CHARACTERISATION OF MATERIAL DEMISABILITY THROUGH PLASMA WIND TUNNEL EXPERIMENTS Adam S. Pagan (1), Bartomeu Massuti-Ballester (1), Georg Herdrich (1) , James A. Merrifield (2) , James C. Beck (3), Volker Liedtke (4), Nils Stelzer (4), Benoit Bonvoisin (5) (1)
Institute of Space Systems, University of Stuttgart, Pfaffenwaldring 29, 70569 Stuttgart, Germany Email:
[email protected] /
[email protected] /
[email protected] (2) Fluid Gravity Engineering Ltd, The Old Coach House, 1 West Street, Emsworth, Hants, United Kingdom, Email:
[email protected] (3) Belstead Research Ltd, 387 Sandyhurst Lane, Ashford, Kent, United Kingdom, Email:
[email protected] (4) Aerospace and Advanced Composites GmbH, Viktor-Kaplan-Straße 2 Objekt F, 2700 Wiener Neustadt, Austria, Email:
[email protected] /
[email protected] (5) ESA ESTEC, Keplerlaan 1, NL-2200 AG Noordwijk, The Netherlands, Email:
[email protected] ABSTRACT In the context of the ESA-funded Characterisation of Demisable Materials project, experimental procedures for the assessment of spaceflight-relevant material demise during atmospheric entry are presented for plasma wind tunnel facilities at the Institute of Space Systems (IRS). A short overview of aerothermal demise mechanisms is given. In light of these, five representative sample materials are selected for initial testing in the Plasma Wind Tunnel 1 (PWK1) facility at IRS. The facility is described together with the intrusive and non-intrusive measurement techniques employed. Criteria are formulated for a generic test procedure, which is then developed accordingly. Finally, the plasma wind tunnel test campaign is put in context to related activities. 1. INTRODUCTION With ever-increasing numbers of artificial satellites in Earth’s orbit becoming a growing concern, attention is increasingly directed towards ensuring a safe and ultimate disposal of satellites that have reached their respective end-of-life. This is typically accomplished through means of a destructive re-entry manoeuvre, wherein the vehicle breaks up and proceeds to demise as a result of the high heat fluxes which occur as the kinetic energy of the vehicle is dissipated by the Earth’s atmosphere. In order to reduce the potential risk to life and property, it is vital that either surviving debris must be ensured to impact in uninhabited areas or that, alternatively, the vehicle burns up to a degree at which no notable debris remains at all [1]. A resource-hungry active deorbit control system would be required for the former option, for which little or no contingency can be made in case of failure. It is thus preferential to anticipate an uncontrolled re-entry event to begin with. Thus, the initial structural vehicle layout and material selection should be performed in such a way that a complete vehicle demise is essentially guaranteed in what is known as the Design-for-Demise approach [2]. To successfully implement such a design philosophy,
comprehensive knowledge of the behaviour of candidate materials subjected to high-enthalpy gas flow environments is vital. At the Institute of Space Systems (IRS), a selection of materials typical for space applications, which includes Al7075 aluminium alloy, titanium grade 5 (Ti6Al4V), stainless steel 316L, sintered silicon carbide (SSiC), and carbon fibre reinforced polymer (CFRP), is to be thoroughly tested in a plasma wind tunnel (PWT) facility within the context of the Characterisation of Demisable Materials ESA project. One objective of this campaign is to develop and subsequently apply standardised test procedures for the assessment of material demisability. A brief overview of the most relevant thermochemical processes and material properties governing the behaviour of different types of materials subjected to reentry-like conditions is presented and discussed. In this context, the experimental setup and the conditions under which the material samples are to be examined are described. A set of test procedures is developed, based on feasibility criteria and experimental data requirements. 2. AEROTHERMAL DEMISE Any spacecraft re-entering Earth’s atmosphere will experience considerable aeromechanical and thermochemical loads, resulting both directly and indirectly from the dissipation of the vehicle’s kinetic energy. A number of factors, including but not limited to the vehicles relative velocity and entrance angle, its size and geometry as well as its material composition, influences the quality and quantity of these loads. At what is known as the fragmentation altitude, vehicles or components that are unshielded by a thermal protection system (TPS) will break up, releasing fragments and internal components, which proceed to decompose individually and at rates that are again highly dependent on their respective shapes and compositions. This event typically occurs at altitudes between 60 and 90 km [1, 3]. Different materials exhibit different response
