The flow field and the energy transport near thermoacoustic couples are simulated using a 2D full Navier-Stokes solver. The thermoacoustic couple plate is maintained at a constant temperature; plate lengths, which are ''short'' and ''long'' compared with the particle displacement lengths of the acoustic standing waves, are tested. Also investigated are the effects of plate spacing and the amplitude of the standing wave. Results are examined in the form of energy vectors, particle paths, and overall entropy generation rates. These show that a net heat-pumping effect appears only near the edges of thermoacoustic couple plates, within about a particle displacement distance from the ends. A heat-pumping effect can be seen even on the shortest plates tested when the plate spacing exceeds the thermal penetration depth. It is observed that energy dissipation near the plate increases quadratically as the plate spacing is reduced. The results also indicate that there may be a larger scale vortical motion outside the plates which disappears as the plate spacing is reduced.

This article reports on the use of a shock tunnel to study the operation of scramjet powered configurations at sub-orbital velocities above 2 km/s. Thrust, as given by a net thrust equation, is used as a figure of merit throughout the study. After a short description of the shock tunnel used and its operating characteristics, experiments on the combustion release of heat in a constant area duct with hydrogen fuel are reviewed. The interaction between heat release in the combustion wake and the walls of the duct produced pressure distributions which followed a binary scaling law, and indicated that the theoretically expected heat release could be realized in practice, albeit with high pressure or long combustion ducts. This heat release, combined with attainable thrust nozzle characteristics and a modest level of configuration drag, indicated that positive thrust levels could be obtained well into the sub-orbital range of velocities. Development of a stress wave force balance for use in shock tunnels allowed the net thrust generated to be measured for integrated scramjet configurations and, although the combination of model size and shock tunnel operating pressure prevented complete combustion of hydrogen, the cruise condition of zero net thrust was achieved at 2.5 km/s with one configuration, while net thrust was produced with another configuration using an ignition promoter in hydrogen fuel. Nevertheless, the combination of boundary layer separation induced inlet choking and limited operating pressure levels prevented realization of the thrust potential of the fuel. This problem may be alleviated by recent increases in the shock tunnel operating pressures, and by promising research involving inlet injection of the fuel. Research on the drag component of the net thrust equation resulted from the development of a fast response skin friction gauge. It was found that existing theories of turbulent boundary skin friction predicted the skin friction when combustion of hydrogen occurred outside the boundary layer, but combustion within the boundary layer dramatically reduced the skin friction. Finally, for the first time in the world, supersonic combustion was produced in a free flight experiment. This experiment validated shock tunnel results at stagnation enthalpies near 3 MJ/kg.

Experiments to measure losses of a linear cascade of transonic turbine blades are reported. Detailed measurements of the boundary layer at the rear of the suction surface of a blade and examination of wake traverse data enable the individual components of boundary layer, shock and mixing loss to be determined. Results indicate that each component contributes significantly to the overall loss in different Mach number regimes. Traverses in the near wake of the blade indicate the way in which the wake develops and facilitate examination of the development of the mixing loss.

Abstract. This paper analyzes different techniques for the calibration of force balances for use in shortduration impulse hypersonic facilities such as shock tunnels. The background to how deconvolution can be used to infer aerodynamic forces on models in impulse hypersonic wind tunnels is presented along with the theory behind the different calibration techniques. Four calibration techniques are applied to a singlecomponent stress-wave force balance. Experiments in the T4 shock tunnel using the balance demonstrate the suitability of the different calibrations for force measurements in an impulse facility. Cross checks between the calibration techniques are used to check their ranges of validity. It is shown that the impulse response derived from tests in which the model and force balance are suspended from a fine wire and the wire cut agree well with impulse responses derived from calibrations made using an impact hammer. The suitability of the balance for measuring dynamic forces is demonstrated by showing that the drag force on a model follows the history of Pitot pressure in the test section in the tunnel shots.

Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg−1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg−1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pilot and static pressure measurements showed that the combustion was supersonic.The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.