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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.

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Studies of the ultraviolet and visible emission from an atmospheric pressure air plasma and its interaction with two carbon-based ablative materials were performed in an inductively coupled plasma torch. These experiments were conducted at a plate power of 40 kW, which produced an air plasma in local thermodynamic equilibrium with a maximum temperature of approximately 6200 K, corresponding to a specific enthalpy of 16.4 MJ∕kg. Three techniques were developed to measure the ablator surface temperature. Recession rates and product species profiles were measured for two different materials (ASTERM™ and carbon-bonded carbon fiber) at a plasma velocity of approximately 20 m∕s. Measured surface temperatures were in the range of 2100 to 2300 K, and corresponding hot-wall heat fluxes were approximately 1.4 MW∕m 2 . Spatially resolved profiles of the main species detected in the boundary layer were recorded, and they showed evidence of strong coupling between the ablated material and the freestream.

This paper reports on a study examining operation of a shockheated secondary driver across the entire operating envelope of a free-piston driven expansion tube. It is found that the secondary driver operating characteristics depend significantly on the specific characteristics of the flow condition. Key trends and characteristics are identified, and theoretical concepts are validated by numerical analysis and experiment.

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