February 2014 

NASA/TM–2014-218169 

Impact Testing and Simulation of Composite 

Airframe Structures 

Karen E. Jackson, Justin D. Littell, Lucas G. Horta, and Martin S. Annett 

Langley Research Center, Hampton, Virginia 

Edwin L. Fasanella 

National Institute of Aerospace, Hampton, Virginia 

Michael D. Seal II 

Analytical Mechanics Associates, Inc., Hampton, Virginia 
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February 2014 

NASA/TM–2014-218169 

Impact Testing and Simulation of Composite 

Airframe Structures 

Karen E. Jackson, Justin D. Littell, Lucas G. Horta, and Martin S. Annett 

Langley Research Center, Hampton, Virginia 

Edwin L. Fasanella 

National Institute of Aerospace, Hampton, Virginia 

Michael D. Seal II 

Analytical Mechanics Associates, Inc., Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not constitute an 

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and Space Administration. 
i

Table of Contents 

 Page 

1.0 ABSTRACT 1

2.0 INTRODUCTION 2

3.0 EXPERIMENTAL TESTING 4

3.1 Specimen Descriptions 4 

3.2 Non-Destructive Evaluation 5 

3.3 Laminate Characterization Testing 5 

3.3.1 Unidirectional Tape 6 

3.3.2 Plain Weave Fabric 8 

3.4 Dynamic Crush Testing of Two I-Beam Specimens 9 

3.4.1 I-Beam Test at 4-ft. Drop Height 10 

 3.4.2 I-Beam Test at 8-ft. Drop Height 13 

3.5 Dynamic Crush Testing of Two T-Section Specimens 16 

 3.5.1 T-Section Test at 2-ft. Drop Height 17 

 3.5.2 T-Section Test at 4-ft. Drop Height 20 

3.6 Dynamic Crush Testing of Two Cruciform Sections 23 

 3.6.1 Cruciform Section Test at 4-ft. Drop Height 25 

 3.6.2 Cruciform Section Test at 8-ft. Drop Height 28 

3.7 Testing of the Subfloor Section 31 

 3.7.1 Modal Testing 32 

 3.7.2 Subfloor Impact Testing 35 

3.8 Testing of the Framed Fuselage Section 40 

 3.8.1 Modal Testing 41 

 3.8.2 Static Testing 42 

 3.8.3 Impact Testing 45 

4.0 MODEL DEVELOPMENT AND TEST-ANALYSIS COMPARISONS 52

4.1 Material Model Development and Comparison with Laminate Property Data 53 

 4.1.1 Description of Mat 54 Material Model 53 

 4.1.2 Description of Mat 58 Material Model 53 

 4.1.3 Comparison of Tensile Laminate Response with Simulation 56 

 (Unidirectional Tape) 

 4.1.4 Comparison of Compressive Laminate Response with 58 

 Simulation (Unidirectional Tape) 

 4.1.5 Comparison of Tensile Laminate Response with Simulation (Fabric) 61 

 4.1.6 Comparison of Compressive Laminate Response with Simulation (Fabric) 61 
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4.2 Simulation of Dynamic Crush Tests of Two I-Beam Specimens 63 

 4.2.1 Description of the Finite Element Model 63 

 4.2.2 Mat 54 Predicted Responses Compared with 4-ft Test Data 64 

 4.2.3 Mat 58 Predicted Responses Compared with 4-ft Test Data 65 

 4.2.4 Mat 54 Predicted Responses Compared with 8-ft Test Data 66 

 4.2.5 Mat 58 Predicted Responses Compared with 8-ft Test Data 66 

4.3 Simulation of Dynamic Crush Tests of Two T-Section Specimens 68 

 4.3.1 Description of the Finite Element Model 68 

 4.3.2 Mat 54 Predicted Responses Compared with 2-ft Test Data 69 

 4.3.3 Mat 58 Predicted Responses Compared with 2-ft Test Data 69 

 4.3.4 Mat 54 Predicted Responses Compared with 4-ft Test Data 71 

 4.3.5 Mat 58 Predicted Responses Compared with 4-ft Test Data 73 

4.4 Simulation of Dynamic Crush Tests of Two Cruciform Sections 76 

 4.4.1 Description of the Finite Element Model 76 

 4.4.2 Mat 54 Predicted Responses Compared with 4-ft Test Data 77 

 4.4.3 Mat 58 Predicted Responses Compared with 4-ft Test Data 77 

 4.4.4 Mat 54 Predicted Responses Compared with 8-ft Test Data 80 

 4.4.5 Mat 58 Predicted Responses Compared with 8-ft Test Data 81 

4.5 Simulation of Longitudinal Impact of a Subfloor Section 83 

 4.5.1 Description of the Finite Element Model 83 

 4.5.2 Mat 54 Predicted Responses Compared with Test Data 84 

 4.5.3 Mat 58 Predicted Responses Compared with Impact Test Data 86

 4.5.4 Multi-Dimensional Model Calibration 89 

4.6 Simulation of Vertical Drop Test of a Framed Fuselage Section 95 

 4.6.1 Description of the Finite Element Model 95 

 4.6.2 Analysis of Static Load Test Data 99 

4.6.3 Comparison of Modal Test Results with Analysis 100 

 4.6.4 Mat 54 Predicted Responses Compared with Impact Test Data 102 

 4.6.5 Mat 58 Predicted Responses Compared with Impact Test Data 104 

 4.6.6 Multi-Dimensional Calibration of the Mat 58 Impact Model 108 

5.0 DISCUSSION OF RESULTS 112

5.1 Experimental Program 112 

5.2 Analytical Program 113 

6.0 CONCLUDING REMARKS 117

7.0 ACKNOWLEDGEMENTS 119

8.0 REFERENCES 120

9.0 APPENDIX A 124
iii

List of Figures

 Page

Figure 1. Post-test photograph of the SARAP TVA prior to the 2008 vertical drop test. 3 

Figure 2. SARAP residual hardware utilization. 5 

Figure 3. Laminate coupon extraction locations with identified fabrication methods. 6 

Figure 4. Sample set of extracted coupon specimens. 6 

Figure 5. Tensile stress-strain curves for laminated coupons taken from three different 8 

locations. 

Figure 6. Compressive stress-strain curves of laminated floor coupons fabricated of 8 

unidirectional tape. 

Figure 7. Fabric coupon responses to tensile and compressive loading. (a) Tension. 9

(b) Compression. 

Figure 8. I-beam specimen configuration. 10 

Figure 9. Image sequence of the I-beam 4-ft drop test. 11 

Figure 10. Drop mass acceleration from the 4-ft I-beam test. 11 

Figure 11. Drop mass crush displacement from the 4-ft I-beam test. 12 

Figure 12. Compressive strain at 1-ms and 7.2-ms after impact of the I-beam. 13 

(a) Time = 1-ms. (b) Time=7.2-ms. 

Figure 13. Post-test photograph of the I-beam tested at 4-ft drop height. 13 

Figure 14. Image sequence of the 8-ft I-beam test. 14 

Figure 15. Drop mass acceleration from the 8-ft I-beam test. 15 

Figure 16. Drop mass crush displacement from the 8-ft I-beam test. 15 

Figure 17. Compressive strain at 1-ms and 5-ms after impact. (a) Time=1-ms. 16 

(b) Time=5-ms. 

Figure 18. Post-test I-beam specimen following 8-ft drop test. 16 

Figure 19. T-section test set-up. 17 

Figure 20. Image sequence of the T-section test at 2-ft drop height. 18 

Figure 21. Drop mass acceleration from the 2-ft T-section test. 18 
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Figure 22. Impact mass crush displacement from the 2-ft T-section test. 19 

Figure 23. PG results for 2-ft drop test of the T-section, 1-ms after impact. 19 

(a) Out-of-plane displacement. (b) Compressive strain. 

Figure 24. Post-test photographs of the 2-ft T-section test specimen. (a) Front view. 20 

(b) Rear view. 

Figure 25. Image sequence of the 4-ft T-section test. 21 

Figure 26. Drop mass acceleration of the 4-ft T-section test. 21 

Figure 27. Drop mass crush displacement of the 4-ft T-section test. 22 

Figure 28. Photogrammetry results for 4-ft T-section test, 0.5-ms after impact. 22 

(a) Out-of-plane displacement. (b) Compressive strain. 

Figure 29. Post-test photographs of the T-section following the 4-ft test. 23 

(a) Front view. (b) Rear view. 

Figure 30. Photograph of a cruciform section test specimen. 24 

Figure 31. Photograph of the cruciform test setup. 24 

Figure 32. Image sequence of cruciform crush during the 4-ft drop test. 25 

Figure 33. Drop mass acceleration from the 4-ft cruciform test. 26 

Figure 34. Vertical strain field at 0.4-ms and 0.65-ms after impact. 26 

(a) Time = 0.4-ms. (b) Time = 0.65-ms. 

Figure 35. Crush displacement of the 4-ft drop of a cruciform section. 27 

Figure 36. Post-test images of cruciform specimen following 4-ft drop test. 27 

Figure 37. Image sequence of cruciform crush during the 8-ft drop test. 28 

Figure 38. Drop mass acceleration from the 8-ft cruciform test. 29 

Figure 39. Crush displacement of the 8-ft cruciform test. 29 

Figure 40. Vertical full-field strain map of the cruciform during the 8-ft drop test. 30 

Figure 41. Post-test images of cruciform specimen following the 8-ft drop test. 30 

Figure 42. SARAP subfloor test article location. 31 
v

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Figure 43. SARAP subfloor test article. 31 

Figure 44. Pre-test photograph of the subfloor section highlighting strain gage locations. 32 

Figure 45. Modal test setup. 33 

Figure 46. Accelerometers mounted on subfloor test article. 33 

Figure 47. Tap excitation in vertical direction. 34 

Figure 48. Horizontal axis reciprocity check. 34 

Figure 49. Vertical axis reciprocity check. 34 

Figure 50. Subfloor longitudinal impact test setup. 35 

Figure 51. North East camera view of subfloor deformation. 36 

Figure 52. North camera view of subfloor deformation. 36 

Figure 53. Drop mass velocity responses from PG measurements of the subfloor. 37 

Figure 54. Drop mass vertical displacement of the subfloor. 38 

Figure 55. Out-of-plane displacement in the subfloor test article at 1.2- and 39 

5-ms after impact. (a) Time = 1.2-ms. (b) Time = 5-ms. 

Figure 56. Vertical strain in the subfloor test article at 1.2- and 5-ms after impact. 39 

(a) Time = 1.2-ms. (b) Time = 5-ms. 

Figure 57. Post-test photographs of the subfloor. (a) Open subfloor deformation. 40 

(b) Base (skinned) side deformation. 

Figure 58. Post-test subfloor stiffener behavior. (a) Subfloor open side failures. 40 

(b) Post-test collected T-stiffeners. 

Figure 59. Montage of composite fuselage test section undergoing modal testing. 41 

Figure 60. Modal test set-up for composite fuselage test section. 42 

Figure 61. Pre-test photograph of the fuselage section in test position (south view). 43 

Figure 62. Discrete tracked target locations used for deflection measurements 44 

during the fuselage section static test. 

Figure 63. Horizontal and vertical displacement fields on the north face of the 45 

fuselage section. (a) Horizontal. (b) Vertical. 
vi

 Page

Figure 64. Image sequence of the fuselage section drop test from southeast camera. 46 

Figure 65. Image sequence of the fuselage section drop test from the north camera. 47 

Figure 66. Close up views of floor failures that occurred during the fuselage section 48 

impact test. (a) Southwest view. (b) Northeast view. 

Figure 67. Portal frame drop mass acceleration and floor acceleration. 48 

(a) Portal frame acceleration. (b) Floor acceleration. 

Figure 68. Velocity time histories from PG of fuselage section floor and portal frame. 49 

Figure 69. Vertical strain fringe plots at 5-ms and 33-ms after fuselage section impact. 49 

(a) Time = 5-ms. (b) Time = 33-ms. 

Figure 70. Post-test photograph of the framed fuselage section test article. 50 

Figure 71. Post-test location of the front frame floor (overhead view). 51 

Figure 72. Post-test photographs showing the floor separation locations. 51 

Figure 73. Post-test photograph showing the interior of the floor. 51 

Figure 74. Typical in-plane tension stress-strain curve used for material within Mat 58. 55 

Figure 75. Finite element model used for simulating in-plane tensile loading. 57 

Figure 76. Comparison of Mat 54 and Mat 58 tensile stress-strain responses with 58 

test data. (a) Mat 54 autoclave properties. (b) Mat 58 properties. 

Figure 77. Results of two Mat 58 parameter studies on the tensile laminate response. 59 

(a) Reduced strength model. (b) ERODS study results. 

Figure 78. Finite element model used for simulating in-plane compressive loading. 59 

Figure 79. Comparison of compression stress-strain responses. 60 

(a) Mat 54 properties. (b) Mat 58 properties.

Figure 80. Results of two Mat 58 parameter studies on the compressive laminate 61 

response. (a) Reduced strength model. (b) ERODS variation study results. 

Figure 81. Comparison of Mat 54 and Mat 58 predicted responses with average 62 

tensile response of fabric laminates. (a) Mat 54. (b) Mat 58. 

Figure 82. Comparison of Mat 54 and Mat 58 predicted responses with average 62 

compression response of fabric laminate. (a) Mat 54. (b) Mat 58. 
vii

 Page

Figure 83. LS-DYNA model of the I-beam component. (a) Front view with drop mass. 63 

(b) Three-quarter view without drop mass. 

Figure 84. Experimental and Mat 54 analytical results for the I-beam crush test 64 

at 189.6-in/s. (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 85. Mat 54 model predicted deformation pattern indicating complete specimen 64 

collapse.

Figure 86. Experimental and Mat 58 analytical results for the I-beam 4-ft. drop test. 65 

(a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 87. Time sequence of I-beam model deformation for Mat 58. 65 

Figure 88. Mat 58 parameter study results for the I-beam (189.6-in/s velocity). 66 

(a) Reduced strength model. (b) ERODS study. 

Figure 89. Experimental and Mat 58 analytical results for the I-beam 8-ft. drop test. 67 

(a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 90. Time sequence of I-beam model deformation for Mat 58 simulation. 67 

Figure 91. Mat 58 parameter study results for the I-beam (265.2-in/s velocity). 68 

(a) Reduced strength model. (b) ERODS study. 

Figure 92. T-section finite element models (drop mass not shown). 69 

(a) 2-ft drop model (two fillets). (b) 4-ft drop model (four fillets). 

Figure 93. Mat 54 T-section model deformation at 0.017-s. 69 

Figure 94. Comparisons of Mat 58 model predicted responses with test data from 70 

the 2-ft drop. (a) Acceleration responses. (b) Displacement responses. 

Figure 95. Maximum deformation of the Mat 58 T-section model at time 0.004-s after 71 

impact. Elements have been deleted where delamination occurred in the test fillets. 

Figure 96. Comparison of reduced strength Mat 58 model with test for the 2-ft drop. 71 

(a) Acceleration comparisons. (b) Displacement comparisons. 

Figure 97. Mat 58 T-section model failure below the fillet for ERODS = 0.1. 72 

Figure 98. Comparison of displacement responses of the drop mass for the 2-ft 72 

T-section test with the Mat 58 models with only the ERODS parameter changed. 

Figure 99. Mat 54 T-section model at time 0.02 seconds. 72 

Figure 100. Comparisons of the 4-ft drop test data with the Mat 54 model predictions. 73 

(a) Acceleration comparisons. (b) Velocity comparisons. 
viii

 Page

Figure 101. Comparisons of Mat 58 model predicted responses with 4-ft T-section test 74 

data. (a) Acceleration comparisons. (b) Velocity comparisons. 

(c) Displacement comparisons. 

Figure 102. Figure 102. Mat 58 predicted damage of the T-section model at time 74 

0.015 seconds after impact (4-ft drop test). 

Figure 103. Comparisons of test and Mat 58 reduced strength material model. 75 

(a) Acceleration responses. (b) Velocity Responses. (c) Displacement responses. 

Figure 104. Comparison of post-test picture with Mat 58 deformed plot. (a) Model 75 

deformation. (b) Post-test photograph. 

Figure 105. Displacement-time histories of the drop mass for the 4-ft T-section test 76 

and Mat 58 models with varying ERODS parameter. 

Figure 106. Mat 58 T-section model deformation for ERODS = 0.1 and 0.2. 76 

(a) Deformed model for ERODS = 0.1 (b) Deformed model for ERODS = 0.2 

Figure 107. Pictures of the cruciform model. (a) Canted front view with drop mass. 77 

(b) View of the model without the drop mass. 

Figure 108. Experimental and Mat 54 analytical results for the cruciform 4-ft drop test. 78 

(a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 109. Mat 54 cruciform model deformation (velocity = 192.5-in/s). 78 

Figure 110. Experimental and Mat 58 analytical results for the cruciform 4-ft drop test. 79 

(a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 111. Cruciform model deformation sequence (velocity = 192.5-in/s). 79 

Figure 112. Mat 58 parameter study results for the cruciform section 80 

(192.5-in/s velocity). (a) Reduced strength model. (b) ERODS study. 

Figure 113. Experimental and Mat 54 analytical results for the cruciform 8-ft. drop test. 81 

(a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 114. Mat 54 cruciform model deformation (velocity = 268.6-in/s) 81 

Figure 115. Experimental and Mat 58 analytical results for the cruciform 8-ft. drop test. 82 

(a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 116. Mat 58 cruciform model deformation sequence (velocity = 268.6-in/s). 8 

 

Figure 117. Mat 58 parameter study results for the cruciform section 83 

(268.6-in/s velocity). (a) Reduced strength model. (b) ERODS study. 
ix

 Page 

Figure 118. Pictures of the subfloor section model. (a) Front view. (b) Rear view. 84 

(c) Side view. (d) Canted view. 

Figure 119. Comparisons of Mat 54 predicted velocity and displacement responses 85 

with test (PG). (a) Velocity. (b) Displacement. 

Figure 120. Pictures of Mat 54 subfloor model deformation. 85 

Figure 121. Comparison of experimental and Mat 58 analytical responses for the 86 

subfloor test. (a) Velocity-time history. (b) Displacement-time history. 

Figure 122. Pictures of Mat 58 subfloor model deformation. 87 

Figure 123. Mat 58 reduced strength parameter study results for the subfloor section. 87 

(a) Velocity-time history. (b) Displacement-time history. 

Figure 124. Mat 58 ERODS parameter study results for the subfloor section. 88 

(a) Velocity-time history. (b) Displacement-time history. 

Figure 125. Pictures of Mat 58 (ERODS=0.1) subfloor model deformation. 88 

Figure 126. PG targets and LS-DYNA evaluation nodes. 91 

Figure 127. Displacement bounds metric M2 from 40 LS-DYNA runs and PG test data. 92 

Figure 128. Orthogonality of impact shapes from the PG data and the Mat 58 model 93 

using the baseline parameter set. 

Figure 129. Sensitivity analysis of the subfloor model. 94 

Figure 130. Norm magnitude comparison of test versus calibrated model. 94 

Figure 131. Orthogonality of impact shapes from PG and LS-DYNA model data using 95 

calibrated parameter set. 

Figure 132. Sikorsky two-frame fuselage model with impact plane. 96 

Figure 133. Sikorsky two-frame fuselage model separated by Material Identification. 96 

Figure 134. Photographs of aluminum blocks used to properly orient the upper mass 97 

on the fuselage section. 

Figure 135. Photographs of raised fasteners and spot face reliefs. 97 

(a) Raised fasteners. (b) Spot face reliefs. 

Figure 136. Final fuselage model updated with 4 lead blocks on the floor plus a 98 

revised upper loading mass that correctly represents the upper mass and contact area 

of the test article. 
x

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Figure 137. Photograph of the lead masses added to the floor of the fuselage section. 98 

Figure 138. Fastener and nodal mass locations on the framed fuselage section. 99 

Figure 139. Mat 54 predicted displacement of the upper portal mass for a slowly 100 

applied gravity loading and side view of PG data from the static test. (a) Predicted

displacement. (b) Side view of PG data. 

Figure 140. Orthogonality of test versus LS-DYNA predicted modes. 101 

Figure 141. Orthogonality of test versus updated LS-DYNA predicted results. 102 

Figure 142. Comparisons of predicted accelaration responses of the upper loading 102 

frame mass and a floor-mounted lead block from the Mat 54 fuselage section model 

with test. (a) Upper mass. (b) Lead block. 

Figure 143. Comparison of Mat 54 predicted velocity of the upper mass with test data. 103 

(a) Data from integrated accelerometer. (b) Data from differentiated photogrammetry. 

Figure 144. Mat 54 predicted time history of the displacement of the upper mass 104 

test portal. 

 

Figure 145. Test and analysis deformations of the fuselage section at time=0.1-s. 104 

(a) Film-capture photo. (b) Mat 54 model prediction. 

Figure 146. Mat 58 (ERODS=0.06) simulation of the fuselage drop test at a time 0.1-s 105 

after impact compared with test. (a) Mat 58 model prediction. (b) Film-capture 

photograph.

Figure 147. ERODS study on the effect of upper mass displacement versus time. 105 

Figure 148. Photogrammetry data collected for target 1200, shown in the upper 106 

left corner. 

Figure 149. Comparison of upper mass displacement predicted with Mat 58 106 

(ERODS=0.075) with photogrammetry of target 1200. 

Figure 150. Test-analysis velocity comparisons. (a) Data from photogrammetry. 107 

(b) Data from integrated acceleration. 

Figure 151. Deformed plot for Mat 58 fuselage model (ERODS=0.075) at 0.1-s. 107 

Figure 152. Comparison of test accelerations with Mat 58 (ERODS=0.075) predictions. 108 

(a) Upper mass location. (b) Lead block on the floor. 

	ͳͷ͵ǤǦǤ   ͳͲͻ
xi

 Page

	ͳͷͶǤʹǦͳʹͲǦǤ   ͳͳͲ

Figure 155. Displacement 2-norm for test, baseline, and calibrated model. 111 

Figure 156. Orthogonality of impact shapes, baseline model versus test. 111 

Figure 157. Orthogonality of impact shapes, calibrated model versus test. 112 

Figure 158. Experimental and Mat 54 analytical results for the I-beam 4-ft drop 114 

(velocity=189.6-in/s) with EFS=0.06. (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 159. Deformation pattern of I-beam model executed with Mat 54 (EFS=0.06). 115 

Figure 160. Fringe plot of History Variable 3 of the T-section model shown just prior 116 

to failure of the vertical flanges (t=0.001-s). 
xii

List of Tables 

 Page

Table 1. Quasi-Static Tension Results. 7 

Table 2. Quasi-Static Compression Results. 8 

Table 3. Target Deflection for the Static Test of the Fuselage Section. 44 

Table 4. Mat 54 Material Model Parameter Definitions. 54 

Table 5. AS4/PEEK Material Properties from Literature [20-22]. 55 

Table 6. Mat 58 Material Properties. 56 

Table 7. Parameters Used in Uncertainty Studies of the Subfloor Model. 92 

Table 8. Parameters Used in Uncertainty Studies of the Fuselage Section Model. 110
xiii

List of Acronyms

AATD – Aviation Applied Technology Directorate 

ANOVA – Analysis of Variance 

ASTM – American Society for Testing and Materials 

CFC – Channel Filter Class 

CPU – Central Processing Unit 

DAS – Data Acquisition System 

FRF – Frequency Response Functions 

ft – unit of length, foot or feet 

g – unit of acceleration, ratio of measured or predicted acceleration to the acceleration of gravity 

Hz – Hertz 

in – unit of length, inch 

IRIG – Inter-Range Instrumentation Group 

kHz – unit of frequency, kilo Hertz 

ksi – unit of pressure, kilo pounds per square inch 

LandIR – Landing and Impact Research 

lb – unit of force, pound 

LSTC– Livermore Software Technology Corporation 

LV – Laser Vibrometer 

mils – 1/1000 of an inch 

min – unit of time, minute 

MPC – Magnitude-Phase-Comprehensive 

ms – unit of time, millisecond (1/1000 of a second) 

mV – unit of voltage, millivolt or 1/1000 of a volt 

MTS – Materials Testing System 

NASA – National Aeronautics and Space Administration 

NDE – Nondestructive Evaluation 

PSD – Power Spectral Density 

psi – unit of pressure, pounds per square inch 

RSVVP – Roadside Safety Verification and Validation Program 

SAE – Society of Automotive Engineering 

SARAP – Survivable Affordable Repairable Airframe Program 

SGC – Sprague & Geers Comprehensive 

SGM – Sprague & Geers Magnitude 

SGC – Sprague & Geers Comprehensive 

SMP – Symmetric Multiprocessing 

SOCIT - System Observer Controller Identification Toolbox 

SPC – Single Point Constraint 

SRW – Subsonic Rotary Wing 

SVD – Singular Value Decomposition 

s – unit of time, second 

TVA – Technology Validation Article 

US – United States 

VPV – Virtual Prototype and Validation 
1

 Impact Testing and Simulation of Composite Airframe Structures 

Karen E. Jackson, Justin D. Littell, Lucas G. Horta, Martin S. Annett 

NASA Langley Research Center 

Hampton, Virginia 

Edwin L. Fasanella Michael D. Seal 

National Institute of Aerospace (NIA) Analytical Mechanics Associates, Inc. 

 Hampton, Virginia Hampton, Virginia 

1.0 ABSTRACT 

Dynamic tests were performed at NASA Langley Research Center on composite airframe 

structural components of increasing complexity to evaluate their energy absorption behavior 

when subjected to impact loading. A second objective was to assess the capabilities of 

predicting the dynamic response of composite airframe structures, including damage initiation 

and progression, using a state-of-the-art nonlinear, explicit transient dynamic finite element code, 

LS-DYNA. The test specimens were extracted from a previously tested composite prototype 

fuselage section developed and manufactured by Sikorsky Aircraft Corporation under the US 

Army’s Survivable Affordable Repairable Airframe Program (SARAP). Laminate 

characterization testing was conducted in tension and compression. In addition, dynamic impact 

tests were performed on several components, including I-beams, T-sections, and cruciform 

sections. Finally, tests were conducted on two full-scale components including a subfloor 

section and a framed fuselage section. These tests included a modal vibration and longitudinal 

impact test of the subfloor section and a quasi-static, modal vibration, and vertical drop test of 

the framed fuselage section. Most of the test articles were manufactured of graphite 

unidirectional tape composite with a thermoplastic resin system. However, the framed fuselage 

section was constructed primarily of a plain weave graphite fabric material with a thermoset 

resin system. Test data were collected from instrumentation such as accelerometers and strain 

gages and from full-field photogrammetry. 

A primary objective of the research program was to assess the capability of LS-DYNA for 

predicting damage initiation and progression of composite airframe structures subjected to crash 

loading. Finite element models of the composite specimens were developed and impact 

simulations were performed. The properties of the composite material were represented using 

both a progressive in-plane damage model (Mat 54) and a continuum damage mechanics model 

(Mat 58) in LS-DYNA. Test-analysis comparisons of time history responses and the location 

and type of damage are presented for each test article. In addition, multi-dimensional model 

calibration techniques were applied to the subfloor and fuselage section models in an attempt to 

reconcile differences between test (photogrammetry data) and analytical predictions. Major 

findings of this research program are listed, as follows. 