behaviours when experiencing such extreme heat loads. These may even increase as some materials encourage exothermic thermochemical reactions at (catalysis) or with (combustion, oxidation) the surface in the energyand oxygen-rich environment of the boundary layer, further contributing to the load. Depending on material properties such as the total emissivity, the heat capacity and transfer coefficient, some of the impinging heat flux is transferred from the surface further into the fragment or returned to the surroundings through thermal radiation and convection. Whatever remains further increases the (surface) temperature of the object, until it is ultimately forced to decompose and thus absorb the energy influx through latent heats associated with material-specific processes such as melt and sublimation, erosion and pyrolysis. The temperature at which an irreversible onset of these auto-destructive processes first occurs is referred to as the critical temperature. 3. MATERIAL SELECTION Five different materials are selected for the initial phase one plasma wind tunnel experiment campaign. This selection includes the three metallic alloys stainless steel 316L, aluminium alloy Al7075-T6, and grade 5 titanium Ti6Al4V, all commonly found in aerospace structures. These alloys have in common that their demise in the face of an impinging superheated gas flow is dominated by melt. As such, their critical temperatures essentially correspond to their respective melting points. Sintered silicon carbide SSiC is selected as representative for ceramic compounds. It finds frequent use in space-borne optical instruments due to its very low thermal expansion coefficient. This material features a high critical temperature at approximately 1700 K, where it gradually begins to succumb to active oxidation [4], and a high total emissivity [5], implying that much of the heat influx is removed quickly by thermal radiation. Its generally high thermal resistance renders its use somewhat of a liability with respect to uncontrolled re-entries, as large ceramic structures with very similar properties on board of returning satellites such as ROSAT are known to have essentially survived to impact [6]. The final sample material selected is carbon fibre
Figure 1. Exemplary standard material sample geometry. Thicknesses vary between materials.
reinforced polymer (CFRP), which is expected to behave much like an ablative thermal protection system material would under intense thermochemical loads. In addition to ablation at the surface dominated by a variety of processes such as sublimation, oxidation and erosion, the binding polymer within the material sample’s volume undergoes endothermic pyrolytic reactions which act as a heat soak and release gases. These syngases diffuse through the surface and into the boundary layer, essentially thickening it and thus reducing the overall heat load to the surface in what is often referred to as the blowing effect [7]. The material fragment’s lifetime is thus prolonged up to the point where it has charred through, i.e. when the volume ablation processes have run their course. Despite all of these processes taking place at comparatively low temperatures, defining a specific critical temperature is difficult due to the compound’s non-linear behaviour. The sum of these five material choices is deemed representative for most typical spaceflight applications, with additional materials to be tested in the course of the second phase of the experiment campaign. A flat, conical standard sample geometry is selected for all specimen (see also Fig. 1), with an exposed surface diameter of 26.5 mm. This, combined with a radiationcooled sample holder (see also chapter 5), allows for the experimental simulation of an effective 1D heat transfer environment and enables the contactless measurement of both the front and rear surface temperatures. In addition, these sample geometries may be subjected to total and spectral emissivity measurements in the Emissivity Measurement Facility (EMF) used at IRS, both in their virgin state and following a plasma wind tunnel test, allowing to anticipate and account for the radiative behaviour of the respective materials with a higher accuracy. The sample thickness varies for each material type between 3mm and 4mm. Tab. 1 presents a comparison of the material properties deemed most relevant for the PWT test campaign. Table 1. Selected phase 1 material properties. Total approximate emissivities have been obtained from [5] (except for CFRP) and were measured near or extrapolated to critical temperatures. Dominant process
