A set of experiments was conducted with the aim of evaluating total heat flux mitigation using magnetohydrodynamic flow control over ionized shock layers surrounding planetary entry vehicles. The University of Queensland's X2 expansion tube was used to generate a flow condition representative of a Mars return trajectory. Spherical permanent magnets were fitted into a scaled test model based on NASA's Stardust capsule. Thermocouples located at the stagnation point and shoulder of the test model, as well as thin-film heat transfer gauges located in the aft section, measured total heat flux for both magnetic and nonmagnetic experiments. High-speed imaging of the shock layer showed a significant increase in shock standoff when using the magnet, and polynomial fits were derived, describing standoff distance along the model's forebody. Despite significant noise affecting the heat flux gauges signals for the magnetic experiments, a simple methodology allowed to compare heat flux during the useful test time. It was found that the applied magnetic field significantly reduces total heat flux at the stagnation point and shoulder of the test model, although heat flux to the aft section was found to significantly increase. Reasonable agreement with the numerical predictions was found for the aft section of the test model, but discrepancies were observed at the other locations.

The aerodynamic heating experienced by capsules entering into the atmospheres of Saturn, Uranus, and Neptune is greatly affected by chemically nonequilibrium processes occurring in the shock layers. There are several reaction schemes available in numerical predictions for hydrogen dissociation and ionization, and more experimental data would assist verifying these existing models. This paper reports the results of electron number density measurements conducted in the X2 expansion tube at the University of Queensland using a condition representative of a proposed Saturn entry, where significant nonequilibrium effects in the shock layer are expected. Electron number density along the stagnation streamline was obtained from Stark broadening. The data presented here provide independent measurements for evaluating the reaction schemes for the conditions created. It was found that the experimental data were qualitatively modeled by a contemporary kinetic model. Quantitative agreement between experimental and numerical data was found by adjusting the ionization rate coefficients from an existing reaction scheme by a factor of 25. This updated reaction rate set was also cross-validated with electron number density measurements in NASA’s shock tube tests. The adjusted rates had better agreement than using the original rates, and quantitative agreement can be found in the high-density cases.

The T6 Stalker Tunnel is a multi-mode, high-enthalpy, transient ground test facility. It is the first of its type in the UK. The facility combines the original free-piston driver from the T3 Shock Tunnel with modified barrels from the Oxford Gun Tunnel. Depending on test requirements, it can operate as a shock tube, reflected shock tunnel or expansion tube. Commissioning tests of the free-piston driver are discussed, including the development of four baseline driver conditions using piston masses of either 36 kg or 89 kg. Experimental data are presented for each operating mode, with comparison made to numerical simulations. In general, high-quality test flows are observed. The calculated enthalpy range of the experimental conditions achieved varies from $$2.7\hbox { MJ kg}^{-1}$$ to $$115.0\hbox { MJ kg}^{-1}$$ .

A set of magnetohydrodynamic (MHD) aerobraking experiments was conducted with the aim of simulating the Earth reentry environment for Mars return trajectories, and to study the effects of MHD flow control on the ionizing shock layer surrounding a planetary entry vehicle. The X2 expansion tube of the University of Queensland was used, which could provide realistic flowfield boundary conditions, where ionization is spontaneously generated within the shock layer as in true flight and the freestream is nonconducting. Suitable flow conditions providing strong MHD interaction at superorbital Earth reentry velocities were developed. Steel ball and neodymium permanent magnet models were tested, both uncoated and coated with electrically insulating high-temperature epoxy. Shock-layer high-speed imaging showed a significant increase in shock standoff distance for the magnetic models. The experimental shock standoff results generally agreed well with analytical and numerical shock standoff predictions. These experiments demonstrate that a strong MHD interaction can be generated for a Mars return trajectory with moderate magnetic field strengths. Furthermore, they provide a promising framework for future MHD aerobraking experiments to investigate MHD drag force and heat-flux mitigation.