(1) Dynamic testing of composite airframe components and full-scale articles demonstrated 

many different modes of structural deformation and material failure behaviors including 

global/local buckling, debonding, material crushing, brittle fracture, layer splitting, and 

delamination. 

(2) Photogrammetry proved useful in providing or augmenting more traditional test data. 
2

(3) Laminate characterization testing and simulations were conducted as a means of 

developing and calibrating the baseline material models. 

(4) In general, simulations that were performed using Mat 54 did a poor job in replicating the 

impact responses of the components and the subfloor. The Mat 54 models tended to fail 

catastrophically in a brittle fashion, absorbing little kinetic energy. The fuselage section 

proved an exception. The Mat 54 model was able to predict the location and amount of 

damage to the fuselage section, even matching the kinematics of the floor following 

separation from the sides of the fuselage structure. 

(4) Simulations performed using Mat 58 to represent the composite material demonstrated 

generally good comparison with test data; however, the level of comparison was highly 

dependent on the ERODS parameter, which defines the ultimate strain for element deletion. 

(5) Neither material model was capable of predicting delamination failure. 

(6) Multi-dimensional calibration techniques were applied to both the composite subfloor and 

fuselage section models to assess model deficiencies and uncertainties, to evaluate parameter 

importance, and to compute required model changes needed to better reconcile test and 

predictions. Results from the vibration test of the composite fuselage section indicated that 

the model needed to have the modulus reduced by 27% from the baseline value. However, 

results using impact data suggested an increase in modulus. These two findings are 

inconsistent; however, both are supported by experimental data. These conflicting results 

highlight the inadequacy of the existing models to be predictive under different loading 

conditions. Although, comparisons of the observed global behavior with analytical 

predictions are reasonable, attempts to use the analytical models to predict other test 

conditions will likely fail. 

2.0 INTRODUCTION 

In 2008, the Survivable Affordable Repairable Airframe Program (SARAP) Virtual Prototype 

and Validation (VPV) Program was initiated between Sikorsky Aircraft Corporation and the 

United States (US) Army Aviation Applied Technical Directorate (AATD). The objectives of 

the program were to validate technology advances in design processes, structural efficiency, 

crashworthiness, materials and manufacturing processes, and reparability of rotorcraft airframe 

structures. Specific objectives of the SARAP VPV program included a 25% weight reduction, 

40% recurring and 40% non-recurring cost reduction, maintaining crashworthiness, and 

repairable structure as compared to a 2002 metallic baseline. The comprehensive SARAP VPV 

program, which is described in Reference 1, included identification, review and evaluation of 

various design, analysis, material, and manufacturing technologies. As part of this program, a 

prototype Technology Validation Article (TVA) was constructed that was representative of the 

center section of a production Black Hawk helicopter, which is a metallic design. 

Following detailed design, manufacturing, and assembly, a full-scale vertical impact test of the 

TVA was performed at the Landing and Impact Research (LandIR) facility at the National 

Aeronautics and Space Administration (NASA) Langley Research Center in August 2008. A 

post-test photograph of the TVA is shown in Figure 1. The purpose of the test was to evaluate 

the performance of a “tilting roof” concept, patented by Sikorsky Aircraft Corporation, that was 

intended to dissipate the kinetic energy of high mass items, such as the rotor transmission, during 

a crash event [1]. The tilting roof concept required controlled and predictable failures of the roof 
3

structure with energy absorption accomplished using aluminum crush tubes. 

Figure 1. Post-test photograph of the SARAP TVA following the 2008 vertical drop test. 

Post-test inspection of the TVA indicated that damage was primarily limited to the roof area, 

with little or no damage found in the subfloor, or the forward framed fuselage section. 

Following the test, the TVA was moved to AATD, located at Ft. Eustis, VA. In 2010, NASA 

Langley obtained the residual SARAP hardware for testing under the NASA Subsonic Rotary 

Wing (SRW) crashworthiness research program. In May 2011, NASA and Sikorsky signed a 

Space Act Agreement [2] allowing for cooperation between the two organizations to pursue 

common research interests, including: composite material and airframe structural testing under 

dynamic loading; development of accurate and robust material models to predict aircraft 

structural response using LS-DYNA; and, validation of analytical models through test-analysis 

correlation. 

Six different types of test specimens of varying complexity were extracted from the SARAP 

residual hardware for testing. Laminate characterization testing was conducted in tension and 

compression. In addition, dynamic impact tests were performed on several components, 

including I-beams, T-sections, and cruciform sections. Finally, several tests were conducted on 

two full-scale components including a subfloor section and a framed fuselage section. These 

tests included a modal vibration and longitudinal impact test of the subfloor section and a quasistatic, modal vibration, and vertical drop test of the framed fuselage section. Most of the test 

articles were manufactured of graphite unidirectional tape composite with a thermoplastic resin 

system. However, the framed fuselage section was constructed primarily of a plain weave 

graphite fabric material with a thermoset resin system. 

Since its inception in 2006, the SRW crashworthiness research program has focused on improved 

prediction of rotorcraft crashworthiness, addressing topics such as occupant modeling and injury 

prediction, multi-terrain impact simulation, model validation studies that focused on probabilistic 

analysis, and development of system-integrated simulation models [3]. Recently, the research 

program was refocused to assess current analytical capabilities used to predict crashworthiness of 

composite airframe structures. Thus, obtaining the SARAP residual hardware was fortuitous and 

allowed testing of fairly simple coupons, more complex components, and complex built-up 

airframe structures for the purpose of model validation. One objective of the research program 
4

was to assess the capability of LS-DYNA [4, 5], a commercial nonlinear, explicit transientdynamic finite element code, for predicting damage initiation and progression of composite 

airframe structures subjected to crash loading. 

Finite element models were developed to represent each of the test articles and simulations were 

conducted using LS-DYNA. Within LS-DYNA, two composite material models were evaluated: 

MAT_ENHANCED_COMPOSITE_DAMAGE (Mat 54), a progressive failure model that uses 

the Chang-Chang failure criterion to simulate ply-by-ply failure and property degradation and 

MAT_LAMINATED_COMPOSITE_FABRIC (Mat 58), a continuum damage mechanics model 

for representing unidirectional tape and fabric composite materials. Sikorsky provided the 

material property values for Mat 54, whereas properties for Mat 58 were determined by assessing 

literature data, by comparing with similar Mat 54 values, and by using a trial and error process in 

which laminated coupons were simulated under both tensile and compressive loading. 

The level of agreement between test and analysis was determined through comparisons between 

experimental and analytical time history responses and comparisons between observed structural 

deformation, damage initiation, and failure with predicted behavior. No specific metrics were 

defined for quantifying test-analysis comparisons, in part because no acceptable metrics have 

been established as standards. However, multi-dimensional model calibration techniques were 

applied to both the composite subfloor and fuselage section models to assess model deficiencies 

and uncertainties, to evaluate parameter importance, and to compute required model changes 

needed to better reconcile test and predictions. 

The report contains several major sections including Experimental Testing, Model Development 

and Test-Analysis Comparisons, Discussion of Results, Concluding Remarks, References, and 

Appendix A. Non-Destructive Evaluation (NDE). The general approach in documenting results 

is to start with simple coupons, then progress to more complicated components, and to end with 

the full-scale structures. 

3.0 EXPERIMENTAL TESTING 

This section of the report describes the experimental program that was conducted using residual 

hardware obtained from the post-drop-test SARAP TVA. There are eight subsections including 

a description of the test specimens and a summary of results from: Nondestructive Evaluation 

(NDE); laminate characterization testing; dynamic crush testing of two I-beams, two T-sections, 

and two cruciform sections; a longitudinal impact test of a subfloor section; and, a vertical drop 

test of a framed fuselage section. 

 

3.1 Specimen Descriptions 

The SARAP residual hardware arrived at NASA in excellent shape. Care was taken to determine 

the most appropriate use for each portion of the residual hardware to optimize the data generated. 

Specimens were chosen such that the hardware could be tested from the coupon level all the way 

to full-scale impact testing. The final layout provided six different types of specimens, with a 

potential for a future seventh specimen. A depiction of the SARAP residual hardware utilization 

is shown in Figure 2. 
It should

unidirect

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9.0 Append

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7

machine were recorded, along with full-field strain photogrammetric measurements. Stress was 

computed by dividing the recorded load by the original cross-sectional area. Strain data were 

obtained from photogrammetry for four discrete points located at four distinct areas on the 

surface of the specimen. Strain data were checked by dividing the photogrammetric displacement 

data, collected at two discrete points located 0.5-in. apart, by the original separation distance. 

This method mimics traditional extensometer readings. All specimens were first scanned using 

NDE techniques to check for possible defects that might be present following the prior TVA 

drop test. No signs of internal damage were discovered and the specimens were deemed suitable 

for static testing. See Section 9.0 Appendix A for NDE results. 

A total of seven laminated coupons fabricated of thermoplastic unidirectional graphite tape 

material were tested in tension. Two coupons were cut from I-beams on the rear subfloor, one 

was cut from a T-stiffener and four were extracted from the floor. Tensile testing was conducted 

at a displacement rate of 0.05 in/min. Table 1 summarizes the results, using averages ± standard 

deviations in the data, where possible. 

Table 1. Quasi-Static Tension Results. 

Coupon from 

rear subfloor 

Manufacture 

Process 

Failure

Strain (%) 

Maximum

stress (ksi) 

Modulus (msi) 

I-beam Autoclave 1.04 ± 0.03 71 ± 4.2 6.8 ± 0.1 

T-stiffener Autoclave 1.12 68 6.1 

Floor In situ 0.95 ± 0.03 59 ± 1.3 6.1 ± 0.1 

Due to the small sample size, statistical measurements could not be performed on the coupons 

extracted from the T-stiffeners. As indicated in Table 1, the I-beam coupons had slightly higher 

moduli than the T-stiffener or the floor coupons. It also shows that the floor coupons had 

slightly lower failure strains and lower maximum stresses, indicating that they are the weakest 

laminates tested, but not by a large amount. The data also suggests that the manufacturing 

methods may factor into the laminate response, as the floor coupons show a much lower failure 

strain and maximum stress than coupons taken from the other two locations. For visualization, 

representative data sets are plotted in Figure 5 for laminated coupons taken from different 

locations. 

The data plotted in Figure 5 reinforces the data shown in Table 1. The blue I-beam coupon data 

follows a slightly steeper slope than the other two curves, while the floor coupon response (green 

curve) shows a slightly softer response. 

Due to limited availability, only coupons removed from the floor section were tested in 

compression. A total of six coupons were tested at a quasi-static displacement rate of 0.05 

in/min. Table 2 summarizes the results. 

Results show that the failure strain and the maximum stress at failure of the composite floor 

coupons in compression are approximately half of the corresponding property values in tension. 

The average modulus is slightly higher for the compressive coupons, but the scatter in the data 

suggests that the modulus is the same in compression as it is for tension. Sample compressive 

stress-strain curves are plotted in Figure 6. 
8

Figure 5. Tensile stress-strain curves for laminated coupons taken from three different locations. 

Table 2. Quasi-Static Compression Results. 

Coupon

extracted from 

Manufacturing

Process 

Failure

Strain (%) 

Maximum

stress (ksi) 

Modulus (msi) 

Floor In situ 0.49 ± 0.05 34.6 ± 2.5 6.3 ± 0.2 

Figure 6. Compressive stress-strain curves of laminated floor coupons fabricated of 

unidirectional tape. 

3.3.2 Plain Weave Fabric 

Tensile and compressive properties of laminated coupons constructed of plain weave graphite 

fabric material were conducted. These specimens were removed from undamaged portions of 

the longitudinal I-beam webs in the subfloor of the rear fuselage section. No NDE examinations 
9

were performed; however, the specimens showed no visible signs of damage. Only eight 

coupons were tested. Specimens were 1-in. wide, 10-in. long and 0.12-in. thick. The 

unsupported gage length in tension was 6-in. 

As with the unidirectional tape coupons, all tests were conducted using a MTS quasi-static load 

test machine operated in displacement control at a rate of 0.05-in/min. Both load and 

displacement values from the test machine were recorded, along with full-field strain 

photogrammetric measurements. 

Four coupons were tested in tension and the stress-strain data are plotted in Figure 7(a). Based 

on analysis of the data, the average Young’s modulus was 8.173e+6-psi, the tensile strength of 

the composite laminate was 99,147-psi, and the strain-to-failure was 1.2%. Likewise, four 

specimens were tested in compression and the stress-strain data are plotted in Figure 7(b). Based 

on data averaging, the Young’s modulus was 7.0e+6-psi, the compressive strength was 47,300-

psi, and the compressive strain-to-failure was 0.64%. Note that the compression specimens were 

less than half the length of the tensile coupons, with an unsupported gage length of 2-in. Finally, 

it is important to note that these properties are for fabric laminates, not fabric material. 

 (a) Tension. (b) Compression. 

Figure 7. Fabric coupon responses to tensile and compressive loading. 

3.4 Dynamic Crush Testing of Two I-Beam Specimens 

The component test series began with a pair of tests on the SARAP I-beam floor supports. Two 

I-beam specimens were extracted from interior regions of the SARAP TVA rear subfloor. The 

test articles were 4.56-in. tall, 6.2-in. wide, and potted into a rigid foundation such that the main 

axis of the I-beam was oriented vertically. Corner notches were cut into the upper ends of the 

web to initiate specimen crushing and to reduce peak accelerations upon initial contact with the 

drop mass. Two vertical tests were conducted at drop heights of 4- and 8-ft using a 14-ft vertical 

drop tower. Instrumentation included two accelerometers on the drop mass and a stochastic 

speckle pattern for use with full-field strain photogrammetry on the web of the specimen. The 
drop mas

shown in

Test data

sampling

used to c

sections. 

3.4.1 I-B

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11

time histories obtained from double integration of the two accelerometers located on the drop 

mass. These curves are in close agreement; however, the photogrammetry data is generally 

considered more accurate due to the fact that acceleration data may contain large offsets or drifts. 

Figure 9. Image sequence of the I-beam 4-ft drop test. 

Figure 10. Drop mass acceleration from the 4-ft I-beam test. 
12

Figure 11. Drop mass crush displacement from the 4-ft I-beam test. 

The photogrammetric data were next examined. The front web was filmed and sampled at 15 

kHz. Compressive strains at two distinct times during the crushing sequence are shown in 

Figure 12. The compressive strain at 1-ms after impact, which occurs on the upward side of the 

acceleration curve is shown in Figure 12(a). The stress wave can be seen in the specimen by 

noting that, near the top where the drop mass is touching the specimen, strains are between -0.8 

and -1.0 percent. The corners appear to have slightly larger strains, which could potentially be 

due to an uneven surface at the free edge of the specimen. However, the fixed end of the 

specimen exhibits between 0.0 and -0.2% strain, indicating that the stress wave has not travelled 

through the specimen and the strain is not fully developed at the fixed end. 

The image shown in Figure 12(b) shows the compressive strain at 7.2-ms after impact, or, more 

specifically, the time of maximum acceleration. Layer delaminations and bending interfere with 

viewing of the uppermost edge of the crush front, so strains cannot be resolved up to the edge. 

However, by comparison of the two images, much of the middle region of the specimen that 

once exhibited -0.5% strain (in green) is now closer to 0.0% strain (in orange). The entire visible 

portion of the strain data on the specimen is exhibiting this low 0.0% value. These results 

indicate that the strain has become localized in the crush front. This finding is confirmed by 

examining the videos and pictures post-test. Visual inspections show that the majority of the 

specimen away from the crush front has stayed intact. Almost all of the damage has occurred at 

the crush front, and it is best characterized by layer delamination and bending with large amounts 

of material crushing. 

A photograph of the specimen post-test is shown in Figure 13. The photogrammetric 

measurements are confirmed when looking at the failure patterns. The impact surface shows 

signs of ply delamination, material crushing, and localized bending, while the bottom of the 

specimen, which is fixed into the rigid base, exhibits no failure. 
13

 (a) Time = 1-ms. (b) Time=7.2-ms. 

Figure 12. Compressive strain at 1-ms and 7.2-ms after impact of the I-beam. 

Figure 13. Post-test photograph of the I-beam tested at 4-ft drop height. 

3.4.2 I-Beam Test at 8-ft Drop Height 

The impact velocity for the 8-ft drop test was 265.2 in/s. Image sequences are shown from the 

front face of the specimen in Figure 14. The upper left image in Figure 14 shows the specimen 

1-ms before impact. The upper right image shows the specimen 1-ms after impact. At this time, 

the drop mass has begun crushing the edge of the web section, and has just contacted the flanges, 

which were approximately 1/16-in. shorter than the web. At 5-ms after impact, the top portion of 

the I-beam web is exhibiting localized bending and delamination with many of the layers 

splitting due to the impact on the free edge. The side flanges are also exhibiting severe 

delaminations. At 10-ms after impact, material crushing from both the web and flanges can be 

seen, by noting that the entire web of the specimen is blocked from the image view by localized 

bending of composite layers. This time also corresponds with the approximate time of the 

maximum displacement of the drop mass. The acceleration response, as measured from the drop 

mass, is shown in Figure 15. The acceleration was filtered with a low-pass Butterworth filter, 

with a cut-off frequency of 1,000-Hz. 
14

Figure 14. Image sequence of the 8-ft I-beam test. 

Even filtered, the drop mass acceleration contains a number of oscillations. Between initial 

impact and 5-ms, the acceleration increases to a peak of 128-g. After 5-ms, the acceleration 

exhibits two subsequent peaks of 115- and 105-g, before decreasing to zero. The maximum 

crush displacement of the drop mass is 1.75 inches, as measured by both photogrammetry and 

double integration of the acceleration response. This value is more than double the displacement 

of the 4-ft test. The crush displacement response is shown in Figure 16. A small rebound is 

evident in the displacement curve, which indicates that very little elastic energy was released. 

The photogrammetric data were next examined. The front web was filmed and sampled at 15-

kHz. Compressive strains at two distinct times during the crush sequence are shown in Figure 

17. The compressive strain at 1-ms after impact, which occurs during the increasing portion of 

the acceleration curve, is shown in Figure 17(a). The stress wave can be seen in the specimen at 

this time. The specimen is exhibiting large strain levels near the top, which are shown in blue. 

The bottom fixed portion of the specimen is yellow and orange in color, indicating that the 

strains at this location are between 0.0 and -0.3%. The corners appear to have slightly larger 

strains, which could potentially be due to an uneven surface at the top. 
15

Figure 15. Drop mass acceleration from the 8-ft I-beam test. 

Figure 16. Drop mass crush displacement from the 8-ft I-beam test. 

Unfortunately, photogrammetric data at 10-ms after impact do not exist due to the large amount 

of ply delamination, material crushing, and localized bending, which interferes with viewing of 

the web. The image shown in Figure 17(b) shows the compressive strain data at 5-ms after 

impact, which is close to the time of peak acceleration. Debris at the uppermost edge of the 

specimen interferes with viewing of the crush front, so strains cannot be resolved up to the edge. 

However, a defined crush front is still visible by noting the transition from high compressive 

strain in blue near the crush front to low compressive strain near the base. The post-test 

specimen is shown in Figure 18. 

The photogrammetric measurements are again confirmed when looking at the failure patterns. 

Even though the specimen crushed approximately double the amount of the 4-ft specimen, the 
16

deformation patterns are the same. Large amounts of ply delamination, localized bending, and 

material crushing are evident toward the upper (impact) end of the specimen. In contrast, the 

bottom of the specimen, which is fixed into the rigid base, exhibits no failure. 

 (a) Time = 1-ms. (b) Time = 5-ms. 

Figure 17. Compressive strain at 1-ms and 5-ms after impact 

Figure 18. Post-test I-beam specimen following 8-ft drop test. 

3.5 Dynamic Crush Testing of Two T-Section Specimens 

Two T-section specimens were extracted from the SARAP TVA rear subfloor, at the location 

where a lateral I-beam terminates at the edge of a longitudinal I-beam. At this termination point, 

the two I-beam sections on the floor were fastened together using doublers. When extracted, 

these specimens resembled a T shape. Two drop tests were conducted on the T-sections using a 

14-ft vertical drop tower, such that the lateral I-beam portion of the specimen is oriented in the 

vertical direction. This orientation ensures that initial contact with the large drop mass occurs at 

the end of the vertical flanges and web. The longitudinal I-beam of the T-section was placed in a 

horizontal orientation, and used to support the vertical I-beam. The side flanges of the horizontal 

I-beams were clamped in four places to a large base fixture, which was fixed. A small 

aluminum bar (approximately 1-in. x 1-in. by 5.56-in.) was placed between the two inner flanges 
17

of the horizontal I-beam and located in the center of the specimen, just beneath the vertical web. 

The bar was intended to react the compressive load applied to the vertical web to promote 

crushing rather than global buckling of the specimen. The specimens were impacted vertically to 

determine crush behavior, failure mechanisms, and energy absorbing characteristics of the 

vertical I-beam of the T-section. Accelerometers were mounted to the drop mass and the vertical 

I-beam web was also painted with a stochastic speckle pattern to collect full-field 

photogrammetric data. The drop mass weight was 204.8 lbs. A picture of a T-section specimen 

in the test setup is shown in Figure 19. The overall dimensions of the specimen were: 9-in. tall, 

6.2-in. wide, and 13-in. long. 

Figure 19. T-section test set-up. 

3.5.1 T-Section Test at 2-ft Drop Height 

The first T-section test was conducted for a drop height of 2-ft. At this height, the drop mass 

obtained an impact velocity of 134.9-in/s. An image sequence taken from the rear, opposite side 

of the pattern, is shown in Figure 20. The top left image is 1-ms before impact. The top right 

image shows the specimen 1-ms after impact, which indicates buckling of the vertical I-beam 

flanges near the transition regions connecting them to the horizontal I-beam flanges. The bottom 

left image shows this buckling progressing to where there are also large delaminations between 

the layers on the flange. The bottom left image also shows buckling at the bottom of the web, 

near the doublers fastening the two webs together. The bottom right image shows the drop mass 

at maximum displacement. The acceleration response recorded from an accelerometer on the 

drop mass is shown in Figure 21. The acceleration trace was filtered using a low-pass 

Butterworth filter with a cut-off frequency of 1,000-Hz. 
18

Figure 20. Image sequence of the T-section test at 2-ft drop height. 

Figure 21. Drop mass acceleration from the 2-ft T-section test. 

The drop mass acceleration shows a large initial spike peaking at 325-g at approximately 1-ms 

after initial impact. This initial spike represents first contact between the drop mass and the 

specimen. Unlike the previous I-beams, the T-sections did not have a trigger to initiate crushing. 

Thus, the initial spike represents the peak load before specimen failure. This peak is followed by 

a second peak of 114-g at 3.2-ms after impact. After 10-ms, the drop mass has reached its 

maximum crush displacement, and the acceleration trace shows no significant acceleration after 

this point. The maximum crush displacement, as measured by photogrammetry, was 0.44 inches, 

as shown in Figure 22. This plot also shows the displacement response obtained by double 

integration of the acceleration data. The integrated acceleration response shows a higher 

maximum crush displacement of 0.48-in. A considerable rebound is evident in Figure 22 
19

indicating that a significant amount of energy was stored and released during the impact event. 

Figure 22. Impact mass crush displacement from the 2-ft T-section test. 

The photogrammetric data are next examined. Cameras were used to film the front web of the Tsection, sampling at 10-kHz. Two fringe plots are shown in Figure 23 at 1-ms after initial 

impact. 

 (a) Out-of-plane displacement. (b) Compressive strain. 

Figure 23. Photogrammetric results for 2-ft drop test of the T-section, 1-ms after impact. 

The photogrammetric data shows large amounts of out-of-plane motion in the specimen web. 

The top portion of the vertical web, shown in Figure 23(a), exhibits displacement toward the 

camera view, while the lower portion of the specimen exhibits out-of-plane displacement away 

from the camera view. This behavior looks like an “oil canning” mode of motion. Compressive 

strain on the surface of the specimen is shown in Figure 23(b). Note that the majority of the web 

is exhibiting low or nearly zero strain. However, the web/doubler junction exhibits large 

variations in strain, as indicated by the blue-green region near the bottom of Figure 23(b). 

Combining the two data sets, the conclusion can be reached that the web exhibits considerable 

out-of-plane motion; however, the primary failure mode is crushing near the base of the web at 
20

the doubler region and delamination of the vertical flanges. This finding is confirmed by 

examining the videos and pictures post-test. Photographs of the post-test specimen are shown in 

Figure 24. 

 (a) Front view. (b) Rear view. 

Figure 24. Post-test photographs of the 2-ft T-section test specimen. 

The post-test photographs shown in Figure 24 confirm photogrammetric data for failure. Both 

the front and back views show a distinct failure line running horizontally right above the doubler 

on the web. Furthermore, these photographs show composite layer delamination and localized 

bending of the vertical flanges, near the transition region between vertical and horizontal. This 

evidence supports the conclusion that the specimen failed through crushing near the boundary 

edges and delamination of the vertical flanges. 

3.5.2 T-Section Test at 4-ft Drop Height 

A second T-section drop test was conducted at a higher impact velocity. The 204.8-lb drop mass 

was released from a height of 4-ft, producing an impact velocity of 192-in/s. An impact 

sequence is shown in Figure 25. 

The upper left image in Figure 25 shows the specimen 1-ms before impact. The upper right 

image shows the specimen 1-ms after impact. By this time, the vertical flanges have failed 

completely above the fillet. At 5-ms after the impact, the vertical I-beam has separated from the 

horizontal I-beam and the energy of the drop mass is crushing the web of the vertical I-beam into 

the doubler region. At 20-ms after impact, the impact mass has reached maximum crush 

displacement. The lower right image shows failure of the vertical I-beam’s web, flange and 

doubler region. The acceleration response, as measured from the drop mass, is plotted in Figure 

26. As with the 2-ft test, the acceleration curve was filtered using a low-pass Butterworth filter 

with a cut-off frequency of 1,000-Hz. 
The drop

This init

second p

Fi

p mass accel

ial spike oc

peak of 71-g

Figure 25.

igure 26. Dro

leration show

curs as the 

g at 3.5-ms 

 Image sequ

op mass acce

ws a large s

drop mass 

after impac

21

uence of the 

eleration of 

spike of 254

impacts the 

ct. At 25-ms

4-ft T-sectio

the 4-ft T-se

4-g occurring

specimen. 

s, the drop m

on test. 

ection test. 

g 1.3-ms aft

 This peak 

mass has re

er initial im

is followed 

ached maxi

mpact. 

by a 

mum 
22

crush displacement, and the acceleration trace shows no significant acceleration after this point. 

The maximum vertical displacement of the drop mass is 1.63 inches, as measured by 

photogrammetry. The maximum value as determined by double integration of the acceleration 

data is 1.67-in. Both displacement curves are shown in Figure 27. 

