.6

MicroStrain

-3585.7


2719

-3238.2 2767-2217.4 2815

-1293.1

66.3

75.5

1023.8

267.7

-422.5

2836

33243307

3292

3703

3703-260.4 3703

998.7 3674

-185.5532.0

292.7

133.6

184.2

-108.5

-549.9

Rib.7 -303.4

Rib.7 -149.6

473 90.0 out Rib.7 211.8

479 0.0 out Rib.7 -299.5480 90.0 out Rib.7 461.0

3918

3918

481 0.0 out482 -45.0 out

483 90.0 out489 0.0

504 0.0504 0.0505 -45.0505 -45.0

506

outaft

aft

aft

aft

-158.8

3918

3960

3960

aft90.0

3960

4167

4348

4348

4348

42504334

Rib.8 -118.6 4475Rib.8

282.8968.9

178.7

0.0

-2456.7

0.0

Rib.8

Rib.8

Ext.Box

Ext.BoxExt.Box

Ext.Box

Ext.Box -319.2

4475

4475

4507

5104

5104

5104

51045104

15

Table II (Continued)


Strain

Gage No

506

510

601

Angle

90.0

0.0

0.0

Surface

aft

int

int

609 - 13.0 int

0.0141 .

Location

Ext.Box

Tr.Str.Up

Upr.SNn

Upr.SNnLwr.SNn

Micro

Strain

0.0

66.3

-4472.7

-4525.4

3338.4ext

144 0.0 ext Lwr.Skin 3288.1

146 0.0 ext Lwr.Skin 3220.3

148 0.0 ext Lwr.Skin 3136.2150 0.0 ext Lwr.Skin 3038.0

152 0.0 ext Lwr.Skin 2926.0

154 0.0 ext Lwr.Skin 2796.6

156 Lwr.Skin0.0 ext 2623.0

NodeNumberin FEM

Model

5104

4865

387

57

1452

1605

1636

16741714

1753

1800

1836

16

Table III. Srain prediction from global model at 50% of ultimate load (includingall gages on external upper skin).

Strain

Gage No

2

4

5

6

7

11

Angle

0.0

0.0

0.0

Surface

ext

extext

45.0 ext

90.0 ext

0.0 ext12 0.0 ext

15 0.0 ext16 0.0 ext

17 0.0 ext18 0.0 ext

20 0.0

22 0.0

25 0.0

26 0.0

ext

ext

ext

ext

0.0

27 45.0 ext

28 90.0 ext32

37 0.0

39 0.0

46 0.0

46 0.0

48 -13.050 0.0

51 0.0

52 45.0

53 90.0

0.058

ext

ext

ext

ext

ext

extext

ext

ext

ext

ext

Location

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_n

Upr.S_nUpr.S_n

Upr.S_n

MicroStrain

-1805.0


196

-1915.2 194

-2428.3 65

-827.9 65

915.4-2430.7

-2430.7

-2268.3

-2268.3

-2239.3

-2339.6

-2266.2-2105.2

-2113.0-2195.9

-397.5

703.7

-2019.1

65

1359

1359

614614

71

414

437

371

101

403403

403

401-2299.2 474

-2258.8-2416.9

-2416.9

523

885

885

-2242.0 47-2342.2 518

-2305.2 877

-376.4 877

738.7

-1934.7

877

125

17

Table III (continued)

Strain prediction from global model at 50% of ultimate load.

Strain

Gage No

62

Angle

0.0

Surface

ext63 0.0 ext

65 0.0 ext

67 0.0 ext

0.069 int

Location

Upr.Skin

Upr.Skin

Upr.SkinUpr.SkinInt.Coast

MicroStrain

!