Approximate critical temperature
Total emissivity
SS316L
Melt
1650 K
0.8
Al7075
Melt
830 K
0.3
Ti6Al4V
Melt
1900 K
0.6
SSiC
Active oxidation
1700 K
0.83
CFRP
Various
NA (charthrough criteria)
0.85
4. FACILITY The plasma wind tunnel 1 (PWK1) facility located at IRS [8] in Stuttgart is employed for the conduction of this test campaign. PWK1 consists of a cylindrical steel vacuum chamber of 6m length and 2m diameter (as pictured in Fig. 2) with double-wall cooling. While the hemispherical back end is connected to the in-house vacuum system and protected against high heat loads through water-cooled copper shielding, the lid at the front end features a conical depression into which any suitable plasma generator may be flanged and subsequently connected to working gas, cooling water, and electrical power feeds. The facility is equipped with a 4-axis positioning system onto which various probes or a specimen support system can be mounted and which facilitates a free movement of these during tests. A range of optical windows located at key positions in the tank body and lid allows experimenters to directly monitor the test and to employ a number of nonintrusive measurement techniques. Self-field magnetoplasmadynamic generators (MPG) of varying sizes have been developed over the past decades at IRS [8]; they are employed successfully in the IRS plasma wind tunnels PWK 1 and PWK 2. The nozzletype MPG plasma generators consist of two coaxial electrodes, separated by water-cooled neutrodes. The nozzle exit, which is also a water-cooled segment, forms the anode. The cathode, made of 2% thoriated tungsten, is mounted in the centre of the plenum chamber. For the MPG RD5 (see Fig. 3), featuring a nozzle exit diameter of 125 mm, mass flow rates between 0.3 g/s and 50 g/s can be attained at current levels between 200 A and 4 kA and power levels from 40 kW up to 1 MW. The average specific enthalpy at the nozzle exit varies between 2 MJ/kg and 150 MJ/kg. The in-house vacuum system simulates pressure environments corresponding to altitudes of up to 90 km in Earth’s atmosphere. By adding or removing
Figure 2. Plasma Wind Tunnel 1 (PWK1)
Figure 3. Cross-sectional drawing of the MPG RD5 individual pumps from the circuit and/or by mixing additional air into the system, the desired tank pressure can be freely adjusted between the highest achievable vacuum at 0.5 Pa and 100 kPa. 5. MEASUREMENT TECHNIQUES 5.1. Intrusive measurement techniques Probes are intrusive measurement devices within the vacuum chamber, to be positioned within the plasma stream. These probes are mounted onto a four-axis moveable platform, enabling full linear freedom of movement in all three dimensions as well as the use of double-headed probes, to be rotated in the course of an experiment. The 50mm diameter probe, to be used in the stagnationpoint flow investigation of the phase 1 material samples is referred to as the sample support system, and features a radiation casing mounted on a water-cooled structure (see Fig. 4). The conical material sample coin is pressed against an insulating ring (not pictured) fitted into the frontal opening of the silicon carbide cap via three spring-held zirconium oxide rods. Minimal physical contact point areas as well as a low temperature gradient between the sample and the cap reduce non-radiative heat transfer of the sample to the support system to a minimum and essentially creates a quasi-1D heat conduction situation. The design of the sample support system has originally been developed for the investigation of catalytic material properties [9], and is of particular interest to this campaign, as it provides means for a contactless measurement of the sample’s rear surface temperature through the miniature linear pyrometer MINI-PYREX (MP3) developed at IRS [4]. The optical head of MP3 is situated to the rear of the sample and is connected to the signal converter via an optical fibre connection. In preparation for material testing, heat flux / enthalpy and Pitot pressure probes made of copper are deployed
