Full report

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NASA/TM--2000-210014 Degradation Factor Approach for Impacted Composite Structural Assessment (MSFC Center Director's Discretionary Fund Final Report, Project No. 96-17) R. Ortega, J.M. Price, and D. Fox Marshall Space Flight Center, Marshall Space Flight Center, Alabama February2000

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The NASA STI Program Office...in Profile Since its founding, NASA has been dedicated to the advancement of aeronautics and space science. The NASA Scientific and Technical Information (STI) Program Office plays a key part in helping NASA maintain this important role. The NASA STI Program Office is operated by Langley Research Center, the lead center for NASA's scientific and technical information. The NASA STI Program Office provides access to the NASA STI Database, the largest collection of aeronautical and space science STI in the world. The Program Office is also NASA's institutional mechanism for disseminating the results of its research and development activities. These results are published by NASA in the NASA STI Report Series, Which includes the following report types: TECHNICAL PUBLICATION. Reports of completed research or a major significant phase of research that present the results of NASA programs and include extensive data or theoretical analysis. Includes compilations of significant scientific and technical data and information deemed to be of continuing reference value. NASA' s counterpart of peer-reviewed formal professional papers but has less stringent limitations on manuscript length and extent of graphic presentations. TECHNICAL MEMORANDUM. Scientific and technical findings that are preliminary or of specialized interest, e.g., quick release reports, working papers, and bibliographies that contain minimal annotation. Does not contain extensive analysis. CO--ACTOR REPORT. Scientific and technical findings by NASA-sponsored contractors and grantees. CONFERENCE PUBLICATION. Collected papers from scientific and technical conferences, symposia, seminars, or other meetings sponsored or cosponsored by NASA. SPECIAL PUBLICATION. Scientific, technical, or historical information from NASA programs, projects, and mission, often concerned with subjects having substantial public interest. TECHNICAL TRANSLATION. English-language translations of foreign scientific and technical material pertinent to NASA's mission. Specialized services that complement the STI Program Office's diverse offerings include creating custom thesauri, building customized databases, organizing and publishing research results...even providing videos. For more information about the NASA STI Program Office, see the following: • Access the NASA STI Program Home Page at http://www.sti.nasa.gov • E-mail your question via the Internet to help @sti.nasa.gov • Fax your question to the NASA Access Help Desk at (301) 621-0134 Telephone the NASA Access Help Desk at (301) 62 !-0390 Write to: NASA Access Help Desk NASA Center for AeroSpace Information 7121 Standard Drive Hanover, MD 21076-1320

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NASA/TM--2000-210014 Degradation Factor Approach for Impacted Composite Structural Assessment (MSFC Center Director's Discretionary Fund Final Report, Project No. 96-17) R. Ortega, J.M. Price, and D. Fox Marshall Space Flight Center, Marshall Space Flight Center, Alabama National Aeronautics and Space Administration Marshall Space Flight Center • MSFC, Alabama,35812 February2000

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TRADEMARKS Trade names and trademarks are used in this report for identification only. This usage does not constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space Administration. Available from: NASA Center for AeroSpace Information 7121 Standard Drive Hanover, MD 21076-1320 (301) 621-0390 ii National Technical Information Service 5285 Port Royal Road Springfield, VA 22161 (703) 487-4650

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TABLE OF CONTENTS 1. INTRODUCTION ....................................................................................................................... 1 2. OBJECTIVE ................................................................................................................................ 2 3. APPROACH ................................................................................................................................ 3 4. BACKGROUND ......................................................................................................................... 4 5. SPECIMEN DEFINITION AND FABRICATION ..................................................................... 6 6. TESTING SETUP DEFINITION ................................................................................................ 7 7. MODEL DEVELOPMENT ......................................................................................................... 9 8. SAMPLE CASE MODEL RESULTS ......................................................................................... 11 9. IMPACTED COMPOSITE STRUCTURAL ASSESSMENT APPROACH ............................. 13 10. CONCLUSIONS .......................................................................................................................... 14 APPENDIX--LISTING OF IMPACT FINITE ELEMENT MODEL ............................................... 15 REFERENCES .................................................................................................................................... 20 111

