The use of composites in primary and secondary structures of aerospace vehicles is important for increased performance with little weight penalty. Determining the response to thermal damage is necessary for a complete understanding of the total use environment of these materials. The objective of the research presented here is to provide a method of quantifying the amount of thermal damage in composite materials. Components which have non-visible damage, but have degraded performance on the order of several percent, are of interest. At this level of damage the safety margin designed into the structure may be compromised

B Tang

DC Leach

H William

IM Daniel

JJ Dim

KE Stansfield

KN Street

MD Seale

SB Dong

The use of composites in primary and secondary structures of aerospace vehicles is important for increased performance with little weight penalty. Determining the response to thermal damage is necessary for a complete understanding of the total use environment of these materials. The objective of the research presented here is to provide a method of quantifying the amount of thermal damage in composite materials. Components which have non-visible damage, but have degraded performance on the order of several percent, are of interest. At this level of damage the safety margin designed into the structure may be compromised.</p

Seale, Michael

Smith, Barry

Digital Repository @ Iowa State University (ISU)

LAMB WAVE PROPAGATION IN THERMALLY DAMAGED COMPOSITES Michael D. Seale* and Barry T. Smith** *Department of Physics **Department of Applied Science The College of William and Mary Williamsburg, VA 23187 INTRODUCTION The use of composites in primary and secondary structures of aerospace vehicles is important for increased performance with little weight penalty. Determining the response to thermal damage is necessary for a complete understanding of the total use environment of these materials. The objective of the research presented here is to provide a method of quan-tifying the amount of thermal damage in composite materials. Components which have non-visible damage, but have degraded performance on the order of several percent, are of inter-est. At this level of damage the safety margin designed into the structure may be compro-mised. Thermal damage of these structures occurs in three different scenarios. The first is damage caused during repair operations of an aircraft component. A thermal cure cycle is used to effect repairs using thermoset or thermoplastic polymers, and the heater blanket may undergo an undocumented or unintentional over temperature excursion leading to a loss of integrity. Second is damage that is incurred in a composite in a heat affected zone of an air-craft, such as near engine exhausts or leading edges. A third is a fire either on board the air-craft or in the immediate vicinity of a parked aircraft. Verification of the structural integrity of composite components is needed to insure safe operation of these aerospace vehicles. Traditional NDE techniques used on metals are not easily adaptable to composites, due to the inhomogeneous and anisotropic nature of these materials. Finding an effective means of non destructively monitoring thermal damage is extremely important to the safety and reliability of such structures. Lamb waves offer a promising method of evaluating damage in composite materials [1-2]. As a material is thermally damaged, the material properties are altered. Since the Lamb wave velocity is directly related to these parameters, an effective tool can be devel-oped to monitor damage in composites by measuring the velocity of these waves. Lamb waves can propagate several tens of centimeters for most thin planar structures and thus, provide a large gauge length measurement. If the measurement is made as a raster scan in two orthogonal directions over the area of interest, a rapid detection technique is possible. McDonnell Douglas Corporation provided several samples which simulate typical thermal damage on carbon epoxy aircraft structures. The experiment and results will be Review of Progress in Quantitative Nondestructive Evaluation, Vol. J 5 Edited by 0,0. Thompson and D.E. Chimenti, Plenum Press. New York. 1996 261 shown, along with a model calculation of the dispersion curves for this material. First a review of some related measurements will be presented. PREVIOUS REPORTS ON FIRE DAMAGE IN COMPOSITES Since the advent of polymer matrix/graphite fiber composite components in aircraft, the fire damage and subsequent property degradation has been important to designers. Sited below are several studies which have examined thermal damage in composite structures. Leach, Briggs, and Carlie [3] investigated a range of material properties for a thermo-plastic PEEK matrix with carbon fibers. The temperature range used was -500 C to 2400 C. It was found that above the glass transition temperature, strength decreased by 50% and mod-ulus fell on the order of 25%. Below the glass transition temperature the modulus only decreased slightly while the strength fell by 20%. Flame resistance of this particular mate-rial WaS noted to be better than that of aluminum. The results for heating in air of AS4/350l-6 composites was examined by Street, Rus-sell, and Bonsang [4]. In this study, the temperature was extended to 3500C and the