·Table of Contents
·Aeronautics and Aerospace
Non destructive testing in space environmentD Simonet,G Ithurralde,JP Choffy, AEROSPATIALE MATRA CCR,
JP Bonnafé AEROSPATIALE MATRA LANCEUR
|Fig 1: Drawing of the Columbus Orbital Facility|
|Fig 2: Top view of a crater after an impact|
3.1.Probe design and working
Figure 3 shows the four superposed elements of the probe: permanent magnet (samarium cobalt disc sustaining high temperature), conductive mask, transmitting spiral coil, receiving spiral coil. Inducing a circular alternative eddy current density in a vertical magnetic field produces radial oscillating Lorentz forces, in a skin depth inferior to 0.1 mm (0.004 in.). These forces act like an ultrasound source inside the material itself. Receiving follows the opposite mechanism.
|Fig 3: Cross view of the probe showing design and principle of shear wave generation||Fig 4: Set-up for thickness measurement under simulated space conditions|
3.2. Implementation and set-up
The installation was composed of:
The EMAT driver generates high voltage square bipolar tone bursts. No cooling was needed.
To maximize signal-to-noise ratio, received signals were electronically filtered, rectified and smoothed. Furthermore, the software averaged 16 acquisitions.
The calibration block is made of the same aluminum alloy as the COF external structure, that is, alloy 2219 T6. Several steps were machined with thickness ranging between 1 mm (0.04 in.) and 5 mm (0.2 in.).
Tests were performed at nine temperatures between -150° C (-238° F) and +170° C (+338° F). Thickness was deduced from the known 20° C (68° F) velocity and the measured time of flight difference between two successive back wall echoes, as shown in figure 5.
Lift-off influence was also studied at 20° C (68° F), by inserting thickness gauges between the receiving coil and the calibration block.
The ultrasound wavelength was too large for the step of 1 mm (0.04 in.) to be evaluated. Larger thickness was measured with satisfactory accuracy (lower than six percent) and signal-to-noise ratio, even at +170° C (+338° F) when electrical conductivity, thus eddy current density, decrease. Relative errors are given in Figure 6. Their calculation took into account theoretical temperature expansion. Thickness is over evaluated at high temperature and is under evaluated at low temperature. These differences may be explained by changes in shear waves velocity due to temperature variations.
|Fig 5: Time of flight measurement on a rectified A-SCAN at 20°C||Fig 6: Relative differences between measured and expansion corrected values between -150 and +170°C|
Lift-off has a dramatic influence on signal amplitude. With a gap between the receiving coil and the calibration block greater than 1 mm (0.04 in.), echoes are not strong enough to exceed noise. This will cause problems for thickness measurement at the center of deep concave craters unless the probe diameter is minimized. A solution may consist in replacing winded coils by flexible circuit coils (obtained by photolithography or similar process) for the active surface to fit with the crater surface.
Temperature variations are not a problem if EMATs are used for wall thickness measurement purpose on the COF, in space environment. Results demonstrate that a probe generating 0° shear wave provides values with relative errors lower than ± 5% between -150 (-238° F) and +170° C (+338° F), for a thickness 1-5 mm (0.04- 0.2 in.).
Development should be continued to optimize the probe, particularly regarding the lift-off effect, and to adapt the whole system to space standards and procedures.
EMATs might also be used for crack detection and sizing, for example with probes generating surface or guided waves. But other NDT tools such as eddy currents could prove to be more efficient.
4.1.Influence of the temperature on 2219 electrical conductivity and penetration depth.