Figure 4. Schematic of an IRS 50mm material probe with integral pyrometer optical fibre: material sample (1); ceramic cap (2); insulator foam (3); sample holding rods x3 (4); compression springs x3 (5); ceramic fixation bolts (6); insulator foam (7); water-cooled metallic adapter (8); water-cooled probe casing (9); ceramic tube enabling optical access to the sample’s rear surface (10); optical fibre head and collimation lens (11); optical fibre (12). Not pictured is an additional ceramic ring placed between the sample and the ceramic cap providing additional insulation towards the characterisation and verification of the plasma flow environment. These probes are designed to exhibit geometric similarity to the sample support probe. 5.2. Non-intrusive measurement techniques The non-intrusive measurement techniques employed for this campaign encompass the use of optical temperature measurement devices as well as Optical Emission Spectroscopy (OES). Pyrometers are used to determine both the front and rear wall temperature of the sample during experiments. The linear pyrometer LP3 80/20, developed with the Physikalisch-Technische Bundesanstalt (PTB) [10], is positioned diagonally to the outside of the vacuum chamber in such a way that the material sample’s surface is in clear view. It features advanced electronics for photocurrent signal processing which allows for reliable real-time measurements of a wide temperature range without the need of additional neutral density filters. The LP3 is a radiance-measuring instrument, the output signal of which is highly proportional to the incident radiation flux, and demonstrates an excellent reliability and precision, provided that the relevant spectral emissivity [5] of the sample material as well as the vacuum tank window’s transmittance are accounted for in the post-processing. The internal interference filter is selected to measure at 958.1nm, a wavelength at which the plasma is essentially transparent. A LumaSense Mikron MCS640 thermographic imaging camera [11] is positioned alongside the LP3 pyrometer in order to provide additional, spatially resolved sample
front surface temperature measurements. The infrared camera is calibrated to measure at 960±5 nm. Within the sample holder, the rear surface temperature of the material sample is monitored using the MiniPYREX (MP3) integral pyrometer developed at IRS [4]. Thermal radiation emanating from the sample is collimated in an optical head and passed on to an InGaAs (P) photodiode via an optical fibre. As the sample’s back surface is enclosed in a quasi-isothermal cavity and is observed through a graphite tube placed between it and the measuring head of the MP3, it is assumed to exhibit optical properties approaching those of a black body radiator. Due to the very delicate nature of the optical fibres connecting the measuring head to the photodiode setup, the MP3 is calibrated both prior to and following its removal from the vacuum chamber in order to determine potential damage to the optical fibres and to adjust post-processing accordingly. To this end, a black body radiator setup at IRS provides a reliable source for reference temperature measurements. A schematic of the spectroscopic setup is depicted in Fig. 6. Through Optical Emission Spectroscopy (OES), atomic and select molecular emission lines associated with distinct species emanating from the sample specimen subjected to the thermochemical loads exerted by the impinging plasma may be observed and ultimately quantified. A monochromator / spectrograph in Czerny-Turner configuration (Acton SpectraPro 2750) is used together with an Andor Newton DU920N-OE EMCCD camera to detect the spectra. Various spectrometer gratings are available for different line number densities and spectral resolutions [12]. Measurements are taken from a thin volume covering an actual length of approximately 55 mm. Aligning this measurement volume vertically with the sample’s surface and positioning it within the prospective boundary layer yields a comprehensive picture of atomic species released by the specimen subjected to the thermochemical loads of the plasma plume. Assuming
Figure 5. Schematic of experimental setup for OES measurements in PWK1
rotational symmetry, a two-dimensional representation can be obtained by performing an Abel inversion during post-processing. In addition, a recently added periscope setup allows to directly rotate the measurement volume by 90° into a horizontal position following the longitudinal axis along the stagnation line. The exact focal point and spatial dimensions defining the measurement volume are fixed prior to the test by using an adequate light source such as an Ulbricht sphere and adjusting the optical setup accordingly. A wavelength calibration is then performed by utilising the known spectrum of a Mercury / Argon lamp, followed by an intensity calibration using the aforementioned Ulbricht sphere. 6. TEST CONDITIONS Three test conditions, differing primarily in the exerted heat flux density as determined through the copper reference heat flux probe, are selected for the experimental plasma wind tunnel campaign. The specific choice of test conditions stems from considerations of various selected destructive re-entry trajectories from Low Earth Orbit (LEO) [3, 13]. It was determined that conditions with reference heat fluxes at 260 kW/m² and 520 kW/m², respectively, would envelop most typical tumble-averaged values experienced by demising space debris. An additional condition with a 1.4 MW/m² heat flux is to account for