-1902.2

NodeNumberin FEMModeli

872

-1984.5 1328-1869.7 864

-1518.6 121

596.8

70 45 int Int.Coast 430.6

71 90 int Int.Coast -257.50.072 ext

74 -13.0 ext

75 0.0 ext

ext0.0 ext

0.0 ext

-13.0

0.0

77

78

79

81 ext

84 0.0 ext

0.0 ext0.0 ext

0.0 ext

0.0 ext

0.0 ext

92

0.0

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.SkinLwr.Skin

Lwr.S_n0.0

int

int

93

94

95

96102

-1552.5

Q3098/99

Q3098/99

Q3098/9968

123

-1402.0 107

-2620.2 117

-810.5 221

-4011.0 43

-3589.1 25

-2889.8 62

-2308.8 63

-2545.0 1234

-2751.8 1292

-2268.3 614

-2268.3 614

-2430.7 1359

1113.9

982.1

14672244

18



Strain

Gage No

125

Angle

0.0

Surface

int

128 0.0 int

136 0.0 int

140 0.0 int

144 0.0 ext

201 -45.0 aft

205 -45.0 aft

209 -45.0 aft

303 45.0 aft

304 90.0 aft

306 0.0 aft

307 45.0 aft

308 90.0 aft

310 0.0 aft

45.0311

312 90.0

444 45.0

445 90.0

448 0.0

454 45.0

455 90.0

464 -45.0

465 90.0

468 0.0

474 -45.0

484

aft

aft

inb

inb

inb

inb

inb

Location

Lwr.Skin

Micro

Strain

1238.6

Node

Number

in FEM

Model

2260

Lwr.Skin 1302.9 2276

Lwr.Skin 1253.9 1371

Lwr.Skin 1554.2 1452

Lwr.Skin 1586.5 1605

-117.9 3007

-63.4 2955

2891.0 302

21.2 3324

176.3 3324

-398.8 3307

60.0 3307

170.3 3307

-94.9 3292

518.7 3292

-198.3 3292

Fwd.spar

Fwd.spar

Fwd.spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.Spar

Rear.SparRIB .4 -189.3 3703

RIB.4 15.0 3651

RIB.4 492.7 3674

RIB.5 289.8 3918

43.6RIB.5 3823

mb Rib.6 111.4 3960

mb Rib.6 57.2 3961

mb Rib.6 -174.2 4167

mb Rib.7 -39.2 4348

Rib.7 65.9inb 4173475 90.0

478 0.0 mb Rib.7 -101.9 4250

-45.0 mb Rib.8 -105.1 4475

inb Rib.8 61.7 4382

inb Rib.8 328.2 4507

fwd 79.5

485 90.0

488 0.0

501 0.0 Ext.Box 5104

19



Strain

Gage NoAngle Surface Location

502

503

507

511

513

602

-45.090.0

0.0

0.0

0.0

0.0

fwd

fwd

Int

ext

ext

ext

604 0.0 ext

606 0.0 ext610 -13.0 ext

611 0.0 ext

612 0.0 ext

613 0.0 ext

614 0.0 ext

615 0.0 ext

616 0.0 ext

0.0617 ext

Ext.Box

Ext.Box

Ext.Box

Tr.Str.Up

Tr.Str.Up

Upr.Skin

Upr.Skin

Upr.Skin

Upr.S_n

Upr.S_n

Upr.S_n

Upr.Skin

Upr.Skin

Upr.Skin

Upr.S_n

Upr.S_n

Micro

Strain

-1240.2-157.4

161.6

27.0

23.2

-2317.1

-2322.1

-2191.3-810.5

-1995.0

-4073.9

-1626.0

-1795.9

-2237.3

-2221.1

-2416.8


51045104

Q41604852

4865

387

421

453221

49

39

120

69

444

449

885

2o

Table IV. Strain prediction from global model at 50% of ultimate load(including all gages on internal upper skin).

Strain

Gage NoAngle

0.0

Surface

int

3 0.0 int

9 0.0 int

10 0.0 int

13 0.0 int

14 0.0 int

30 90.0 int

31 0.0 int

33 0.0 int

34 45.0 int

35 90.0 int

36 0.0 int38 0.0 int

40 0.0 int

41 45.0 int

42 90.0 int

43 0.0 int

44 0.0 int

47 - 103.0 int

64 0.0 int68 0.0 int

80 0.0 int

87 0.0 int

91 0.0 int

0.0101

Location

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.Skin

Upr.Skin

Upr.SkinUpr.Skin

Upr.Skin

Upr.S_n

Upr.S_n

Upr.S_nLwr.S_n

MicroStrain

-2001.4


196

-2206.4 194

-2358.6 1359

-2358.6 1359

-2172.0 614

-2172.0 614

-2119.2 399

-2014.1 401

137.0 4336

154.1 4336

259.7 4336

-2134.3 474

-2185.8 523

245.7 4388

109.7298.1

4388

4388

-2535.9 885

-2535.9 885890.6

-1761.1

-1613.7-2970.6

-2547.4

4931328

121

25

1292

-2570.4 12341291.5ext 1496

103 0.0 ext Lwr.Skin 1402.1 155045.0 247.8ext104 Lwr.Skin 1550

21

Table IV (continued)


