Proceedings of the XIth International Congress and Exposition June 2-5, 2008 Orlando, Florida USA ©2008 Society for Experimental Mechanics Inc.
In-service TSA of composite structures using transient loading
S. Quinn*, J.M. Dulieu-Barton and R.K. Frühmann School of Engineering Sciences University of Southampton Highfield, Southampton SO17 1BJ UK *
[email protected] ABSTRACT At present Thermoelastic Stress Analysis (TSA) is a laboratory based technique as it is limited by the requirement to use a cyclic load. TSA has been proven effective in the monitoring, detection and characterisation of damage in composite components in the laboratory. The purpose of this paper is to demonstrate that it is possible to obtain quantitative data from a transient excitation provided by an impulse load. Test specimens of woven glass reinforced epoxy material are manufactured and impact damage produced in the specimens by drop tests. A Cedip Infrared Systems, solid state cooled, array detector (320 by 256 elements) was used to collect the thermal data from the specimens. A specially designed rig was used to clamp the specimens and provide a controlled single impulse. The paper contains a detailed description of the rig. Specimens were tested that contained different levels of damage and it is demonstrated that the different damage severities can be identified by the TSA during the transient loading.
Introduction Thermoelastic Stress Analysis (TSA) is a well-established stress analysis technique [1], based on the measurement of the small temperature change developed in a material as a result of elastic cyclic loading. The temperature change is directly proportional to the change in the sum of the principal surface stresses (∆(σ1 + σ2)) [2]. For an orthotropic material the small thermoelastic temperature change, ∆T, is given as [3]:
∆T = -
T (α1∆σ1 + α 2 ∆σ 2 ) ρCp
(1)
where T is the absolute temperature, ρ is the density, Cp is the specific heat at constant pressure, α is the coefficient of linear thermal expansion and the subscripts 1 and 2 denote the principal material directions of the surface lamina. Although there are further issues to resolve when using TSA to study orthotropic laminated composite materials, rather than homogeneous, isotropic materials, there are many examples of the successful application of thermoelastic techniques in the quantitative assessment of composites, e.g. [4, 5]. Much of the work is centred on damage assessment, e.g. [6, 7], where the non-contact, full-field nature of the technique is particularly attractive. Other advantages are the ability to collect high resolution data and to monitor damage progression in near real time since the advent of focal plane array detectors in the mid 1990’s. The TSA system used in the current work is the Silver 480M infra-red system, manufactured by Cedip Infrared Systems. This is a solid state cooled InSb focal plane array detector with 320 by 256 elements. As it is
radiometrically calibrated the thermoelastic output can be in temperature difference, ∆T, as well as uncalibrated thermoelastic signal, which is denoted as DL (Digital Level). The standard laboratory set-up for TSA requires a cyclic load to achieve the necessary adiabatic conditions to conduct stress/strain analysis. Thus publications in the open literature that have successfully applied TSA to problems when only a transient, or random, load is available are limited; to the authors knowledge there are only two examples [8, 9]. These show that the difficulties are not insurmountable but TSA is still regarded as a technique that requires laboratory conditions. The aim of the current work is to obtain quantitative thermoelastic stress data from woven glass reinforced epoxy composite specimens subjected to transient loading. The transient loading is provided by a controlled mechanical impulse. A rig has been designed that allows the application of a controlled impulse to a coupon of material via a swinging pendulum tup arrangement. Coupon specimens of glass reinforced woven epoxy composite material are mounted in the rig so that one end is fully clamped and the impulse applied to the other, i.e. in a cantilever arrangement. A photograph of the rig, specimen and the infra-red camera is shown in Figure 1. This experimental arrangement has been chosen because the stress distribution in a cantilever is known and therefore the accuracy of using a transient response can be assessed with a simple analysis. Specimens that contain different damage severities are used to demonstrate that the damage can be identified by the TSA during the transient loading. The key issue is adiabatic behaviour, which is governed by the test specimen material thermal conductivity and in composites this material property is favourable.
Figure 1: Photograph of transient loading rig and infra-red camera
Specimens and material calibration The material being considered in this work is glass/epoxy composite. A 380 x 500 mm E-glass/epoxy composite panel was manufactured by resin infusion, from which the specimens used in this work were cut. The epoxy is Prime 20 LV resin with a corresponding fast hardener from Gurit. The glass fibre is in the form of a plain weave fabric with a weight of 280 gm-2, which was laid up to produce a symmetric laminate of 8 plies with a stacking sequence of [0, 25, -25, 0]s. It has been shown [10, 11] that when materials are manufactured in this manner a resin rich isotropic surface layer is produced and that the thermoelastic response originates from this layer and not the orthotropic material. As all the specimens used in this work are of identical lay-up then it is possible to use a simplified form of Equation (1) to obtain the thermoelastic constant, K, for the material from a strip specimen subjected to a uniaxial tension, ∆σ, as follows: ∆T = −KT ∆σ
where K = α/ρCp.