the effects of stagnation heating, which was estimated at up to fourfold the tumble-averaged values. A definition of the three selected baseline test conditions is presented in Tab. 2. Based on these recommendations, a set of three wellcharacterised conditions are selected for the RD5 MPD generator, which have demonstrated both excellent reproducibility and stability in past experimental campaigns [8, 9]. These conditions have been characterised using copper heat flux probes, as is standard procedure at IRS (see also section 5.1). Taking into consideration that copper exhibits strong catalytic behaviour as compared to many other materials [9], the actual heat flux experienced by the respective sample specimen differs significantly more between both one another and the reference value than the test conditions would initially suggest. 7. PREVIOUS EXPERIENCE WITH PM1000 Previous PWT campaigns at IRS have seen extensive testing of materials of potential and proven relevance both to radiation-cooled and ablative thermal protection systems, including Graphite, CFRP, Silicon Carbide as well as a range of super alloys such as PM1000. In most cases, the nature of these tests have resulted in a significant alteration, decomposition and mass loss to the respective specimen. Experiences gained in the course of subjecting samples of the nickel-chromium super alloy PM1000 to
Table 2. Selected plasma wind tunnel test conditions for material demisability test campaign Heat Flux Condition
260 kW/m²
520 kW/m²
1400 kW/m²
Parameter
Unit
xmean ± σ
xmean ± σ
xmean ± σ
Nitrogen mass flow
[g/s]
1.53 ± 0.01
1.6 ± 0.01
6.4 ± 0.01
Oxygen mass flow
[g/s]
0.47 ± 0.01
0.4 ± 0.01
1,6 ± 0.01
Argon mass flow rate
[g/s]
0,3 ± 0.01
0.1 ± 0.01
0,3 ± 0.01
Total mass flow
[g/s]
2.3 ± 0.03
2.1 ± 0.03
8.3 ± 0.03
Ambient pressure Specific enthalpy (effective) Heat flux
[hPa]
0.5 ± 0.2
2.9 ± 0.2
4.9 ± 0.2
[MJ/kg]
31.5 ± 0.1
28 ± 0.1
12 ± 0.1
[MW/m²]
0.263 ± 0.04
0.49 ± 0.05
1.3 ± 0.13
Pitot pressure
[hPa]
0.5 ± 0.2
4.35 ± 0.5
8…9 ± 0.2
Local mass spec. enthalpy (y = 0 mm)
[MJ/kg]
34
22
42
Reference sample material
-
SSiC
SSiC
Reference temperature
[K]
1353 (expected)
C 1543 (database value)
Reference mass loss rate
[kg/(m²h)]
< 0.1
-
nozzle
outlet
ca. 1950 ≈5
comparable test conditions in the IRS plasma wind tunnel PWK3 are considered as most instructive concerning the planned testing of metal alloys in particular. Fig. 6 depicts a selection of PM1000 specimen subjected to various test conditions in pure O2 and N2 plasma flows. Melting at the surface was observed at a conveniently slow pace, implying that sample materials of roughly comparable melting points will provide an ample window of time for the conditions presented in chapter 6, within which measurements may be conducted. The option of visual inspection allows for a timely manual termination of the test prior to compromising the sample’s integrity. An additional indication with regards to the specimen’s integrity is given by the thermal imaging camera and the linear pyrometer, which provide surface temperature measurements in real time. The onset of melt occurs at the point at which the measured temperature stabilises at the expected melting point, indicating an earliest convenient point in time at which the test may be terminated. For PM1000, this reference point is situated at 1681 K. This approach is deemed reasonably safe for materials with roughly similar properties as compared to PM1000 such as stainless steel or titanium alloy, while leaving ample time to conduct OES measurements. More care should be taken with aluminium due to its very low melting point. An optional procedure aimed at determining the survivability of fast-melting materials prior to an actual test is suggested in section 8.3.
Figure 6. PM1000 subjected to different test conditions in PWK3 facility. Clockwise from top-left: N2 plasma at 100 Pa ambient pressure, N2 plasma at 500 Pa, virgin sample, O2 plasma at 500 Pa.