The presence of strong non-linear interactions and inherent uncertainties associated with hypersonic flight makes vehicle design particularly challenging. As such, the use of different aerodynamic modelling techniques and analysis tools can have a profound impact on vehicle performance. Oftentimes, computational restrictions make it desirable to use computationally efficient tools, which are likely to introduce more modelling uncertainty compared to higher-fidelity, computationally expensive tools. This paper examines the impact uncertainties in vehicle aerodynamic characteristics, which arise from using different analysis tools, have on system performance when attempting to fly an optimal flight trajectory. A delta wing configuration of the North American X-15 is used as a candidate vehicle to perform a 15-25 km minimum time to climb manoeuvre at a nominal flight speed of Mach 6. Two aerodynamics databases were developed to each represent 'design' and 'truth' aerodynamic characteristics. The 'design' aerodynamics database was built using the inviscid flow solver Cart3D, and the 'truth' database adds fidelity to this database by approximating viscous drag effects. Using these two databases, flight simulations with and without model mismatch were run to investigate how the differences between the 'design' and 'truth' databases impact open-loop performance. This study showed that there a significant impact on performance, despite a relatively small change in the aerodynamics of the vehicle. Since uncertainties are inherent to hypersonic flows, the results emphasise the need for co-design of vehicle aerodynamics and control systems to ensure near-optimal performance in flight.

A promising technique for protecting the sharp leading edges of a high-speed vehicle from the intense heating of hypersonic flight is Electron Transpiration Cooling. This paper develops a simplified theoretical model of the energy balance at the stagnation point of a blunted sphere-cone, and computes the adiabatic temperature of the leading edge over a range of altitudes from 10km to 60km and flight velocities from 1 km/s to 6 km/s, for various cone radii between r=1mm to r=100mm. A pair of leading edge temperatures is computed at each condition, one with the incoming heat from the flow balanced by surface radiation, and a second with the incoming heat balanced by both radiation and electron transpiration cooling. The ETC augmented case is significantly cooler, opening up most of the flight space compared to the radiation only case, which is mostly unavailable assuming a nominal material limit of 2000K. The second part of the paper extends these conclusions using numerical CFD modelling of the r=10mm case, finding that the radiation only cases are approximately 500-1000 K cooler than the simple model predicts, but that the effect of ETC is still significant.

Experiments were performed using the University of Queensland's X2 expansion tube with the aim of measuring stagnation point radiative heat flux to a blunt reentry body in superorbital Earth reentry air test flow. The selected flow condition produced a flight equivalent velocity of 14.9 km/s, representative of a Mars return trajectory. A radiation gauge was developed which consisted of a thin-film heat flux gauge placed behind an optical window, enabling isolation of the radiative heat flux component. Magnesium fluoride and B270 windows were used which made it possible to isolate the vacuum ultraviolet and infrared contributions, which are the main radiating wavelength regions at Mars return conditions. Absorbance calibration of the gauges had to account for the spectral dependence of the Nickel sensing element, and relied on radiation simulations using NEQAIR to predict the spectral distribution of the radiative heat flux. Experimental measurements yielded an average calibrated radiative heat flux of 10.7 MW/m2 over the 120 - 7000 nm wavelength range, of which 60% was predicted to be in the vacuum ultraviolet. Experimental radiative heat flux was around four times higher than the NEQAIR prediction, which, based on CFD did not, however, account for precursor radiation and excitation of the high temperature freestream.

ABSTRACTHypersonic air-breathing propulsion can improve cost and flexibility of Low Earth Orbit (LEO) satellite launch missions. However, at the high flight Mach numbers required for access-to-space, performance margins are extremely tight. Techniques to improve mixing efficiency can push this technology forward. However, these are required to produce a minimal increase in losses and heat loads to be viable. The use of inlet-generated vortices in scramjets for mixing enhancement was previously studied. These vortices interact with the injected fuel plume, stretching it and increasing its effective surface for mixing. Moreover, these vortices are intrinsic to the flowfield. Therefore, contrary to other methods, when using inlet vortices mixing is enhanced without producing additional heat loads or losses. This work studies the vortex-injection interaction through numerical RANS simulations. A non-dimensional variable defining the quality of the plume shape for mixing purposes is proposed. This parameter is used to assess the effect of vortex intensity and injector location on fuel plume shape. The results show the ability of inlet vortices to modify fuel plume shape significantly increasing fuel mixing rate with minimal impact on losses.

ABSTRACT The eddy dissipation model (EDM) is analysed with respect to the ability to address the turbulence–combustion interaction process inside hydrogen-fuelled scramjet engines designed to operate at high Mach numbers (≈7–12). The aim is to identify the most appropriate strategy for the use of the model and the calibration of the modelling constants for future design purposes. To this end, three hydrogen-fuelled experimental scramjet configurations with different fuel injection approaches are studied numerically. The first case consists of parallel fuel injection and it is shown that relying on estimates of ignition delay from a 1D kinetics program can greatly improve the effectiveness of the EDM. This was achieved through a proposed zonal approach. The second case considers fuel injection behind a strut. Here the EDM predicts two reacting layers along the domain which is in agreement with experimental temperature profiles close to the point of injection but not the case any more at the downstream end of the test section. The first two scramjet test cases demonstrated that the kinetic limit, which can be applied to the EDM, does not improve the predictions in comparison to experimental data. The last case considered a transverse injection of hydrogen and the EDM approach provided overall good agreement with experimental pressure traces except in the vicinity of the injection location. The EDM appears to be a suitable tool for scramjet combustor analysis incorporating different fuel injection mechanisms with hydrogen. More specifically, the considered test cases demonstrate that the model provides reasonable predictions of pressure, velocity, temperature and composition.