damage assessed by measuring interlaminar fracture toughness, interlaminar shear strength, and hardness. The data suggested that for overheating at 3500C, less than 5 minutes is sufficient to seriously degrade all the parameters measured. A combined time temperature profile of 3000 C for 15 minutes led to a loss of interlaminar fracture toughness. Up to 30 minutes at temperatures of 2250 C caused insignificant damage, while a temperature of 3000 C for the same period resulted in a 50% loss of interlaminar shear strength. Barcol hardness tests of the 3000 C damage only dropped by 20% for a 30 minute exposure. The sensitivity of the Barcol hardness test as an indicator of thermal damage seems somewhat low as illustrated by this study [4]. Operating conditions can cause what are labeled as "thermal spikes", or temperature fluctuations in aircraft. Damage generated under these conditions was studied by Stansfield and Pritchard [5]. For composites that undergo these thermal fluctuations in service where moisture ingress is possible, the matrix dominated mechanical properties can be severely limited. It was concluded that matrix cracking and fiber debonding were the main damage mechanisms. SAMPLES The heat damaged samples were provided by McDonnell Douglas Aerospace. These 6 x 6 inch samples were fabricated as 16 ply, uniaxial plates of AS4/977-3. All samples had received different levels of heat damage, had been C-scanned, and had Barcol hardness readings performed. The exposure temperatures ranged from 550°F to 10000F for 3 minutes and included a sample exposed to 6000p for 10 minutes. CONTACT MEASUREMENT Lamb wave modes can be generated by placing a longitudinal wave contact type trans-ducer on a plate structure [6]. Mode conversion of a part of the wave to shear waves and subsequent mixing of SV and L waves form Lamb wave modes. In these studies Lamb waves were generated by using a 5 cycle, 500 kHz sine wave to excite the sending trans-ducer. The signal was received at various positions and digitized. The transmitting trans-ducer was a broadband, longitudinal wave transducer with a nominal center frequency of 1 MHz and a diameter of 0.5 inches. The receiving transducers were of the pinducer type, 262 ~ ~ ~ ~ Power Amplifier Function Generator Switch Out Sync. Box Out In I I I Computer Digitizer Amp Trig. Ch.l Out In I Figure 1. Experimental arrangement for velocity measurement. with active elements of 0.1 inches. A coupling gel was used between the transducers and the plate. Shown in Figure 1 is a schematic of the experimental arrangement. The time differences were measured by imposing a computational delay on the first signal to overlap the signals received at greater distances. The leading part of both waves is identified as the lowest order symmetric Lamb wave, which is not very dispersive. The trail-ing part of the waves is the dispersive antisymmetric wave. For these measurements, a least-squares fit from a plot of time and distance was performed to obtain the velocity of the prompt So mode. RESULTS Figure 2 shows the phase velocity of the So mode for a set of specimens that were exposed to the plotted temperatures for a duration of 3 minutes. One exception, plotted at 6000 F, saw a 10 minute duration. The measurements were done with the propagation direc-tion in the fiber direction. The error bars represent a 5% absolute error, mainly due to uncer-tainty in the digital channel alignment, and are quite liberal. The decrease in velocity with what is assumed to be increasing thermal damage is quite dramatic. The variation in velocity is on the order of 18% for the two extremes of temperature exposure. This variation is comparable to the modulus and velocity changes measured for fatigue damage of one million cycles reported in previous research [1]. The correlation with extended temperatures is excellent and provides a method to assess thermal damage. The Barcol hardness numbers showed little variation over this temperature range. A model calculation to determine the dispersion curves for this material system was computed using material parameters in Table 1, which includes not only the engineering constants supplied, but the corresponding elastic constants. The engineering values can be used to calculate most of the elastic constants using various techniques [7]. Since a value for G23 was not available, the value of C44 (where C44 =(C22 - C23)/2) must be estimated using common values for polymer matrix/graphite fiber composites [8]. The density of the mate-rial was taken to be 1560 kg/m3. The model calculation was the one dimensional model of Dong and Huang [9] and is shown in Figure 3. 