Figure 27. Drop mass crush displacement of the 4-ft T-section test. 

The photogrammetric data are next examined. Cameras were used to film the front web of the Tsection, sampling at 10-kHz. Because of the faster impact velocity, the two fringe plots are 

shown at 0.5-ms after initial impact in Figure 28. 

 (a) Out-of-plane displacement. (b) Compressive strain. 

Figure 28. Photogrammetry results for 4-ft T-section drop test, 0.5-ms after impact. 

As with the lower velocity impact test, the photogrammetry data again shows large amounts of 

out-of-plane motion in the vertical web of the specimen. The out-of-plane behavior resembles an 

“oil canning” mode of motion with opposite out-of-plane displacement on the top and bottom 
23

separated by a zero datum in the middle. A contour plot of compressive strain on the surface of 

the specimen is shown in Figure 28(b). The majority of the web is exhibiting low (between 0 

and -0.25%) compressive strain. However, as with the lower velocity test, the web/doubler 

junction exhibits large variations in strain, as indicated by the blue-green region near the bottom 

of Figure 28(b). Combining the two data sets, the conclusion can be reached that the web 

exhibits considerable out-of-plane motion; however, primary failures occur near the base of the 

web at the doubler region and at the vertical flanges located just above the fillets. Unlike the 

previous lower velocity test, delaminations of the vertical flanges created a complete line of 

fracture, separating the vertical flange into two parts. This separation allowed the upper flange 

to become wedged inside the lower structure. This data is confirmed by examining the videos 

and pictures post-test, as shown in Figure 29. 

 (a) Front view. (b) Rear view. 

 Figure 29. Post-test photographs of the T-section following the 4-ft test. 

The photogrammtric data are confirmed by examining the post-test photographs shown in Figure 

29. Both the front and back views show complete failure of the doubler region between the 

vertical and horizontal I-beam webs. Furthermore, it shows a complete shear failure of the 

composite layers on the vertical flanges, near the transition region between vertical and 

horizontal. This evidence supports the conclusion that the specimen failed near the boundary 

edges, at the fasteners, and at regions of transition in the geometry. These two T-section tests 

proved to be less of a demonstration of a pure material failure, but rather an examination of the 

influence of test constraints and boundary conditions. 

3.6 Dynamic Crush Testing of Two Cruciform Sections 

Two cross-shaped cruciform sections located at the junctions of the longitudinal and lateral Ibeams were extracted from the post-test SARAP TVA rear subfloor. These components were 

designated cruciform sections. The length of each quadrant was approximately 6 in., which 

made the entire specimen a 12-in. by 12-in. cross. Each cruciform was approximately 6 in. tall. 

A photograph of one cruciform test specimen is shown in Figure 30. 
24

Figure 30. Photograph of a cruciform section test specimen. 

Two vertical drop tests were conducted on identical cruciform specimens. The specimens were 

impacted using a 14-ft vertical drop tower to determine their crush behavior, failure patterns and 

mechanisms, and energy absorbing characteristics for two different impact velocity conditions. 

During the test, the top flange was impacted by the drop mass, while the bottom flange remained 

fixed. The cruciform sections were instrumented with fifteen strain gages mounted on the Ibeam webs located around the various quadrants, with nine outboard gages located near the free 

edge of the web and six inboard gages located near the intersections of the webs. Two quadrant 

faces were painted with a stochastic speckle pattern to collect full-field photogrammetric data. 

A picture of the cruciform in the drop tower is shown in Figure 31. 

Figure 31. Photograph of the cruciform test setup. 
25

The drop mass weight was 215 lbs. Two tests were conducted, one from a drop height of 4-ft 

and one from a drop height of 8-ft. Strains were measured from the strain gages and from fullfield photogrammetric techniques. Accelerometers were mounted on the drop mass to measure 

the impact acceleration responses. 

 

3.6.1 Cruciform Section Test at 4-ft Drop Height 

The first specimen was impacted at a velocity of 192.5-in/s. An image series of the impact is 

shown in Figure 32. The upper left image shows the cruciform specimen immediately before 

impact. The left side of the picture shows the unpainted region and locations of strain gages on 

the cruciform. The right side of the picture shows one of the painted faces of the cruciform. 

Also note that there is an edge of a face of the cruciform perpendicular to the camera view in the 

middle of the image. The upper right image shows the cruciform immediately following impact. 

The three visible faces of the cruciform are wavy, as noticed by the reflections in the light 

patterns. At this time, the maximum acceleration has occurred. At 2-ms after impact, the lower 

left picture shows the first signs of failure. The webs have split near the bottom of the cruciform. 

At 15-ms after impact, the lower right image shows the drop mass at the time of maximum crush 

depth. At this time, the cruciform webs exhibit signs of large deformations and global buckling. 

Crushing in the lower webs is still present; however, by this time the web layers have 

delaminated along the free edges due to the large amounts of deformation present. 

Figure 32. Image sequence of cruciform crush during the 4-ft drop test. 

The acceleration time history responses, as measured from two accelerometers on the drop mass, 

are plotted in Figure 33. The data are filtered with a low-pass Butterworth filter with a 1,000-Hz 

cut-off frequency. The peak acceleration is 540-g and occurs 1.2-ms after impact. The 

extremely high initial spike can be attributed to the lack of a “trigger” or intentional defect 

designed into the specimen whose main function is to minimize initial peak loads and to initiate 
crushing

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26

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27

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28

The failure near the bottom of the specimen is clearly shown in both images in Figure 36. The 

early onset of this failure was the same as that tracked through photogrammetric techniques. The 

delamination seen on the right side of the photogrammetric results seems to be absent from the 

right image in Figure 36. When the drop mass was removed from the specimen, the wall 

appeared to spring back to an almost pre-test condition. The separations from the delamination 

were visible when the mass was resting on the cruciform; however, the damage seemed to close 

up when the mass was removed. Both the base and impact flanges of the specimen were 

undamaged. This finding confirms that the energy absorbed from the cruciform specimen came 

primarily from global buckling, crushing, and delamination of the web sections. 

3.6.2 Cruciform Section Test at 8-ft Drop Height 

A second cruciform specimen was tested at an increased velocity of 268.6-in/s. An image 

sequence of the impact is shown in Figure 37. The image in the upper left of Figure 37 shows 

the specimen immediately before impact. This view is similar to the previous cruciform view: it 

shows one cruciform quadrant on the left side, one on the right side and an edge of one face in 

the middle, perpendicular to the camera view. The upper right image shows the specimen 1-ms 

after impact. At this time, failure has already occurred through crushing on the left side 

quadrant, and delamination in the perpendicular face. The lower left image shows the cruciform 

2-ms after impact. Failures are present in all three visible quadrants. Finally, the lower right 

image in Figure 37 shows the cruciform 15-ms after impact at which point the drop mass has 

reached maximum crush displacement. Acceleration responses from two accelerometers 

mounted to the drop mass are plotted in Figure 38. These curves were filtered using a low-pass 

Butterworth filter with a cut-off frequency of 1,000-Hz. 

Figure 37. Image sequence of cruciform crush during the 8-ft drop test. 
29

Figure 38. Drop mass acceleration from the 8-ft cruciform test. 

Maximum accelerations ranged from 400- to 550-g and occurred at a time of 1.2-ms after initial 

impact. Two smaller peaks ranging in magnitude from 100- to 145-g occurred after the initial 

spike. Following these two smaller peaks, the acceleration stabilizes at approximately 40-g and 

remains at that level for approximately the next 10-ms. At 17-ms after impact, the acceleration 

response decays and reaches zero at 20-ms. This acceleration response is very similar to the 

cruciform test which was conducted at a lower impact velocity. Even though the energy is two 

times greater in this case, both specimens required approximately the same peak acceleration to 

initiate failure and similar trends are observed in the acceleration time histories after initial 

failure. However, the extra kinetic energy is absorbed through specimen crushing, as evidenced 

by the higher maximum displacement of the drop mass, shown in Figure 39. 

Figure 39. Crush displacement of the 8-ft cruciform test. 
A maxim

obtained 

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31

specimen, it was possible to remove entirely the top flange away from the rest of the specimen 

due to the failures near the crush front. The greater amount of damage in this test, as compared 

with the previous 4-ft drop test, is attributed to the higher kinetic energy associated with the 8-ft 

drop test condition. 

3.7 Testing of the Subfloor Section 

A fairly large subfloor section was removed intact from the post-test SARAP TVA, with the 

intent of conducting a longitudinal impact test. The location of the removed subfloor section is 

highlighted in Figure 42. A longitudinal test was conducted to replicate a horizontal impact of a 

helicopter with a large amount of forward velocity into a rigid wall, which is one of the test 

conditions specified in the military crashworthiness requirement MIL-STD-1290A(AV) [8]. 

This test condition is infrequently evaluated; however, given that the subfloor was available, the 

longitudinal impact test was planned and conducted. To replicate this test condition, the subfloor 

section was positioned vertically and crushed dynamically using the 70-ft vertical drop tower 

located at the LandIR facility. The impact occurred such that the subfloor specimen remained 

stationary while a large instrumented portal frame drop mass was released from a height of 91-

inches and impacted the specimen at a nominal velocity of 22-ft/s (264-in/s). The subfloor 

section test article is shown in Figure 43 in a vertical configuration. 

Figure 42. SARAP subfloor test article location. 

Figure 43. SARAP subfloor test article. 
32

The subfloor I-beams were instrumented at various locations with 12 strain gages. A stochastic 

speckle pattern was painted onto the base (smooth, skinned side of the subfloor) and the I-beams 

for use with photogrammetric imaging techniques. The bottom of the specimen was potted to 

achieve a fixed condition on the aft end. The impact end was free, but attached to a flat plate, 

which would provide a large flat surface for the drop mass to impact. A photograph of the 

subfloor after potting, instrumenting, and painting is shown in Figure 44. A green circle in 

Figure 44 represents a strain gage mounted to the I-beam flange in the location noted. A yellow 

circle represents a strain gage mounted to the I-beam web in the location noted. Overall 

dimensions of the test article were: 35-in. tall, 62-in. wide, and 6-in. deep. 

3.7.1 Modal Testing 

A modal test was performed on the composite subfloor in position while under the 70-ft vertical 

drop tower at NASA’s LandIR facility prior to the scheduled impact test. Analysts involved with 

the project were interested in performing a modal test with the 4,279-lb-impact mass resting on 

top of the test article in order to estimate the first plunging mode. A photograph of the test set-up 

is shown in Figure 45. From the bottom to the top, the test set-up consisted of concrete blocks, 

two aluminum plates, composite subfloor, aluminum impact mass, and orange portal frame. 

Figure 44. Pre-test photograph of the subfloor section highlighting strain gage locations. 

Accelerometers were mounted to the test article at selected locations. Eighteen accelerometers 

were mounted in the X-axis (normal to the test article floor surface) using wax. Following the 

first set of tap tests with an impact hammer at several locations, the accelerometers were rotated 

to the Z-axis (vertical) and all taps were repeated. An overall view of the test article with 

accelerometers mounted in the Z-axis is shown in Figure 46. 

Taps were performed at four seleted locations on the test article with a 1-lb hammer, as depicted 

in Figure 47. Each location was tapped six times and the results were averaged to obtain Power 

Spectral Density (PSD), Frequency Response Functions (FRF), and coherence data. After 
33

obtaining poor results with the 1-lb hammer, additional datasets were obtained with a 3-lb 

hammer at two locations in the Z-axis. The 1-lb hammer was used to excite with ~30 lb peak 

forces in the X-axis and ~100 lb peak forces in the Z-axis, and the 3-lb hammer increased the 

forces in the Z-axis to ~200 lb peak. 

Figure 45. Modal test setup. 

Figure 46. Accelerometers mounted on subfloor test article. 

FRF and coherence results are shown in Figures 48 and 49. A reciprocity check between the Xaxis excitation locations is shown in Figure 48. Poor coherence below 120 Hz indicates that the 

measurement results cannot be trusted below that frequency. The FRFs indicate that reciprocity 

is satisfied for these measurements above 120 Hz, but the peaks are not well defined and are 

noisy. A similar reciprocity check is shown in Figure 49 for the Z-axis excitation locations with 

the 3-lb hammer, with the frequency resolution reduced from 0.5 Hz to 2.0 Hz. While the 

coherence of these measurements is improved, they are still noisy and reciprocity is not 

exhibited. 
Figure 

Figure

Figur

47. Tap exc

 48. Horizon

e 49. Vertic

34

itation in ve

ntal axis reci

cal axis recip

rtical directi

iprocity chec

procity check

ion. 

ck. 

k. 
35

The reasons behind the poor measurement results could not be effectively diagnosed and fixed 

due to the limited amount of time allotted for the modal test. Improvements in coherence with 

higher force inputs suggest that exciting the structure with more energy, such as with a modal 

test shaker, would have been an appropriate follow-up step to obtain better measurements. 

However, this approach was beyond the scope of the “simple” modal test that was agreed upon. 

Attempts to use experimental modal analysis methods on the measurements resulted in complex 

mode estimates and poor agreement between measured and regenerated FRF curves. 

Unfortunately, none of the mode shape estimates obtained were satisfactory enough to present 

without being misleading. 

3.7.2 Subfloor Impact Test Results 

The longitudinal impact test of the subfloor was conducted on June 27, 2012. Photogrammetric 

tracking targets were placed on the portal frame such that velocities and impact angles could be 

obtained. Aluminum honeycomb stops were placed underneath the portal frame feet to serve as 

safety barriers should a test anomaly occur. A picture of the test set-up with the test article under 

the 70-ft drop tower is shown in Figure 50. The weight of the test article including the portal 

frame drop mass was 5,100-lbs, and the impact condition was nominally 22-ft/s (264-in/s). The 

specimen was crushed to failure. 

Figure 50. Subfloor longitudinal impact test set-up. 

Image sequences of the test are shown in Figures 51 and 52. These images show the time history 

of the subfloor failure. The upper left pictures show the specimen immediately before impact. 

The upper right pictures show the specimen at 3-ms after impact. In these pictures, the Tstiffeners, which are attached to the subfloor base between the supported I-beam flanges on the 

floor, have already debonded from the test article. These pictures also show the first indication 

of I-beam crushing. The lower left pictures are taken at 6-ms after the impact and show the Ibeam sections post failure. The lower right pictures show the specimen 25-ms after initial 
36

impact. In these pictures, the specimen has completely failed and the residual I-beam 

components are absorbing energy through debonding, shearing or crushing. 

Figure 51. North east camera view of subfloor deformation. 

Figure 52. North camera view of subfloor deformation. 
37

Five accelerometers were mounted on the portal frame drop mass, three of which were oriented 

in the vertical direction and two were oriented in the two normal directions to the vertical. The 

vertical accelerometers ranged between 750- and 2000-g while the upper range of the normal 

accelerometers was 200-g. The sample rate for the accelerometers was 25 kHz using the T-DAS 

Pro DAS. The results from the accelerometers were inconclusive. Of the three vertical 

accelerometers used, two accelerometers (750- and 2000-g maximum range) immediately failed 

upon first contact between the drop mass and the test article. The center accelerometer recorded 

an acceleration time history, but after further inspection, a large unexplained offset was present 

in the data immediately after first contact. Because of this anomaly, data from this accelerometer 

were not used. Strain gages were intended to provide load path information during impact. 

However, the results were also inconclusive since most of the strain gages debonded upon first 

contact of the portal frame and the test article. 

Photogrammetric imaging was conducted only on the base (skinned) side of the test article due to 

camera malfunctions on the open side. The base side was filmed at 600 x 800 resolution at 10 

kHz, which captured the motion of the drop mass, along with the strain field around the bottom 

of the specimen. Through examination of the high-speed video and the photogrammetric data, 

the majority of the damage and failure occurred in the first 1- to 6-ms after impact. The velocity 

time history of the drop mass was determined based on photogrammetric data analysis and is 

plotted versus time in Figure 53(a). Vertical red lines are included in the plot to highlight the 

time of initial impact, the time at which the portal impacts the honeycomb stops, and the time of 

motion stop. The time duration between initial impact and contact with the honeycomb stops is 

important to the analytical correlation study. The data shown in Figure 53(a) was obtained by 

differentiating photogrammetric displacement data, which can produce a noisy response. 

Consequently, the photogrammetric velocity response was filtered using a Society of Automotive 

Engineering (SAE) Channel Filter Class (CFC) 60-Hz low-pass filter, as shown in Figure 53(b). 

 (a) Velocity data from photogrammetry. (b) Velocity data filtered. 

Figure 53. Drop mass velocity responses from photogrammetric measurements of the subfloor. 

The velocity at impact was 21.98-ft/s (263.76-in/s). The curves plotted in Figure 53(a) represent 

four different tracked targets on the drop mass. All four show good agreement, which gave 

further confidence in the photogrammetric results. The velocity curves show an almost linear 

trend with some slight oscillations as they approach the time of honeycomb impact. Kinematic 
38

equations were used to determine the average acceleration during this time period by dividing the 

change in velocity over the change in time. With this method, the average acceleration between 

initial and honeycomb impact is approximately 4.9-g. The data ends partially between the 

honeycomb impact and the motion stop because the targets being tracked on the drop mass fell 

out of the camera view and were obstructed by the base of the subfloor at this time. The motion 

stop datum was determined qualitatively by examining the high-speed videos. 

The displacement of the drop mass was also measured using the photogrammetric techniques to 

investigate crush displacement of the test article, as plotted in Figure 54. The displacement 

response shows that the subfloor crushed approximately 13 inches before the honeycomb stops 

engaged, which occurred at 0.071-s. The displacement of the subfloor base was next examined 

using full-field strain photogrammetric techniques. Out-of-plane displacement is shown in 

Figure 55. Note that the lower left portion of the base could not be tracked. 

 

Figure 54. Drop mass vertical displacement of the subfloor. 

The out-of-plane displacement at 1.2- and 5-ms after impact are shown in Figures 55(a) and (b), 

respectively. The large green areas are the areas on the base between the I-beam structural 

members, which are only supported by post-bonded stiffeners. However, due to the immediate 

failure of the post-bonded T-stiffeners, the areas between the I-beams are unsupported. The 

structure between the areas of green corresponds to the locations of the I-beams on the test 

article. The displacement shows an “oil-canning” mode of motion until 5-ms when a shear 

failure near the top of the base causes it to crush instead of deform. The failure shape can be 

observed by examining the top of the tracked region in Figure 55(b). Note the convex shaped 

curve, above which data is not tracked. The strains in the specimen were very low, as shown in 

Figure 56. 
39

At 1.2-ms after impact, the subfloor base is intact and exhibits a strain variation between -0.4 and 

-1.0%, as shown in Figure 56(a), indicating that the subfloor is absorbing the initial impact 

energy by deforming. No clear pattern occurs which quantifies the differences in straining 

between the unsupported base and the supported I-beam sections. The failure location near the 

top of the base does begin to show by noting the large blue horizontal lines near the top, 

indicating that the strain is approximately -1% in compression. However, at 5-ms after impact, 

the average full-field strain ranges from 0 to -0.2%, as shown in Figure 56(b). This result 

indicates that the impact has progressed past structural deformation and moved toward material 

failure. 

 (a) Time = 1.2-ms. (b) Time = 5-ms. 

Figure 55. Out-of-plane displacement in the subfloor test article at 1.2- and 5-ms after impact. 

 (a) Time = 1.2-ms. (b) Time = 5-ms. 

Figure 56. Vertical strain in the subfloor test article at 1.2- and 5-ms after impact. 

The test article was damaged beyond repair from the impact. Photographs of the specimen are 

shown in Figure 57 immediately after the impact. The open side of the subfloor specimen, shown 

in Figure 57(a), shows that two of the I-beams actually crushed to absorb the impact energy and 

two of the I-beam flanges delaminated from the I-beam webs to absorb energy. The photograph 

of the base side of the subfloor section, shown in Figure 57(b), shows that the unsupported 

section of the base of the floor remained relatively intact. However, the base separated from the 
40

I-beams, evidence of which is supported from the displacement and strain data. The photograph 

of Figure 57(b) also indicates that there was a large shear failure near the top of the base of the 

subfloor. 

 (a) Open subfloor deformation. (b) Base (skinned) side deformation. 

Figure 57. Post-test photographs of the subfloor. 

Further inspections showed that the T-stiffeners debonded cleanly off of the subfloor base and 

contributed almost no energy absorption. A photograph showing a portion of the base in which 

both the T-stiffeners and the I-beam flanges have come apart from the base section (circled) is 

shown in Figure 58(a). The largely intact T-stiffeners, which debonded immediately upon 

impact, were collected and are shown post-test in Figure 58(b). 

 (a) Subfloor open side failures. (b) Post-test collected T-stiffeners. 

Figure 58. Post-test subfloor stiffener behavior. 

3.8 Testing of the Framed Fuselage Section 

This section of the report describes modal, quasi-static, and impact testing of the forward framed 

fuselage section that was extracted from the post-test SARAP TVA. The extraction location is 

depicted in Figure 2. The fuselage section dimensions were: 17-in. deep, 69-in. tall, and 97-in. 

wide. The height of the subfloor was 6-in. NDE was not performed on the test article; however, 

some areas of localized delamination were present near the top right-side outer skin. Epoxy was 

inserted into the delaminated area during attachment of the upper mass. Two nondestructive 

tests were performed including modal vibration and quasi-static testing, prior to vertical drop 

testing of the fuselage section. 
41

3.8.1 Modal Testing 

Modal testing of the SARAP forward fuselage section concentrated on extracting two critical 

modes that were identified by pre-test analysis as important contributors to the fuselage impact 

response (the 4th vibration mode at 73-Hz and the 8th mode at 144-Hz). From the pre-test 

analysis, it was also determined that supporting the fuselage at two points underneath the floor 

would not constrain these two critical modes and therefore this configuration was selected for the 

modal test. A montage of pictures from the modal test is shown in Figure 59. A front-left view 

of the entire fuselage test section is shown in the center of Figure 59. Close-up photographs of 

the left, right, top, and floor accelerometers are shown including a view of the bottom support. 

For all testing PCB Piezotronics Model T333B42 accelerometers were used with nominal 

sensitivity of 500 mV/g. For excitation, a PCB Piezotronics 086D20 impact hammer with a hard 

tip was used with an in-line load cell to measure the input force. Because only nine 

accelerometers were available, testing was conducted by sections (i.e., left, top, etc.). For each 

data set, the fuselage was hit with the impact hammer in all 3 directions. Although most of the 

accelerometers were placed to measure accelerations normal to the inner surface of the fuselage, 

1 to 2 accelerometers were placed to measure the out-of-plane accelerations (i.e. x-direction). 

Acceleration data were collected from 0- to 256-Hz with a frequency resolution of 0.5-Hz. A 

total of 34 acceleration measurements were obtained. 

Figure 59. Montage of composite fuselage test section undergoing modal testing. 

 

Because the modal test is nondestructive, it is a convenient way to collect data for an initial 

assessment of the LS-DYNA model. During the drop test, only certain vibration modes play a 

role in the test article deformation. To determine what vibration modes need to be targeted, an 

impact simulation was conducted and the time histories were decomposed into time-invariant 

impact shapes. Then, the impact shapes were compared to the vibration modes of the structure 

using orthogonality. Only vibration modes with high orthogonality values were targeted during 

the modal test. Based on this procedure, only two vibration modes were deemed important for 

impact: the 4th vibration mode at 73 Hz and the 8th mode at 144 Hz. The mode at 144 Hz is 
42

primarily the fuselage floor moving in the vertical direction, whereas the mode at 73 Hz involves 

motion of the entire cross section. 

For modal testing, the fuselage section was simply supported at the bottom. The fuselage section 

is shown in Figure 60 prior to modal testing. In the modal test configuration, the test article 

weighed 1,610-lb, including 368-lb for the fuselage section, 423.5-lb for the floor ballast, and 

818.5-lb for the upper ballast. Two different tests were conducted: 1) the first test used a Laser 

Vibrometer (LV) to measure velocity at different points inside the fuselage due to an excitation 

from an overhead shaker positioned to impart loads in the vertical direction, and 2) the second 

test used an impact hammer and accelerometers. Based on model predictions, the bandwidth for 

the modal test was from 10- to 250-Hz. Data from the LV was collected at 38 locations on the 

top, bottom, and both sides of the fuselage. However, all data from the LV only measured 

velocity perpendicular to the laser head and therefore no forward/backward motion of the 

fuselage was recorded. In contrast, 34 accelerometers were also placed at top, bottom, and both 

sides; however, in this case, 2 accelerometers were also placed to measure forward/backward 

motion. In addition, the hammer tests included hits in all 3 directions while measuring 

acceleration. Data sets from tests using a shaker with an LV system and data using 

accelerometers with an impact hammer were both analyzed. Surprisingly, data with the LV 

system resulted in mode shapes that were inconsistent with the acceleration data and the LSDYNA model. For this reason, results using the LV data will not be presented and efforts to use 

this dataset were dropped. 

Figure 60. Modal test set-up for composite fuselage test section. 

Frequency response functions from the vibration tests with 34 accelerometers were analyzed 

using the MATLAB® SOCIT (System Observer Controller Identification Toolbox) [9] and the 

Frequency Domain Identification Toolbox [10]. Frequency data were analyzed to determine the 

experimental mode shapes and frequencies used to compare with analysis. The results of this 

analysis are reported in Section 4.6.3 of this report. 

3.8.2 Static Testing 

Prior to static and impact testing, the fuselage section was modified to include ballast mass and 
43

instrumentation. Due to the slanted roof line of the test article, two 5.5-in. wide, 84-in. long 

aluminum wedge shaped blocks were mounted via epoxy to the top of the north (rear) and south 

(front) edges of the test article to create a flat surface. An aluminum plate was then fastened 

onto the two wedges to tie them together. The two wedges and plate were known as the upper 

ballast mass. Once together, the entire fixture was clamped to the top of the portal frame. The 

portal frame also included an aluminum impact mass, which was installed for the previous 

subfloor test. This impact mass remained attached to the portal frame to provide extra 

accelerometer mounting locations as well as adding extra ballast weight. The extra mass, plates 

and wedges were all needed to achieve the appropriate clearance for the test article to fit within 

the portal frame’s bracing legs and also to increase the total drop mass to 5,884 lbs. Ballast mass 

was added to the floor in the form of four 100-lb lead weights to simulate seats and occupants. A 

photograph of the test article is shown in Figure 61. 

Figure 61. Pre-test photograph of the framed fuselage section in test position (south view). 

Prior to static testing, a stochastic speckle pattern was painted on both sides of the fuselage 

section for collecting full-field photogrammetric data. The static test was first conducted on the 

test article to determine vertical deflection while in the test condition. The applied static load 

was the overhead mass used for the impact test. The static test was conducted by raising the 

portal frame a slight amount such that the test article was approximately 1-in. off of the ground 

and the weight of the overhead ballast mass was offloaded. Then, a set of images was acquired 

from both sets of photogrammetric cameras in this unloaded state. Next, the test article was then 

lowered back onto the impact surface such that the entire overhead ballast mass was being 

supported by the test article. Finally, a second set of images was acquired with both 

photogrammetric cameras. The images were then loaded into the photogrammetry software, 

compared, and deflections were computed. 
44

Three data sets were used to determine the test article deflection for the static test. The three 

data sets were the north face stochastic pattern and both the north and south face target locations. 

