Neil Mudford

Radical farm ignition processes in two-dimensional supersonic combustion

A two-dimensional numerical study was performed of the ignition processes associated with the concept of radical farming for supersonic combustion. In a preliminary parametric study, a range of freestream conditions attainable in a hypersonic shock tunnel was investigated and mapped according to whether or not the behavior known as radical farming is present: a combustion-induced pressure rise in second or subsequent hot pockets rather than the first Two such cases were analyzed in detail, both having mean conditions across the combustion-chamber entrance that would result in extremely long ignition lengths. The initiation, branching cycle, and heat release reactions in the combustion process become active in the radical farm, and H and OH radicals are produced. Their rate of production slows in the regions of flow expansion, but does not approach chemical freezing until toward the end of the localized expansion zones. Simultaneously, heat release elevates the local temperature. When the mixture flows through the shock at the second hot pocket, the elevated temperature and the presence of radicals enable the branching cycle and three-body recombination heat release reactions to accelerate, and significant pressure rise due to heat release is then able to occur. The extent to which this is completed in the second hot pocket depends on the inflow conditions.

The interaction of a shock wave with a laminar boundary layer at a compression corner in high-enthalpy flows including real gas effects

The high-enthalpy, hypersonic flow over a compression corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg -1 to 19 MJ kg -1 , with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7 x 10 4 to 2.7 x 10 5 . The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the compression corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction.

OH PLIF Imaging of Supersonic Combustion using Cavity Injection

Planar laser-induced fluorescence of the hydroxyl radical is used to investigate a cavitybased fuel injection system for scramjet operation at a flight Mach number of 11.5. Hydrogen and ethylene fuels are compared over a range of fuel-lean global equivalence ratios. Both fuels show evidence of combustion at each fuel injection pressure. Hydrogen burns diffusively in the shear layer and further downstream, with penetration distance increasing linearly with injection pressure. The combustion in the cavity shear layer sheds regular vortices. Ethylene behaves similarly to hydrogen at the lower equivalence ratios but shows evidence of intense, localized combustion at the highest equivalence ratio. In all cases combustion occurs in the shear layer above the cavity rather than in the recirculating cavity flow.

Mole-Fraction-Sensitive Imaging of Hypermixing Shear Layers

A theoretical model that determines the optimum excitation frequency for obtaining a e uorescence signal with a strong dependence on fuel mole fraction is presented for supersonic fuel ‐air compressible mixing studies. The challenge associated with this is to maintain a high sensitivity to fuel mole fraction with minimal sensitivity to temperature and pressure in a e ow with large temperature variations and pressure gradients. The results of the modelareappliedtothemixingregion behindvariousscramjetfuelinjectorsinashocktunnelto measurefuelmole fraction. Hydrogen fuel at a Mach numberof 1.7 is injected into a mostly N 2 freestream at Mach 4.8. Experimental e uorescence images are presented in streamwise and spanwise planes.

Laminar boundary layer separation at a fin-body junction in a hypersonic flow

Abstract. Planar laser-induced fluorescence (PLIF) imaging was performed to visualize the fin bow shock, separation shock, viscous shear layer and recirculation region of the flowfield at the junction of a blunt fin and a flat plate. Making use of the temperature dependence of the PLIF technique, images were made sensitive to temperature to provide qualitative information on the flowfield. The PLIF technique was also used as the basis for a flow-tagging technique, making it possible to measure a velocity component and to demonstrate the reverse flow of the separated region. Flow visualisation of the plane of symmetry allowed determination of the point of boundary layer separation, the angle of the separation shock and the bow shock standoff distance. These parameters were compared with predictions made by computational fluid dynamic simulations of the flowfield. Good agreement between theory and experiment was achieved. Comparisons between theoretical and experimental velocity measurements showed good agreement.

Observation of an ablating surface in expansion tunnel flow

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.

A comparison of two hypermixing fuel injectors in a supersonic combustor

An experimental study has been undertaken to evaluate the performance of two hypermixing injectors designed for supersonic combustion ramjet (scramjet) applications. Supersonic mixing and combustion studied in a free-piston driven shock tunnel is examined using surface pressure measurements and shadowgraphy. Tests were conducted at two inlet Mach numbers: M=2.5 and M=3.7.

The boundary layer on a flat plate in hypervelocity flow

The high enthalpy and high Mach number effects on flat plate boundary layer flow are examined. In particular, the behaviour of the density profile under high enthalpy, real gas conditions is considered. Mach-Zehnder interferograms of the flow over a flat plate are analysed using a two-dimensional Fourier transform procedure in order to obtain the density field. The comparison between theory and experiment is only fair. It must be pointed out, however, that the real gas effects on the density profiles were found to be small. Estimation of the velocity boundary layer thickness as obtained from a measurement of the thermal boundary layer thickness is seen to compare well with a theoretical prediction.

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.

CFD calculations for intake-injection shock-induced-combustion scramjet flight experiments

Computations are reported of the combusting flows within a hydrogen fuelled model supersonic combustion ramjet (scramjet) operating in a shock tunnel flow. The work is in support of future launches in the HyShot free flight scramjet program. The scramjet simulates all features of a free flight scramjet including inlet compression ramps, combustion chamber and thrust surfaces. The device employs ramp injection and its ignition is thought to be shock induced. The purpose of the investigation is to enhance understanding of the detailed processes which operate within the device. The computational results are compared with pressure measurements for cases of no fuel injection and fuel injection into nitrogen and air coflows. Significant combustion pressure rise has been observed in the latter experiment. The comparison shows that the computations are physically sound and quite accurately predict many features of the flows. However, they fail to predict significant combustion in the fuel + air flow. Detailed examination is made of the computed solution from which it is inferred that the temperatures within the flow are too low for combustion. Therefore, unless the real flows are quite unlike the prediction, combustion would not be expected within them. No convincing explanation for this disparity between experiment and theory has yet been found and investigation is continuing.