(2)
Two 25 mm wide strip specimens were cut from the panel of material manufactured for this work, see above. One of these specimens had end tabs bonded to each end so that it could be used to experimentally derive the material’s elastic properties and for thermoelastic calibration. The exact dimensions of this specimen were 237 x 25.1 x 1.58 mm and the mild steel end tabs were 30 x 25 x 1 mm in size. The other strip specimen was used for the cantilever beam subjected to an impulse load using the swinging tup rig described in the introduction section. The Young’s modulus and Poisson’s ratio of the material were measured in a quasi-static tensile test using an Instron 5569 servo-mechanical test machine. The strains were measured using clip-gauge extensometers with gauge lengths of 50 mm and 12.5 mm for the measurement of the longitudinal and transverse strains respectively. This test gave measured values for the stiffness, E, of 15.6 GPa and Poisson’s ratio, ν, of 0.205. The thermoelastic constant, K, of the material was measured in a series of dynamic tensile tests, at a variety of mean loads, load amplitudes and loading frequencies, using an Instron 8802 servo-hydraulic test machine. A summary of the tests is given in Table 1, although it should be noted that not all mean stress and applied stress ranges were tested at all loading frequencies. TSA measurements were taken from the smooth, resin rich, surface by viewing the specimen at a slight angle (approximately 15°) to counteract reflections. There was no need to apply a paint coating to any of the specimens in this work as epoxy has a naturally high and uniform emissivity in the infra-red wavelength range. The applied tensile stress, ∆σ, in the laminate was calculated by dividing the applied dynamic load range by the cross-sectional area. The average ∆T measurement from a central area of the specimen, approximately 45 pixels by 140 pixels in size, was used in combination with the average surface temperature, T, from the same area to calculate K using Equation (2). The average of 25 measurements was -1 used to give a calibration constant K of 0.011 GPa with a standard deviation of 5.54%. This calibration constant is used to quantify the stresses in the specimens subjected to transient loads later in the paper. Table 1: Summary of test parameters for the measurement of the thermoelastic constant, K Mean stress (MPa) 25.2 25.2 37.8 50.4
32.8 x
Stress range, ∆σ (MPa) 37.8 50.4
x x
x
x x
63
75.6
10
x x
x
x
Loading frequency (Hz) 20 30 x x x x x x
40 x x
TSA utilising a transient impulse load The rig designed to apply a controlled transient load to a test coupon was based upon a modified version of an Izod impact tester. A pendulum system allows the specimen to be vertically fixed and provides low velocity impacts for an impactor that rebounds off the test piece, rather than passing through it. The major components of the rig are an impact hammer, starting and stopping mechanisms, a vice to hold the test piece and a stand. An isometric drawing of the assembled impact rig is given in Figure 2. Masses can be added at the end of the lever arm of the impact hammer to vary the impact energy and the impact tup can be changed. The starting mechanism ensures consistent impacts by fixing the start angle at a specific point. To ensure that the test piece is subjected to only a single impact a sprag clutch is used in the stopping mechanism to catch the impact hammer and lever arm at their maximum rebound angle. The vice has been designed to allow for either fixed or simple support end conditions and is adjustable to account for samples of different lengths and thicknesses. The stand of the rig was designed with stability and stiffness in mind so that the rig does not move or vibrate during an impact event on a specimen. Figure 1 shows the test set-up for the transient load TSA tests, for which a 27 mm lens and a frame rate of 383 Hz were used by the Cedip system. The second strip specimen, identical to the first that was used to measure the relevant mechanical properties and thermoelastic constant, except that no end tabs were required, was used in these tests. This specimen was clamped in the vice of the rig to create a fixed end and a cantilever beam. This allowed the stress to be calculated using the following equation for the deflection, δ, of a cantilever beam:
δ=
Px 3 3EI
(3)
where P is the applied force, x is the distance from the fixed end and I is the second moment of area of the beam. By measuring the deflection of the beam the applied force was derived and thus by application of simple bending theory the moment and therefore the bending stress in the beam at different distances from the fixed end can be calculated. This will be used to validate the TSA measurements.
= Stopping Mechanism
= Impact Hammer
= Starting Mechanism
= Vice
= Stand Figure 2: Isometric view of the impact rig
The pendulum was released from three controlled positions, 90, 100 and 110 mm from the edge of the specimen, as marked on the base of the stand in Figure 1. At each of these distances the specimen was firstly viewed from above and the impact videoed with the Cedip system to quantify the displacement, δ, using the average of 3 readings. Then the camera was positioned as indicated in Figure 1 and the thermal response of the beam to the transient load was measured, again using a video triggered by hand as the pendulum was released. The camera was positioned to view the specimen from a slight angle in order to avoid the Narcissus effect. Sand paper was
used to provide a dull and uniform background with which to block out reflections from other heat sources present in the laboratory, see Figure 1. The measurement from each release position was repeated six times. A sample of the thermal response from the specimen surface is given in Figure 3, for a spherical tup and a release position of 90 mm. Measurements were taken from seven line profiles positioned between 10 and 70 mm from the fixed end of the cantilever in 10 mm increments, as indicated in Figure 3. The plot shown in Figure 3 is a time history of the mean temperature value of line profile 1. The large temperature spike after the moment of impact is clear, followed by an oscillating temperature profile where the cantilever is vibrating freely after the impact.