The south face stochastic pattern was too small for the south cameras to analyze. The discrete 

target locations analyzed on both the south and north faces are depicted in Figure 62. The 

differences in the purely vertical positions for the targets located on each of the faces, along with 

a column for the lateral motion of the targets located on the sides of the test article are shown in 

Table 3. 

Figure 62. Discrete tracked target locations used for deflection measurements during the fuselage 

section static test (north view). 

Table 3. Target Deflection for the Static Test of the Fuselage Section. 

Camera 

view

East deflection (mil) Middle 

deflection

(mil)

West

deflection

(mil)

Lateral

deflection

(mil)

South 59.9 57.6 48.3 6.4 

North 52.7 68.6 54.1 14.1 

The average for all of the deflections was approximately 56.9 mils, while the standard deviation 

was 7.0-mil. It should be noted that the resolution of the photogrammetric system in this 

configuration was approximately 5-mil, and thus the spread in the data can be partially attributed 

to the resolution. However, there are also potential physical causes for discrepancies in the data. 

Two major factors noted were the possibility of misalignment in attaching the test article to the 

portal frame and the addition of a layer of composite underneath the north face of the test article. 

This additional layer was most likely added during the original fabrication of the frame section to 

help support the subfloor. When the framed fuselage section was cut away from the TVA for the 

facilitation of this test series, the extra layer remained. The addition of this layer did make the 

test article slightly asymmetric and caused the north side to be loaded slightly higher than the 

south side. 

The full-field north side photogrammetric data confirms the results in Table 3. Horizontal and 

vertical displacement maps from the north side of the test article are shown in Figure 63. 
45

Though individual values were not examined from the displacement fields, general trends were 

noted. The maximum horizontal displacement of the right edge, depicted in Figure 63(a), shows 

approximately 15- to 18-mils of displacement, whereas the centerline shows almost no horizontal 

displacement. The vertical displacement map is shown in Figure 63(b). The bottom of the test 

article was selected as the data reference plane and, as such, this region of the test article showed 

very little displacement. Starting at this almost zero datum, the displacement increased from the 

bottom to the top and reached a maximum of approximately 50 mils at the top of the test article. 

Due to imaging issues, the data on the left side of the test article spanning up to the upper left 

corner was deemed suspect and not included in the analysis. 

 (a) Horizontal. (b) Vertical. 

Figure 63. Horizontal and vertical displacement fields on the north face of the fuselage section. 

3.8.3 Impact Testing 

The framed fuselage section was drop tested using the 70-ft. vertical drop tower at the LandIR 

facility. The test article was attached to the drop tower portal frame, raised to a nominal height 

of 91-in. and dropped to achieve a nominal 22-ft/s (264-in/s) velocity. The total drop weight, 

which included the portal frame, the upper ballast mass, lower floor ballast mass and the test 

article was 5,884 lbs. Instrumentation was limited to seven accelerometers at locations on the 

portal frame and two accelerometers on the lead weights on the test article floor. The sample 

rate for the accelerometers was 25 kHz using the T-DAS Pro DAS. Photogrammetric data were 

acquired from both sides of the test article. Sampling rate for the south photogrammetric 

cameras was 10-kHz, while the sampling rate for the north photogrammetric cameras was 1-kHz. 

The main purpose of the south cameras was to measure the impact conditions along with time 

histories for displacement and velocity through the discrete targets, while the north cameras were 

to capture full-field deformations of the test article through analysis of the stochastic speckle 

pattern. However, the targets and the stochastic pattern were applied to both sides in an attempt 

to collect extra data, if possible. Aluminum honeycomb was placed underneath both portal legs 

to stop the portal frame should a test anomaly occur. The impact test set-up is depicted in Figure 

61. 

The measured impact velocity was 20.75-ft/s (249-in/s) and the total vertical crush of the test 

article was 7.25-in., as measured from photogrammetric data analysis. The pitch angle at impact 

was 0.42° leaning towards the south, and the roll angle was 0.29° leaning towards the west, as 

determined from the photogrammetric data. Failure occurred at the floor through shearing and 
46

brittle fracturing from the sides near the lower corners. Delaminations occurred near the crush 

front on the lower sides. An image sequence of the test is shown in Figure 64 from the southeast 

view and in Figure 65 from the north view. Note that the honeycomb stops did not engage during 

the test. 

Figure 64. Image sequence of the fuselage section drop test from the southeast camera. 

Four still images extracted from the high-speed videos are shown in Figures 64 and 65. The first 

image in the upper left shows the test article immediately before impact. The upper right image 

shows the specimen 10-ms after impact. At this time, the bottom of the test article has contacted 

the ground and the sides have bowed out. Localized failure due to shear and buckling of the 

individual layers has begun; however, the test article has largely retained its structural integrity. 

The lower left pictures show the test article 25-ms after impact. At this point in time, large-scale 

failures due to shearing and brittle fracturing in the floor structure were present in the lower 

corners. The lower right images show the test article 50-ms after impact. The floor structure has 

completely broken away from the rest of the test article. During the time history of the impact, 

the floor structure rebounds before coming to rest skewed toward the southwest corner of the test 

article. Photographs of the southeast and northwest corners of the test article are shown in 

Figure 66 during floor separation. 

Acceleration responses from the five accelerometers mounted on the portal frame and drop mass 

are similar, so only a representative acceleration response is plotted. Similarly, the accelerations 

from the two accelerometers located on the floor ballast mass are similar to each other, but 
47

different from the other five, so only one of the floor accelerations is examined. Thus, there are 

two distinct acceleration responses from the test. The portal frame, drop mass, and upper test 

article represent the first, while the floor of the fuselage section represents the second. The 

acceleration responses from accelerometers located on the middle of the portal drop mass and on 

the lower floor ballast mass are shown in Figure 67 as representative traces for both locations. 

The raw acceleration curves at both locations were filtered using a low-pass Butterworth filter 

with cut-off frequencies of 300- and 1,000-Hz. The lower filter frequency was needed to capture 

the underlying pulse, due to the large amount of oscillations in the data. 

Figure 65. Image sequence of the fuselage section drop test from the north camera. 

Based on the 300-Hz filtered data, the portal frame exhibited an oscillatory acceleration 

response, achieving an initial 100-g peak for the first 1-ms of the response, then immediately 

plunging to approximately -100 g. Thereafter, the response oscillated at 26-Hz around a mean 

value of 2.6-g starting after 30-ms. Based on the 300-Hz data, the floor acceleration, shown in 

Figure 67(b), also exhibited a very large spike of 475-g during the first 1-ms of impact. This 

large acceleration resulted from the initial impact of the test article with the ground. After the 

initial spike, three subsequent acceleration peaks of 60-, 47- and 57-g occurred at times of 15-, 

30-, and 38-ms, respectively. These peaks all occurred during the dip in the drop mass 

acceleration. After the four initial acceleration spikes, the floor accelerations settled out and 

eventually decayed to zero by 0.05-s. At this point, the 26-Hz oscillation arises in the upper drop 

mass acceleration. The 26-Hz oscillation mode was attributed to the portal frame legs flexing 

during the impact. 
48

 (a) Southwest view. (b) Northeast view. 

Figure 66. Close up views of floor failures that occurred during the fuselage section impact test. 

 (a) Portal frame acceleration. (b) Floor acceleration. 

Figure 67. Portal frame drop mass acceleration and floor acceleration. 

The south face photogrammetric data was next examined for the velocity time histories, shown 

in Figure 68. The differentiated photogrammetric data shows two distinct velocity traces, which 

are similar to what was seen with the acceleration data. Before impact, the traces overlay, 

reaching a maximum of 20.75-ft/s. Immediately after impact, the curves diverge. While the first 

visible evidence of failure at the floor occurred at approximately 5-ms after impact, the curve for 

the test article floor suggested that the beginning of separation between the floor and the rest of 

the test article occurred almost immediately after impact. The floor velocity goes positive 

between 20- and 160-ms, indicating a velocity in the upward direction, which also agreed with 

the test videos. The test videos showed that the west side of the floor rebounded over 22-in. 

Meanwhile, the other curve, which represents the response of the portal frame and upper mass, 

showed a gradual decrease from the impact velocity toward zero. There were still noticeable 

oscillations in the portal frame and upper test article data after impact. These oscillations were 
49

absent from the floor data because the floor separated from the rest of the test article early in the 

impact event. However, after 160-ms, the two curves began their convergence toward zero, 

which was an indication that the impact event had concluded. 

Full-field photogrammetric data was captured from the north side of the test article only. The 

vertical strain on the test article is shown at 5- and 33-ms after impact in Figures 69(a) and (b), 

respectively. These two particular times were chosen to examine the structural deformation 

before failure (5-ms) and after failure (33-ms) of the test article. Due to debris interfering with 

the camera view during the impact, the floor deformation was not tracked. 

Figure 68. Velocity responses from photogrammetry of the fuselage section floor and portal 

frame. 

 (a) Time = 5-ms. (b) Time = 33-ms. 

Figure 69. Vertical strain fringe plots at 5-ms and 33-ms after fuselage section impact. 
50

Care must be taken when interpreting the strain results, as the low resolution in the data allows 

for potentially invalid and misinterpreted data due to edge effects in the pattern. An example is 

the area of red ( >1% strain) near the upper left corner of the test article. This data is invalid 

because the software is grouping the deformation data from the flange with the deformation data 

from the web. The data is physically nonsensical and must be ignored. However, after 

examining the remainder of the test article, the results show that both tensile and compressive 

strain were present at 5-ms. The majority of the tracked locations on the top of the test article 

show a tensile strain, while the sides of the test article show compressive strain in the 0 to -0.2% 

range. At 33-ms after impact, the results show more compressive strain near the top of the test 

article than near the bottom, where the failures occurred. The top strains approach -0.4 to -0.6% 

in compression, while strains near the floor failures hover around 0.0%. This finding indicates 

that the bottom sides of the test article have become unloaded due to the failure of the floor. 

However, the data resolution as acquired from photogrammetry is one data point every 5/8-in., 

and the data sampling rate is 1-kHz, which are both low to capture any highly localized behavior 

which develops during composite failures, and thus these areas of highly localized strain that 

develop quickly before failure are not captured. Instead, photogrammetric data represent more 

of an overall structural deformation, which is intended to acquire and display macro-level results 

for use with computer modeling correlations. 

The framed fuselage section came to rest shortly after impacting the surface. A post-test 

photograph is shown in Figure 70. 

Figure 70. Post-test photograph of the framed fuselage section test article. 

After the test, the fuselage section was examined thoroughly for locations of failure. The lower 

sides were damaged beyond repair due to the large amounts of buckling, folding and 
51

delaminations that occurred near the impact. The floor, despite being heavily damaged near the 

edges where it detached from the sides, appeared to be undamaged near the middle and under the 

lower ballast mass. Post-test photographs of the test article are shown in Figures 71-73. The 

upper portion of the test article was in excellent shape after the test. This undamaged condition 

was most likely due to the rigidization of this area when attaching the aluminum wedges over the 

outer I-beam flanges. 

Figure 71. Post-test location of the front frame floor (overhead view). 

Figure 72. Post-test photographs showing the floor separation locations. 

Figure 73. Post-test photograph showing the interior of the floor. 
52

4.0 MODEL DEVELOPMENT AND TEST-ANALYSIS COMPARISONS 

A primary objective of this project is to evaluate the capabilities of the commercial nonlinear, 

explicit transient dynamic code, LS-DYNA [4, 5], for simulating the impact response of 

composite airframe structures, focusing on the prediction of damage initiation and progression 

leading to ultimate failure. This section of the report will document the material model 

development and provide test-analysis comparisons for: laminate characterization tests; dynamic 

crush tests of I-beam, T-section, and cruciform section components; the longitudinal impact test 

of the subfloor section; and, the vertical drop test of a two-frame fuselage section. 

Finite element models were developed to represent each of the test articles and simulations were 

conducted using LS-DYNA. Two composite material models were evaluated: 

MAT_ENHANCED_COMPOSITE_DAMAGE (Mat 54), a progressive failure model that uses 

the Chang-Chang failure criterion to simulate ply-by-ply failure and property degradation and 

MAT_LAMINATED_COMPOSITE_FABRIC (Mat 58), a continuum damage mechanics model 

for representing unidirectional tape and fabric composite materials. Sikorsky provided the 

material property values for Mat 54, whereas properties for Mat 58 were determined by assessing 

literature data, by comparing with similar Mat 54 values, and using a trial and error process in 

which laminated coupons were simulated under both tensile and compressive loading. 

The level of agreement between test and analysis was determined through comparisons between 

experimental and analytical time history responses and comparisons between observed structural 

deformation, damage initiation, and failure with predicted behavior. Several correlation metrics 

are available for use in quantifying the level of agreement between two curves as part of the 

validation process of a numerical model. These include the Roadside Safety Verification and 

Validation Program (RSVVP) [11, 12] that allows pre-processing of two curves, including 

filtering, phasing, and timing adjustments, etc. Sixteen different metrics are included that are 

classified into 3 categories: (1) magnitude-phase-comprehensive (MPC) metrics, (2) single-value 

metrics, and (3) analysis of variance (ANOVA) metrics. This methodology was applied 

successfully in Reference 13. An MPC metric was developed by Sprague and Geers [14, 15], 

which provides a measure of the “goodness of fit” between two curves. Three parameters are 

calculated over a specified time interval: Sprague and Geers Magnitude (SGM), Sprague and 

Geers Phase (SGP), and Sprague and Geers Comprehensive (SGC), which is a combination of 

magnitude and phase. Generally, a value of less than 40 for SGM, SGP, or SGC is considered 

passing the criteria. 

No specific metrics were defined for quantifying test-analysis comparisons, as presented in this 

report, in part because the impact dynamics community at large has not established acceptable 

standards. Consequently, the level of test-analysis agreement is determined by visual inspection 

of the time history responses. However, multi-dimensional model calibration techniques were 

applied to both the composite subfloor (see Section 4.5.4) and fuselage section (see Section 

4.6.6) models to assess deficiencies and uncertainties, to evaluate parameter importance, and to 

compute required model changes needed to better reconcile test and predictions. 
53

 4.1 Material Model Development and Comparison with Laminate Property Data 

4.1.1 Description of Mat 54 Material Model 

The LS-DYNA material property designated MAT_ENHANCED_COMPOSITE_DAMAGE or 

(Mat 54) [5], was selected for evaluation in this study. Mat 54 material properties were obtained 

from existing finite element models, provided by Sikorsky. This material model is used to 

represent orthotropic materials, such as composite laminates fabricated of unidirectional tape 

material, and should be limited to thin shell construction. Composite lamination theory is used 

to determine ply stresses and strains in the laminate. Two different failure criteria can be 

specified including Chang-Chang [16, 17] or Tsai-Wu [18], by setting the CRIT parameter to 54 

for Chang-Chang, or 55 for Tsai-Wu in the Mat 54 card. For all of the Mat 54 simulations 

presented in this document, CRIT was set to 54. The Chang-Chang failure criterion is a 

progressive in-plane damage model that accounts for three in-plane failure modes: matrix 

cracking, fiber-matrix shearing, and fiber breakage. The nonlinear shear stress-strain behavior is 

based on the theory described in References 19 and 20. For fiber failure and/or fiber-matrix 

shearing, the level of property degradation within the damaged area depends on the size of 

damage predicted by the fiber failure criterion [21]. Note that most progressive failure models 

are strength-based. When the strength of a ply within the laminate is exceeded, the properties are 

degraded. Progressive failure is achieved through ply-by-ply failure within the laminate. Once 

all plies in the laminate have failed, the element is deleted. Mat 54 input parameters are defined 

in Table 4; however, specific input values are not included to protect Sikorsky data rights. 

Mat 54 incorporates a set of model-specific parameters that influence material failure behavior, 

but that cannot be measured in a laboratory. Some of these parameters include TFAIL, SOFT, 

ALPH, YCFAC, FBRT, BETA, and EFS, as defined in Table 4. If utilized, these parameters 

must be determined through analytical calibration studies, or by trial and error. Reference 22 

provides an excellent description of these parameters, along with the results of parametric studies 

showing how changes in these parameters can influence model responses. 

Several different Mat 54 models were defined and used in finite element models of the SARAP 

residual hardware. For example, two Mat 54 models were based on autoclave and in situ 

properties of unidirectional graphite tape materials. However, since the autoclave and in situ 

properties were so similar, only the autoclave properties were used in simulations of all test 

articles, with one exception. The fuselage section model introduced three new Mat 54 material 

models, used to represent both thermoplastic unidirectional graphite tape and thermoset plain 

weave graphite fabric materials. Note that fabric materials were only used in the fuselage section 

model. All other models used the autoclave Mat 54 properties. The Mat 54 property values for 

thermoplastic tape and thermoset fabric composites that were provided by Sikorsky are 

considered baseline properties and they were not changed for any of the simulations conducted. 

4.1.2 Description of Mat 58 Material Model 

MAT 58 is a continuum damage mechanics material model based on the Matzenmiller-LublinerTaylor theory [23] and is intended for use with shell elements to simulate composite tape 

laminates and woven fabrics. The model requires input of material properties in tension, 

compression, and shear to define stress-strain behavior within the lamina or laminate. The user 

specifies the in-plane elastic modulus and Poisson’s ratio in two primary directions, designated A 
54

and B in LS-DYNA. Maximum strength values in tension, compression, and shear are also 

specified at corresponding strain values. A representation of the stress-strain curve for in-plane 

tension is illustrated in Figure 74. The tensile response is initially linear elastic with the modulus 

specified by EA. Stress increases nonlinearly until XT, the maximum strength, is reached. The 

nonlinear portion of the response is defined internally by LS-DYNA based on a continuum 

damage mechanics theory [23]. Once XT is reached, the stress is reduced based on the “stress 

limiting” factor SLIMT1, and is then held constant at the reduced value until elements reach a 

strain specified by the ERODS parameter in the material model, at which point the elements are 

deleted and removed from the solution. Through the appropriate selection of the SLIM 

parameters and ERODS, it is possible to incorporate plastic-like behavior in the model and avoid 

pre-mature element failure. Similar stress-strain responses are defined for in-plane compression 

and shear. Reference 24 provides an excellent description of continuum damage mechanics 

based composite failure models and gives an excellent assessment of Mat 58 for predicting 

energy absorption of composite materials. Additional information on this material model can 

also be found in Reference 5. 

Table 4. Mat 54 Material Model Parameter Definitions. 

Property Description Symbol 

Density RO 

Longitudinal Young’s modulus EA 

Transverse Young’s modulus EB 

Poisson’s ratio, BA PRBA 

Shear modulus, AB GAB 

Shear modulus, BC GBC 

Shear modulus, CA GCA 

Material axes option AOPT 

Max matrix strain, for element deletion DFAILM 

Max shear strain for element deletion DFAILS 

Time step size criteria for element del. TFAIL 

Shear stress parameter for nonlinear term ALPH 

Softening reduction factor for material strength in the crash front SOFT 

Softening for fiber tensile strength FBRT 

Reduction factor for comp fiber strength after matrix comp failure YCFAC 

Max strain for fiber tension DFAILT 

Max strain for fiber comp DFAILC 

Effective failure strain EFS 

Longitudinal comp strength, psi XC 

Longitudinal tensile strength, psi XT 

Transverse comp strength, psi YC 

Transverse tensile strength, psi YT 

Shear strength, psi SC 

Failure criteria (54-Chang) CRIT 

Weighting factor for shear in fiber BETA 

A literature search was conducted and several references [25-27] were found that contained 

fairly detailed material property data for AS4/PEEK, which is a commonly used thermoplastic 
composit

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56

58 can be determined through laboratory testing, especially the SLIM parameters and ERODS. 

These parameters must be based on calibrating the simulation results to best match the test data. 

Finally, it should be noted that, once the baseline Mat 58 properties were established, they were 

not changed for any of the simulations conducted. However, separate parameters studies were 

performed to study the influence of variations in ERODS, the parameter that controls element 

deletion; and in strength values XT, XC, and E11T. As with Mat 54, actual property values for 

Mat 58 are not included in Table 6 to protect Sikorsky data rights. 

 

Table 6. Mat 58 Material Model Property Definitions. 

Material Property Description, units Symbol 

Density, lb-s2/in4 RO 

Young’s modulus longitudinal direction, psi EA 

Young’s modulus transverse direction, psi EB 

Poisson’s ratio in the BA direction PRBA 

Stress limit of nonlinear portion of shear curve, psi TAU1 

Strain limit of nonlinear portion of shear curve, in/in GAMMA1 

Shear modulus in the AB direction, psi GAB 

Shear modulus in the BC direction, psi GBC 

Shear modulus in the CA direction, psi GCA 

Min stress factor for limit after max stress (fiber tension) SLIMT1 

Min stress factor for limit after max stress (fiber compression) SLIMC1 

Min stress facto for limit after max stress (matrix tension) SLIMT2 

Min stress factor for limit after max stress (matrix compression) SLIMC2 

Min stress factor for limit after max stress (shear) SLIMS 

Material axes option (=0.0 locally orthotropic) AOPT 

Time step for automatic element deletion TSIZE 

Maximum effective strain for element layer failure ERODS 

Softening reduction factor for strength in the crash front SOFT 

Failure surface type FS 

Strain at longitudinal compressive strength, in/in E11C 

Strain at longitudinal tensile strength, in/in E11T 

Strain at transverse compressive strength, in/in E22C 

Strain at transverse tensile strength, in/in E22T 

Strain at shear strength, in/in GMS 

Longitudinal compressive strength, psi XC 

Longitudinal tensile strength, psi XT 

Transverse compressive strength, psi YC 

Transverse tensile strength, psi YT 

Shear strength in the A-B plane, psi SC 

4.1.3 Comparison of Tensile Laminate Response with Simulation (Unidirectional Tape) 

The in-plane tension model is shown in Figure 75. The 1-in. x 1-in. tension model contains 

approximately 300 quadrilateral shell elements. Single Point Constraints (SPCs) were applied to 

fully constrain the left edge nodes, i.e. translations and rotations about the x-, y-, and z-axes were 

fixed. In addition, an SPC was defined to fix translations and rotations for the right edge nodes, 
57

except that translation in the x-direction was allowed. The right edge nodes were assigned a 

BOUNDARY_PRESCRIBED_MOTION_SET specifying that at the termination time (0.3 

seconds) the maximum x-displacement is 0.03-inches, which corresponds to a maximum strain in 

the x-direction of 3%. The fairly long termination time of 0.3 seconds ensured that quasi-static 

loading was applied and that the kinetic energy of the simulation was minimal. 

A unidirectional graphite tape laminate was simulated using the tension model shown in Figure 

75. A local material direction was specified for this laminate such that 0°-layers were oriented in 

the x-direction. The ply orientations were input using the PART_COMPOSITE card in LSDYNA with a total laminate thickness of 0.096-inches. The baseline Mat 54 and Mat 58 

unidirectional tape material models were assigned. The tensile model was executed using the 

default Belytschko-Tsay shell element formulation (ELFORM=Type 2). Output from the model 

included SPC forces in the x-direction for the left edge nodes, which were summed and then 

divided by the cross-sectional area to obtain an average stress. Strain was obtained by scaling 

the time response. Models were executed using LS-DYNA Symmetric Multi Processing (SMP) 

Version 971 R6.0.0 with double precision and required up to 78 minutes of Central Processing 

Unit (CPU) on a single processor for an end time of 0.3 seconds. 

 

Figure 75. Finite element model used for simulating in-plane tensile loading. 

Comparisons of experimental and analytical stress versus strain responses are shown in Figures 

76(a) and (b) for Mat 54 and Mat 58 material models, respectively. For Mat 54, the predicted 

response is slightly stiffer than the test and both curves are essentially linear elastic to failure. 

However, the Mat 54 analytical response under predicts the experimental tensile strength and the 

strain-to-failure. For example, the test coupon fails at a tensile strain of 0.0107-in/in and at a 

tensile strength of 73,400-psi, whereas the model fails at a lower strain of 0.007-in/in and at a 

lower stress of 53,056-psi. Following failure, the laminate no longer has load-bearing capability. 

One explanation for the premature failure in the Mat 54 model is that DFAILM is set at a very 

low value. As indicated in Reference 22, DFAILM, which is the maximum strain in the matrix, 

must be greater than or equal to the maximum value of YT/EB or YC/EB. In this case, the 

minimum value was selected to provide a conservative response. 

For Mat 58, the model matches the stiffness of the test response well; however, the strength of 

the laminate is over predicted. The Mat 58 model obtains maximum strength at a strain of 0.016-

in/in and a stress of 93,600-psi. Based on the fact that SLIMT1 is set at a high value (close to 

1.0), the value of stress at failure is maintained in the elements until the strain corresponding to 

the value of ERODS is achieved. Note that since element deletion produces holes in the model 
58

resulting in stress concentrations, it is prudent to keep ERODS as high as possible to prevent 

numerical problems and subsequent pre-mature failure of the model. 

 (a) Mat 54 autoclave properties. (b) Mat 58 properties. 

Figure 76. Comparison of Mat 54 and Mat 58 tensile stress-strain responses with test data. 

To improve the Mat 58 comparisons with the tensile test results, two parametric studies were 

conducted. In the first, the values of XC, XT, and E11T were reduced. XC was reduced by 

25%; XT was reduced by 25%; and, E11T was reduced by 35%. This Mat 58 model is 

designated the reduced strength model. Predictions from this model are plotted in Figure 77(a) 

with the tensile test data, as well as the predicted response from the baseline Mat 58 model. The 

reduced strength model accurately predicts the value of tensile strength; however, due to the 

large value of ERODS, the stress level is maintained. A second study was conducted to 

determine the influence of changes to the ERODS parameter. Simulations were conducted using 

the baseline Mat 58 material model for unidirectional tape material; however, ERODS was 

varied between 0.01 and 0.05 in increments of 0.01. In the baseline Mat 58 material model, a 

high value of ERODS was specified. The results of the ERODS parameter study, shown in 

Figure 77(b), indicate that best agreement with test data is for an ERODS value between 0.02 

and 0.03. 

4.1.4 Comparison of Compressive Laminate Response with Simulation (Unidirectional Tape) 

The in-plane compression model is shown in Figure 78. This model has a height of 1-in. and a 

width of 0.5-in. The reduced width, as compared with the tension model, was needed to 

minimize the tendency of the model to buckle out-of-plane during the simulations. The 

compression model contains 50 quadrilateral shell elements. The compression models were 

executed using the default Belytschko-Tsay shell element formulation (ELFORM= Type 2). 

SPCs were applied to fix the left edge nodes. In addition, SPCs were used to fix translations and 

rotations for the right edge nodes, except that translation in the x-direction was allowed. The 

BOUNDARY_PRESCRIBED_MOTION_SET card was defined for the right edge nodes 
59

specifying that at the termination time (0.3 seconds) the maximum x-displacement is -0.015-

inches, which corresponds to a maximum x-direction strain of -3%. As before, the fairly long 

termination time of 0.3 seconds ensured that quasi-static loading was applied and that the kinetic 

energy of the simulation was minimal. 