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LIST OF FIGURES ° Impact test setup ....................................................................................................................... 7 2. Bend test setup .......................................................................................................................... 8 3. Impact model mesh ................................................................................................................... 9 4. Composite model impact displacement results ........................................................................ I 1 5. Maximum principle stress versus time results at the center of the panel ................................. 12 LIST OF TABLES 1. Material property data .............................................................................................................. 10 V

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TECHNICAL MEMORANDUM DEGRADATION FACTOR APPROACH FOR IMPACTED COMPOSITE STRUCTURAL ASSESSMENT (MSFC Center Director's Discretionary Fund Final Report, Project No. 96-17) 1. INTRODUCTION The use of composite laminates in aerospace structures has increased extensively over the last several years. These composites offer larger strength-to-weight ratios over metals, making them very attractive to aerospace applications. Over the years, many analysis tools have been developed to predict the behavior of the composite structures under load. However, a need remains for a tool that can predict, in a reasonable period with minimal costs, the load-carrying capacity of a composite structure that has received damage due to an impact force. Real-life examples of unintended impacts on composite structures range from a workman dropping a wrench to an in-flight impact of a bird. Currently most of the research in this area is concentrated in predicting the damage in the composite material and its subsequent damage tolerance. This approach involves several steps which include: (1) Determining the geometry of the impactor and impact force; (2) determining the damage caused by the impact including the extent of fiber breakage, delamination, and matrix cracking which involves large amounts of research in nondestructive evaluation, and impact damage progression; (3) estimating the damaged material properties to be used in the damage tolerance analysis; and establish inspection criteria. For real-life situations, and (4) conducting the damage tolerance analysis the above steps can represent long delays on a program that might not be able to afford schedule slips and cost increases.

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1. OBJECTIVE There were three main objectives of this study. First, conduct a literature search on the residual strength of impacted composites to identify promising Second, prepare a plan for conducting impact testing approaches for more expedient analysis tools. on two laminates to obtain data for the development of a strength degradation factor to be used with the undamaged elastic material properties. Third, develop a concept for assessing the structural integrity of impacted composite structures using the strength degradation factor in conjunction with available finite element analysis tools. 2

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1. APPROACH The basic approach was to conduct a literature search of testing methods for the impact of composite structures and the analysis tools used in assessing the residual strength of the structure. An impact test method was then chosen. This impact test method consists of dropping weights on composite panels. The three-point bend test was then chosen to test the damaged impacted panels to assess residual strength. Therefore, three-point bend specimens made of 00/90 ° and quasi-isotropic AS4/3501-6 composite laminates were designed and fabricated. This material was chosen because some was left over from previous programs and was available at no cost. The only costs incurred in obtaining the specimens were the material layup and the specimen machining. Once the test method was identified from the literature search and specimens were fabricated, a testing plan was developed. This plan included impact testing, three-point bend testing of impacted panels, and several tests on nonimpacted specimens to develop a baseline set of material properties for anchoring the results. The next step was to develop a process for the analytical evaluation of the residual strength of an impacted composite. This included the development of finite element models of the impact process as well as the three-point bend tests. Finally, the results of the finite element models were to be correlated and compared to the three-point bend tests of the damaged and undamaged panels.

