Super-orbital re-entry in Australia - laboratory measurement, simulation and flight observation

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]. JAXA’s ‘Hayabusa’ sample return capsule provides a recent super-orbital reentry example through which we illustrate contributions in three areas: (1) physical simulation of super-orbital re-entry conditions in the laboratory; (2) computational simulation of such flows; and (3) remote acquisition of optical emissions from a super-orbital re entry event

Application of photogrammetry at USQ hypersonic wind tunnel

This paper discusses an investigation into the use of photogrammetric techniques at the University of Southern Queensland's Ludwieg tunnel facility. The aim was to show that photogrammetric techniques and compact cameras could be used to achieve sub-millimetre precision in hypersonic speed tracking. The paper provides details on the setup of the imaging cameras, video capture, photogrammetric computation and the results of two aerodynamic experiments. A free-flying projectile constructed from a hollow metal cylinder with a round-nose and flat-tail was tracked using two low-cost Casio high-speed cameras. A set of retro-photogrammetric targets of known coordinates was used for camera relative orientation purposes. Corresponding frames were extracted from the video clips using VirtualDub software and a photogrammetric adjustment was applied to these frames to extract a set of 3D coordinates of the tracked targets. The results showed that the photogrammetric technique gave a 3D object-space measurement accuracy of 0.064 mm when the projectile was in a static position and a precision of 0.11 mm during its flight in a pressurized airflow

Comparison of Computational and Experimental Studies on Shock Induced Ignition in Scramjets

Due to the problems associated with rapid mixing of fuel in scramjet engines, attempts are now being made to inject the fuel on the intake ramp so that it is well mixed by the time it enters the combustor. This however requires a reduced compression on the intake to ensure the flow does not ignite before it reaches the combustor. Once in the combustor, the flow is then ignited via shock waves generated from the engine cowl. This paper initially examines the ignition characteristics of a simple 2-D premixed hydrogen/air flow through a duct of constant height to determine the mechanisms responsible for radical production and heat release. Results are then reported for computed 2D scramjet combustion chamber flows of premixed hydrogen and air. Results of a 3D CFD calculation without fuel injection are also presented, and compared with experimental data obtained from the T3 shock tunnel at the Australian National University. Experimental data is also presented for a scramjet flow with combustio

Expansion tunnel radiation experiments to support Hayabusa re-entry observations

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. Copyright © 2010 by Buttsworth, D'Souza, Potter, Eichmann, Mudford, McGilvray, McIntyre, Jacobs, and Morgan

13th Australasian Fluid Mechanics Conference

This paper presents a numerical investigation that provides a detailed study of heat release effects on the flow for two hypermixing fuel injectors designed for a supersonic combustor. Three injector configurations are compared: a swept compression-expansion ramp (SCER); a castellated blunt trailing edge, and a plane base injector. All injector configurations are strut mounted with hydrogen base-injection. A one-step reaction is used for all reacting cases. Details on nearfield flow and mixing mechanisms with effects of heat release are discussed. It is found that heat release decreases mixing significantly in the near-field