 (a) Reduced strength model. (b) ERODS study results. 

Figure 77. Results of two Mat 58 parameter studies on the tensile laminate response. 

Figure 78. Finite element model used for simulating in-plane compressive loading. 

As with the tensile loading case, a unidirectional graphite tape laminate was simulated. A local 

material direction was specified for this laminate such that 0°-layers were oriented parallel to the 

x-axis. The ply orientations were input using the PART_COMPOSITE card with a total 

laminate thickness of 0.096-inches. In the first simulation, the Mat 54 unidirectional tape 

material model representing the autoclave fabrication technique was assigned. In the second, 

Mat 58 properties for unidirectional graphite tape material were assigned. Output from the 

compression model included SPC forces in the x-direction for the left edge nodes, which were 

summed and then divided by the cross-sectional area to obtain an average stress. Strain was 

obtained by scaling the time response. The model was executed using LS-DYNA SMP Version 
60

971 R6.0.0 with double precision and required 13 minutes of CPU on a single processor for an 

end time of 0.3-s. 

Comparisons of experimental and analytical stress versus strain responses are shown in Figures 

79(a) and (b) for Mat 54 and Mat 58 material models, respectively. For the Mat 54 simulation, 

the predicted response exhibits a linear elastic response to failure. Note that the compression 

data deviates from linearity at a low value of strain. However, the analytical curve over predicts 

both the failure stress and strain-to-failure of the test coupon. As with the tension model, when 

failure occurs in compression, the laminate loses the capability to carry load. 

 

As shown in Figure 79(b), the Mat 58 model matches the initial stiffness of the test response 

well. However, the test response deviates from linearity very early and fails at a strain of 

0.0055-in/in and at a stress of 33,900-psi. The model exhibits a linear elastic response and 

reaches maximum strength at a strain of 0.0066-in/in and a stress of 48,500-psi. At that point, 

the stress at “failure” is reduced by the SLIMC1 factor, which leads to the observed stair step 

response. Elements are removed from the model once a strain corresponding to the value of 

ERODS is achieved. 

 (a) Mat 54 properties. (b) Mat 58 properties. 

Figure 79. Comparison of compression stress-strain responses. 

As with the tensile loading case, two separate parametric studies were performed including the 

Mat 58 reduced strength model, in which XC, XT, and E11T were lowered; and the ERODS 

variation study. Results are shown in Figures 80(a) and (b), respectively. The reduced strength 

model matched the stiffness of the current Mat 58 material model; however, both models were 

stiffer than the test, which deviated from linearity at low strain values. However, the reduced 

strength model achieves maximum strength at a significantly lower stress (36,667-psi) and lower 

strain (0.005-in/in) than the baseline Mat 58 model, which initially fails at a stress of 48,750-psi 

and a strain of 0.0066-in/in. This change is attributed to the 25% reduction in XC, the 

longitudinal compressive strength in Mat 58. The ERODS parameter variation shows that best 

agreement with test is obtained using a low value of ERODS of 0.01. 
61

 (a) Reduced strength model. (b) ERODS variation study results. 

Figure 80. Results of two Mat 58 parameter studies on the compressive laminate response. 

4.1.5 Comparison of Tensile Laminate Response with Simulation (Fabric) 

The finite element model, shown in Figure 75, was modified to represent fabric material. The 

modifications included changing the total laminate thickness to 0.12-in., and changing the 

laminate stacking sequence to match the test specimen. In addition, the material models for Mat 

54 and Mat 58 were updated to reflect the fabric construction. All other model details remained 

the same, as described in Section 4.1.3. 

Predicted stress-strain results are shown in Figures 81(a) and (b) for Mat 54 and Mat 58, 

respectively, compared with the average tensile response of the fabric coupons. The Mat 54 

response exhibits excellent agreement with the stiffness of the tensile data; however, the strength 

is under predicted. The model fails at approximately 75,000-psi and 0.0085-in/in strain, 

compared with the average test coupons that failed at 100,000-psi and 0.012-in/in strain. For 

Mat 58, the model also predicted the initial stiffness of the test response well. Whereas the test 

response is linear elastic to failure, the Mat 58 predicted response deviates from linearity at 

approximately 60,000-psi, and eventually fails at a stress of 92,500-psi and a strain of 0.015-

in/in. Note that the Mat 58 fabric model incorporates a much lower value of ERODS indicating 

that element deletion will occur at a lower value of strain. 

4.1.6 Comparison of Compression Laminate Response with Simulation (Fabric) 

The finite element model, shown in Figure 78, was modified to represent compressive loading of 

fabric coupons. The modifications included changing the total laminate thickness to 0.12-in., 

and changing the laminate stacking sequence to match the test specimen. Mat 54 and Mat 58 

fabric material properties were used in the model. All other model details remained the same, as 

described in Section 4.1.4. 

Predicted stress-strain results are shown in Figures 82(a) and (b) for Mat 54 and Mat 58, 

respectively, compared with the average compressive response of the fabric coupons. Both 

models exhibit a stiffer response than the test, which is linear elastic to failure. For the test 
62

coupons, failure occurs at a stress of approximately 46,700-psi and a strain of 0.0064-in/in. The 

Mat 54 predicted curve fails at a stress of 91,000-psi and a strain of 0.01-in/in. Following 

failure, the laminate loses load-bearing capability. Conversely, the Mat 58 predicted curve 

reaches maximum strength at a stress of 57,000-psi and a strain of 0.0072-in/in. However, due to 

the SLIMC1 parameter, the compressive stress is reduced to 30,000-psi and elements continue to 

carry that load until the strain defined by ERODS is reached. 

 (a) Mat 54. (b) Mat 58. 

Figure 81. Comparison of Mat 54 and Mat 58 predicted responses with average tensile response 

of fabric laminates. 

 (a) Mat 54. (b) Mat 58. 

Figure 82. Comparison of Mat 54 and Mat 58 predicted responses with average compression 

response of fabric laminates. 
63

4.2 Simulation of Dynamic Crush Tests of Two I-Beam Specimens 

 

4.2.1 Description of the Finite Element Model 

A finite element model, depicted in Figure 83, was developed to simulate two dynamic crush 

tests of I-beam specimens. The model consisted of 24,300 nodes; 3,050 Belytschko-Tsay 

(ELFORM=Type 2) quadrilateral shell elements representing the I-beam with a nominal element 

edge length of 0.125-in.; and 16,400 solid elements representing the drop mass. Mat 54 or Mat 

58 were assigned to the shell elements representing the I-beam, whereas Mat 20 (*MAT_RIGID) 

was assigned to the drop mass. The density of the rigid material was adjusted to ensure that the 

weight of the drop mass matched that of the test (204.8-lb). Seven different parts were used 

including one part for the drop mass, and six parts representing different regions of the I-beam. 

Several different PART_COMPOSITE cards were created to represent the upper edges of the Ibeam flanges that incorporated ply drop-offs to mimic the reduced thicknesses produced by 

chamfering. I-beam models were executed using LS-DYNA SMP Version 971 R6.0.0 with 

double precision and required 5 hours and 51 minutes of CPU on four Linux-based processors 

for an end time of 0.02-s. 

 

Two SPC definitions were used, one to fix the bottom nodes of the I-beam, and the second to 

ensure that the drop mass could move only in the vertical direction. Finally, the nodes forming 

the drop mass were assigned an initial velocity to match the two test conditions, 189.6- and 

265.2-in/s. These two impact velocities correspond to the 4- and 8-ft drop heights used during 

the tests. As mentioned previously, both Mat 54 and Mat 58 material models were evaluated and 

input properties were based on unidirectional graphite tape composite material. 

Output from the model included acceleration-, velocity-, and displacement-time histories of a 

central node on the drop mass, as well as pictures of model deformation. For the I-beam, Tsection, and cruciform section component tests, acceleration data from the accelerometer(s) 

attached to the drop mass are compared with analytical predictions. Predicted velocity and 

displacement responses are also compared with experimental velocity and displacement data 

obtained from single- and double-integration of the acceleration data, respectively. In addition, 

displacement data obtained from photogrammetry were included for comparison. 

 (a) Front view with drop mass. (b) Three-quarter view without drop mass. 

Figure 83. LS-DYNA model of the I-beam component. 
64

4.2.2 Mat 54 Predicted Responses Compared with 4-ft Test Data 

Poor results were obtained using the Mat 54 material model. Essentially, the model failed in a 

brittle fashion, exhibiting very little energy absorption. Numerous parameter studies were 

conducted in an attempt to obtain stable crushing of the model. However, all models provided 

similar results to those shown in Figure 84. Predicted acceleration responses, shown in Figure 

84(a), consisted of a series of spikes representing repeated contact and release of the drop mass 

onto the I-beam. These short-duration spikes occur early in the pulse. Only 50-in/s of velocity is 

removed during the simulation through failure of the I-beam, as shown in Figure 84(b). As 

indicated in Figure 84(c), the vertical displacement of the drop mass is not alleviated and it 

continues to displace linearly without abatement. The model fails catastrophically with no sign 

of stable crushing, as shown in Figure 85. The model loses the bottom nodal constraints, as 

elements are failed and removed. 

 (a) Acceleration. (b) Velocity. (c) Displacement. 

 

Figure 84. Experimental and Mat 54 analytical results for the I-beam crush test at 189.6-in/s. 

 0.0000-s 0.0008-s 0.0016-s 

 0.0024-s 0.0032-s 0.004-s 

Figure 85. Mat 54 model predicted deformation pattern indicating complete specimen collapse. 
65

4.2.3 Mat 58 Predicted Responses Compared with 4-ft Test Data 

The baseline Mat 58 material model for unidirectional tape was used to represent the I-beam. 

The predicted responses show good agreement with the overall shapes, magnitudes, and pulse 

durations of the experimental curves, as shown in Figure 86. The experimental and predicted 

acceleration responses, shown in Figure 86(a), were filtered using a 1000-Hz Butterworth lowpass filter. The displacement results, shown in Figure 86(c), include the Mat 58 predicted 

response, the responses based on double integration of the acceleration data, and 

photogrammetry. The maximum crush displacement predicted by the model (0.85-in.) falls 

between the experimental results (0.82- to 0.86-in.). A sequence of model deformation is shown 

in Figure 87 for four time steps. These depictions show that the I-beam model exhibits stable 

crushing from the top edge, matching the deformation patterns of the test article, depicted in 

Figures 9 and 13. No global buckling is evident, though some folding of the webs is observed. 

 (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 86. Experimental and Mat 58 analytical results for the I-beam 4-ft. drop test. 

 (a) Time=0.0024-s. (b) Time = 0.0048-s. 

 (c) Time=0.0072-s. (d) Time = 0.01-s. 

Figure 87. Time sequence of I-beam model deformation for Mat 58. 
66

As with the laminate characterization simulations, the I-beam model was executed using the 

reduced strength Mat 58 material model. In addition, a parameter study to evaluate the influence 

of ERODS was conducted. For this evaluation, ERODS was varied from 0.1 to 0.5 in increments 

of 0.1. Displacement-time history results for these two studies are shown in Figures 88(a) and 

(b), respectively. The reduced strength model resulted in over prediction of the maximum crush 

displacement by approximately 0.18-in., or 22.5%. The ERODS study indicates that a value of 

0.5 provides best agreement with test data. 

 

 (a) Reduced strength model. (b) ERODS study. 

Figure 88. Mat 58 parameter study results for the I-beam (189.6-in/s velocity). 

4.2.4 Mat 54 Predicted Responses Compared with 8-ft Test Data 

The finite element model of the I-beam, shown in Figure 83, was re-executed with a higher 

impact velocity of 265.2-in/s corresponding to the 8-ft drop test condition. All other model 

details were kept the same as for the lower velocity impact, including the Mat 54 autoclave 

material properties. Similar difficulties were found with the Mat 54 simulation at the higher 

impact velocity. The predicted acceleration responses consisted of a series of spikes representing 

repeated contact and release of the drop mass onto the I-beam as elements are deleted, until 

ultimate model collapse. Very little velocity is removed and the vertical displacement of the 

drop mass continues to increase linearly without abatement. As with the lower velocity case, 

many parametric simulations were executed in an unsuccessful attempt to generate a model that 

would predict stable crushing, rather than collapse. Consequently, no results are shown for the 

Mat 54 model. 

4.2.5 Mat 58 Predicted Responses Compared with 8-ft Test Data 

The I-beam model was executed with Mat 58 unidirectional tape properties for the higher 

velocity condition. Comparisons of predicted and experimental acceleration-, velocity- and 

displacement-time histories are shown in Figure 89. As before, the acceleration responses shown 

in Figure 89(a) were filtered using a 1000-Hz Butterworth low-pass filter. In addition, 

displacement data are shown based on double integration of the acceleration data and 

photogrammetry. The model generally exhibits a stiffer response than the experiment. For 

example, analytical acceleration peaks are higher than the experimental peaks, the predicted 

velocity curve is steeper than the test curve and crosses zero velocity earlier than the test, and the 
67

model under predicts the maximum crush displacement (1.4-in. versus 1.78-in. for the test). 

A sequence of model deformation is shown in Figure 90 for four time steps. These depictions 

show that the model I-beam web exhibits stable crushing from the top edge, thus mimicking the 

deformation behavior of the test article. No global buckling is evident, though some folding of 

the web section is observed. The predicted deformation simulates the observed test response, as 

shown in Figure 18. 

 (a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 89. Experimental and Mat 58 analytical results for the I-beam 8-ft. drop test. 

 (a) Time=0.0024-s. (b) Time = 0.0048-s. 

 (c) Time=0.0072-s. (d) Time = 0.0096-s. 

Figure 90. Time sequence of I-beam model deformation for Mat 58 simulation. 

The I-beam model (velocity=265.2-in/s) was re-executed using the reduced strength Mat 58 

material model. In addition, a parameter study to evaluate the influence of ERODS was 
68

conducted. For this evaluation, ERODS was varied from 0.1 to 0.5 in increments of 0.1. Note 

that for the ERODS study, the baseline Mat 58 material properties were used. Displacementtime history results for these two studies are shown in Figures 91(a) and (b), respectively. The 

reduced strength model provided much better agreement with the experimental maximum crush 

displacement as compared with the original Mat 58 model. The reduced strength model 

predicted a crush displacement of 1.73-in. compared with 1.78-in. for the test. The ERODS 

study shows that a value of 0.3 provides best agreement with test data. 

 (a) Reduced strength model. (b) ERODS study 

Figure 91. Mat 58 parameter study results for the I-beam (265.2-in/s velocity). 

4.3 Simulation of Dynamic Crush Tests of Two T-Section Specimens 

4.3.1 Description of the Finite Element Model 

Finite element models, shown in Figure 92, were developed to represent the two dynamic impact 

tests of T-section components. Models were developed to represent two different configurations 

of the T-sections. For example, the 2-ft drop test was conducted on a T-section with two fillets 

only, whereas the higher velocity test (4-ft drop) was performed on a T-section with four fillets. 

The fillets are located at the intersection of the vertical and horizontal flanges. The 4-fillet 

model consists of: 15,367 nodes; 13,746 quadrilateral shell elements; 512 solid elements to 

represent the drop mass; 9 parts; 3 material models; 2 constraint definitions; and, one contact 

definition. Models were executed using both Mat 54 and Mat 58, in which material properties 

for unidirectional graphite tape were input. For all simulations, Belytscko-Tsay shell elements 

were used (ELFORM=Type 2). Mat 54 and Mat 58 were assigned to the shell elements 

representing the T-section, whereas Mat 20 (*MAT_RIGID) was assigned to the drop mass. The 

density of the rigid material was adjusted to ensure that the weight of the drop mass matched that 

of the test (204.8-lb). 

Two SPC definitions were used, one to fix the bottom nodes of the horizontal I-beam flanges, 

and the second to ensure that the drop mass could move only in the vertical direction. Finally, 

the nodes forming the drop mass were assigned an initial velocity to match the two test 

conditions, 134.9- and 192-in/s. These two impact velocities correspond to the 2- and 4-ft drop 
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heights used during the tests. Output from the model included acceleration-, velocity-, and 

displacement-time histories of a central node on the drop mass, as well as pictures of model 

deformation. The nominal element edge length was 0.125-in. Simulations were executed for 

0.04-s, which required 8 hours of CPU on four Linux-based processors. 

 (a) 2-ft drop model (two fillets). (b) 4-ft drop model (four fillets). 

Figure 92. T-section finite element models (drop mass not shown). 

4.3.2 Mat 54 Predicted Responses Compared with 2-ft Drop Test Data 

The model for the 2-ft drop test is shown in Figure 92(a). As observed in the test article, the 

fillets on the right side of the vertical flange have been cut off, leaving only the two fillets on the 

left side. Mat 54 properties representing unidirectional graphite tape composite material were 

assigned to shell elements representing the T-section. During the simulation, the model rapidly 

disintegrates as shown in Figure 93, which depicts the model deformation at time 0.017 seconds. 

Since the model does not represent the observed damage, no further test-analysis comparisons 

are presented. 

Figure 93. Mat 54 T-section model deformation at 0.017-s.

 

4.3.3 Mat 58 Predicted Responses Compared with 2-ft Drop Test Data 

Comparisons of the measured acceleration- and displacement-time histories for the 2-ft drop test 

versus the Mat 58 simulation predictions are shown in Figures 94(a) and (b), respectively. The 

test and analysis acceleration curves, shown in Figure 94(a), were filtered using a low-pass 

Butterworth 1000-Hz digital filter. Test data shown in Figure 94(b) is from double integration of 
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the acceleration data and from photogrammetry. The initial peak accelerations are comparable at 

270-g for the model and 320-g for the test, indicating that the initial stiffness, initial conditions, 

and mass of the model match those of the test article. However, after the initial spike, the model 

accelerations are higher than the test from 0.002- until 0.0075-s. 

The damage observed from the 2-ft drop test indicated that both flanges on the vertical I-beam 

delaminated and buckled outward at the top of the fillets. The LS-DYNA model was not set up 

to account for ply delamination. Consequently, once delamination initiated, the actual test 

specimen became much weaker than the model and vertical displacement is greater in the test 

than in the model. From Figure 94(b), the maximum vertical test displacement from 

photogrammetry was measured to be 0.44-in. as compared to 0.27-in. in the LS-DYNA model. 

 (a) Acceleration responses. (b) Displacement responses. 

Figure 94. Comparisons of Mat 58 model predicted responses with test data for the 2-ft drop. 

The deformed model of the T-section is shown in Figure 95 at 0.004-s. The figure indicates that 

a line of elements half way across the flange and located just above the fillet have failed and 

been deleted. However, other surrounding elements have not failed and these elements limit the 

amount of displacement. In the test, delamination and buckling allowed more deflection to 

occur, resulting in greater damage in the web than was seen in the model. 

The same T-section model was re-executed with the reduced strength Mat 58 material model and 

test-analysis results are compared in Figure 96. The 1000-Hz low-pass filtered peak acceleration 

predicted by the model in Figure 96(a) is a little over 200 g’s as compared to 320 g’s from the 

test accelerometer. Thus, the peak acceleration of the reduced strength model does not compare 

as well as the baseline Mat 58 model. From Figure 96(b), the displacement for the reduced 

strength Mat 58 simulation is now closer to the test value. The final deformed plot looks very 

similar to Figure 95 and hence is not shown. 

The T-section model representing the 2-ft drop test with baseline Mat 58 material properties was 

executed in which ERODS was varied from 0.1 to 0.5, in increments of 0.1. A depiction of 

model deformation is shown in Figure 97 for the Mat 58 simulation with ERODS = 0.1. 

Elements were deleted along the vertical flange and below the fillet for this small ERODS value. 

Comparisons of displacement-time histories for the different ERODS values are shown in Figure 
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98 along with data from the 2-ft drop test. Based on these results, an ERODS value between 0.3 

and 0.4 shows best agreement with the test data. 

Figure 95. Maximum deformation of the Mat 58 T-section model at time 0.004-s after impact. 

Elements have been deleted where delamination occurred in the test. 

 (a) Acceleration comparisons. (b) Displacement comparisons.

Figure 96. Comparison of reduced strength Mat 58 model with test for the 2-ft drop. 

4.3.4 Mat 54 Predicted Responses Compared with 4-ft Test Data 

The model representing the 4-ft drop test of the T-section is shown in Figure 92(b). Mat 54 

material properties were assigned to shell elements representing the T-section. The model 

rapidly disintegrates as shown in Figure 99, which depicts model deformation at time 0.02 

seconds. Multiple element failures are observed in different regions of the model, most of which 

do not match the test response. However, the simulation does capture the failure of the two 

vertical flanges above the fillets. The acceleration- and displacement-time histories of the dropmass are compared with test data in Figure 100. The test-analysis comparisons are poor. 
72

Figure 97. Mat 58 T-section model failure below the fillet for ERODS = 0.1. 

Figure 98. Comparison of displacement responses of the drop mass for the 2-ft T-section test 

with the Mat 58 models with only the ERODS parameter changed. 

Figure 99. Mat 54 T-section model at time 0.02 seconds. 
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 (a) Acceleration comparisons. (b) Displacement comparisons. 

Figure 100. Comparisons of the 4-ft drop test data with the Mat 54 model predictions. 

4.3.5 Mat 58 Predicted Responses Compared with 4-ft Test Data 

The initiation of failure for the 4-ft drop test of the T-section occurred as buckling and 

delamination in the vertical I-beam flanges above the fillet. Delaminations caused complete 

failure of the flange, allowing the flanges on both sides above the fillets to become wedged 

inside the horizontal flanges below the fillets. Comparisons of the test and Mat 58 predicted 

acceleration, velocity, and displacement responses are shown in Figures 101(a), (b), and (c), 

respectively. Both test and analytical acceleration responses have been filtered using a 1000-Hz 

low-pass Butterworth filter. The Mat 58 model shows three peak accelerations whereas the test 

only shows two. However, the peak accelerations are comparable in magnitude to the test. 

Again, the displacement is much lower for the model than the test as is evident in Figure 101(c). 

The displacement disparity can be explained by observing the failure mechanism in the test and 

comparing with the simulation. Although the vertical flanges in both the test and model 

completely failed above the fillet, in the test the vertical flanges wedged inside the horizontal 

flanges below the fillet. However, in the model (see Figure 102), the vertical flange rotated 

slightly after failing, which allowed edge-on-edge contact with the horizontal flange after failure. 

Since single surface contact was used, this contact allowed the upper flange to cut into the lower 

flange. The higher force generated prevented further displacement from occurring after 0.01 

seconds, whereas the actual test specimen continued displacing up to 0.02 seconds. The final 

deformation pattern of the model is illustrated in Figure 102. 
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 (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 101. Comparisons of Mat 58 model predicted responses with 4-ft T-section test data. 

Figure 102. Mat 58 predicted damage of the T-section model at time 0.015 seconds after impact 

(4-ft drop test). The vertical flanges have delaminated and failed above the fillets. 

The same T-section model was executed with the reduced strength Mat 58 material model and 

test-analysis results are compared in Figure 103. The acceleration peak predicted by the model 

in Figure 103(a) is less at 170 g’s as compared to 250 g’s for the test. From Figure 103(c), the 

displacement of the simulation is much greater than the test value. Both the photogrammetry data 

and the displacement data obtained from double integration of the acceleration data were used 

for comparison with predicted displacement. Photogrammetry provides direct displacement data 

and does not require careful reduction of accelerometer data required to account for zero offsets 

and constants of integration; however, both experimental displacement curves are similar. 
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 (a) Acceleration responses. (b) Velocity Responses. (c) Displacement responses. 

Figure 103. Comparisons of test and Mat 58 reduced strength material model. 

A picture of model deformation is shown along with a photograph of the post-test article, in 

Figures 104(a) and (b), respectively. The vertical flanges of the model break just above the fillet 

as can be seen in Figure 104(a). However, the upper flanges slide over the lower flanges instead 

of wedging inside as occurred during the test, see Figure 104(b). Also, the rivets were not 

included in the model. From Figure 104(b), rivet-on-rivet contact contributed to lowering the 

displacement of the test article. 

 (a) Model deformation. (b) Post-test photograph. 

Figure 104. Comparison of post-test picture with Mat 58 deformed plot.

The T-section model of the 4-ft drop test was executed with the baseline Mat 58 properties; 

however, the ERODS parameter was varied from 0.1 to 0.5 in increments of 0.1. The 

deformation and failure of the model varied considerably with the ERODS value. Consequently, 

there is no uniformity of results and there is no convergence, as shown in Figure 105. Although 

the displacement response for an ERODS value of 0.1 most closely matches the test data, the 

actual deformed model, shown in Figure 106(a), failed catastrophically and does not compare 

favorably with the post-test damage. The best deformation results are obtained with an ERODS 

of 0.2. The deformed shape shown in Figure 106(b) is similar to the observed damage with both 

vertical flanges failing above the fillets. However, the damage near the top of the vertical web 

did not occur in the test. 
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Figure 105. Displacement responses of the drop mass for the 4-ft T-section test and Mat 58 

models with varying ERODS parameter. 

 (a) Deformed model for ERODS = 0.1 (b) Deformed model for ERODS = 0.2 

Figure 106. Mat 58 T-section model deformation for ERODS = 0.1 and 0.2. 

4.4 Simulation of Dynamic Crush Tests of Two Cruciform Sections 

4.4.1 Description of the Finite Element Model 

Two views of the cruciform section model are depicted in Figure 107. The model consists of 

16,916 nodes; 15,750 Belytschko-Tsay (ELFORM=Type 2) quadrilateral shell elements 

representing the cruciform; and, 512 solid elements representing the drop mass. Shell elements 

had a nominal element edge length of 0.125-in. Ten parts were defined, one part for the drop 

mass and nine parts for the cruciform, including four flange parts, two web parts, and three 

separate parts defining the center clips. Five material cards were defined including Mat 20 

(MAT_RIGID) for the impact mass. Four different Mat 54 and Mat 58 cards were required to 

ensure consistent and correct local material axis directions. These were input by setting AOPT = 
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2.0, which specifies that the local material axis directions are defined using vector input [5]. The 

density of the rigid material was selected to match the weight (215-lb) of the experimental drop 

mass. 

 (a) Canted front view with drop mass. (b) View of the model without the drop mass. 

Figure 107. Pictures of the cruciform model. 

Two SPCs were defined, one to allow the drop mass to move in the vertical direction only and 

one to fix the nodes of the bottom flanges. Finally, the nodes forming the drop mass were 

assigned an initial velocity to match the two test conditions, 192.5- and 268.6-in/s. These two 

impact velocities correspond to the 4- and 8-ft drop heights used during the tests. Output from 

the model included acceleration-, velocity-, and displacement-time histories of a central node on 

the drop mass, as well as pictures of model deformation. Models were executed using LSDYNA SMP Version 971 R6.0.0 with double precision and required up to 1 hour and 54 minutes 

of CPU on four processors for an end time of 0.02 seconds. 

