A ground-based tracking camera and coaligned slitless spectrograph were used to measure the spectral signature of visible radiation emitted from the Hayabusa capsule as it entered into the Earth's atmosphere in June 2010. Good quality spectra were obtained, which showed the presence of radiation from the heat shield of the vehicle and the shock-heated air in front of the vehicle. An analysis of the blackbody nature of the radiation concluded that the peak average temperature of the surface was about (3100 +/- 100) K. Line spectra from oxygen and nitrogen atoms were used to infer a peak average shock-heated gas temperature of around (7000 +/- 400) K.

Vacuum Ultraviolet (VUV) emission spectroscopy radiation measurements were conducted for Titan and Mars atmospheric entry conditions using the Centre for Hypersonics X2 expansion tube. The VUV measurements were taken while viewing downstream through the shock layer of a scaled model. UV-Visible emission spectroscopy measurements were conducted in parallel, viewing the shock layer through a side window. The spectra was dominated by the CN violet molecular bands and the main atomic lines identified were C, N and Al. For a Titan 8.5km/s condition, sample of the calibrated spectra for the VUV is presented and discussed, alongside uncalibrated continues spectra from the VUV through to the UV-Visable.

The expansion tube is our tool of choice for producing radiating, high-enthalpy flows relevant to the study of re-entry aerothermodynamics. First, we describe the operating principles of a free-piston-driven expansion tube and two example facilities that have been constructed at The University of Queensland. These have been applied to a range of aerothermodynamic studies and a large effort has also been put into the development of instrumentation, including surface measurements, optical and spectrometric instruments. Two situations are described where radiating and ablating flows have been created, using both preheated hot walls and walls heated by aerodynamic processes during the test.The detailed flow in these expansion tube facilities turns out to be quite complex and computer models are needed to completely and accurately describe the experimental flow properties and to diagnose the gas dynamic behaviour of the machine when developing new operating conditions. We describe two of the specialized flow simulation codes that we have written to assist us in our experimental work and look at the application of these codes to the expansion tube experiments.

At high flight speeds, radiation becomes an important component of aerodynamic heat transfer, and its coupling with the flow field can significantly change the macroscopic features of the flow. As radiating flight conditions are typically encountered in re-entry trajectories, the associated flight regimes range from rarefied to continuum, and may have many levels of thermal, chemical and electronic nonequilibrium. Accurate estimates of the nonequilibrium radiation involved in high speed operations such as reentry are essential in order to more efficiently design thermal protection systems.

Hayabusa, an unmanned Japanese spacecraft, was launched to study and collect samples from the surface of the asteroid 25143 Itokawa. In June 2010, the Hayabusa spacecraft completed it’s seven year voyage. The spacecraft and the sample return capsule (SRC) re-entered the Earth’s atmosphere over the central Australian desert at speeds on the order of 12 km/s. This provided a rare opportunity to experimentally investigate the radiative heat transfer from the shock-compressed gases in front of the sample return capsule at true-flight conditions. At these conditions, the total heat transfer to the vehicle has a significant radiative component and this can be estimated by studying the radiation emitted from the shock layer and the hot surface. Such measurements can be compared with numerical simulations of the flow and with results from ground-based testing in shock tunnels and expansion tubes. This in turn leads to a better understanding of the complex thermochemistry occurring within the shock layer and aids in the design of more efficient thermal protection systems for future spacecraft.

There are large uncertainties in the aerothermodynamic modelling of super-orbital re-entry which impact the design of spacecraft thermal protection systems (TPS). Aspects of the thermal environment of super-orbital re-entry flows can be simulated in the laboratory using arc- and plasma jet facilities and these devices are regularly used for TPS certification work [5]. Another laboratory device which is capable of simulating certain critical features of both the aero and thermal environment of super-orbital re-entry is the expansion tube, and three such facilities have been operating at the University of Queensland in recent years [10]. Despite some success, wind tunnel tests do not achieve full simulation, however, a virtually complete physical simulation of particular re-entry conditions can be obtained from dedicated flight testing, and the Apollo-era FIRE II flight experiment [2] is the premier example which still forms an important benchmark for modern simulations. Dedicated super-orbital flight testing is generally considered too expensive today, and there is a reluctance to incorporate substantial instrumentation for aerothermal diagnostics into existing missions since it may compromise primary mission objectives. An alternative approach to on-board flight measurements, with demonstrated success particularly in the ‘Stardust’ sample return mission, is remote observation of spectral emissions from the capsule and shock layer [8].