For turbulent simulations in hydrodynamic applications, a fine mesh close to wall boundaries is required to correctly predict the friction and heat fluxes. This results in a remarkable increase of computational cost due to the high aspect ratio cells and reduced allowable time step. To reduce the computational cost, a compressible wall function is introduced and validated with the representative flows. Moreover, high aspect ratio cells lead to simulation instabilities. These spurious oscillations are smoothed through the addition of a fourth-order artificial dissipation term. To verify that the solution accuracy is not affected, the method of manufactured solution is applied. These two additions result in a fast and stable solver for turbulent simulations of foil thrust bearings.

A spacecraft returning to Earth requires an effective thermal protection system (TPS) to withstand the hypervelocity flow and extreme heat loads encountered during an aerobrak-ing descent. The potential to optimise TPS designs while ensuring a successful return depends on confidence in ground-based prediction methods, which can be achieved by verifying such methods with flight data. The infrared (IR) and ultraviolet (UV) spectroscopic data collected by University of Queensland researchers during the Hayabusa spacecraft re-entry has been the basis for numerical studies using the in-house compressible flow computational fluid dynamics (CFD) code eilmer3 and experimental studies in the X2 expansion tube. Numerical spectra generated from CFD simulations of the full scale vehicle at the selected trajectory point will be compared to experimental spectra captured from the binary scaled and enthalpy matched radiating shock layer around a Hayabusa model, and both sets of ground testing data will be compared to flight.

This paper demonstrates the use of the method of manufactured solutions to verify the implementation of tightly coupled conjugate heat transfer for fluid–solid solvers. The interface conditions in the prescribed manufactured solutions were implemented to mimic real effects such as no-slip, and temperature/heat flux match between the solid and fluid domains. The newly developed solid heat transfer solver was verified in standalone mode using this prescribed manufactured solution and was found to have no apparent coding errors. Our pre-existing in-house compressible fluid solver (Eilmer) was used to demonstrate the conjugate heat transfer implementation. Both the fluid and solid solvers showed an expected spatial order of convergence of 2.0 in the standalone mode. The coupled conjugate heat transfer mode also showed no coding errors and demonstrated that the spatial order of convergence was again 2.0. The one-sided spatial discretisation utilised to enforce the tight coupling for the interface conditions were effectively equivalent to a central difference. Hence, the overall spatial order of the error convergence for the entire domain, including the interface, was 2.0. The method prescribed in this work can be extended for verification of other conjugate heat transfer solvers, in particular for compressible flow scenarios where analytical solutions may not be readily available.

This paper examines the effect of modelling choices on the numerical simulation of premixed methane air combustion in narrow channels. Knowledge on standard and well-accepted numerical methods in literature are collected in a cohesive document. The less well-established modelling choices have been thoroughly evaluated and discussed. A systematic method of computing the grid convergence index (GCI) has been presented for refining the computational grid. Two types of inflow boundary conditions have been tested and compared in terms of their wave-damping characteristics. The effect of different reaction schemes on simulation results have been examined and an appropriate mechanism (DRM-19) has been selected. Various types of ignition strategies to initiate the flame have been tested and compared. The transient ignition process which has not been discussed extensively in existing literature has been quantitatively described in this paper. (C) 2016 Elsevier Ltd. All rights reserved.

The effects of hydrogen and carbon monoxide addition on premixed methane/air flame dynamics in a heated narrow channel are numerically investigated using a time accurate, compressible flow solver along with the DRM-19 reaction mechanism. By adding a small amount of either H2 or CO, flame instabilities present for the pure CH4/air combustion in the form of flame-extinction and re-ignition could be effectively suppressed. This suppression can be attributed to a few important elementary reactions, that play a dominant role in contributing to the heat release rate, getting enhanced with H2 or CO addition. This leads to a higher flame propagation speed and a shorter flame-extinction period, and eventually leads to flame stabilisation after a few cycles of spatial oscillations.