4.4.2 Mat 54 Predicted Responses Compared with 4-ft Drop Test Data 

Results are shown in Figure 108 comparing the 192.5-in/s drop test data with the Mat 54 

predicted responses. The experimental and analytical acceleration responses shown in Figure 

108(a) were filtered using a low-pass Butterworth 1000-Hz filter. Also, displacement data are 

shown based on double integration of the acceleration response and photogrammetry. As with 

previous Mat 54 comparisons, the level of agreement between test and analysis for the Mat 54 

predicted responses is poor. The velocity responses in Figure 108(b) indicate that the model is 

removing some kinetic energy during the simulation; however, the predicted response is not 

close to the test and it never reaches zero velocity during the 0.02-s simulation time. Likewise, 

the model never achieves a maximum crush displacement during the 0.02-s simulation time. 

Deformation plots are illustrated in Figure 109 for the Mat 54 simulation for four time steps. 

The model fails in a brittle fashion, with no bending or plasticity shown in the webs. By 0.015-s, 

large pieces of the model including entire web sections have been separated from the upper and 

lower flanges. The predicted deformation pattern does not resemble the observed response, 

which is depicted in Figure 32. During the test, the webs exhibited bending and global bucking; 

however, no separations occurred between the flanges and webs. 

4.4.3 Mat 58 Predicted Responses Compared with 4-ft Drop Test Data 

Results are shown in Figure 110 comparing the 192.5-in/s drop test data with the Mat 58 

predicted responses. The acceleration responses shown in Figure 110(a) have been filtered using 
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a Butterworth 1000-Hz low pass filter. Also, displacement data are shown in Figure 110(c) from 

double integration of the acceleration data, as well as direct photogrammetric measurements. 

The test accelerations exhibit an initial oscillatory response with a high-magnitude peak of 

approximately 540-g. The predicted response matches the timing of the initial test peak, but is 

much lower in magnitude (167-g). The velocity responses, shown in Figure 110(b), indicate that, 

after 0.004-s, the model is removing velocity more quickly than the test. The model response 

crosses zero velocity 0.0035-s sooner than the test response. A sudden dip in the velocity 

responses occurs in both traces and the analysis curve shows a marked change in slope during the 

same time period. In Figure 110(c), the test data indicate a maximum crush depth ranging 

between 0.97- and 1.02-inches that occurs at approximately 0.0127-s. Photogrammetric results 

show a slightly higher maximum displacement of 1.07-in. However, the model predicts a lower 

maximum crush displacement of 0.74-in. that occurs earlier in time at 0.0094-s. 

 (a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 108. Experimental and Mat 54 analytical results for the cruciform 4-ft drop test. 

 (a) Time = 0.005-s (b) Time = 0.01-s 

 (c) Time = 0.015-s (d) Time = 0.02-s 

Figure 109. Mat 54 cruciform model deformation (velocity = 192.5-in/s). 

A sequence of model deformation is shown in Figure 111 for five time steps. The cruciform 

webs in the model exhibit bending and buckling. Few element failures are observed, except at 
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the interface of the vertical webs and the lower flanges. Though not highly evident in this 

sequence, the web buckling behavior produces a slight rotation of the top flanges near the end of 

the pulse. The model deformation compares well with the test response, which is highlighted in 

Figure 32. 

 (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 110. Experimental and Mat 58 analytical results for the cruciform 4-ft drop test. 

 (a) Time = 0.002-s. (b) Time = 0.004-s. (c) Time = 0.006-s. 

 (d) Time = 0.008-s. (e) Time = 0.01-s 

Figure 111. Mat 58 predicted cruciform model deformation sequence (velocity = 192.5-in/s). 

The cruciform section model was re-executed using the reduced strength Mat 58 material 

properties to simulate the 4-ft drop test (velocity=192.5-in/s). In addition, an ERODS study was 

conducted in which ERODS was varied from 0.1 to 0.5 in increments of 0.1. The results for 

these two parametric studies are shown in Figure 112. The reduced strength model shows better 

agreement with the test maximum crush displacement, 0.98-in. for the model compared with 

0.97- to 1.02-in. for the test, than the baseline Mat 58 material. The ERODS study indicates that 

a value of 0.4 provides the best agreement with test data. 
80

 (a) Reduced strength model. (b) ERODS study. 

Figure 112. Mat 58 parameter study results for the cruciform section (192.5-in/s velocity). 

4.4.4 Mat 54 Predicted Responses Compared with 8-ft Drop Test Data 

Results are shown in Figure 113 comparing the 268.6-in/s drop test data with the Mat 54 

predicted responses. The acceleration responses shown in Figure 113(a) have been filtered using 

a Butterworth 1000-Hz low pass filter. Also, displacement data are shown in Figure 113(c) from 

double integration of the acceleration data, as well as direct photogrammetry measurements. As 

with the lower velocity test, the experimental acceleration response exhibits a high spike 

initially, followed by a somewhat stable crushing phase. However, the model is not able to 

capture any feature of the test response. The velocity comparison, shown in Figure 113(b), 

indicates that kinetic energy is removed from the test and the drop mass crosses zero velocity at 

0.015-s. Conversely, the model shows that during the simulation time of 0.02-s, only 120-in/s of 

velocity was removed from the simulation. The displacement plot, shown in Figure 113(c) 

indicates that while the test curve exhibits a maximum crush stroke of 1.63-in., the predicted 

response continues to displacement vertically without abatement. 

Four depictions of the Mat 54 model deformation are shown in Figure 114 for the simulation of 

the 268.6-in/s vertical drop test of a cruciform specimen. As was observed for the lower velocity 

condition, the model fails in a brittle fashion, with no evidence of bending or buckling. By 0.01-

s, the interface elements between the web and flanges have failed and been removed from the 

simulation. These element failures allow the webs to separate from the flanges, and as the 

simulation continues, part-to-part contact forces the broken webs to rotate sideways. The 

predicted deformation pattern does not match the observed experimental results, which are 

depicted in Figure 37. 
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 (a) Acceleration. (b) Velocity. (c) Displacement. 

Figure 113. Experimental and Mat 54 analytical results for the cruciform 8-ft. drop test. 

 (a) Time = 0.005-s. (b) Time = 0.01-s. 

 (c) Time = 0.015-s. (d) Time = 0.02-s. 

Figure 114. Mat 54 cruciform model deformation (velocity = 268.6-in/s). 

4.4.5 Mat 58 Predicted Responses Compared with 8-ft Drop Test Data 

The cruciform model (see Figure 107) was re-executed with Mat 58 material properties for the 

higher impact velocity of 268.6-in/s, representing the 8-ft drop test condition. All other model 

details remained the same as for the lower velocity simulation. Experimental and analytical 

acceleration-, velocity-, and displacement-time histories are compared in Figure 115. Note that 

the acceleration curves shown in Figure 115(a) were filtered using a Butterworth 1000-Hz low 

pass filter. As with the lower velocity case, the test acceleration exhibits an oscillatory response 

with a high initial peak acceleration of approximately 550-g. The predicted response also 

exhibits an initial peak acceleration (163-g), which is much lower in magnitude than the test 

response. Following the initial spike, the two curves are close in magnitude and duration. 

Further verification of the level of agreement is seen in the velocity and displacement responses, 

shown in Figures 115(b) and (c), respectively. The sudden dip that was observed in the velocity 

response of the 4-ft drop test is also seen here. Once again, the analytical curve exhibits a 
82

significant change in slope early in the response and essentially averages the oscillations in the 

test response. The maximum predicted crush displacement is 1.68-in., which compares well with 

the test value of 1.63-in. Also, note that the photogrammetric results indicated a consistent 

maximum crush displacement of 1.63-in. 

 (a) Acceleration responses. (b) Velocity responses. (c) Displacement responses. 

Figure 115. Experimental and Mat 58 analytical results for the cruciform 8-ft. drop test. 

Model deformations are shown in Figure 116 for six time steps. These depictions indicate severe 

buckling and bending of the webs; however, few if any elements are removed from the model. 

Also, the integrity of the web-flange interface is maintained. The predicted deformation pattern 

is a reasonably good match to the test response, which is shown in Figure 37. 

 (a) Time = 0.002-s. (b) Time = 0.004-s. (c) Time = 0.006-s. 

 (d) Time = 0.008-s. (e) Time = 0.01-s. (f) Time = 0.012-s. 

Figure 116. Mat 58 cruciform model deformation sequence (velocity = 268.6-in/s). 

The cruciform section model was re-executed using the reduced strength Mat 58 material 

properties to simulate the 8-ft drop test (velocity=268.6-in/s). In addition, an ERODS study was 

conducted in which ERODS was varied from 0.1 to 0.5 in increments of 0.1. The results for 
83

these two parameter studies are shown in Figures 117(a) and (b), respectively. In this case, the 

reduced strength material shows a much higher maximum crush displacement than the test, e.g. 

2.25-in. for the model compared with 1.63-in. for the test. The ERODS study indicates that a 

value of 0.5 provides the best agreement with test data. 

 (a) Reduced strength model. (b) ERODS study. 

Figure 117. Mat 58 parameter study results for the cruciform section (268.6-in/s velocity). 

4.5 Simulation of the Longitudinal Impact of a Subfloor Section 

4.5.1 Description of the Finite Element Model 

Pictures of the finite element model of the subfloor section are shown in Figure 118. The model 

consists of 109,393 quadrilateral shell elements; 1,500 solid elements; 111,607 nodes; 15 parts; 

and, 5 material cards. A single contact definition was used to define contact between the impact 

mass and the subfloor, and self-contact within the subfloor, which is designated 

CONTACT_AUTOMATIC_SINGLE_SURFACE. Of the 15 parts specified in the model, all 

but two were defined as PART_COMPOSITE, which allows input of ply orientations, 

thicknesses, and material designations for each layer in a multi-layered composite laminate. A 

Mat 20 (MAT_RIGID) card was used for the impact drop mass and a Mat 24 

(MAT_PIECEWISE_LINEAR_PLASTICITY) card was used for the upper and lower aluminum 

support plates. All other parts were assigned either Mat 54 or Mat 58 material properties. For 

Mat 54 and Mat 58, autoclave material properties for unidirectional graphite tape were evaluated. 

Both models were executed using the fully integrated shell element formulation 

(ELFORM=Type 16). Models were executed using LS-DYNA SMP Version 971 R6.0.0 with 

double precision and required up to 129 minutes of CPU on four processors for an end time of 

0.1 seconds. The nominal element edge length is 0.25-inch. 

Nodes forming the drop mass were assigned an initial velocity of 264-in/s, matching the test 

condition. Two SPCs were defined, one to fix the bottom nodes of the lower support plate and a 

second to ensure that the drop mass could only move in the vertical direction. Model output 

included acceleration-, velocity-, and displacement-time histories of a center node on the drop 

mass, as well as pictures of model deformations. As indicated in Section 3.7.2, the only reliable 

data collected for the subfloor drop test was from photogrammetry of the base skinned panel, 

including vertical velocity and displacement responses. 
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 (a) Front view. (b) Rear view. (c) Side view. 

(d) Canted view. 

Figure 118. Pictures of the subfloor section model. 

4.5.2 Mat 54 Predicted Responses Compared with Test Data 

Comparisons of the Mat 54 predicted velocity- and displacement-time history responses with 

photogrammetric data obtained during the subfloor section impact test are shown in Figure 119. 

The photogrammetric velocity response, shown in Figure 119(a), was filtered using an SAE 

CFC60 low-pass filter. The Mat 54 analysis indicates that only a small amount of kinetic energy 

is removed during the simulation. The predicted response started at a velocity of -264-in/s and 

ended at a velocity of -200-in/s. Likewise, the predicted vertical displacement increases in a 

linear fashion, without abatement, throughout the duration of the simulation. 

The predicted deformation pattern of the subfloor model that was executed with Mat 54 material 

properties is shown in Figure 120. As with previous simulations conducted using the Mat 54 

material model, failure occurs in a brittle fashion, with large pieces of the structure breaking 

away as the model shatters like glass. The test article failed catastrophically through debonding 

of the T-stiffeners, tearing of the rear skin, and crushing of the I-beams, as described in Section 

3.7.2. However, in the test, sufficient energy absorption occurred to limit the displacement of the 

upper mass to approximately 13-inches by the time of contact with the honeycomb stops (0.071-

s). In the simulation, the drop mass displaced approximately 17-inches by 0.071-s. The 
85

honeycomb stops were not included in the model and the drop mass continued to displace 

vertically with no abatement, as brittle failure of the structure continued. 

 (a) Velocity responses. (b) Displacement responses. 

Figure 119. Comparisons of Mat 54 predicted velocity and displacement responses with PG. 

 (a) Time=0.018-s. (b) Time=0.036-s. 

 (c) Time=0.054-s. (d) Time=0.072-s. 

(e) Time=0.09-s. 

Figure 120. Pictures of Mat 54 subfloor model deformation. 
86

4.5.3 Mat 58 Predicted Responses Compared with Impact Test Data 

Comparisons of Mat 58 pre-test predicted velocity- and displacement-time histories with 

photogrammetric test data are shown in Figure 121. The predicted velocity response indicates 

that the model is too stiff and removes velocity much more quickly than the test. Likewise, the 

model predicts a maximum crush displacement of only 7-inches. Conversely, the test article 

displaces 13-inches, at which point the honeycomb stops are contacted. The pre-test simulation 

was executed using the baseline material properties for unidirectional graphite tape. The value 

of ERODS used in this material model proved to be too high to match the experimental 

deformations and failures. Pictures of model deformations are shown in Figure 122. 

Whereas the Mat 54 simulation predicted too much damage to the subfloor, the pre-test Mat 58 

simulation predicts too little damage. In addition, the type of predicted damage, which is mostly 

uniform crushing, does not match the test. The rear skin, I-beams, and T-stiffeners in the upper 

half of the section exhibit some out-of-plane bending and localized crushing; however, the lower 

half of the section remains undamaged. 

 

 (a) Velocity-time history. (b) Displacement-time history. 

Figure 121. Comparison of experimental and Mat 58 analytical responses for the subfloor test. 

As with previous simulations, two parametric studies were conducted. The first involved reexecuting the simulation using the reduced strength Mat 58 material model in which XC, XT, 

and E11T were lowered. As a reminder, XC was reduced by 25%; XT was reduced by 25%; 

and, E11T was reduced by 35%. Next, a study was conducted using the baseline Mat 58 material 

properties in which the ERODS parameter was varied from 0.1 to 0.5 in increments of 0.1. The 

results of these two studies are shown in Figures 123 and 124. 

The reduced strength model was not successful in generating improved prediction of the velocity 

response or the maximum crush displacement, as shown in Figures 123(a) and (b). The results, 

shown in Figure 124, indicate that ERODS values of 0.5, 0.4, and 0.3 provide similar velocity 

and displacement responses with maximum displacement values ranging from 7- to 8-inches. 

However, the model executed with ERODS=0.1 shows excellent agreement with test data up to a 

crush displacement of 10-inches. 
87

 (a) Time=0.02-s. (b) Time = 0.04-s. 

 (c) Time=0.06-s. (d) Time = 0.08-s. 

Figure 122. Pictures of Mat 58 subfloor model deformation. 

 (a) Velocity-time history. (b) Displacement-time history. 

Figure 123. Mat 58 reduced strength parameter study results for the subfloor section. 
88

 (a) Velocity-time history. (b) Displacement-time history. 

Figure 124. Mat 58 ERODS parameter study results for the subfloor section. 

Since the Mat 58 model with ERODS=0.1 gave much improved prediction of both the velocityand displacement-time history responses obtained during the subfloor impact test, the 

deformation pattern for this simulation is shown in Figure 125. This simulation exhibits 

considerably more damage than the model that was executed with baseline properties that 

include a relatively high value of ERODS (see Figure 122). Pieces of the model are failed and 

separated from the main model, as crushing initiates from the upper edge. Damage appears to 

progress from the upper edge of the subfloor to the mid-section in a stable and fairly uniform 

fashion. By 0.1-s, the upper half of the section has been completely removed from the model. 

  (a) Time = 0.02-s. (b) Time = 0.04-s. (c) Time = 0.06-s. 

  (d) Time = 0.08-s. (e) Time = 0.1-s. 

Figure 125. Pictures of Mat 58 (ERODS=0.1) subfloor model deformation. 
89

Once again, the predicted behavior is not truly representative of the damage seen during the test, 

though this model (Mat 58 with ERODS=0.1) comes as close as any subfloor simulation that was 

executed. Part of the explanation is that neither the Mat 54 or Mat 58 models were setup to 

capture debonding or delamination. One of the first events that occurred during the test is that 

the T-stiffeners debonded from the base skin. These parts were almost completely undamaged, 

as they popped off of the subfloor. Other localized areas of delamination were noted. In general, 

the effect of debonding and delamination is to lower the stiffness of the structure in either a 

global or localized manner. For example, the loss of the T-stiffeners reduced the overall 

longitudinal stiffness of the subfloor, while delaminations reduce the stiffness and strength of 

parts in the localized region where they occur. While T-stiffener debonding could be simulated 

fairly easily through model modifications, it is a much more challenging task to prepare the 

model to accurately represent the location and amount of delamination. Note that stiffener 

debonding simulations were not attempted due to constraints on time and resources. 

4.5.4 Multi-Dimensional Model Calibration 

One of the research areas pursued under the SRW Rotorcraft Crashworthiness program has been 

multi-dimensional model calibration [28-31]. The process of model calibration involves 

reconciling differences between test and analysis. Most calibration efforts combine both 

heuristics and quantitative methods to assess model deficiencies, to consider uncertainty, to 

evaluate parameter importance, and to compute required model changes. Calibration of large 

structural models presents particular challenges because the computational burden limits the 

number of solutions obtainable in a timely manner. Oftentimes, efforts are focused on predicting 

responses at critical locations as opposed to assessing the overall adequacy of the model. As 

documented in this report thus far, calibration efforts typically rely on scalar deterministic 

metrics, such as test-analysis comparison of acceleration-, velocity-, and displacement-time 

history responses. 

Multi-dimensional calibration of the subfloor model concentrated in two areas: 1) processing of 

photogrammetric images to augment the number of evaluation points for model calibration, and 

2) conducting parametric studies of the LS-DYNA subfloor model. As a reminder, all 

accelerometer data from the test were lost, as well as the photogrammetric data from the camera 

covering the open side of the subfloor. Consequently, only photogrammetric data from cameras 

viewing the base skinned side of the subfloor were collected. As noted previously, Mat 58 pretest simulations grossly under predicted the vertical crushing of the impact mass and the amount 

of damage seen in the test. Consequently, this model was a prime candidate for application of 

multi-dimensional model calibration. 

One complication with calibration efforts of nonlinear models in general is the lack of 

universally accepted metrics to judge model adequacy. References 32 and 33 are two 

noteworthy efforts that provide users with metrics to evaluate nonlinear time histories. However, 

these metrics are seldom used to assess model adequacy. In addition, the metrics, as stated in 

References 32 and 33, do not consider the multi-dimensional aspect of the problem explicitly. A 

more suitable metric for multi-dimensional calibration used in the present work exploits the 

concept of impact shapes, as proposed in Reference 34 and demonstrated in Reference 35. The 

approach used to perform subfloor model calibration focused on parameter uncertainty 

propagation and quantification. 
90

For calibration purposes, certain parameters are assumed to be uncertain and are arbitrarily 

assigned a uniform distribution function. This uncertainty model is used to create a family of N

equally probable parameter vectors, where N is arbitrarily selected. The analyst needs to know 

the probability of being able to reconcile measured data with predictions, given a particular 

model for the structure and parameter uncertainty. To this end, let ܳሺݐǡ ݌ሻ ൌ ԡݒሺݐǡ ݌ሻԡଶ be a 

scalar time varying function, in which the response vector v is used to compute the 2-norm at 

time t, using parameter vector p. Furthermore, let ߪ ൌ ݉݅݊׊௣ܳሺݐǡ ݌ሻbe the minimum value over 

all parameter variations, and let ߪ ൌ ݉ܽ׊ݔ௣ܳሺݐǡ ݌ሻ be the maximum value. Using these 

definitions and N LS-DYNA solutions corresponding to equally probable parameter vectors, a 

calibration metric is established to bound the probability of predicting different values of ܳሺݐǡ ݌ሻ

as; 

ܯଵሺݐሻ ൌ ܲ݋ݎܾ൫ߪ ൏ ܳሺݐሻ ൏ ߪ൯ ب ேିଵ

ே (1) 

A corollary to this probability statement indicates that the probability of the model to explain 

experimental observationsܳ௘ሺݐሻ (in terms of 2-norm of responses) outside the analysis range is 

that less than 1/N. Note that N controls the tightness of the bounds and also the number of LSDYNA solutions required. 

The use of norms, although convenient, tends to hide the spatial relationships that exist between 

responses at different locations in the model. In order to study this spatial multi-dimensional 

dependency explicitly, a different metric must be established. 

Spatial multi-dimensional dependency of models has been studied in classical linear dynamic 

problems in terms of mode shapes or eigenvectors resulting from a solution to an eigenvalue 

problem. An efficient and compact way to study the spatial relationship is by using a set of 

orthogonal basis vectors. These basis vectors, referred to as impact shapes, are computed by 

decomposing the time histories using singular value decomposition (SVD). For example, time 

histories from analyses and experiments are decomposed using SVD as: 

ݕሺݔǡ ݐሻ ൌ σ ߪ௜߮௜݃௜ሺݐሻ ௡

௜ୀଵ (2) 

In this form, the impact shape vector ߮௜ sized m x 1, contains the spatial distribution information 

for m sensors, g(t) contains the time modulation information, ı contains scalar values with shape 

participation factors, and n is the number of impact shapes to be included in the decomposition, 

often truncated based on allowable reconstruction error. Although Eq. (2) is written in 

continuous time form, for most applications, time is sampled at fixed intervals such that ݐ ൌ ݇οܶ

where the integer k=0,...,L and οܶis the sample time. From Eq.(2), the fractional contribution of 

the i

th impact shape to the total response is proportional to ߜ௜, defined as: 

 

௡ ௜ߪ σ Ȁ௜ߪ ൌ௜ ߜ

௜ୀଵ (3) 

Similar to the approach used in classical dynamic problems, impact shapes can now be used to 

compare models using orthogonality. Orthogonality, computed as the dot product operation of 

vectors (or matrices), quantifies the similarities or projection between two vectors. If the 

projection is zero, vectors are orthogonal, i.e., dissimilar. This same idea applies when 
91

comparing test and analysis impact shapes. Numerically, the orthogonality metric is computed 

as: 

ܯଶ ൌ Ȱෙ ்Ȳෙ (4) 

where Ȱෙ is sized m x l with l impact shapes at m locations computed from test data and Ȳෙ, sized 

m x l, are shapes computed using simulation data. Note that both Ȱෙ and Ȳෙ are normalized 

matrices such that Ȱෙ ்Ȱෙ ൌ ܫ and Ȳෙ ்Ȳෙ ൌ ܫ, where I is the identity matrix. Because individual 

impact shape vectors are stacked column-wise, metric M2 is a matrix sized l x l with diagonal 

values corresponding to the vector projection numerical value. If vectors are identical then their 

projection equals 1. Consequently, when evaluating models, multi-dimensional closeness with 

experiment is judged based on similarity of impact shapes and shape contributions. Application 

of multi-dimensional calibration techniques to the subfloor is discussed next. 

Based on previous efforts, it was concluded that using a limited number of target points (25 

targets) from photogrammetry was not sufficient to understand the behavior of the subfloor. 

Because the entire back panel of the subfloor was measured using photogrammetry, a task was 

undertaken to process the images to produce a large subset of measured points. This process 

evaluated images and removed targets that were not in all images throughout the time span 

needed, transformed the photogrammetric data into the LS-DYNA coordinate system, conducted 

interpolation of time histories to correct for anomalies due to photogrammetric drop-outs, and 

mapped the resulting target displacements to nodes in the LS-DYNA model. A y-z grid overlay 

is shown in Figure 126 of the subfloor back panel photogrammetry targets (dot-blue) versus LSDYNA evaluation nodes in (+ red). The algorithm to map targets to LS-DYNA nodes only kept 

photogrammetric targets within 1 inch of LS-DYNA node locations. Voids in the images 

showed up as white spaces in Figure 126. After the mapping process was completed, the 

photogrammetric displacement time histories had to be screened for anomalies manually. When 

the processing was completed, 3.6-ms of data was retained for 214 targets. This expanded data 

set allowed for a significantly better understanding of the panel behavior. 

Figure 126. Photogrammetric targets and LS-DYNA evaluation nodes. 
92

With the expanded photogrammetric data set at hand, calibration of the subfloor model 

proceeded by conducting parametric studies of the subfloor model to determine parameters to 

reconcile model with test. As discussed previously, this process uses bounds of the subfloor 

displacement response as a metric to assess the ability to reconcile the model. Many parametric 

studies were completed to evaluate the effect of varying thicknesses of several components, 

impact angles, and longitudinal/transverse modulus. The latest set of parameters, shown in Table 

7, included thicknesses of various parts, longitudinal and transverse moduli, and the pitch angle. 

These parameters are assumed to be equally valid within the bounds defined in the table. 

Combinations of these parameters were used to conduct 40 LS-DYNA runs, using Mat 58 

baseline material properties for unidirectional graphite tape. For each LS-DYNA run, the 

displacement magnitude norm was computed for the 214 targets matched from the 

photogrammetric data. The upper and lower displacement bounds are shown in Figure 127 from 

40 LS-DYNA runs (dashed blue) plotted along with the displacement magnitude computed from 

photogrammetric data (solid red). For regions where the analysis bounds encompass the test 

there is a probability greater than 1/40 that a parameter set exists that would reconcile test with 

analysis. However, finding such set is a very difficult task that requires optimization. 

Table 7. Parameters Used in Uncertainty Studies of the Subfloor Model. 

Variable 

Number 

Parameter 

Description 

Nominal 

Value 

Lower 

Bound 

Upper 

Bound 

Calibrated 

Value 

1 EA (psi) 19,000,000 14,000,000 19,000,000 17,792,075.6

2 EB and EC (psi) 1,400,000 840,000 1,400,000 1,042,798.33

3 Thickness of beam cap 

and T-stiffener (in.) 0.006 0.004 0.007 0.004 

4 Thickness of beam 

web, frame (in.) 

0.006 0.004 0.007 0.006 

5 Pitch angle, (degree) 0 -0.8 -0.3 -0.7 

Figure 127. Displacement bounds metric M2 from 40 LS-DYNA runs and photogrammetric test 

data. 