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1. BACKGROUND Over the last 30 yr, the residual strength of impacted composites has received much research attention. Several techniques for assessing the damage tolerance of impacted composites have evolved. Husman et al. l derived a relationship for residual strength, o R, in terms of the kinetic energy imparted to the specimen, WKE: _ R = O'u Ws - WsKdWKE " , (1) where WKE _ WKE, o"u is the undamaged static strength, W s is the work per volume required to break an t undamaged specimen, t is the laminate thickness, and K a is defined as an effective damage constant. The K d factor relates the kinetic energy imparted to the specimen to the difference of the energy necessary to break undamaged and impacted specimens. In addition, K d is assumed invariant to geometry and impact energy level. However, it may depend upon boundary experimental tests are required to express the residual conditions, composite material, and layup. Two strength in terms of the imparted kinetic energy. The two tests are one tension test on an undamaged specimen and one tension test on an impacted specimen without through penetration. Husman's work showed experimental data that agree well with equation (1) for in-plane tension loads of 00/90 ° laminates. The ratio aR/a u is equal to the ratio of modulii, Ed/E, of the damaged and undamaged tension specimens for pure 0 ° laminates by following a procedure described by Mallick. 2 Furthermore, aR/a,, and Ed/E is equal to the ratio of load-carrying, cross-sectional area of the damaged and undamaged tension specimens, Ad/A. These relationships do not hold true for laminates of varying orientations. Nevertheless, more involved expressions can be obtained for multidirectional laminates relating the ultimate strength ratio to the modulii of the damaged and undamaged tension specimens. As an example, an expression relating the ultimate strengths and modulii of damaged and undamaged 00/90 `, laminates can be derived as ( KoAoEL + K9oA9oET )etu __ KoAoEL (elu - en,) O-_____R=[. KoA0 + K90A90 where E L = longitudinal (0 °) ply modulus A 0 = undamaged load-carrying, cross-sectional KoA0 + K90A90 , (2) area of 0 ° plies K 0 = ratio of the damaged to undamaged load-carrying, cross-sectional area of 0 ° plies 4

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E T = transverse (90 °) ply modulus A90 = undamaged load-carrying ,cross-sectional area of 90 ° plies K90 = ratio of the damaged-to-undamaged load carrying, cross sectional area of 90 ° plies A = total specimen load-carrying, cross-sectional area (Ao+A90) elu = longitudinal (0 °) ply failure strain etu = transverse (90 °) ply failure strain. A relationship between K d, K 0, and Kgo can be obtained by combining equations (1) and (2). Therefore, for the 00/90 ° laminate, an expression relating the kinetic energy imparted in the specimen during impact and the fraction of damage caused by the impact process can be obtained by conducting tensile tests of damaged and undamaged specimens. However, equation (2) only applies to the 00/90 ° laminate under uniaxial tensile loading. Different and more complicated expressions are required for varying layups. Also, the expression is not valid for bending or multiaxial loads under varying levels of specimen constraint. In addition, K 0 and K90 cannot be resolved without additional information from the impact event or through the use of nondestructive techniques. One can conclude that the usefulness of developing expressions such as equation (2) for individual cases is limited or nonexistent except to point out that there is a direct relationship between the kinetic energy imparted during impact, the damage caused, and the residual strength of laminated composites. The impact damage can be evaluated by the use of nondestructive techniques such as ultrasonic, acoustic emission, and, x-radiography techniques as discussed by Agarwal and Broutman. 3 However, the use of nondestructive techniques requires a great deal of effort and is not always feasible on a given piece of hardware. Another possible way of evaluating the damage caused by impact is to model the event using finite element modeling. Sun 4 used finite element models to estimate the amount of energy that causes damage in the area of impact. Hackett 5 used finite element modeling to simulate the process by which the internal damage propagates. This was accomplished by modeling the individual constituents of the composite and including the effects of matrix creeping, and the statistical nature of the fiber strength, fiber spacing, and manufacturing flaw distributions. Choi and Chang 6-8 developed finite element programs to estimate the impact damage zone including ply delamination and matrix cracking. Whereas Husman et al. 1 used an empirical relationship relating residual strength to the kinetic energy imparted to the specimen, Kutlu 9 and Shahid l° predicted failure by using damage accumulation criteria for matrix cracking, fiber-matrix shearing failure, fiber breakage for tensile loads, and buckling instability criteria for compression loads. The initial damage was modeled by representing the reduced effective stiffness as a function of matrix-crack density for matrix cracking and fiber-matrix shearing failure, and fiber failure area and fiber interaction length for fiber-breakage failure. In addition, ply transverse tensile strength and shear strength are calculated as functions of crack density in the individual plies. Damage progression is accomplished by continually recalculating the effective stiffness and effective strengths as the plies meet the required damage level for a mode of ply failure. The final failure load is predicted once the laminate can no longer sustain additional loads. 5