The foil bearing is an enabling technology for turbomachinery systems, which has the potential to enable cost efficient supercritical CO2 cycles. The direct use of the cycle's working fluid within the bearings results in an oil-free and compact turbomachinery system; however, these bearings will significantly influence the performance of the whole cycle and must be carefully studied. Moreover, using CO2 as the operating fluid for a foil bearing creates new modeling challenges. These include highly turbulent flow within the film, non-negligible inertia forces, high windage losses, and nonideal gas behavior. Since the flow phenomena within foil bearings is complex, involving coupled fluid flow and structural deformation, use of the conventional Reynolds equation to predict the performance of foil bearings might not be adequate. To address these modeling issues, a three-dimensional flow and structure simulation tool has been developed to better predict the performance of foil bearings for the supercritical CO2 cycle. In this study, the gas dynamics code, eilmer, has been extended for multiphysics simulation by implementing a moving grid framework, in order to study the elastohydrodynamic performance of foil bearings. The code was then validated for representative laminar and turbulent flow cases, and good agreement was found between the new code and analytical solutions or experiment results. A separate finite difference code based on the Kirchoff plate equation for the circular thin plate was developed in Python to solve the structural deformation within foil thrust bearings, and verified with the finite element analysis from ansys. The fluid-structure coupling algorithm was then proposed and validated against experimental results of a foil thrust bearing that used air as operating fluid. Finally, the new computational tool set is applied to the modeling of foil thrust bearings with CO2 as the operating fluid.

To provide data for the validation of computational fluid dynamics models, measurements of the shock standoff distance on spheres in hypervelocity flows have been made. Test flows of air at 8.7 and 9.7 km/s were generated in the X2 expansion tunnel fitted with a Mach 10 nozzle. High-speed video images were analysed with a least-squares shape-fitting algorithm. Assuming a spherical shock shape near the nose enabled increased resolution measurements beyond the native pixel size. Normalised shock standoff distances, \(\Delta \)/\(D\), in the range 0.03–0.04 were measured, with sphere diameters, \(D\), of 40, 60 and 80 mm.

SUMMARYWe describe the formulation of the gas dynamics and high‐temperature thermochemical modules of the Eilmer code, an open‐source Navier–Stokes solver for transient compressible flow in two and three dimensions. The core gas dynamics formulation is based on finite‐volume cells, and the thermochemical effects are handled with specialised updating schemes that are coupled into the overall time‐stepping scheme. Verification of the code is explored via a number of case studies that use analytic and semi‐analytic solutions as comparison. These include both smooth and shocked flows and are used to demonstrate the order of spatial accuracy of the code. Cases include manufactured solutions for rather abstract inviscid and viscous flow, an idealised detonation wave supported by a curved body, and the transient flow of an idealised but high‐performance shock tube. Validation of the inviscid gas dynamics and thermochemical models is then explored using data from a selection of experimental studies. These studies include ballistic range experiments with chemically‐inert noble gases and high‐temperature chemically‐reacting air. These comparisons show that the code performs well and they provide a lesson in considering a range of experimental data rather than relying upon isolated data points for validation. These verification and validation cases are described in full detail and will be useful for other code developers of high‐temperature compressible flow solvers. Copyright © 2013 John Wiley & Sons, Ltd.

Experimental pitot pressure measurements in impulse facilities have typically had large scale harmonic fluctuations associated with them. A combined experimental and numerical approach is used to investigate if a Helmholtz resonance is created from the shrouding that protects the pressure sensor from particle impact, rather than disturbances being present in the free stream. To verify this experimentally, hydraulic oil was used to change the sound speed in the cavity. Numerical calculations of the pitot probe were used to show that both the steady and the expected transient inflow cause a Helmholtz resonance of similar period. However, a 5% isothermal level of free stream noise was required to match the experimental fluctuation levels. Also, viscous eects were shown to be significant during the initial transient response of the pressure measurement. A new pitot probe design was successfully tested and shown to be able to reduce the magnitude and period of the fluctuations.

We describe a numerical modelling technique used to simulate the gas flow in the complete X2 facility in non-reflected shock tube mode. The technique uses a one-dimensional model to simulate piston dynamics and diaphragm rupture and couples this to an axisymmetric simulation of the shock tube which captures viscous and finite-rate chemistry effects. This technique is used to simulate a nonequilibrium radiation condition relevant to a Titan atmospheric manoeuvre. The condition is a 7 km/s shock propagating into a N-2/CH4 Mixture at 80 Pa. The results show that the shock remains relatively planar at the exit of the shock tube such that there should be little difficulty for the optics. In terms of modelling, the finite-rate chemistry gas performs better than the equilibrium gas for these flows with regards to flow property estimates.