0 0.5 1 1.5 2 2.5 3 3.5

0

2

4

6

8

10

Time (msec)

Displacement 2-Norm (in)

Test

Upper Bound 40C

Lower Bound 40C
93

To evaluate metric M2 for the subfloor, the orthogonality is computed based on the magnitudenormalized impact shape vectors from test and analysis. For the subfloor problem, each impact 

shape vector is sized 642 x 1, which corresponds to data from 214 target locations in 3 

directions. Orthogonality values computed for the first 10 impact shapes for both test and the 

nominal model are shown in Figure 128. Along the abscissa are values corresponding to test 

impact shapes, top labels are the individual shape contribution to the response, whereas the 

ordinate shows the corresponding predicted analysis contribution. The numerical values for the 

colors used are shown on the color bar to the right with the highest orthogonality value less than 

0.6. For the case shown, the 1st impact shape predicted using LS-DYNA explains 58% of the 

data whereas for test, the 1st impact shape explains 45%. Although the orthogonality values are 

low, these results are significantly better than previous results. In particular, with the addition of 

more photogrammetric targets the dominant impact shape appears first in both LS-DYNA and 

test. The next step in the calibration process is to use optimization. 

Figure 128. Orthogonality of impact shapes from the PG data and the Mat 58 model using the 

baseline parameter set. 

Before proceeding with the optimization, it is important to understand and quantify the 

contribution from each parameter to the displacement variance. To judge parameter importance, 

a sensitivity analysis was conducted, using a variance-based approach, which produced the 

results shown in Figure 129. Parameter variance contributions are shown at each time step for 3-

ms of simulation. Parameters labeled P1-P5 in Figure 129 correspond to those listed in Table 7. 

At each time, the sum total of the parameter variance contribution adds up to a value near 1. 

From Figure 129, note that P5 (pitch angle) had the highest contribution to the variance followed 

by the longitudinal modulus EA. 

Optimization of the subfloor model proceeded by using the five parameters in Table 7 to search 

for a reconciling solution. The optimizer sought to minimize the prediction error between 

displacements from LS-DYNA and test. Because execution time of LS-DYNA models is large, 

the optimizer was only allowed to compute 25 additional runs. Parameters for the best solution 

1 2 3 4 5 6 7 8 9 10

0.58

0.13

0.08

0.03

0.02

0.02

0.02

0.01

0.01

0.01

Test 1

Case 1

MAC 

 

Frequency (Hz)

0.450.170.060.050.030.030.020.020.020.02

0.1

0.2

0.3

0.4

0.5




Ǧ








94

are also listed in Table 7. In addition, a comparison of the norm magnitude using calibrated 

model parameters versus test is plotted in Figure 130. Although a region still exists where the 

calibrated model over predicts the measured values, results are significantly better than when 

using the baseline parameters. Note that the norm values for the 214 targets compared indicate 

that the largest error within the first 3-ms is less than 1-in., which is the maximum difference 

between the curves shown in Figure 130. On the other hand, the orthogonality of the impact 

shapes using the calibrated parameter set is shown in Figure 131. When compared to Figure 128, 

orthogonality values are higher for the dominant impact shapes. Furthermore, the 1st impact 

shape from LS-DYNA explains 50% of the response as compared to 45% from test, as shown in 

Figure 131. 

 

Figure 129. Sensitivity analysis of the subfloor model. 

Figure 130. Norm magnitude comparison of test versus calibrated model. 

0 0.5 1 1.5 2 2.5 3 3.5

0

1

2

3

4

5

6

7

Time (msec)

PV Bounds

Test

Calibrated

Ǥ

ʹǦ


95

Figure 131. Orthogonality of impact shapes from photogrammetry and LS-DYNA model data 

using calibrated parameter set. 

4.6 Simulation of a Framed Fuselage Section 

4.6.1 Description of the Finite Element Model 

Sikorsky provided an updated 2-frame fuselage model to NASA in September 2012 that is 

illustrated in Figure 132. Although some of the test article was constructed of unidirectional 

tape, the primary material used for construction of the framed fuselage sections and subfloor was 

a plain weave thermoset graphite fabric. The finite element model consisted of two frames plus 

associated skin and beam structure in the upper and lower frame sections. The actual section 

was cut from the forward end of the post-test SARAP TVA that was drop tested at NASA 

Langley in 2008 (see Figure 2). 

Sixty-nine PART_COMPOSITE cards were used to define the layer-by-layer thickness, 

orientation, and material assigned to plies within each composite laminate in the model. Some of 

the composite layups had over 100 plies. The light green solid volume attached to the top of the 

upper fuselage structure is intended to simulate the entirety of the upper loading mass (~5100 

lbs.) and facility hardware (portal frame). This part was originally connected to the upper 

fuselage structure using CONTACT_TIEBREAK_SURFACE_TO_SURFACE. The empty 

weight of the fuselage was 368 pounds. An illustration is provided in Figure 133 showing the 

individual parts that are assigned to the material models (unidirectional tape and fabric) used in 

the framed fuselage section. 

In order to securely attach the fuselage structure to the portal frame, aluminum blocks were 

machined to fit the contour of the upper surface along the front and rear cross member. These 

aluminum fixtures are shown in Figure 134 prior to and during the bonding process. Highstrength adhesive was used as the bonding agent. Due to the existence of the Hy-Lok fasteners 

protruding from the top of the fuselage structure a series of spot-face reliefs (holes) were made in 

the aluminum beams to accommodate the rivet heads. These reliefs were intended to provide for 

1 2 3 4 5 6 7 8 9 10

0.50

0.19

0.07

0.04

0.03

0.03

0.02

0.02

0.01

0.01

Impact Shape Number

Calibrated Model Contribution

Test Impact Shape Contribution

0.450.170.060.050.030.030.020.020.020.02

0.1

0.2

0.3

0.4

0.5

0.6




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100

noted previously, a strip of material was discovered on the bottom of the fuselage section that 

was located on one side only. This extra material caused asymmetry of the fuselage section and 

may have contributed to the rigid body rotation observed in the static test results. From the data, 

the rotation angle was computed to be 0.44 degrees. If, in fact, the measured deformation was 

due to rigid body rotation, the elongation computed between any two points on the fuselage 

should be within the photogrammetric variability. To evaluate the amount of fuselage 

elongation, all 539 targets were used to compute elongation values between any two points. 

From the photogrammetric data, the mean elongation value across any two pair of targets was 

0.0184-inches and the maximum value was 0.0635-inches. Because no rotation of the fuselage 

under static load was predicted, it is suspected that differences in the boundary conditions 

between test and analysis caused the fuselage to rotate. In particular, if the fuselage was loaded 

while initially tilted, as the photogrammetric data seems to suggest, a static load would cause an 

additional rotation. Without further analysis, data from the static test cannot be used to estimate 

fuselage stiffness. As a result, the static test was not simulated using the Mat 58 material model. 

 (a) Predicted displacement. (b) Side view of PG data. 

Figure 139. Mat 54 predicted displacement of the upper portal mass for a slowly applied gravity 

loading and side view of photogrammetric data from the static test. 

4.6.3 Comparison of Modal Test Results with Analysis 

As described in Section 3.8.1, modal vibration tests were performed on the fuselage section prior 

to impact testing. Test-analysis results are shown in Figure 140 as orthogonality between the 

identified vibration mode shapes and the LS-DYNA predicted mode shapes. Readers should 

note that although orthogonality is used again, in this case the model comparisons are in terms of 

vibration modes and not impact shapes, as described earlier. This comparison is a validation of 

the linear behavior of the fuselage when using Mat 58 material properties. No results for Mat 54 

are shown. Listed along the ordinate of Figure 140 are the frequencies for the identified modes 

and along the top abscissa are the predicted frequencies using the baseline model parameters. 

Circled in red are the two target modes based on pre-test predictions. All orthogonality results 

are weighted using the reduced mass matrix M such that the orthogonality matrix is computed as 

ܱൌȰௗ

்ܯȰ௧, where Ȱௗare the mode shapes from an explicit solution from LS-DYNA and Ȱ௧

are the mode shapes from the modal test. Pairing of test and analysis modes was done based on 
101

orthogonality values (e.g. modes with orthogonality values near 1 are similar). In some cases, a 

single analytical mode matched several experimental modes. This finding is often due to too 

small a sensor count or problems with the data processing. For this case, if an analysis mode is 

paired to an identified mode(s), the analytical mode exists but perhaps was not well identified. 

As shown in Figure 140, the first analysis target mode at 73-Hz appears to be similar (i.e. high 

orthogonality value) to identified modes 39.32- and 39.78-Hz. Furthermore, the LS-DYNA 

predicted mode at 408-Hz appears to be similar to identified modes in the 220-230-Hz range. 

Similarly, the second target mode at 144-Hz appears to be similar to an identified mode at 74.5-

Hz. Based on these results, it should be clear that the LS-DYNA model is consistently overpredicting the frequencies of the fuselage vibration modes and therefore changes to the model 

should seek to reduce all predicted frequencies. One obvious choice to reduce the model 

frequencies is to reduce the modulus of the material. 

Figure 140. Orthogonality of test versus LS-DYNA predicted modes. 

Because the second target mode had the highest orthogonality value, it was used as a reference to 

estimate changes required to get the frequencies to agree. Specifically, to get a frequency 

reduction from 144- to 74.5-Hz (~ 52%) one would need to reduce the modulus to about 27% of 

its original value. To test this assumption, the Mat 58 unidirectional tape and fabric material 

properties EA, EB, GAB, GBC, and GCA were all adjusted accordingly based on the required 

values. The orthogonality results using modes predicted with the updated LS-DYNA model 

versus test are shown in Figure 141. With this modulus adjustment, the first analysis target mode 

is now at 43.4-Hz as compared to the measured modes in the 39.32- to 39.78-Hz range. 

Similarly, the second target mode is now at 80.3-Hz as compared to a measured mode at 74.5-

Hz. Interestingly, the LS-DYNA mode at 230.9-Hz appears to be similar to modes in the 220- to 

230-Hz range. Consequently, a modulus reduction as initially suggested will greatly improve 

frequency agreement not only for the target modes but also for other system modes. Certainly, 

the modulus can be tweaked again to get a better frequency matching; however, that was not the 

objective of this exercise. Although the orthogonality values with the reduced stiffness are 
102

slightly smaller than those from the baseline model, frequency agreement is significantly better 

and thus provides a good starting point for model calibration. 

Figure 141. Orthogonality of test versus updated LS-DYNA predicted results. 

4.6.4 Mat 54 Predicted Responses Compared with Impact Test Data 

The Mat 54 predicted acceleration of a typical node on the upper mass, shown in Figure 142(a), 

exhibits very high amplitude, high-frequency oscillations even when filtered with a 300-Hz lowpass Butterworth filter. The acceleration response from an accelerometer located on the portal is 

plotted in Figure 142(a) for comparison. The magnitudes of the initial peaks are comparable. 

Likewise, the Mat 54 predicted acceleration of a node located on a 100-lb lead block is shown in 

Figure 142(b) along with data from an accelerometer mounted to the floor mass. The simulation 

overpredicts the magnitude of the initial peak acceleration; however generally good agreement is 

shown after the initial peak. 

 (a) Upper mass. (b) Lead block. 

Figure 142. Comparisons of predicted accelaration responses of the upper loading frame mass 

and a floor-mounted lead block from the Mat 54 fuselage section model with test. 
103

The velocity and displacement curves are much easier to compare as low-pass filtering is not 

required. The velocity response obtained by integrating the acceleration data from an 

accelerometer mounted on the upper mass is compared with the computed velocity from the Mat 

54 simulation in Figure 143(a). Unfortunately, the integrated velocity is hard to interpret as a 

small drift or offset in acceleration can produce a very large difference in the integrated response. 

Consequently, it is instructive to compare with velocity that is obtained by differentiating fullfield photogrammetric displacement data. However, since differentiation of digital data can 

produce large oscillations, it is best to smooth the curve before differentiating. A comparison of 

the predicted velocity from the Mat 54 fuselage section model with differentiated 

photogrammetric displacement data (velocity) is shown in Figure 143(b). In general, the Mat 54 

predicted velocity response agrees well with the test data and tends to fair through the test 

ocsillations. Note that the analysis was conducted to 0.12-seconds termination time. 

The static displacement of the portal frame was measured post-test to be approximately 7.25 

inches. As illustrated in Figure 144, the Mat 54 model predicted that the maximum dynamic 

displacement of the portal was slightly over 8.5 inches. A possible explanation for the difference 

is that there was likely some rebound of the portal from the maximum. In addition, the rollers 

may have bound slightly causing energy loss. However, the results shown in Figure 144 indicate 

that the model was more flexible and weaker than the test article. 

 

Finally, some qualitative assessments can be performed by comparing photographs and motion 

picture analysis with predictions of model kinematic behavior. A photograph that was captured 

from a high-speed video is shown in Figure 145(a). This picture shows the rebound of the floor 

section after it had completely broken loose from the fuselage. The Mat 54 predicted fuselage 

damage at a time slightly over 100-ms after impact is shown in Figure 145(b). The kinematic 

motion ans structural deformations of the test and analysis are nearly identical. 

 (a) Data from integrated accelerometer. (b) Data from differentiated photogrammetry. 

 

Figure 143. Comparison of Mat 54 predicted velocity of the upper mass with test data. 
104

Figure 144. Mat 54 predicted time history of the displacement of the upper mass test portal. 

 (a) Film-capture photo. (b) Mat 54 model prediction. 

Figure 145. Test and analysis deformations of the fuselage section at time=0.1-s. 

4.6.5 Mat 58 Predicted Responses Compared with Impact Test Data 

The fuselage section model, shown in Figure 136, was modified by changing the unidirectional 

tape and fabric material properties from Mat 54 to Mat 58, and the model was re-executed. A 

low value of ERODS (0.06) was used in the baseline model for both material types. A picture of 

the deformed fuselage model at a time of 0.15 seconds after impact is shown in Figure 146(a). 

This figure shows that the deformation and rebound of the floor closely match the actual test 

scenario, shown in Figure 146(b). 

The maximum displacement of the upper mass predicted by the Mat 58 model with 

ERODS=0.06 is slightly over 9 inches, which is greater than the displacement of 7.25 inches 

measured after the test. Consequently, a parametric study was performed to determine the 

influence of ERODS on the displacement of the upper mass. The results are shown in Figure 

147. ERODS was varied from a high of 0.1 to a low of 0.06. The results show that the 

displacement of the upper mass is bounded by ERODS values of 0.06 and 0.09. The value of 

0.075 provided the best match to photogrammetric displacement data. The displacement of the 

upper mass was measured post-test to be 7.25 inches. However, there is always some rebound 
105

and settling of a structure after an impact. Note that the photogrammetry data was collected 

from target number 1200, which is shown in Figure 148. 

 (a) Mat 58 model prediction. (b) Film-capture photograph. 

Figure 146. Mat 58 (ERODS=0.06) simulation of the fuselage drop test at a time 0.1 seconds 

after impact compared with test. 

Figure 147. ERODS study on the effect of upper mass displacement versus time. 

As a result of the ERODS study, the fuselage model was re-executed in which the only change 

was to increase ERODS from 0.06 to 0.075. The predicted displacement of the upper mass now 

agrees better with photogrammetric data, as shown in Figure 149. In addition, comparisons of 

predicted velocity with test data based on differentiated photogrammetric data and with the 

integrated accelerometer data are shown in Figures 150(a) and (b), respectively. The 

differentiated displacement (velocity) response in Figure 150(a) is for the target 1200 shown in 

Figure 148 at the upper left. Differentiation of the photogrammetric displacement to produce a 

velocity plot does produce considerable oscillations in the velocity. It is better to first smooth 

the photogrammetric displacement before differentiating. However, the velocity produced from 

photogrammetric data analysis is typically more accurate than the velocity produced from 

integrating accelerometer data due to large offsets or drifts that can occur. An acceleration offset 
or drift g

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107

free. Note that, since the floor does not completely break away, the amount of rebound for the 

ERODS=0.075 model is not as great as for the model with ERODS=0.06. 

A comparison of Mat 58 (ERODS=0.075) predicted and test acceleration-time histories for the 

upper mass and for a lead block mounted to the floor are shown in Figures 152(a) and (b), 

respectively. All acceleration curves were filtered using a 300-Hz Butterworth low-pass filter. 

As shown in Figure 152(a), the predicted and experimental acceleration responses of the upper 

mass are highly oscillatory from initial impact through 0.03-s. The Mat 58 simulation slightly 

over predicted the initial acceleration spike, but generally matched the test response quite well. 

Likewise, the initial peak acceleration of the lead mass response was slightly over predicted, as 

shown in Figure 152(b). 

 (a) Data from photogrammetry. (b) Data from integrated acceleration. 

Figure 150. Test-analysis velocity comparisons. 

Figure 151. Deformed plot for Mat 58 fuselage model (ERODS=0.075) at 0.1-s. 
108

 (a) Upper mass location. (b) Lead block on the floor. 

Figure 152. Comparison of test accelerations with Mat 58 (ERODS=0.075) predictions. 

4.6.6 Multi-Dimensional Calibration of the Mat 58 Impact Model 

Multi-dimensional model calibration was used exclusively with photogrammetric data to get a 

broad perspective of the overall behavior of the system. This work concentrated in three areas: 

1) processing of photogrammetry images to map targets to LS-DYNA, 2) conducting parametric 

studies of the LS-DYNA model to evaluate metrics M1 and M2, and 3) to determine a parameter 

set to reconcile impact model predictions with test. Although the number of photogrammetric 

targets collected was large, not all targets mapped to LS-DYNA nodes. In addition, not all 

targets were tracked during the entire test. For model multi-dimensional calibration, only the 

first 9-ms of data were selected, since the floor was completely attached during this time period. 

Note that this time window is significantly shorter than the 100-ms of test data that was 

compared with simulation results in the previous section. Processing of the photogrammetric 

images required removal of targets that were not in all images throughout the time span needed, 

transforming the photogrammetric data into the LS-DYNA coordinate system, conducting 

interpolation of time histories to correct for anomalies due to dropouts, and mapping the 

resulting target displacements to nodes in the LS-DYNA model. A y-z grid overlay of the LSDYNA nodes (dot-blue) versus photogrammetric targets in (o red) is shown in Figure 153. The 

algorithm mapped photogrammetric targets to LS-DYNA nodes within 2-in. of each other. 

With the photogrammetric data set in hand, calibration of the fuselage model proceeded by 

conducting parametric studies to determine parameters to reconcile the model with test. Instead 

of looking at individual target displacements, the process used the displacement metric M1 to 

assess the ability to reconcile the model. Specifically, the reconciliation distance between the 

model and test is gaged using a vector 2-norm. Simply stated, displacements from 584 targets in 

3 directions are used to construct a response vector with 1752 entries corresponding to the x, y, z

displacements as a function of time, i.e, ܳሺݐǡ ݌ሻ. At each time step, differences between 

photogrammetry and LS-DYNA model predictions were compared in terms of a vector 2-norm. 
109

Figure 153. Photogrametry targets mapped to LS-DYNA nodes. 

The first step in the calibration process was to determine a parameter set that when adjusted 

would reconcile the model with test observations. To find such a set, many parametric studies 

were completed which included the effects of varying the longitudinal modulus, transverse 

modulus, shear modulus, impact angles, impact velocity, and ERODS parameters. The final 

parameter set used in the study is shown in Table 8. Parameters were assumed to be uniformly 

distributed within the bounds defined in the table. A random sampling of these parameters was 

used to execute 120 LS-DYNA runs. For each LS-DYNA run, the displacement magnitude 

norm was computed for all 584 targets. Figure 154 shows the displacement upper and lower 

bounds from 120 LS-DYNA runs (dashed blue) plotted along with the displacement norm 

computed with photogrammetric data (solid red). Test and analysis reconciliation is only 

probable if the analysis bounds encompass the measured data. 

At this point, although the bounds are reasonably close, they do not encompass the entire 

measured data set. A parameter set to fully encompass the test results has not been found. 

Nonetheless, because the bounds are relatively close to the measured data, a reconciling solution 

using optimization was computed and reported in Table 8 under “Cal Solution”. Additionally, 

the displacement 2-norms are shown in Figure 155 for test (solid red), the baseline model 

(dashed-dot black), and the calibrated solution (dashed blue). At any point in time the maximum 

displacement error in x, y, or z is less than or equal to the difference in the norm between test and 

analysis. 

A final check of the model examined the orthogonality metric M2 (see Eq. 4) to assess model 

similarity in terms of impact shapes from test and analysis. A plot is shown in Figure 156 of the 

metric M2 using 5 impact shapes computed using photogrammetric data and those corresponding 

to the analysis when using the baseline parameters. The ordinate shows the fractional shape 

contributions for analysis, as defined by Eq. 3, and the test results are shown across the top. 

Note that the 1st impact shape from test contributes 73% to the response in contrast to 87% for 

-50 0 50

190

200

210

220

230

240

250

260

270 

Y (in)

Z (in)

LS-DYNA

PG targets
110

the analysis. Similarly, the orthogonality results are shown in Figure 157 using the calibrated 

parameters instead. It should be apparent that differences between the two models, i.e., baseline 

versus calibrated, are small indicating that the relative contribution of the impact shapes is not 

improved. More importantly, the model was only able to predict the dominant impact shape. 

Therefore, the more complex structural behavior measured by photogrammetry was not well 

predicted. 

Table 8. Parameters Used in Uncertainty Studies of the Fuselage Section Model.

Param. 

No. Properties Nominal Lower 

Bound 

Upper 

Bound 

Cal 

Solution 

1 EA (lb/in2) 1.00E+07 5.00E+06 2.50E+07 1.64E+07

2 EB (lb/in2) 1.00E+07 3.00E+05 2.50E+07 3.00E+05

3 GAB, GBC, GCA (lb/in2

) 7.40E+05 1.90E+05 1.05E+06 2.73E+05

4 ERODS 0.15 0.02 0.60 0.56 

5 Tx roll (degrees) 0.00 -0.50 0.00 -0.10 

6 Ty pitch (degrees) 0.00 0.00 0.50 0.00 

Figure 154. Displacement 2-norm for test and 120 LS-DYNA runs. 

0 2 4 6 8 10

0

10

20

30

40

50

Time (msec)

Displacement 2-Norm (in)

Test

Upper Bound 120C

Lower Bound 120C

Ǥ

ʹǦ

ȋȌ
111

Figure 155. Displacement 2-norm for test, baseline, and calibrated model. 

Figure 156. Orthogonality of impact shapes: baseline model versus test. 

1 2 3 4 5

0.87

0.07

0.02

0.01

0.01

Impact Shape Number

Analysis Impact Shape Contribution

Test Impact Shape Contribution 

0.730.080.050.040.03

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 1 2 3 4 5 6 7 8 9

0

5

10

15

20

25

30

35

40

45

Time (msec)

Disp. Norm (in)

Test

Calibrated

Baseline Model

 

Disp. 

2-Norm 

(in) 

 

 

 

 
112

Figure 157. Orthogonality of impact shapes: calibrated model versus test. 

5.0 DISCUSSION OF RESULTS 

5.1 Experimental Program 

One of the realities of conducting impact tests is that repeated tests on the same type of specimen 

are rarely possible. Consequently, most tests are a one-time only event, making quantification of 

experimental uncertainties difficult. As reported in this document, the only opportunity for 

repeated tests were using the laminated coupons, as described in Section 3.3, in which as many 

as seven coupons were tested. For the component tests, photogrammetric measurements assisted 

in bounding the uncertainty in the experimental responses. For example, target-tracking 

measurements enabled accurate determination of the velocity and attitude of the drop mass or the 

test article at the moment of impact. In addition, photogrammetric techniques were used to 

collect full-field strain contour plots, when strain gages failed. In the case of the subfloor test, all 

data were lost except for photogrammetric data from the skinned side of the subfloor. 

Fortunately, that information provided accurate velocity- and displacement-time histories that 

were used in conducting test-analysis comparisons. Finally, full-field strain and out-of-plane 

displacement data generated using photogrammetric techniques were able to capture localized 

regions of high strain and oil-canning modes of failure. 

A variety of composite failure modes were evident during testing including brittle fracture, 

crushing of the material, ply delamination, layer splitting, and localized bending. The I-beam 

specimens exhibited uniform crushing of the material from the top edge that was aided by the 

presence of a crush initiator. Conversely, the T-section components exhibited delamination and 

localized brittle fracture, with very little crushing of the material. In addition to material failures, 

structural failures occurred such as stiffener debonding and global/local buckling. As an 

1 2 3 4 5

0.92

0.03

0.02

0.01

0.01

Impact Shape Number

Anaysis Impact Shape Contribution

Test Impact Shape Contribution

0.730.080.050.040.03

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9
113

example, the T-stiffeners that were bonded onto the base of the subfloor separated almost 

immediately upon impact of the specimen. Likewise, the cruciform webs exhibited global 

buckling and crushing under impact loading. The ability to accurately model so many different 

structural deformation and composite failure modes is a major challenge for the current 

generation of nonlinear, explicit transient dynamic finite element codes. 

 

5.2 Analytical Program

One great disadvantage in conducting the analytical simulations was the lack of a complete set of 

material property data. Laminate characterization tests were performed using coupons that were 

cut from the SARAP hardware. Typically, the longitudinal modulus (EA) of a composite 

material is determined from tensile tests on 0° unidirectional tape coupons. Likewise, the 

transverse modulus (EB) is determined from tensile tests on 90° unidirectional tape coupons. 

For this research effort, the procedure used to determine Mat 58 material properties involved 

assessment of literature data, comparison with similar Mat 54 properties, and a trial and error 

process in which laminated coupons were tested and simulated under both tensile and 

compressive loading. Reverse engineering had to be used to derive fundamental material 

property values from tests on laminates. Often, one set of material properties that provided 

excellent comparison with coupon test data was too weak or perhaps too stiff when simulating 

more complex structures. Consequently, an iterative process was required. However, once a 

baseline set of material properties was established, the properties were fixed. Thus, the Mat 58 

material properties represent compromises, meaning that they are not optimized for any one 

simulation. In the end, the lack of complete material property data remains a major source of 

modeling uncertainty. 

Tests of all of the SARAP components were simulated using LS-DYNA shell-based finite 

element models. The shell elements were assigned either Mat 54 or Mat 58 material models to 

represent the behavior of the composite structure. In general, poor levels of agreement were seen 

when the Mat 54 material was used. Models that were assigned Mat 54 tended to fail 

catastrophically in a brittle fashion, absorbing little kinetic energy. As mentioned previously, 

many parametric studies were conducted to find a set of parameters that would predict stable 

crushing, as opposed to brittle failure and structural collapse. 

Reference 22 describes the application of Mat 54 in modeling the quasi-static crushing response 

of a composite sinusoidal-shaped component. The authors of Reference 22 demonstrate 

excellent prediction of composite crushing. Consequently, the Mat 54 model that was reported 

in Reference 22 was incorporated into the I-beam model and executed. This Mat 54 model had 

the DFAILM, DFAILS, DFAILT, and DFAILC parameters specified, as well as the TFAIL, 

SOFT, ALPH, FBRT, YCFAC, and BETA parameters. Unfortunately, even this Mat 54 

simulation failed to produce stable crushing of the I-beam component under dynamic 

compressive loading. 

