·Home ·Table of Contents ·Aeronautics and Aerospace

Non destructive testing in space environment D Simonet,G Ithurralde,JP Choffy, AEROSPATIALE MATRA CCR, JP Bonnafé AEROSPATIALE MATRA LANCEUR Contact

1. Introduction

Since the launching of Sputnik 1 on October 4, 1957, more than 11000 satellites were placed in orbit. Each one as of these launchings was accompanied by a certain number of debris (bolts, caps...). The total mass of these debris is estimated at 3.106 kg with an average density of 2,8 g/cm3 and an average speed of 11 km/s. The collision risks with spacecraft such as future orbiting station ISS must be taken into account, these debris could cause damage by impact on the shell structure of the module. To ensure a lifetime of about thirty years «the C.O.F.»( Columbus Orbital Facility) of the ISS, it would be necessary to be able to locate, quantify and repair in orbit the areas damaged by impact. Within the framework of the Columbus project, AEROSPATIALE-MATRA had made an inventory of the techniques of non destructive testing (NDT) and their potential applications in a space environment. Among the conclusions of this study, the possibility had been planned to use three techniques, the computer-assisted visual inspection, Ultrasounds EMAT and Eddy currents. It then remained to demonstrate the capacity of these techniques to answer the detection criteria of potential defects. Each one of these techniques was studied for a particular application. Eddy currents, for the thickness measurement of thin walls and the detection of defects like craters or cracks. Ultrasounds EMAT for the thickness measurement of thicker walls and the computer-assisted visual inspection for the dimensional measurements of shell deformations.

Fig 1: Drawing of the Columbus Orbital Facility

2. Requirements

Fig 2: Top view of a crater after an impact

Impacts between debris and space structure such as Columbus were simulated during a test campaign performed at EMI in Freiburg, Germany. The objective of these tests was to define the MDPS [Meteorite Debris Protection System] configuration and to test it under high velocity impacts. The samples impacted were some plates in aluminum alloys (300 mm x 300 mm x 5 mm in 2219 T87), one of the result of this test is shown on the figure 2
The plate is stretched out of shape under the impact and the thickness is locallyreduced. Some holes, cracks, craters are present on affected impact area. The first objective on NDI tools is to give the main dimensions of the plastic deformation (diameter, depth, ...) and remaining thickness in the affected area.The second objective is to detect and to evaluate the surface defect like cracks,holes and craters. All these elements are necessary to evaluate the damage and the method to repair the structure.
The space environment can have many influences on NDI tools. The temperature between -150° C to +170° c is certainly the main problem for the systems, no couplant can be used, some physical properties like conductivity are modified, the geometry of probe, encoding system and other mechanical parts are deformed.
The other main environment requirements are:
• Resistance to cosmic radiation with high energy, able to damage equipment in case of long exposure
• Resistance to electromagnetic emissions
• Usable in vacuum (between 10^-10 and 10^-11 Atm )
• Low gravity
NDI use in space needs special specification (ESA, NASA) such as
• No use of aerosol, liquid, ionizing radiation.
• Limited electromagnetic contamination
• Long service stability and easy maintenance
• Limited complexity to use
• High level of automation
Some special technical specification have been added
• The power consumption is limited to 20 W
• The weight of system must not exceed 20 Kg
• The system must be remote-controlled from the COF or from the ground in order to limit the space manipulation by the astronaut.

3. EMAT (ElectroMagnetic, Acoustic Transducer)

Conventional UT with piezoelectric transducers is not applicable for thickness measurement in space environment because the vacuum and wide variations in temperature do not permit the use of any liquid couplant. For this reason electromagnetic acoustic transducers (EMATs) are a good alternative. Our objective was to prove their efficiency under space conditions.
A probe was specially designed and manufactured to generate and receive 0° shear waves vibrating between 3 MHz and 4 MHz. It was tested in a homemade temperature regulated vessel, between -150° C (-238° F) and +170° C (+338° F).
The results were encouraging since thickness was measured down to 1 mm (0.04 in.) with satisfactory accuracy on a calibration block with steps made of aluminum alloy (2219 T6). On the one hand, the influence of temperature was not as important as expected. On the other hand signal amplitudes were very sensitive to the lift-off. Miniaturized flexible circuits instead of wound coils should solve the lift-off problem.
Space standards and other requirements such as maximum power consumption, procedure, etc. have to be considered and integrated into the design of the probe and the driver

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:

• Homemade temperature regulated vessel, permitting temperature variation between -150° C (-238° F) and +170° C (+338° F).
• R/D Tech Tomoscan EMAT driver,
• Personal computer supporting R/D Tech Tomoview software,
• Impedance analyzer to follow changes in the coils electrical characteristics according to the temperature.

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.

3.3. Results
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.

3.4. Conclusions
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. Eddy Current

Eddy current is a method widely used for different kind of applications on aluminum alloy such cracks and other surface defect detection, conductivity measurement. The main problem for a space application is the temperature ranging from -150 °C to 170° C. It is important to evaluate the influence of the temperature on the parameter method such the variation of probe impedance combined with the conductivity variation of 2219 aluminum alloy.

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

One of the consequences of this variation is the modification of the penetration depth of eddy current waves according to the temperature. If we apply this conductivity variation to formula of standard penetration , we obtain for the 2219 for 60 kHz the following chart

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

4.3. Results
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

4.4. Conclusion
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. Visual inspection tools

Visual inspection is the first natural way to have a realistic view of the damage due to an impact. But due to the environmental limitations, and due to characterisation requirements, this way is not sufficient and has to be improved.
From internal or external side, in current areas, there is no problem to visualise the impact with naked eyes or with a video camera. The only requirement is to have a light source device because of the darkness that can be encountered inside as well as outside of the module. For the same areas, dimensioning of the impact damage will not be a major problem because there are plenty of places to position a measurement device in order to characterise damage dimensions in diameter, plastic deformation depth, holes, cracks.. Of course, even if visual inspection can detect cracks, it will be necessary to complete the information with other NDI tools. For other areas, the problem is not the same. Geometrical contingencies are going to limit visibility and accessibility to the damaged area. On the external side, the visibility of the impact area on the cover itself will be easy but the visibility of the damage under cover will be limited. For the same reason, the evaluation of the damage could be very restrictive.

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.

6. Conclusion

The goal of this study was to present an evaluation of the capabilities of NDI tools selected in the previous phase to characterise the defects due to an impact: Eddy currents, EMATs and visual inspection. Each of these tools has been studied for a specific use, Eddy currents for the measurement of depth of the defects, EMATs for thickness measurements and visual inspection for surface dimensions measurement. Except for visual inspection, all these tests have been done in representative space environment according to temperature requirements. They have shown that all these three tools were complementary to obtain a complete characterisation of the defects.
Concerning Eddy current tools, this method presents a high potential for the detection of cracks and holes in a space environment.
Concerning EMATs tools, this method presents a high potential for thickness measurements in a space environment. The limitation of use in the considered range of temperature has been validated down to a thickness equal to 0.80 mm with the use of a transducer with a working frequency equal to 4 MHz. The decrease of measurement precision due to curvature of the plastic deformed area, and due to local roughness of the impact site, has also been enlighted.
Concerning visual inspection tool, it can be said 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 its done, demonstration has been made that we could obtain the same level of automation than for others NDI tools studied before.

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.

© AIPnD , created by NDT.net

|Home|    |Top|