The following paper outlines the importance of vacuum ultraviolet (VUV) aerothermodynamic heating during re-entry, the current level of uncertainty in the models used to predict radiative heating load, and the challenges associated with reducing this uncertainty. A method has been developed at The Centre for Hypersonics that allows the capture of VUV emission spectra from across and through the surface of a blunt model. Two superorbital conditions were generated to match high-speed points on a re-entry trajectory and validated computationally and experimentally. The proposed system has been constructed and a calibration procedure developed. There have now been over one hundred experiments at superorbital velocities conducted with this system and some sample uncalibrated spectral images are presented.

This paper presents an image-based visual servoing system that was used to track the atmospheric Earth re-entry of Hayabusa. The primary aim of this ground based tracking platform was to record the emission spectrum radiating from the superheated gas of the shock layer and the surface of the heat shield during reentry. To the author's knowledge, this is the first time that a visual servoing system has successfully tracked a super-orbital re-entry of a spacecraft and recorded its spectral signature. Furthermore, we improve the system by including a simplified dynamic model for feed-forward control and demonstrate improved tracking performance on the International Space Station (ISS). We present comparisons between simulation and experimental results on different target trajectories including tracking results from Hayabusa and ISS. The required performance for tracking both spacecraft is demanding when combined with a narrow field of view (FOV). We also briefly discuss the preliminary results obtained from the spectroscopy of the Hayabusa's heat shield during re-entry.

Thermochemical relaxation behind a normal shock in Mars and Titan gas mixtures is simulated using a CFD solver, DPLR, for a hemisphere of 1 m radius; the thermochemical relaxation along the stagnation streamline is considered equivalent to the flow behind a normal shock. Flow simulations are performed for a Titan gas mixture (98% N2, 2% CH4 by volume) for shock speeds of 5.7 and 7.6 km/s and pressures ranging from 20 to 1000 Pa, and a Mars gas mixture (96% CO2, and 4% N2 by volume) for a shock speed of 8.6 km/s and freestream pressure of 13 Pa. For each case, the temperatures and number densities of chemical species obtained from the CFD flow predictions are used as an input to a line-by-line radiation code, NEQAIR. The NEQAIR code is then used to compute the spatial distribution of volumetric radiance starting from the shock front to the point where thermochemical equilibrium is nominally established. Computations of volumetric spectral radiance assume Boltzmann distributions over radiatively linked electronic states of atoms and molecules. The results of these simulations are compared against experimental data acquired in the X2 facility at the University of Queensland, Australia. The experimental measurements were taken over a spectral range of 310-450 nm where the dominant contributor to radiation is the CN violet band system. In almost all cases, the present approach of computing the spatial variation of post-shock volumetric radiance by applying NEQAIR along a stagnation line computed using a high-fidelity flow solver with good spatial resolution of the relaxation zone is shown to replicate trends in measured relaxation of radiance for both Mars and Titan gas mixtures.

The University of Queensland (UQ) is currently developing high Mach number, high total pressure scramjet flow conditions in its X2 and X3 expansion tube facilities. These conditions involve shock-processing a high-density air test gas followed by its unsteady expansion into a low-pressure acceleration tube. This relatively slow shock-processing requires the driver to supply high pressure gas for a significantly greater duration than normally required for superorbital flow conditions. One technique to extend the duration is to operate a tuned free-piston driver. For X2, this involves the use of a very light piston at high speeds so that, following diaphragm rupture, the piston displacement substitutes for vented driver gas, thus maintaining driver pressure much longer. However, this presents challenges in terms of higher piston loading and also safely stopping the piston. This article discusses the tuned driver concept, the design of a very lightweight but highly stressed piston, and details the successful development of a new set of tuned free-piston driver conditions for X2.