All attempts to modify the Mat 54 material model to improve the crushing behavior failed, with 

one exception. An I-beam model was executed using Mat 54 autoclave properties, except that all 

failure strain inputs (DFAILM, DFAILS, DFAILT, and DFAILC) were undefined. The 

parameter EFS (Effective Failure Strain) was varied from 0.01 to 0.06 in increments of 0.01. 
114

The EFS parameter can be used instead of DFAILT, DFAILC, DFAILS and DFAILM. At each 

time step, LS-DYNA calculates an effective scalar strain for each integration point within an 

element and then compares the scalar strain to the EFS value. If the scalar strain is higher than 

EFS, then the integration point is deleted, regardless of loading condition. When all integration 

points have failed, the element is removed. Thus, EFS controls element erosion/deletion, and 

does not represent progressive failure as is governed by Chang-Chang in Mat 54. Using ChangChang, the elastic properties of plies within the laminate stacking sequence are degraded as part 

of the ply-by-ply failure process. When all plies within an element are deleted, then the element 

is removed from the simulation. 

Predicted acceleration-, velocity, and displacement-time history responses are compared with test 

data in Figure 158 for the I-beam model executed with EFS=0.06. The impact velocity for this 

simulation was 189.6-in/s. This model was able to generate predicted responses that matched the 

test data reasonably well, especially when compared with the original Mat 54 results, shown in 

Figure 84. The experimental and predicted acceleration responses, shown in Figure 158(a), were 

filtered using a 1000-Hz Butterworth low-pass filter. 

 (a) Acceleration. (b) Velocity.

 

(c) Displacement. 

Figure 158. Experimental and Mat 54 analytical results for the I-beam 4-ft drop 

(velocity=189.6-in/s) with EFS=0.06. 
115

However, despite the reasonably good comparison with time history responses, the predicted 

deformation pattern, shown in Figure 159, does not match the test. The model begins to crush 

initially from the top edge. By 0.0064-s, elements have failed along the web-flange intersection 

almost completely separating these two parts on both sides. Also, a large “crack” has opened 

running from the middle of the top edge of the web to the side flange, which separates a large 

chunk of the web from the specimen. By 0.024-s the failed portion of the web is removed from 

the model and no additional damage is seen. Conversely, the test article exhibits fairly uniform 

crushing of the specimen from the top edge, as shown in Figure 13. 

As described in Reference 22, the success of the analytical predictions using Mat 54 were based 

on utilizing a contact definition called CONTACT_RIGID_NODES_TO_RIGID_BODY. This 

contact allows direct input of a load versus penetration curve. The user-defined load-penetration 

curve can be adjusted to provide improved predictive responses. In discussions with the authors 

of Reference 22, they recommended selecting the rigid nodes to rigid body contact, and then 

conducting a parameter study using SOFT in the Mat 54 material model. The I-beam model was 

modified to include the new contact definition with input of a load-penetration curve similar to 

the one described in Reference 22. Unfortunately, no improvement in the predicted response 

was obtained and a SOFT parameter study was not attempted. 

 (a) Time=0.00-s (b) Time=0.0016-s (c) Time=0.0032-s (d) Time=0.0048-s 

 (e) Time=0.0064-s (f) Time=0.008-s (g) Time=0.0096-s (h) Time=0.024-s 

Figure 159. Deformation pattern of I-beam model executed with Mat 54 (EFS=0.06). 

Unlike Mat 54, Mat 58 incorporates parameters such as SLIMs and ERODS, which allow the 

analyst to incorporate some degree of plastic-like behavior into the model. This feature allows 

failed elements to carry small stresses, yet remain in the model. As seen when using Mat 54, 

once elements are removed from the simulation, holes are created in the model, resulting in 

stress concentrations. The model can quickly collapse. The only case in which the brittle failure 

predicted by Mat 54 matched the test response was for the fuselage section vertical drop test. 

The fuselage section failed catastrophically in the lower corners, which allowed the floor to 

completely separate from the sidewalls, as shown in Figure 145(a). The fuselage model with Mat 

54 was successful in predicting this response. 
116

In general, the capability of LS-DYNA to accurately predict all of the composite failure modes is 

inconsistent. Material crushing of the I-beam components was successfully simulated; however, 

debonding and delamination were not predicted. Current composite material models such as Mat 

54 and Mat 55 are not capable of predicting delamination based on either fracture mechanics or 

energy–based approaches. However, if the location of a potential delamination or debond is 

known, a model could be modified to account for this failure mode using either tiebreak contact 

definitions or spot weld connections [36, 37]. 

Within LS-DYNA, Mat 54 and Mat 58 offer opportunities to view “hot spots” in the model 

through output of history variables. In general, the history variables are calculated as the 

simulation progresses and are typically associated with a particular failure mode. The values can 

range from 0 (no failure) to 1 (imminent failure). Mat 54 allows output of 16 history variables 

and Mat 58 allows for 15. As an example, a fringe plot of History Variable 3 is shown in Figure 

160 for the T-section model that was executed with Mat 58 material properties. History Variable 

3 is an interactional strain based on maximum values of two in-plane strains and the shear strain. 

Figure 160. Fringe plot of History Variable 3 of the T-section model shown just prior to failure 

of the vertical flanges (t=0.001). 

Three “hot spots” shown in the plot of Figure 160 indicate the locations of imminent failure, 

which are the vertical flanges just above the fillets and the top edge of the vertical web. This plot 

was taken from a d3plot file using LS-Prepost. The very next d3plot file showed that element 

failures had occurred across the vertical flanges on both sides. In the future, using the History 

Variable plots as a guide, the hot spot regions could be modified to incorporate delamination. 

Recent advances in composite material model development have focused on three-dimensional 

ply failure models that utilize solid element representations of composite materials, as described 

in Reference 21. Also, Livermore Software Technology Corporation (LSTC), the company that 

markets LS-DYNA, has recently incorporated a user-defined material model (Mat 162) that can 

be purchased and used in LS-DYNA to represent composite materials as solid elements [38]. 

While these approaches were not evaluated in the present study, they offer some potential for 

future simulations. In addition, a joint ASU-FAA-LSTC-NASA-GWU research project was 

initiated in 2012 with the goal of producing a new material model in LS-DYNA that has the 

flexibility to simulate a wide range of properties, accumulated damage behavior, and different 

types of failure, which exist in composite materials. This model should successfully predict 
117

aircraft turbine engine containment threshold velocities based upon mechanical property tests, 

without relying on post-test correlation to large-scale system tests [39]. 

Both Mat 54 and Mat 58 contain parameters that cannot be measured in the laboratory. These 

parameters must be tuned through repeated test-analysis comparisons. Over time, an analyst 

may gain experience in selecting these parameters based on prior simulations. However, without 

relying on test data for comparison, it can be difficult to choose the correct values. For example, 

pre-test simulations for the subfloor longitudinal impact were executed using the baseline Mat 58 

material model. However, the baseline ERODS parameter for unidirectional tape was set at a 

relatively high value (close to 1.0), meaning that elements would not be removed until a high 

value of strain had been achieved. Consequently, the pre-test simulation of the subfloor impact 

predicted only 7-in. of vertical displacement of the drop mass, whereas the actual displacement 

was 13-in. The value of ERODS was too high, and a post-test parameter study was performed. 

The actual value that gave best correlation with test was ERODS=0.1. Consequently, it is fair to 

say that, while the capabilities of the current generation of explicit transient dynamic codes, such 

as LS-DYNA, may not always provide accurate pre-test predictions of the impact response of 

composite airframe structures, the test response can be bounded using analysis. 

Finally, multi-dimensional calibration techniques were applied to both the composite subfloor 

and fuselage section models in an attempt to assess model deficiencies and uncertainties, to 

evaluate parameter importance, and to compute required model changes needed to better 

reconcile test and predictions. The need for additional subsystem tests to reduce the uncertainty 

in joints, aggregate material properties, and geometry is extremely important if model validation 

is required for more than just one test condition. As an example, the fuselage section was tested 

under quasi-static loading, modal vibration, and vertical drop test conditions. These tests were 

conducted to provide additional opportunities for calibration of the global model. While results 

from the static test proved to be unusable, results from the vibration test of the composite 

fuselage indicated that the model needed to have the modulus reduced by 27% from the baseline 

value. However, results using impact data suggested an increase in modulus. Clearly, these two 

requirements are inconsistent; however, both are supported by experimental data. These 

conflicting results highlight the inadequacy of the existing model to be predictive under different 

loading conditions. Although, comparisons of the observed global behavior with analytical 

predictions are reasonable, attempts to use the analytical model to predict other tests conditions 

will likely fail. 

6.0 CONCLUDING REMARKS 

This report documents a research study to assess the capabilities of a nonlinear, explicit transient 

dynamic code, LS-DYNA, for predicting the impact response of composite airframe structures. 

The test specimens were extracted from a previously tested composite prototype fuselage section 

developed and manufactured by Sikorsky Aircraft Corporation under the US Army’s Survivable 

Affordable Repairable Airframe Program (SARAP). Laminate characterization testing was 

conducted in tension and compression. In addition, dynamic impact tests were performed on 

several components, including I-beams, T-sections, and cruciform sections. Finally, tests were 

conducted on two full-scale components including a subfloor section and a framed fuselage 

section. These tests included a modal vibration and longitudinal impact test of the subfloor 
118

section and a quasi-static, modal vibration, and vertical drop test of the framed fuselage section. 

Most of the test articles were manufactured of graphite unidirectional tape composite with a 

thermoplastic resin system. However, the framed fuselage section was constructed primarily of a 

plain weave graphite fabric material with a thermoset resin system. 

Due to the varying complexity of the test articles, several different composite failure modes were 

evident during testing including brittle fracture, delamination, material crushing, and localized 

bending. In addition, structural failure modes occurred such as stiffener debonding and 

global/local buckling. The ability to accurately simulate so many different failure responses is a 

major challenge for the current generation of nonlinear, explicit transient dynamic finite element 

codes. 

Typically, the test articles were instrumented with accelerometers and strain gages. Based on the 

results of the test program, strain gages tended to separate from the specimens immediately 

following impact and provided no information. In the longitudinal impact test of the subfloor, 

which was a severe impact event, even the accelerometers failed. In addition to more traditional 

instrumentation, all of the test articles were sprayed with a stochastic speckle pattern that was 

used in conjunction with photogrammetric techniques to capture full-field strain contour plots. 

During the subfloor test, all data were lost except for photogrammetric data from the skinned 

side of the subfloor. Fortunately, that information provided accurate velocity- and displacementtime histories that were used in conducting test-analysis comparisons. Photogrammetry was also 

able to provide accurate information on impact conditions, such as drop mass velocity and offnominal attitudes. Thus, photogrammetry proved to be an invaluable tool for this test program. 

Finite element models were developed to represent the laminate characterization tests; the impact 

tests of I-beam, T-section, and cruciform components; the longitudinal impact test of the 

subfloor; and the vertical drop test of the framed fuselage section. Sikorsky Aircraft 

Corporation, the manufacturer of the SARAP hardware, provided most of the models to NASA 

as part of a cooperative research agreement. Once received, the models were updated and 

modified to best represent the test conditions. Simulations were conducted using LS-DYNA 

version 971 R6.0.0. Within LS-DYNA, two composite material models were evaluated: Mat 54 

and Mat 58. Mat 54 is a progressive failure model that uses the Chang-Chang failure criterion to 

simulate ply-by-ply failure and property degradation, while Mat 58 is a continuum damage 

mechanics model for representing unidirectional tape and fabric composite materials. Mat 54 

material properties were provided by Sikorsky, whereas material properties for Mat 58 were 

determined by assessing literature data, comparing with similar Mat 54 property values, and 

using a trial and error process in which laminated coupons were simulated under both tensile and 

compressive loading. 

In general, the Mat 54 simulations did a poor job in simulating the impact responses of the 

components and the subfloor. These models tended to fail catastrophically in a brittle fashion, 

absorbing little kinetic energy. Numerous parametric studies were conducted to find a set of 

parameters that would predict stable crushing, as opposed to brittle failure and structural 

collapse. None of these attempts were successful. The only case in which the brittle failure 

predicted by Mat 54 matched the test response was for the fuselage section vertical drop test. 

The fuselage section failed catastrophically in the lower corners, which allowed the floor to 
119

completely separate from the sidewalls. The fuselage model with Mat 54 was successful in 

predicting the location and amount of damage to the fuselage section, even matching the 

kinematics of the floor following separation from the sides of the fuselage structure. 

Unlike Mat 54, Mat 58 includes parameters, which allow the analyst to incorporate plastic-like 

behavior into the model. This feature allows failed elements to carry small stresses, yet remain 

in the model. As seen when using Mat 54, once elements are removed from the simulation, holes 

are created and stress concentrations are generated, leading to numerical instabilities and element 

failures. The model can quickly collapse. Mat 58 models of the I-beam were able to accurately 

predict material crushing. Likewise, Mat 58 models of the cruciform sections predicted global 

buckling behavior of the webs under dynamic loading. Consequently, excellent agreement was 

seen for the Mat 58 model predictions with the I-beam and cruciform responses. 

However, during the impact test, the T-section component failed primarily through delamination. 

Likewise, the first failure event that occurred with the subfloor longitudinal impact test was 

debonding of the T-stiffeners from the base. Current composite material models such as Mat 54 

and Mat 55 cannot predict delamination or debonding without a priori knowledge of where it 

will occur. If the location of a potential delamination or debond is known, the model can be 

modified to account for this failure mode using tiebreak contact or spot welds. Whether this is a 

reasonable or practical approach for modeling debonding or delamination was not evaluated in 

this study. 

Because both Mat 54 and Mat 58 contain parameters that cannot be measured in the laboratory 

these parameters must be tuned through repeated test-analysis comparisons. Over time, an 

analyst may gain experience in selecting these parameters based on prior simulations. However, 

without relying on test data for comparison, it can be difficult to choose the correct values. As a 

result, these analysis methods cannot be considered truly predictive. However, it is fair to say 

that, while the capabilities of the current generation of explicit transient dynamic codes such as 

LS-DYNA may not provide accurate pre-test predictions of the impact response of composite 

airframe structures, the test response can be bounded using analysis. 

Finally, in this report, model calibration efforts relied primarily on scalar deterministic metrics, 

such as test-analysis comparisons of acceleration, velocity-, and displacement time-histories at 

critical locations (e.g. comparisons of test-analysis acceleration responses of the drop mass 

during a component impact test). In addition, multi-dimensional model calibration techniques 

were applied to the subfloor and fuselage section models in an attempt to reconcile differences 

between test (photogrammetry data) and analytical predictions. Two metrics were evaluated for 

both simulations including the displacement 2-norm and the orthogonality of impact shapes. 

These techniques were successful in assessing the overall adequacy of the model based on 

parameter uncertainty propagation and quantification, and in generating a calibrated model that 

minimized differences in test and predicted responses. 

7.0 ACKNOWLEDGEMENTS 

The authors of this report gratefully acknowledge Charles Clarke, Brian Wilson, and Joe Shen of 

Sikorsky, who helped establish the Space Act Agreement between NASA and Sikorsky, 
120

provided finite element models, and offered invaluable assistance in planning and executing the 

test and simulation program. We also recognize the work of Michael A. Polanco, who 

performed many preliminary simulations. In addition, we thank the NASA Aeronautics Rotary 

Wing Program Office for supporting this work financially. Also, Joe Zalameda, Jim Bly, and Pat 

Johnston of the Nondestructive Evaluation Sciences Branch of NASA Langley performed the 

Non Destructive Evaluation. Finally, we would like to acknowledge the team of technicians and 

contractors at the LandIR Facility who provided technical support for conducting the tests, 

mounting instrumentation and operating the data acquisition system, and setting up cameras and 

high-speed video systems. These technicians included: Pat Quander-Haas, Doug Prochnow, Paul 

Bagby, Nelson Seabolt, Richard Wheeler, Dave White, and Steve Sieder. 

8.0 REFERENCES

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Prototype Airframe Design as Part of the Survivable Affordable Repairable Airframe Program,” 

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2. Non-Reimbursable Space Act Agreement between Sikorsky Aircraft Corporation and NASA 

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of Composite Airframe Structures, SAA1-1122, signed May 5, 2011. 

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Aeronautics Research Program in Rotorcraft Crashworthiness,” Journal of Aerospace 

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971, Livermore Software Technology Company, Livermore, CA, August 2006. 

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8. Military Standard, MIL-STD1290A(AV),Light Fixed- and Rotary-Wing Aircraft Crash 

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109039, Sept. 1996. 

11. Mongiardini M., Ray M.H., Anghileri M., “Development of a Software for the Comparison 

of Curves During the Verification and Validation of Numerical Models,” Proceedings of the 7th

European LS-DYNA Conference, Salzburg, Austria, May 14-15, 2009. 

12. Marzougui D., Kan C-D., and Opiela K. S., “Comparison of the Crash Test and Simulation of 

an Angle Impact of a 2007 Chevrolet Silverado Pick-Up Truck into a New Jersey-Shaped 

Concrete Barrier for MASH Conditions,” National Crash Analysis Center NACA 2010-W-001 

Working Paper, July 2012. 

13. Jackson K. E., Fasanella E. L., Annett M. S., and Polanco M. P., “Material Model Evaluation 

of a Composite Honeycomb Energy Absorber,” Proceedings of the 12th International LS-DYNA 

Users Conference, Dearborn, MI, June 3-5, 2012. 

14. Sprague M.A., and Geers T. L., “Spectral Elements and Field Separation for an Acoustic 

Fluid Subject to Cavitation,” Journal of Computational Physics, Vol. 184, 2003, pp: 149-162. 

15. Schwer L.E., “Validation Metrics for Response Histories: Perspectives and Case Studies.” 

Engineering with Computers,” Engineering with Computers (2007), 23:295-309. 

16. Chang, Fu-Kuo, and Chang, Kuo-Yen, “Post-Failure Analysis of Bolted Composite Joints in 

Tension or Shear-out Mode Failure Mode,” Journal of Composite Materials, Vol. 21, No. 9, 

September 1987, pp. 809-833. 

17. Chang, Fu-Kuo, and Chang, Kuo-Yen, “A Progressive Damage Model for Laminated 

Composites Stress Concentrations,” Journal of Composite Materials, Vol. 21, No. 9, September 

1987, pp. 834-855. 

18. Tsai, S. W. and Wu, E. M., “A General Theory of Strength for Anisotropic Materials,”

Journal of Composite Materials, Vol. 5, 1971, pp. 58–80. 

19. Hahn H., and Tsai S.W., “Nonlinear Elastic Behavior of Unidirectional Composite 

Laminate,” Journal of Composite Materials, Vol. 7, 1973, pp. 102–110. 

20. Hahn H., and Tsai S.W., “Nonlinear Behavior of Laminated Composites,” Journal of 

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21. Donadon M.V., de Almeida S.F.M., Arbelo M.A., and de Faria A.R., “A Three-Dimensional 

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22. Feraboli P., Wade B., Deleo F., Rassaian M., Byar A., “LS-DYNA MAT54 Modeling of the 

Axial Crushing of a Composite Tape Sinusoidal Specimen,” Composites: Part A, Vol. 42 

(2011), pp. 1809-1825. 

23. Matzenmiller A., Lubliner J., and Taylor R. L., “A Constitutive Model for Anisotropic 

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24. Xiao X., “Modeling Energy Absorption with a Damage Mechanics Based Composite 

Material Model,” Journal of Composite Materials, Vol. 43, No. 5, 2009, pp. 427-444. 

25. Maekawa Z., Hamada H., Lee K., and Kitagawa T., “Reliability Evaluation of Mechanical 

Properties of AS4/PEEK Composites,” Composites, Vol. 25, No. 1, 1994, pp. 37-45. 

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27. Hsu S. Y., Vogler T. J., Kyriakides S., “Inelastic behavior of an AS4/PEEK Composite 

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Impact Models," Aeronautics and Astronautics, edited by Max Mulder, ISBN 978-953-307-

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31. Horta, L. G., Reaves, M. C., Annett, M. S., and Jackson, K. E., “Multi-Dimensional 

Correlation of Impact Dynamic Models,” Proceedings of the IMAC-XXIX Conference & 

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Impact Analysis, 2012. 

 

 
124

9.0 APPENDIX A - NON DESTRUCTIVE EVALUATION 

Non Destructive Evaluation (NDE) methods including transmission flash thermography and 

ultrasonic C-scans were used to determine the integrity of the SARAP residual hardware. This 

evaluation was considered extremely important given the fact that the SARAP hardware was 

obtained from a fuselage section that had already experienced a severe impact test. The 

Nondestructive Evaluation Services Branch of NASA Langley Research Center conducted the 

tests and their contributions are gratefully acknowledged. Only the smaller sized specimens 

were evaluated such as the laminate tension and compression coupons, the I-beams, the T-beams, 

and the cruciform sections. No NDE was performed on the larger subfloor section or the framed 

fuselage section. 

Results are divided into two categories based on evaluation method. For the first category, 

results are shown for the baseline ultrasonic NDE, which were conducted using 2.25 MHz. For 

the second, results are shown for flash thermography. 

A-1. Baseline Ultrasonic NDE 

Ultrasonic C-scan images of the laminate coupons, including back wall reflections, are shown in 

Figure A-1. These scans indicate no flaws present in the coupons. 

Figure A-1. Ultrasonic C-scan and back wall reflection of laminate coupons. 

Next, the flanges of the I-beam specimens were examined, as shown in Figure A-2. Both 

ultrasonic C-scan images and back wall reflections are shown. Once again, no indications of 

flaws are seen, except possibly in the F D area. 
125

Figure A-2. Ultrasonic C-scans of the I-beam flanges. 

The I-beam webs were also scanned and the images and back-wall reflections are shown in 

Figure A-3. Most of these images indicate that flaws are present near the ends of the specimens. 

However, the nature of the flaws is not determined. 

Figure A-3. Ultrasonic C-scans of the I-beam webs, indicating flaws near the ends of the 

specimens. 

Ultrasonic C-scans were performed on the bottom and top of cruciform 1’s flanges, as well as its 

four webs. These scans are shown in Figures A-4 through A-6, respectively. These images 

show no strong indication of any damage. 
126

Figure A-4. Ultrasonic C-scale of the bottom flanges of cruciform 1. 

Figure A-5. Ultrasonic C-scan of the top flanges of cruciform 1. 
127

Figure A-6. Ultrasonic C-scan of the webs of cruciform 1. 

Ultrasonic C-scans were performed on the bottom and top of cruciform 2’s flanges, as well as its 

four webs. These scans are shown in Figures A-7 through A-9, respectively. These images 

show no strong indication of any damage. 

Figure A-7. Ultrasonic C-scan of the bottom flanges of cruciform 2. 
128

Figure A-8. Ultrasonic C-scan of the top flanges of cruciform 2. 

Figure A-8. Ultrasonic C-scan of the webs of cruciform 2. 
129

A-2. Flash Thermography Inspection 

Single-sided flash thermography was used to conduct follow-on inspections of the cruciform 

specimens. The test set-up is depicted in Figure A-9. Thermal Wave Imaging, Inc. developed 

the system that was used for the evaluation, which incorporates a 3-5 micron infrared camera 

with a sample rate of 60 Hz. Data acquisition time is approximately 10 seconds (600-720 

frames). 

 (a) Thermal inspection system photo. (b) Close up of the thermal hood. 

Figure A-9. Photographs of the thermal inspection system. 

Thermal inspection photographs of cruciform 1 are shown in Figures A-10 through A-37. Areas 

of possible damage are highlighted in red circles. 

Figure A-10. Thermal inspection results for cruciform 1, rivet 1. 
130

Figure A-11. Thermal inspection results for cruciform 1, rivet 1 bottom. 

Figure A-12. Thermal inspection results for cruciform 1, rivet 1 top. 

Figure A-13. Thermal inspection results for cruciform 1, rivet 2. 
131

Figure A-14. Thermal inspection results for cruciform 1, rivet 2 bottom. 

Figure A-15. Thermal inspection results for cruciform 1, rivet 2 top. 

Figure A-16. Thermal inspection results for cruciform 1, rivet 3. 
132

Figure A-17. Thermal inspection results for cruciform 1, rivet 3 bottom. 

Figure A-18. Thermal inspection results for cruciform 1, rivet 3 top. 

Figure A-19. Thermal inspection results for cruciform 1, rivet 4. 
133

Figure A-20. Thermal inspection results for cruciform 1, rivet 4 bottom. 

Figure A-21. Thermal inspection results for cruciform 1, rivet 4 top. 

Figure A-22. Thermal inspection results for cruciform 1 top. 
134

Figure A-23. Thermal inspection results for cruciform 1 bottom. 

Figure A-24. Thermal inspection results for cruciform 2, rivet 1. 

Figure A-25. Thermal inspection results for cruciform 2, rivet 1 bottom. 
135

Figure A-26. Thermal inspection results for cruciform 2, rivet 1 top. 

Figure A-27. Thermal inspection results for cruciform 2, rivet 2. 

Figure A-28. Thermal inspection results for cruciform 2, rivet 2 bottom. 
136

Figure A-29. Thermal inspection results for cruciform 2, rivet 2 top. 

Figure A-30. Thermal inspection results for cruciform 2, rivet 3. 

Figure A-31. Thermal inspection results for cruciform 2, rivet 3 bottom. 
137

Figure A-32. Thermal inspection results for cruciform 2, rivet 3 top. 

Figure A-33. Thermal inspection results for cruciform 2, rivet 4. 

Figure A-34. Thermal inspection results for cruciform 2, rivet 4 bottom. 
138

Figure A-35. Thermal inspection results for cruciform 2, rivet 4 top. 

Figure A-36. Thermal inspection results for cruciform 2, top. 

Figure A-37. Thermal inspection results for cruciform 2, bottom. 
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Impact Testing and Simulation of Composite Airframe Structures

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Jackson, Karen E.; Littel, Justin D.; Horta, Lucas G.; Annett, Martin S.; 

Fasanella, Edwin L.; Seal, Michael D., II

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NASA Langley Research Center

Hampton, VA 23681-2199

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14. ABSTRACT

Dynamic tests were performed at NASA Langley Research Center on composite airframe components of increasing complexity to evaluate their energy 

absorption behavior when subjected to impact loading. A second objective was to assess the capabilities of predicting the dynamic response of composite 

airframe structures, including damage initiation and progression, using a state-of-the-art nonlinear, explicit transient dynamic finite element code, LS-DYNA. 

Specimens were extracted from a composite prototype fuselage section manufactured by Sikorsky Aircraft under the US Army's Survivable Affordable 

Repairable Airframe Program (SARAP). Laminate characterization testing was conducted and dynamic impact tests were performed on I-beam, T-section, and 

cruciform section components. Full-scale tests were conducted including a longitudinal impact test of a subfloor section and a vertical drop test of a framed 

fuselage section. Most of the test articles were manufactured of graphite tape composite with a thermoplastic resin. However, the framed fuselage section was 

constructed of a plain weave graphite fabric with a thermoset resin. Finite element models were developed and impact simulations were executed. The 

composite materials were represented using both a progressive in-plane damage model and a continuum damage mechanics model. Test-analysis comparisons 

of time history responses and structural damage development are presented for each test article. 

15. SUBJECT TERMS

Aircraft structures; Airframes; Composite structures; Finite element method; element modeling; impact tests; nonlinearity

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