Strain

Gage No

105

Angle

90.0

Surface

ext

Location

Lwr.Skin

MicroStrain

-540.1

Node

Numberin FEMModel

1550

107 0.0 ext Lwr.Skin 1330.0 1667


112 45.0 ext Lwr.Skin 490.9 1574

113 90.0 ext Lwr.Skin -522.7 1574

115 0.0 ext Lwr.Skin 1486.0 1446

116 0.0 ext Lwr.Skin 1278.0 1691


120 90.0 ext Lwr.Skin -361.1 1784

122 13.0 ext Lwr.Skin 312.2 1385

124 13.0 ext Lwr. Skin 694.3 1374

126 0.0 ext Lwr.Skin 1181.3 2273

129 0.0 ext Lwr.Skin 1315.3 2287

131 0.0 ext Lwr.Skin 1323.5 1808

132 45.0 ext Lwr.Skin 217.2 1808

133 90.0 ext Lwr.Skin -416.9 1808137 13.0 ext Lwr.Skin 1057.0 1371

138 0.0 ext Lwr.Skin 1469.8 2276

139 0.0 ext Lwr.Skin 1533.8 2291

fwd202 0.0

203 -45.0 fwd

204 90.0 fwd

206 0.0 fwd

207 -45.0 fwd

208 90.0 fwd

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.sparfwd0.0

-333.0 3007

-165.7 3007

-290.8 2007

-253.6 2955

-66.6 2955

44.5 2955210 -158.0 2891

22



Strain

Gage No

211

Angle

-45.0

Surface

fwd

Location

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

Fwd.spar

212 90.0 fwd

213 0.0 fwd

214 0.0 fwd

215 0.0 fwd

216 0.0 fwd

217 0.0 fwd

218 0.0 fwd

301 45.0 fwd Rear.Spar

305 45.0 fwd Rear.Spar

309 45.0 fwd Rear.Spar441 0.0 out RIB.4

442 45.0 out RIB .4

443 90.0 out RIB.4

449 0.0 out RIB.4

451 0.0 out RIB.5

452 45.0 out RIB.5

453 90.0 out RIB.5

461 0.0 out Rib.6

462 -45.0 out Rib.6

463 90.0 out Rib.6

469 0.0 out Rib.6

471 0.0 out Rib.7

472 -45.0 out Rib.7

473 90.0 out Rib.7

479 0.0 out Rib.7

480 90.0 out Rib.7

481 0.0 out Rib.8

MicroStrain

17.9

Node

Number

in FEM

Model

2891

15.4 2891

-2038.8 3172

-2017.9 2671

-1712.9 2719

-1526.5 2767

-1087.3 2815

-618.9 2836

34.1 3324

54.8 3307

521.4 3292

125.3 3703

-196.2 3703

-32.2 3703

515.8 3674

-89.3 3918

294.0 3918

170.6 3918

84.5 3960

111.4 3960

-39.4 396O

-300.4 4167

-77.0 4348

-34.7 4348

106.8 4348

-161.3 4250

206.8 4334

-57.2 4475

23



Strain

Gage No

482

483

489

504

504

5O55O5

5O6506

510

601

609141

144

146

148

150

152

154156

Angle

-45.0

90.0

0.0

0.0

0.0

-45.0

Surface

out

outout

aft

aft

aft

Location

Rib.8

Rib.8

Rib.8

Ext.Box

Ext.Box

Ext.Box

MicroStrain

-95.9

154.2

569.8

86.1

0.0

-1233.4


4475

4475

4507

5104

51045104

-45.0

90.0

aft

aft

Ext.Box

Ext.Box

0.0

-158.9

90.0 aft Ext.Box 0.0

0.0 Tr.Str.Up0.0

int

int Upr.Skin

Upr.Skin-13.0 int

32.6

-2222.2

-2495.9

5104

5104

5104

4865

387

570.0 Lwr.Skin 1654.5ext

0.0 ext Lwr.Skin 1648.80.0 ext Lwr.Skin 1615.9

0.0 ext Lwr.Skin 1573.7



0.0 ext Lwr.Skin 1405.90.0 Lwr.Skinext 1315.9

1452

1605

16361674

1714

1753

1800

1836

24

Table V Strain estimation from local model at ultimate loadof 166,000 lbs.

Strain

Gage No2223243334

r

35

43

44

45

46

48

49

50

51

52

53

54

55

56

57

5859

6O

61

Angle Surface

0 Ext

0 IntINB

0 IntOUT

0 Int

0 Int

0 Int

0 Int

0 Int

0 Int

0 Ext

-13 Ext

0

MicroStrain

Node

-4876.50 18

-147.00 E2000

77.50 E2000

977.50 E2415

-889.50 E2415

-983.00 E2415

-4366.50 1184

-5001.50 1203-4743.25 1214

-6061.00 1212

-5105.34 239

INT -2271.07 506

0 Ext -4257.25 504

0 Ext -4248.00 40245 Ext

90 Ext

0 INT

0 Int

0 Int

0 Int

0 Ext0 Int

0 Int

-2437.13

1269.75

-2027.92-5002.50

402

0

402

5O6

2643

-6058.50 2613

-5410.00 751-3949.00 498

-3914.75 1886

-4370.50 676

Int -4668.25 1891

62 0 Ext -3997.00 679

63 0 Ext -4312.00 15120 Int

Ext0-3233.25 1512-4410.50 1296

0 IntlNB -930.00 E2130

0 InflNB 1699.00

64

65

66

607

608

609

617

E2070

0 Ext

-13 Int

0 Ext

-4680.00 468

-3072.17 242

-5320.50 2148

25

Table VI Strain estimation from local model at 83,000 lbs (50%Loading)