8. TEST PROCEDURE 8.1. Generic procedure An initial generic procedure for the characterisation of material demisability is developed under consideration of criteria related to maintaining the feasibility of the required measurements as well as the desire to maintain the basic structural integrity of the specimen. Preventing violent deformation or disintegration is required in order to maintain an effectively uniform and reproducible heat load distribution throughout the test. In addition, further testing of the specimen, e.g. in an emissivity measurement facility (EMF) [5], is made possible. Thermomechanical failure is to be avoided accordingly by implementing stringent test termination criteria. Both the mass and the thickness of the respective sample specimen are determined prior to and following the plasma wind tunnel test. The pre- and post-test states of the samples are photographed and documented. Following the setup of the sample support system housing the specimen, the MP3 Mini-PYREX pyrometer is connected and the Optical Emission Spectroscopy setup is aligned and focused on the region of interest in close proximity to the sample’s front surface. The vacuum tank is closed and evacuated while the thermographic imaging camera and LP3 linear pyrometer are set up and focused at the sample’s front surface centre, whereby the sample initially rests at the position required for the desired test conditions. The sample support structure is then moved to the side, until a radial distance of 180mm of the plasma generator’s axis of symmetry is attained. Once the experimental setup is completed and the tank has been evacuated, the RD5 MPD generator is activated. The plasma flow conditions are gradually and carefully established, before the sample support structure is moved into the centre of the plasma plume. During this phase, the specimen suffers some preheating, the extent of which is primarily determined by the target condition and duration of the initialisation phase as well as by the sample’s relative position to the generator. Having reached the central position, the test commences and is finally terminated as per the material-specific conditions described in the following subsection. Whereas front surface temperature monitoring as well as OES measurements cannot begin before the sample has attained its test position, the MP3 pyrometer may already record the back surface temperature during the preheating phase. Termination is effected by deactivating the RD5 generator, upon which a cold gas flow briefly succeeds the plasma stream. Prior to pressurising the tank and removing the sample for post-test analysis, the behaviour of the specimen is observed for a further three minutes. As some of the materials tested may suffer from fast degradation, e.g. through melt, potentially
compromising the experimental setup, an optional pretest procedure is developed, designed to ascertain the survivability of a material in question prior to subjecting it to the actual test procedure. This additional procedure is described in subsection 8.4. 8.2. Procedural variations and termination criteria Mainly due to differences in their respective composition and durability against heat loads, the test procedure is varied for each individual sample material type. These variations are reflected mainly in the respective test termination criteria, which are naturally more stringent for materials with a low life expectancy, as well as in the wavelength spectrum selected for observation with the Optical Emission Spectroscopy (OES) setup. As these measurements benefit from attaining near-steady-state conditions and typically require multiple iterative manual attempts, in which the exposure time is varied, the observed spectra are selected in such a way that atomic emission lines correlated with as many of the most relevant species as possible may be recorded in a single shot. These lines must be situated within a common spectrum of approximately 120 nm. Exceptions from this restriction can be made for durable materials such as SiC, where extended test durations make multiple exposures in different wavelength regions possible, which allows the entirety of the spectrum observable with the OES setup (approximately 200 to 1010 nm) to be recorded. With the possible exception of the Al7075 aluminium alloy, previous experiences with metallic alloys such as PM1000 demising in the course of PWT tests at IRS suggest that the time required for a metal specimen to fully melt is sufficiently long to allow for a termination of the experiment based on constant visual inspection of the sample’s surface. Sufficient time exists to conduct the required OES measurements. Whereas this likely applies to Stainless Steel 316L and Titanium Ti6Al4V, an experimental pre-assessment of the behaviour of Al7075 is encouraged, as described in subsection 8.3. Extensive testing experience at IRS and other institutions with compounds identical or similar to those constituting the silicon carbide [13] and carbon fibre reinforced polymer [14] samples gives reason to believe that the basic structural integrity of both material sample types will be maintained for extensive timespans in all test conditions presented in this article. For SSiC this implies that a wider atomic emission spectrum can be covered by OES through the successive performance of varying measurements during a single test shot. Concerning CFRP, it is of potential interest to perform measurements for identical spectra at different stages of the respective specimen’s demise, as its decomposition is known to be a highly non-linear process (see also chapter 3). While the degradation of all specimen will be monitored closely, the impending complete demise of SSiC and