The Hayabusa sample return capsule is scheduled for re-entry near Woomera, Australia in June 2010 and expansion tube experiments are being performed to support the planned re-entry observation campaign. Initial experiments using a 1/10th scale model of the Hayabusa forebody have been performed in the X2 expansion tunnel facility at The University of Queensland to simulate aerothermal elements of the anticipated re-entry. Experiments have been performed at an effective flight speed of around 9.8 km/s using steel models, and steel models coated with a layer of epoxy to simulate pyrolysis gases associated with heat shield ablation. Spectral emissions from the stagnation region of the capsule have been acquired using a spectrograph system. Two dimensional maps of the luminous emissions from the shock heated flow have also been acquired using a high speed camera. Deduction of flow conditions generated in the X2 expansion tunnel is achieved using quasione-dimensional simulations coupled to an axisymmetric simulation of the flow through the expansion tunnel nozzle. The effects of the ablative epoxy material are observed in the data from both the spectrograph system and the high speed camera. Both systems register strong emissions in the ablative layer, and the strength of the spectral peaks associated with CN emissions are shown to be enhanced by the presence of the epoxy. Further measurement and analysis is required to confidently define the flow conditions produced by the expansion tunnel, and to quantify results from the spectrograph and high speed camera measurements.

The observation of an ablating surface in expansion tunnel flow was reported. A one-dimensional (1-D) semi-infinite analysis was performed using an empirical estimation of the stagnation-point heat flux. A surface temperature change for epoxy of 178 K in 50 μs was calculated, which is sufficient for the epoxy coating to commence ablation during the steady test period. The 10% temperature penetration depth is 6 μm in 50 μs. The luminosity from substantial portions of the model and shock-layer flow was visualized using a Shimadzu HPV1 high-speed charge-coupled device (CCD) video camera recording at 500 kfps with a 1 μs exposure time. The uniform image response over the axisymmetric model implies the shock-layer gas irradiance is high immediately behind the shock and decreases rapidly as the model surface is approached. The measurements showed that use of an epoxy coating results in greatly increased CN and C2 radiation, much greater than when no coating is employed.

: Shock tunnels and expansion tubes are our tools of choice for producing high-enthalpy flows in aerothermodynamic studies relevant to hypersonic flight. The detailed flow in these machines turns out to be quite complex and computer models are needed to completely and accurately describe the experimental flow properties and to diagnose the gas dynamic behaviour of the machine when developing new operating conditions. We describe a couple of the flow simulation codes that we have written to assist us in our experimental work. The first is a quasi-one-dimensional flow code that is capable of modelling entire free-piston driven facilities, albeit with some significant limitations. The second is a finite-volume flow code that can more accurately capture the strong viscous and thermochemical interactions that affect strongly expanding flows that are an operating characteristic of expansion tube facilities. Some application studies are then described.

It is envisaged that future Lunar and Mars return spacecraft will re-enter Earth's atmosphere at speeds above 11 km/s. The heating environment is extreme at these speeds, necessitating well-designed thermal protection systems (TPS). Ablative TPS have demonstrated superiority over other forms of TPS under these conditions. As the TPS is heated by the shock-compressed gas in front of the spacecraft, the ablator undergoes pyrolysis and erodes, emitting gases and solid particles. This material is injected into the boundary layer and is known to shield convective heat transfer to the spacecraft; however little is known of the effect this material has on radiative heat transfer. Recent computational studies have found the effect of ablation on radiation varies throughout the electromagnetic spectrum. This paper presents the results of experiments measuring radiation in an ablating shocklayer over a Stardust model at 9.8 km/s in air, nitrogen and Mars atmospheres in the X2 expansion tunnel at The University of Queensland. Copyright