Figure 3: Thermal data from the undamaged specimen showing the time history of the mean temperature along line profile 1 for impact release position of 90 mm
By considering the time taken for the thermal data to rise from a position of zero stress to the maximum stress in Figure 3, which is approximately 0.1 seconds, an equivalent loading frequency of 2.5 Hz can be defined. Videos 383 frames long, i.e. 1 second of data, were sufficiently long enough to provide enough data to analyse. To define the thermoelastic temperature change, ∆T, due to the impulse load the average temperature from 200 frames of data after the impact was taken away from the maximum temperature, i.e. the spike. The average temperature from the 200 frames after impact also represents the absolute temperature of the specimen surface, T, which in conjunction with the calibration constant for the material allows the stress to be quantified through application of Equation (2). Data of this type is given in Figure 4, for a release position of 100 mm, which shows the comparison of the experimental data to the theoretical stress distribution in the beam from consideration of the deflection of the tip of the cantilever. The results of these tests are repeatable, as shown by the six sets of data in Figure 4. However, the differences between the experimental results and the theoretical prediction are huge, the experimental results are between 60 and 75% greater than theory would suggest. This clearly warrants further investigation and will form the focus of future work.
Figure 4: Stress distribution along the beam for an impact release position of 100 mm
Introduction of damage
Impact damage was introduced into the second strip specimen using a drop weight with a spherical tup and a mass of 0.665 kg. The centre of the specimen was positioned at a 75 mm hole in a thick steel plate with a second similar thick steel plate bolted to the first plate, clamping the specimen between them. The drop weight was then released from a height of 0.3 m to introduce impact damage into the specimen. Three different levels of damage were introduced, with transient TSA tests performed between each damage level (as described in the previous section). Damage level 1 is after the first drop, damage level 2 is after the third drop. From visual inspection only minimal damage was being induced in the specimen at these two damage levels and in order to localise the damage the bottom clamping plate was replaced by one with only a 25 mm hole before conducting the fourth drop to create damage level 3. The damage from the first three drops was introduced into the specimen 40 mm from the fixed end of the cantilever beam, whilst for the fourth drop this distance was 32 mm. Figure 5 shows the comparison between the measured data for the beam in an undamaged state and after each of the 3 levels of damage have been introduced, as well as the theoretical stress distribution for an impact release position of 110 mm. The ‘Undamaged average’ line is the linear line of best fit drawn through the 6 sets of data for the beam in its undamaged condition and the data points represent a single set of readings for each of the damage levels. There is no discernable difference between the ‘Undamaged average’ line and damage levels 1 and 2, which matches the observation that only minimal damage severity was introduced in these two stages. A difference only becomes apparent for damage level 3 and the dashed line in Figure 5 is a linear line of best fit drawn through this data, which is noticeably different to the undamaged data, illustrating the potential of the current approach to detect damage. Previous research [12] has shown that this lay-up, [0, 25, -25, 0]s, is useful for encouraging delamination damage to occur at the ±25o interface. A rig has been designed that introduces fatigue damage to initiate growth of interlaminar shear damage from a central point during a simple bending test, as illustrated in Figure 6. The rig promotes interlaminar shear between the plies when a laminate panel is cyclically deflected. The deflection is applied at the free end of the panel via a clamp that allows the panel to rotate between rollers, ensuring that the loading models a point load at the end of the panel. A full description of this fatigue rig is given in Ref. [13]. A small panel of the current woven material, 295 by 100 mm, was cut from the material manufactured from this work, so that it fitted in this fatigue rig. In preliminary tests this panel was subjected to approximately 35, 000
fatigue cycles. However, the delaminations that were produced in the work of Ref. [12] were not reproduced in this woven laminate, where only fibre/matrix debonding was observed from visual inspection. This is thought to be due to the use of woven fabric, whilst that of Ref. [12] used unidirectional material. This will be investigated further in future work and it is also intended to make a rig similar to that shown in Figure 6 in which larger panels can be fatigued. After the introduction of fatigue damage these larger panels will be subjected to transient loads, perhaps in four-point bend configuration, with the impulse load introduced by an instrumented hammer, which could also be used to trigger collection of the thermal data.
Figure 5: Stress distribution along the beam for an impact release position of 110 mm
Half sine clamp
Damage
Plan view of specimen clamping arrangement
Specimen Clamp stand
Test machine bed
Figure 6: Schematic of fatigue rig
Closure
This work has produced the following conclusions: • •
It is possible to obtain the stresses from a transient load using TSA Damage changes the TSA response
In general the initial work has provided an exciting indication that it will be possible to apply TSA in the field and there are many possible future avenues of investigation.
References
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