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1. SPECIMEN DEFINITION AS4/3501-6 composite material was selected AND FABRICATION for specimen fabrication. The material was obtained from another program and was supplied at no cost. Two specimen layups were chosen for this project. Eighty 9.75x3 in. specimens were fabricated for each layup. The layups are as follows: Panel I: 16-ply-[0, 45, 90, -45, 0, 45, 90,-45] symmetric (quasi-isotropic composite) Panel 2: 16-ply-[0, 90, 0, 90, 0, 90, 0, 90] symmetric (00/90 ° composite). Each panel, measuring approximately 48x49 in., was laid up and vacuum cured to a matted finish. Each specimen was then milled from these panels. A total of 160 specimens, 80 from each panel, are available for testing. 6

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1. TESTING SETUP DEFINITION A drop weight apparatus is envisioned for the impact tests of the specimens. This method is preferred because of its simplicity and there is no kinetic energy loss on the impactor due to guide rail friction. However, the rebound velocity of the impactor and the maximum rebound height are difficult to determine. This makes it more difficult to establish the transfer of energy onto the specimen. The basic test setup is shown in figure 1. "__ .._ Composite Specimen \ I 7.5 in. Solenoid j Impactor Weight= 0.375-2Ib 18 in.<y < 72 in, _ ClampingRollers Figure 1. Impact test setup. Four stainless steel 316 ball sizes (1.375, 1.75, 2, and 2.5 in.) were chosen for the impact tests. These sizes were chosen based on the kinetic energy available at impact as compared to the kinetic energy at impact of the tests conducted by Husman et al. 1 Given the differences of test setup, the proposed impactor sizes for this project are much larger than those used by Husman et al. I However, the chosen sizes seem reasonable for real-life examples of the handling of composite structures. The height variation from 18-72 in. allows variation of kinetic energy for each given ball size, thus, allowing some assessment of impactor-size variation. It is proposed that four heights be chosen and tested with four specimens for each ball size. This gives 64 impact specimens for each composite layup. The 16 remaining samples will be used for the nonimpacted bend tests and serve as spares. 7

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Impactedandnonimpactedspecimensareto betestedin a three-pointbendapparatus.The 3-in.-widespecimenswith anunsupportedspanof 7.5 in. will be testedon anMTS® SystemsCorporation tensiletestingmachinewith a 5,000-1bloadcapacity.The tensionside(bottom)of the specimenwill be testedwith a straingaugewhenappropriate.Strokedisplacement,load,andstrainwill be recorded. The testsetupis shownin figure 2. 7.5 in. MaximumLoad 5,000-1b A Figure 2. Bend test setup. 8

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1. MODEL DEVELOPMENT A parametric defined finite element model of a steel ball impacting the 00/90 ° composite was developed. The three-dimensional finite element model is composed of the finite element program ANSYS ® 5.3,11 layered structural shell elements modeling the AS4/3501-6 composite plate, threedimensional 10-node tetrahedral structural solids for the steel ball impactor, and three-dimensional point-to-surface contact elements. The input parametric variables include the steel ball diameter, ply thickness, support span width for the plate, width of the plate, and drop height. The composite plate has pinned boundary conditions at the span supports. The contact elements attach to the steel ball and the composite plate to keep track between the two parts until contact is established. When contact is achieved, the load is transferred from one part to the other. ANSYS uses a normal contact stiffness value used to determine contact forces. The value of normal value needs to be as high as possible to ensure contact contact stiffness is a user-assigned value. The without the parts going through one another. However, arriving at this value is an iterative, time-consuming process. A sample case for the model is shown in figure 3. In this sample case, the steel ball diameter is 2.25 in. and the drop height is 72 in. The input deck for this model is enclosed in the appendix. The quasi-isotropic composite case can be obtained by modifying the real card section of the input deck. I I L 41 i |!IliL - "J "G -'- i Symmetry Figure 3. Impact I , ..... I ANSYS 5.3 SEP 22 1999 10:19:43 I-'-_-- -- ELEMENTS TYPE NUM YV =1 DIST =4.125 XF =3.75 YF =0.75 ZF =1.175 A-ZS =-0.854E-06 Z-BUFFER model mesh. 9