CFRP may not necessarily constitute an immediately relevant test termination criterion at the heat flux densities presented. A complementary termination criterion is thus introduced, wherein the test is terminated after a quasi-steady-state condition has been reached and maintained for a limited period of time. This state is indicated e.g. by the attainment of a constant surface temperature for SSiC and by the apparent suspension of outgassing from a CFRP specimen. 8.3. Pre-assessment of material survivability For materials prone to early melting such as aluminium, an additional test procedure is presented to be conducted in preparation of the actual investigation described beforehand. As the actual heat flux density experienced by the sample is difficult to predict with any accuracy for a given reference heat flux, one of the specimens of the material in question is used to determine the approximate reference value at which melt sets in. To this end, any one specific plasma source setting with a well-defined axial heat flux profile is selected from the three baseline conditions. The sample is mounted and moved to the furthest possible rear position of the generator, minimising the expected heat flux density. Once the generator conditions have been attained, the sample is moved in predefined intervals towards the source. At the onset of melt, the position and surface temperature are documented and the reference heat flux is determined from axial profiles known from earlier heat flux probe measurements. Using this information, a decision regarding the feasibility and adjustment of the tests under the conditions envisaged can be made. As the dynamic nature of this test variant forbids the use of a delicate optical fibre connection, the rear surface may not be monitored using the MP3 pyrometer. Similarly, OES measurements require a fixed sample position and are thus equally impossible under these conditions. Front surface temperature recordings are possible, but require a thoughtful initial setup of both the LP3 and the thermographic imaging camera and/or a dynamic repositioning thereof. In any case, the onset of melt should be determined redundantly through careful visual inspection. 9. SUMMARY AND OUTLOOK A preliminary procedure for the testing of spacerelevant materials has been developed for plasma wind tunnel experiments. This procedure considers the adequate representation of destructive atmospheric reentries through appropriate test conditions, individual material properties, the respectively expected material responses, the feasibility of the desired measurements, and incorporates previous experiences with demising specimen. In the following, a short overview of ongoing and future related activities is presented.
At Aerospace and Advanced Composites GmbH (AAC), preliminary experiments complementary to the PWT tests are ongoing in the Re-Entry Chamber static test facility. This facility provides additional flexibility, allowing for a full simulation of pressure and heat flux profiles over the course of an entire reference re-entry trajectory. In addition, mechanical effects may be instigated in-situ, providing for a more complete assessment of material demisability. Pending the completion of the initial PWT test series, static test conditions will be recalibrated depending on the prevalent demise mechanisms identified for each material in question, thus refining the static tests further. A methodology enabling a reliable scientific read-across between the static and dynamic facilities is to be established. Following the completion of the phase 1 material testing, the lessons learnt will to be utilised towards refining the test procedures developed, ultimately establishing standardised generic procedures to be applied to a second set of as of yet unspecified selection of sample materials. Numerical simulations conducted by Fluid Gravity Engineering Ltd. (FGE) using the Simplified Aerothermal Models (SAM) end-to-end software suite [16] will complement the data gained in the course of the experimental campaigns and may serve to verify and improve modelling approaches. The output of the experimental and modelling efforts is to culminate in the construction of a design for demise database as a tool assisting in the design and analysis of demisable spacecraft. These activities are further supported by Belstead Research Ltd. (BRE). 10. ACKNOWLEDGEMENTS All investigations referred to are conducted in the context of the Characterisation of Demisable Materials project. The plasma wind tunnel experiments are performed at the Institute of Space Systems at the University of Stuttgart in the state of BadenWürttemberg, Germany. The authors would like to gratefully acknowledge funding of these research efforts by the European Space Agency (ESA) under contract 4000109981/13/NL/CP. 11. REFERENCES 1. Patera, R.P. & Ailor, W.H. (1998). The realities of reentry disposal. In Proc. AAS/AIAA Space Flight Mechanics Meeting, pp9-11. 2. Kelley, R.L. (2012). Using the Design for Demise Philosophy to Reduce Casualty Risk Due to Reentering Spacecraft. In Proc. 63rd International Astronautical Congress. 3. Beck, J. & Holbrough, I. (2014). Test Conditions and Facility Selection for Representation of Uncontrolled Re-Entry Events, Technical Report,
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