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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
NodeNumberin FEMModel
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 prediction from global model at ultimate load of 166,000 lbs.
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 prediction from global model at ultimate load of 166,000 lbs.
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
NodeNumberin FEMModel
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 prediction from global model at ultimate load of 166,000 lbs.
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
NodeNumberin FEMModel
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 prediction from global model at ultimate load of 166,000 lbs.
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
NodeNumberin FEMModel
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 prediction from global model at ultimate load of 166,000 lbs.
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
NodeNumberin FEMModel
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
Table III (continued)
Strain prediction from global model at 50% of ultimate load.
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
Table III (continued)
Strain prediction from global model at 50% of ultimate load.
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
NodeNumberin FEMModel
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
NodeNumberin FEMModel
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 prediction from global model at 50% of ultimate load.
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
109 0.0 ext Lwr.Skin 1670.2 1558111 0.0 ext Lwr.Skin 1450.6 1574
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
118 0.0 ext Lwr.Skin 1157.6 1784119 45.0 ext Lwr.Skin 290.1 1784
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
Table IV (continued)
Strain prediction from global model at 50% of ultimate load.
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
Table IV (continued)
Strain prediction from global model at 50% of ultimate load.
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
NodeNumberin FEMModel
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 1524.6
0.0 ext Lwr.Skin 1469.5
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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