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The materialpropertiesusedin this modelareshownin table 1.The compositepropertieswere obtainedfrom Tsail2 and the DOD_ASA Advanced Composites Design Guide. 13 It was assumed that Ez=Ey, nuxz=nuxy=nuyz, and Gxy=Gxz. Additional research or material characterization is needed to determine the adequacy of the properties listed. Table 1. Material property data. Property Modulus, Ex (psi) Modulus, Ey (psi) Modulus, Ez(psi) Poisson's Ratio, nuxy Poisson's Ratio, nuxz Poisson's Ratio, nuyz Shear Modulus, Gxy (psi) Shear Modulus, Gxz(psi) Shear Modulus, Gyz(psi) Density (lb sec2/in.) AS4/3501-6 StainlessSteel 2.07E+07 3.00E+07 2.00E+06 3.00E+07 2.00E+06 3.00E+07 0.3 (prxy) 0.33 0.3 (prxz) 0.33 0.3 (pryz) 0.33 1.03E+06 1.03E+06 7.69E+05 1.47E-04 7.61E-04 A composite failure criterion needs to be incorporated in the model to determine the failed layers and area of the composite. ANSYS has three built-in failure criteria that can be used. These are the maximum strain failure criterion, the maximum stress failure criterion, and the Tsai-Wu failure criterion. A fourth criterion based on the maximum shear stress criterion can also be included as a user-defined input. This fourth criterion is highly recommended in a paper by Hart-Smith. 14 Finally, the bend test model will be obtained by taking the composite panel section of the impact model mesh and applying a load to the nodes at midspan. Both, the impacted and nonimpacted specimens are to be modeled. The area of damage will be modeled by modifying the properties of those elements to eliminate the composite layers that have been damaged. Failure in the model will be an iterative process of increasing the load until a limit maximum load is reached. 10

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1. SAMPLE CASE MODEL RESULTS The displacement results of the sample case are shown in figure 4. ANSYS 5.3 SEP 29 1999 13:35:33 DISPLACEMENT STEP =1 SUB =1 TIME =0.424E-05 RSYS =0 DMX =0.001 SEPC =90.623 (a) Prior to impact ANSYS 5.3 I SEP29 1999 J 13:39:04 DISPLACEMENT STEP =4 SUB =1 TIME =0.441E-03 Z RSYS =0 I DMX =0.103999 SEPC =58.012 (b) 6 p.secafter impact ANSYS 5.3 SEP 29 1999 12:22:47 DISPLACEMENT STEP =5 Z SUB =39 TIME =0.001726 RSYS =0 DMX =0.389645 SEPC =54.889 (c) 1.28 msec after impact Figure 4. Composite model impact displacement results. 11

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The stressresultshaveshownconsiderableoscillationover theperiodobservedthusfar.Additional effort is neededin understandingthereasonfor theoscillation andin makingappropriatechanges asnecessary.Someof this additionaleffort shouldincludevariationsin meshsizeat the immediate contactareaandincreasingthe observationperiod.The time-dependentprinciplestressesareshownin figure5. 1,250 1,125 1,000 875 e_ ,_ 750 625 - .E.x 500 - 375 250 125 0 ' I ANSYS 5.3 SEP30 1999 15:05:28 POST26 ZV =I DIST =0.75 *YF =0.75 *YF --0.5 *ZF =0.5 Z-BUFFER I ! 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Time(msec) Figure 5. Maximum principle stress versus time results at the center of the panel. 12

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1. IMPACTED COMPOSITE STRUCTURAL ASSESSMENT APPROACH Although the work has not been completed, the envisioned process of assessing the structural integrity of impacted composite structures is as follows: ° For the given layup, build a finite element model of the composite structure. Run a load case of a structural test (such as a proof test) that has been carded out on the structure where data are available. This step is necessary to anchor the model's elastic properties, geometry, and boundary conditions. 2. Determine the impact object geometry, mass, speed, acceleration, and location of impact. 3, Model the impact process to determine the damage zone. Due to computational limitations, a submodel of the structure might be needed for the given structure. Determine the degradation of the individual plies in the damage zone. Anchor to any available impact data. 4. Run the model under operating loads with degraded properties for the plies and location that sustained damage. 5. Determine whether failure occurs. In this study, the composite plate represents the composite structure. The bend test of the undamaged panel is to be modeled to check properties and boundary conditions. The impact test and subsequent three-point bend tests will generate the data to verify the procedure of first determining the damage zone, and then predicting failure under load. 13