263 9.5~~----~------~----~------~----~----~ 9 7.5 1810f 1806 t f I 1804 1809 \ I1805 1813 (10 minute duration) I 1807 I f 1802 7 ~~------~----~------~------~------~------~ 500 600 700 800 900 1000 1100 Exposure Temperature (oF) For 3 Minute Duration Figure 2. Plot of So velocity for a range of samples exposed to different temperatures for three minutes. Note: one exposure of ten minutes at 6oo°F. Table 1. Engineering constantsl and associated elastic stiffness constants. Constant Value (MSI) Constant Value (GPa) EI 19.3 Cll 135.0 E2 1.27 Cl2 3.20 GI2 0.73 Cll 9.09 V12 0.30 C23 1.59 C44 (see text) 3.75 (GPa) C55 5.03 1 Provided by McDonnell Douglas Theoretical dispersion curves were also calculated for AS4/977-3 using reduced stiff-ness constants. The effect of reducing different Cij'S by 20% on the low frequency, non-dis-persive region of the lowest order symmetric mode (0.0 :s; fd :s; 1.0 MHz·mm) was investigated. The only parameter which altered the velocity of the So mode in this region was Cll. Figure 4 shows a plot of the experimentally determined normalized velocity (velocity in the damaged sample divided by the velocity in an undamaged control sample) versus exposure temperature. Also plotted is the normalized velocity of the So mode with reduced values of Cll. Here, the normalized velocity is the velocity with Cll reduced divided by the velocity with the full values of the elastic stiffness constants. Both velocities were taken at a constant frequency-thickness product of 0.5 MHz·mm on the dispersion curve. 264 20 ~----~----~----~----~~---.~----~------~~--~ • 15 5 • • • • • , • • • • S2 • .~ • • • • • • -\ •• : S3 • • • • • • e. • • e •• . -:. .-. .-. · -. . .. • e. \.. · ... ... • •••• •• e. • • S • ••• Al I • • •••• • • •••• ........ .... -............ . .......... ...... ........... ~~I'" • • • 1 1.5 2 2.5 fd(MHz-mm) 4 Figure 3. Theoretical dispersion curve for an AS4/977-3 unidirectional composite. 1.05 r--....,..-...,.---"""T"-----,---.,----....,..-----, full parameters cll reduced by 10% j 1 c- ~ 1 ~ -t -f -------------f·95 ~ - ~ ----[ i ---------l 0.9 b: - - - - - - - r ------- -: ~ 0.85: _______________ t cll reduced by20% - cll reduced -- by30% 0·~0~0:----:6~00~--:7:-!0:-:::-0---:::-800~-~9~OO~--:1~OO~0:---~1100 Exposure Temperature (oF) For 3 Minute Duration Figure 4. Plot of the normalized So velocity of samples exposed to different temperatures. Also shown (dashed lines) is the theoretically determined normalized So velocity with reduced values of cll , as described in the text. DISCUSSION Presented above are the results of Lamb wave velocity measurements in thermally damaged composites. The results of the experimental measurement show that the velocity of 265 the lowest order symmetric Lamb mode dropped significantly for extended thermal damage. Theoretical calculations with reduced values of Cll show that a 25-30% reduction in this parameter lowered velocities of the So mode by 15%. This value corresponds to the velocity reduction of the sample which saw an exposure temperature of l000°F. Barcol hardness tests did not show a significant change for the samples studied. This result is consistent with the observations reported by Street, Russell, and Bonsang [4]. The Lamb wave velocity is a quantitative measurement and can be used to verify the integrity of a composite structure over the service life of the component. This is an impor-tant measurement for flight qualified composite hardware that may have undergone some measure of thermal damage. Since the Lamb wave velocity depends on the material proper-ties, an ideal technique can be developed to monitor composites as damage is incurred. Finally, because Lamb waves can propagate a long distance, global information about the integrity of the material can be obtained. ACKNOLEDGEMENTS The authors would like to thank Kevin Slattery of McDonnell Douglas Aerospace for providing the samples studied. This work was supported in part by NASA Grant NAG-l-1063. REFEERENCES l. M.D. Seale, B.T. Smith, W.H. Prosser, and I. E. Masters, "Lamb Wave Response of Fatigued Composite Samples," Proceedings of 1993 Review of Progress in QNDE, eds. D.O. Thompson and D.E. Chimenti, Vol. 13B, Plenum Press, New York, 1994, pp.1261-1266. 2. B. Tang and E.G. Henneke, II, "Lamb-Wave Monitoring of Axial Stiffness Reduc-tion of Laminated Composite Plates," Materials Evaluation, 47, August 1989, pp. 928-934. 3. D.C.Leach, P.I. Briggs and D.R.Carlile, "Mechanical and Fire Properties of Aro-matic Polymer Composites," Proceedings of the Third European Symposium on Spacecraft Materials in Space Environment, Noordwijk, The Netherlands, October 1-4,1985 (ESASP-232,Nov.1985),pp.153-160. 4. K.N. Street, A.I. Russell and F. Bonsang, "Thermal Damage Effects on Delamina-tion Toughness of a GraphitelEpoxy Composite," Composites Science and Technol-ogy, Vol. 32, (1988), pp. 1-14. 5. K.E. Stansfield and E. Pritchard, "Damage Generation and Healing During Compos-ite Thermal Spikes," Proceedings of the International SAMPE Technical Confer-ence, 21st, Atlantic City, NI, September 25-28, 1989, pp. 120-129. 6. I.I. Ditri, A. Pilarski, B. Pavlakovic, J.L. Rose, "Generation of Guided Waves in a Plate By Axisymmetric Normal Surface Loading," Proceedings of 1993 Review of Progress in QNDE, eds. D.O. Thompson and D.E. Chimenti, Vol. 13A, Plenum Press, New York, 1994, pp. 133-140. 7. Isaac M. Daniel and Ori Ishai, Engineering Mechanics of Composite Materials, Oxford University Press, Inc., New York, 1994. 8. William H. Prosser, "Ultrasonic Characterization of the Nonlinear Elastic Properties of Unidirectional GraphitelEpoxy Composites," NASA Contractor Report 4100, NASA Langley Research Center, October, 1987, p. 65. 9. S.B. Dong and K.H. Huang, "Edge Vibrations in Laminated Composite Plates," Journal of Applied Mechanics, Vol. 52, Iune 1985, pp. 433-438. 266 

Lamb Wave Propagation in Thermally Damaged Composites

Abstract

Similar works

Full text

Available Versions

Digital Repository @ Iowa State University (ISU)

Crossref