The electrical conductivity variation of 2219 according to the temperature is given by the following formula:
|sT = Conductivity at Temperature in MS/m|
s réf = reference Conductivity at reference30
temperature in MS/m
q réf = reference temperature
q = control Temperature
a = 0,0025 (for 2219)
Fig 7: Example for 2219 with 19.7 MS/m for 20°C for reference value
|Fig 8:standard penetration for 60kHz according temperature||Fig 9: impedance Variation | Z | according the temperature|
4.2.Influence of the temperature on the probe characteristic
Different types of eddy current probes were designed for this study in order to compensate the temperature effect on the impedance. All elementary coil used during this study were tested in temperature ranging from -150 ° C to + 170° C in order to evaluate the variation of the impedance and their durability. The following chart is an example of impedance variation measured on coil according the temperature.(figure 9) The other question was to find some materials capable to resist at low and high temperature and to use it to build some probes. The ferrite, the cooper wire, the silver soldering demonstrated a good ability to work at theses type of temperatures. We had used some Teflon to build the body of the probe. Finally a good surprise was that common and chip material could be used to make our probe. A nominal probe was designed (figure 10) in order to be tested on artificial defect with the influence of temperature.
Fig 10: Probe designed for temperature application
In order to validate the possibility to perform some mapping in a space temperature condition, we have designed a reference block with different type of artificial surface defect corresponding to the minimum defect requirement. (Figure 11). We have also used the same temperature regulated vessel used for the EMAT Study (Figure 4) and Eddy current mapping system commonly used. We have performed some mapping from -150 ° C to 170 ° C to evaluate the influence of the temperature on the mapping result (see figures 12 and 13).
All the artificial defects were detected but the set up must be adjusted for each mapping.
|Fig 11: Reference plate||Fig 12: +170°C Map Figure 13 :-150°C Map|
To conclude on eddy current tools, we can say this method has a great potential to be used in a space environment without large development. This method can work between -150 °C to +170 °C in respect of detection limit criteria. The prototype probe specially designed for this application use common and chip materials and the pollution due to this method are very limited according the requirements. The next step is to design a remote controlled eddy current system an to demonstrate its ability to work on a damaged area on shell Columbus space structure.
5.1. Visual tools device description
To dimension the damage due to an impact in a fully accessible area, the device used is based on a video camera but this camera needs to be maintained in a stabilised position and localisation in order to give quantitative information. So the solution is to place the camera on a support, which allow movements around the vertical position to perform measurements. In that case, this support is placed on the surface to examine and a small laser is used to realise the space localisation by telemetry. Small electric engines provide all the degrees of freedom necessary for the measurements. The support is equipped with removable handles and is linked to the command and control device by a connection cable providing power in one way and electrical information in the other way, or can be self-powered and exchange data through a radio link. The astronaut (inside or outside of the module) places this device then the procedure is automatically performed. The following figure shows the device principle with main equipment.
Fig 14: prototype of computer-assisted visual inspection
A working mock-up has then been made in order to validate every technology that we expect to use. The following drawings gives the main dimensions of the visual inspection device, which is about 300mm x 300mm (the size of a typical pocket on cylindrical panels).
5.2. Visual inspection tools capability
The concept of visual inspection tool is quite simple, using mainly mechanical devices available separately and placed together on the same support. Taking into considerations the experimental database concerning tools to be used in space environment, it is clear it will be easy to find solutions to make this tool compatible with vacuum and high and low temperatures. So, this point has not been developed here.
The main work that has been done was to verify that the camera was able to supply a picture of the damaged zone (coming from the camera on support) that would be able to be analysed to give information about damage dimensions. Once the mechanical movements of the camera have been developed and made available, the camera has been linked to a computer through a video acquisition extension card working in real time, and the recorded pictures have been analysed through a picture processing software. A laser emitter (1 mW for a wavelength equal to 633 nm) has been added to perform a simple localisation record. The adjustment tests have been performed on an impacted plate from EMI high velocity impact (shown at the beginning of the document). The results of theses test have been demonstrated the ability of computer-assisted visual inspection to
5.3. Conclusion for visual inspection Tools
To conclude on visual inspection tool, we can say that this method can be applied for defect characterisation because of the possibility to make the diagnostic in an automatic way. The positioning of the control device (camera on support) is still made by the astronaut but when it is done, demonstration has been made that we could obtain the same level of automation than for others NDI tools studied before. In that case, all surface dimensions can be obtained with a good precision, sufficient enough to complete the defect database description.
All these tools now need complementary studies in order to be specifically adapted to space environment and contingencies but we know now that they are applicable for such a use.
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