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1. CONCLUSIONS A finite element approach for assessing the integrity of impacted composite structures using a material degradation factor was documented. The work effort to successfully complete this program was initially estimated to be 33 man-months. Of the requested 33 man-months, only 4-5 man-months were actually spent working on this project over a span of 4 yr. During this time, a literature search was conducted, composite specimens were designed and procured, a finite element model of the impact process was built, and the required reports for the program were submitted as necessary. However, much work is needed to complete the goals of the program. The requested amount of work could not be accomplished because of responsibilities to other programs with higher priority. Given the current workload, it is impractical to continue this program at this time. Perhaps, the intent of this effort can be renewed in future Center Director's Discretionary Fund proposals. 14

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APPENDIX--LISTING OF IMPACT FINITE ELEMENT MODEL The following are the line commands for the finite element model in ANSYS, release 5.3: /filename,composite +i /prep7 spdiam=2.25 sprad=spdiam/2 plythk=.00625 span=7.5 width=3 vel=30 et, l,99 keyopt, 1,2,0 keyopt, 1,3,0 keyopt, 1,4,0 keyopt, 1,5,1 keyopt, 1,6,1 keyopt, 1,8,1 r,l,16,0 rmore,,,,,,, rmore, 1,0,plythk, 1,90,plythk rmore, 1,0,plythk, 1,90,plythk rmore, 1,0,plythk, 1,90,plythk rmore, 1,0,plythk, 1,90,plythk rmore, 1,0,plythk, 1,90,plythk rmore, 1,0,plythk, 1,90,plythk rmore, 1,90,plythk, 1,0,plythk rmore, 1,90,plythk, 1,0,plythk rmore, 1,90,plythk, 1,0,plythk rmore, 1,90,plythk, 1,0,plythk rmore, 1,90,plythk, 1,0,plythk rmore, 1,90,plythk, 1,0,plythk mp,ex, 1,20.7e6 mp,ey, 1,2e6 mp,ez, 1,2e6 mp,prxy, 1,.3 mp,prxz, l,.3 mp,pryz, 1,.3 mp,gxy, 1,1.03e6 15

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mp,gxz, 1,1.03e6 mp,gyz, 1,.769e6 rap,dens, 1,1.474e-4 et,2,92 mp,ex,2,3Oe6 mp,nuxy,2,.33 mp,dens,2,7.61 e-4 et,3,49 keyopt,3,7, i r,3,60e3 .... 1 et,4,49 keyopt,4,7,1 r,4,60e3 .... 1 real, 1 type, 1 mat, 1 rectng,O,span,O,width/2 esize,sprad/8 amesh, 1 nsel,s,loc,x,O d,all,ux,O .... uy, uz nsel,a,loc,x,span d,all,uz,O lsel,s,loc,y,O dl,all, 1,symm local, 12,0,span/2,0,sprad+. 1,,,90 wpcsys, 1,12 i sphere,sprad,,O, 180 J wpcsys, i ,0 j J csys,O eshape, 1,2 esize,sprad/4 type,2 mat,2 real,2 vmesh, 1 type,3 mat,3 real,3 local, 13,2,span/2,0,. 1+sprad nsel,s,loc,x,sprad- 1e-4,sprad+ I e-4 csys,O nsel,r, loc,z, .099,. 1O01 ÷sprad cm,sphereo,node 16