Strain Angle Surface Micro NodeGage No Strain

22 0 Ext -2533.00 18

23 0 IntlNB -22.50 E2000

24 0 IntOUT -62.50 E2000

33 0 Int 457.00 E2415

34 0 Int -442.50 E2415

35 0 Int -480.00 E2415

43 0 Int -2454.75 1184

44 0 Int -2784.50 1203

45 0 Int -2585.00 1214

46 0 Ext -3292.00 121248 -13 Ext -2807.77 239

49 0 INT -1231.50 506

50 0 Ext -2275.25 504

51 0 Ext -2277.25 402

52 45 Ext -1212.63 4025'3 90 Ext 694.25 402

54 0 INT -1436.50 506

55 0 Int -2670.00 2643

56 0 Int -3227.00 2613

57 0 Int -2928.00 751

58 0 Ext -2248.50 498

59 0 Int -2150.25 1886

60 0 Int -2553.50 676

61 0 Int -2434.25 189162

6300

ExtExt

-2276.00i-2271.25-1999.501

679

151264 0 Int 1512

65 0 Ext -2148.50 129666 0 IntlNB -867.50 E2130

607 0 In_NB -332.00 E2070

608 0 Ext -2246.25 468

609 -13 Int -1665.24 242

617 0 Ext -2867.50 2148

26

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REPORT DOCUMENTATION PAGE I Form ApprovedOMB No. 0704-0188

,._,L,_,_ _ __=_n=_.jnauer_.._j____p._or rec_g t_= Du_. to wa¢'_ Hea0quaneaSen,_. D_c¢_= e _or_om_on O_,=,o_ _ _=. 121SJet.on D_"Pqjmqw. oum ]acuA. Amngtan, VA Z_2-Aucr_. wta to me _11=¢eo_ M_I and B_II_, P_k FleduoJon Protect I0704-0188), Wal.hi_. DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

March 1996 Technical Memorandum4. TITLE AND SUBTITLE : S. FUNDING NUMBERS

Global and Local Stress Analyses of McDonnell Douglas Stitched/RFI 510-02-12-04Composite Wing Stub Box

6. AUTHOR(S)

John T. Wang

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Langley Research CenterHampton, VA 23681-0001

g. SPONSORING / MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronauticsand Space AdministrationWashington, DC 20546-0001

11. SUPPLEMENTARY NOTES

8. PERFORMING ORGANIZATIONREPORT NUMBER

10. SPONSORING IMONrrORINGAGENCY REPORT NUMBER

NASA TM-110171

12a. DISTRIBUTION I AVAILABILITY STATEMENT

Unclassified - UnlimitedSubject Category 39

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 _j

This report contains results of structural analyses performed in support of the NASA structural testingof an all-composite stitched/RTI (resin film infusion) wing stub box. McDonnell Douglas AerospaceCompany designed and fabricated the wing stub box. The analyses used a global/local approach. Theglobal model contains the entire test article. It includes the all-composite stub box, a metallicload-transition box and a metallic wing-tip extension box. The two metallic boxes are connected to theinboard and outboard ends of the composite wing stub box, respectively. The load-transition box wasattached to a steel and concrete vertical reaction structure and a load was applied at the tip of theextension box to bend the wing stub box upward. The local model contains an upper cover regionsurrounding three stringer runouts. In that region, a large nonlinear deformation was identified by theglobal analyses. A more detailed mesh was used for the local model to obtain more accurate analysisresults near stringer runouts. Numerous analysis results such as deformed shapes, displacements atselected locations, and strains at critical locations are included in this report.

14.SUBdF.CTTERMS

Stitched Composites, Resin Film Infusion,Wing Box Analysis,Finite ElementAnalysis, Nonlinear Analysis

17. _JECURITY CLASSlRCATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATIONOF THIS PAGE

Unclassified

lg. SECURITY CLASSIFICATIONOF ABSTRACT

Unclassified

15. NUMBER OF PAGES

71

16. PRICE CODE

A04

20. LIMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2.89)Prlmcrl_d by ANSI Std. Z39-18298-1(]2

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