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nsel,s,loc,x,span/2-sprad,span/2+sprad nsel,r, loc,y,-.0001,1.1 *sprad nsel,r, loc,z,-.05,.05 cm,shello,node cmsel,s,sphereo *get,numb,node,0,count *do,in, 1,numb, 1 *get,nv, node,0,num,min *get,xv, node,nv, loc,x *get,yv, node,nv, loc,y nsel,u,node,,nv cm,dummy, node nsel,s,node,,nv cm,contact,node cmsel,s,shello nt=node(xv, yv,0) :_get,xxv, node,nt,loc,x *get,yyv, node,nt,loc,y nsel,r, loc,x,xxv nsel,r, loc,y, yyv nsel,r, loc,z,0 esln esel,u,type,,3 nsle cm,target,node cmsel,a,contact gcgen,contact,target I-" cmsel,s,dummy enddo C**End loop 1 ,- type,4 mat,4 real,4 cmsel,s,shello,node *get,numb,node,0,count *do,in, 1,numb, 1 _ *get,nv, node,0,num,min *get,xv, node,nv, loc,x *get,yv, node,nv, loc,y nsel,u,node,,nv cm,dummy, node nsel,s,node,,nv cm,contact,node 17

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cmsel,s,sphereo xf=xv-span/2 if,xfxf+yvyv, gt,spradsprad,then cmsel,s,dummy cycle endif fr=. 1+sprad-sqrt(spradsprad-yvyv-xfxf) nt=node(xv,yv,fr) nsel,s,node,,nt esln eseI,u,type,,3 nsle csys,13 nsel,r,loc,x,sprad-I e-4,sprad+Ie-4 csys,0 cm,target,node J cmsel,a,contact j4 gcgen,contact,target cmsei,s,dummy enddo m C**End Loop 2 allsel nsel,s,loc,z'05,20 nsel,r,loc,y,0 d,all,uy,0....rotx,rotz allsel sbctran r,3,50e6....1 r,4,50e6....1 save ! fini /solu antype,transient ti=O outress,all,all autots,on timint,off nsel,s,loc,z,.05,10sprad d,all,uz,-.O01 nsel,r, loc,x,.4987span,.50133span d,all,ux,O ti=.001/velo time,ti 18

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nail nsubs,l,l,l,on acel,,,O solve yes timint,on nsel,s,loc,z,.05,lOsprad ddel,all,uz nail eall td=(sqrt(velovelo+2az,l)-velo)/az ti=ti+.OOOOl*.99td time,ti acel,,,az nsubst,l,l, l,on \ k. solve yes ti=ti+.99td time,ti nsubst,lO,lO,l,on solve yes ti=ti+l.6td time,ti nsubst,60,60,40,on /nerr,40,600000 solve yes /solu antype,transient,rest ti=ti+ 1.6td z time,ti - nsubst,60,60,40,on /nerr,40,600000 solve yes fini 19

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REFERENCES o Husman, G.E.; Whitney, J.M.; and Halpin, Laminated Composites Subjected to Impact J.C.: "Residual Strength Characterization of Loading," Foreign Object Impact Damage to Composites, ASTM STP 568, American Society of Testing and Materials, Philadelphia, PA, pp. 92-113, 1975. , Mallick, EK.: Fiber-Reinforced Composites: Materials, Manufacturing, and Design, 2nd ed., Marcel Dekker, Inc., New York, NY, pp 201-359, 1993. ° Agarwal, B.D.; and Broutman, L.J.: Analysis and Perfo_vnance of Fiber Composites, 2nd ed., John Wiley & Sons, New York, NY, 387-393, 1990. o r i I Sun, C.T.: "An Analytical Method for Evaluation of impact Damage Energy of Laminated 7 Composites," Composite Materials: Testing and Design (Fourth Conference), ASTM STP 617, r American Society for Testing and Materials, , Hackett, R.M.; and Slattery, K.T.: "Modeling Philadelphia, PA, pp. 427-440, 1977. Stiffness Degradation in Filamentary Composite Materials," Journal of Materials in Civil Engineering, Vol. 4, No. 2, pp. 196-211, May 1992. , Choi, H.Y.; Downs, R.J.; and Chang, E K.: "A New Approach Toward Understanding Damage Mechanisms and Mechanics of Laminated Composites Due to Low-Velocity Impact: Part I- Experiments," JounTal of Composite Materials, Vol. 25, pp. 992-1011, August 1991. ° Choi, H.Y.; Wu, H.Y.T.; and Chang, EK.: "A New Approach Toward Understanding Damage Mechanisms and Mechanics of Laminated Composites Due to Low-Velocity Impact: Part II- Analysis," Journal of Composite Materials, o Vol. 25, pp. 1012-1038, August 199 i. Choi, H.Y.; Wang, H.S.; and Chang, F.K.: "Effect of Laminate Configuration and Impactor's Mass on the Initial Impact Damage of Composite Plates Due to Line-Loading Impact," Journal of Composite Materials, Vol. 26, No. 6, pp. 804-827, 1992. . Kutlu, Z.; and Chang, EK.: "Modeling Compression Failure of Laminated Composites Containing Multiple Through-the-Width Delaminations," Journal of Composite Materials, Vol. 26, No. 3, pp. 350-387, 1992. 20

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1. Shahid, I.; Sun, H.-T.; and Chang, F.K.: "Accumulated Damage and Fracture of Composite Laminates Under In-Plane Loads," Use of Plastics and Plastic Composites, Proceedings of the 1993 ASME Winter Annual Meeting, New Orleans, LA, ASME Publication, MD-Vol. 46, pp. 469-479, 1993. 11. "ANSYS Release 5.3 Online Documentation," ANSYS Inc., Ist ed., June 1996. 12. Tsai, S.W." "Theory of Composites Design," Think Composites, Dayton, OH, C 1-C4, 1992. 13. DOD/NASA Advanced Composites Design Guide, I st ed., Structures/Dynamics Division, Flight Dynamics Laboratory, Air Force Wright Aeronautical Laboratories, Wright-Patterson Air Force Base, OH, Prepared Under Contract F33615-78-C-3203 by Rockwell International Corporation, section 3.16.A, pp. 1-8, July 1983. -.. 14. Hart-Smith, L.J.: "How to Calculate In-Plane Strengths of Fiber-Polymer Composite Laminates," SAMPE Journal, Vol. 28, No. 6, pp. 25-35, November/December 1992. 21

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Form Approved REPORT DOCUMENTATION PAGE OMBNo.0704-0188 Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for revlew_g instructions, searching existing data sources, gartering and maintaining the data needed, and complating and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of i_'lformation, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operafian and Repods, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503 1. AGENCY USE ONLY (Leave Blank) 2. REPORT DATE 3. REPORTTYPE AND DATES COVERED February 2000 Technical Memorandum 4. TITLE AND SUBTITLE Degradation Factor Approach for Impacted Composite Structural Assessment 5. FUNDING NUMBERS (MSFC Center Director's Discretionary Fund Final Report, Project No. 96-17) 6. AUTHORS R. Ortega, J.M. Price, and D. Fox ADDRESS(ES) 8. PERFORMING ORGANIZATION 7.PERFORMINGORGANIZATIONNAMES(S)AND George C. Marshall Space Flight Center Marshall Space Flight Center, AL 35812 REPORT NUMBER M-967 ADDRESS(ES) 10. SPONSORING/MONITORING 9. SPONSORING/MONITORINGAGENCYNAME(S)AND National Aeronautics and Space Administration Washington, DC 20546-0001 11.SUPPLEMENTARYNOTES AGENCY REPORT NUMBER NASA/TM--2000-210014 Prepared for Structures, Mechanics, and Thermal Department, Engineering Directorate 12a, DISTRIBUTION/AVAILABILITY STATEMENT Unclassified-Unlimited Subject Category 39 Nonstandard Distribution 13. ABSTRACT (Maximum 200 words) 12b, DISTRIBUTION CODE iThis technical memorandum documents the results of the research to develop a concept for assessing the structural integrity of impacted composite structures using the strength degradation factor in conjunction with available finite element tools. For this purpose, a literature search was conducted, a plan for conducting impact testing on two laminates was developed, and a finite element model of the impact process was created. Specimens for the impact testing were fabricated to support the impact testing plan. 14.SUBJECTTERMS 15. NUMBER OF PAGES 28 composites, impact testing, finite elements, strength degradation 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION OF REPORT OF THIS PAGE Unclassified Unclassified NSN7540-01-280-5500 16. PRICE CODE A03 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF ABSTRACT Unclassified Unlimited Standard Form 298 (Rev. 2-89) p rescdbed by ANSI Sld. 239-18 298-102