We investigate a new type of flow-tagging velocimetry technique for hypersonic flows.

Two-dimensional rotational and vibrational temperature measurements were made at the nozzle exit of a free-piston shock tunnel using Planar Laser-Induced Fluorescence. The Mach 7 flow consisted predominantly of nitrogen with a trace quantity of nitric oxide. Nitric oxide was employed as the probe species and was excited at 225 run. Non-uniformities in the distribution of nitric oxide in the test gas were observed, and concluded to be due to contamination of the test gas by driver gas or cold test gas. The nozzle-exit rotational temperature was measured and is in reasonable agreement with computational modelling. Non-linearities in the detection system were responsible for systematic errors in the measurements. The vibrational temperature was measured to be constant with distance from the nozzle exit, indicating it had frozen during the nozzle expansion.

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A sensor based on tunable diode laser absorption spectroscopy was constructed for time-resolved temperature and water vapor concentration measurements in a scramjet combustor. The sensor probed two absorption lines near 1390 nm with two time-multiplexed lasers used to measure temperature and water vapor concentration at up to 20 kHz. A demonstration experiment was performed in the supersonic, expanding exhaust region of the combustor, showing the measurement to be repeatable, able to resolve temporal trends during tunnel operation, and sensitive to changes in combustor operating conditions.

A theoretical model that determines the optimum excitation frequency for obtaining a fluorescence 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 flow with large temperature variations and pressure gradients. The results of the model are applied to the mixing region behind various scramjet fuel injectors in a shock tunnel to measure fuel mole-fraction. Hydrogen fuel at a Mach number of 1.7 is injected into a mostly N2 free stream at Mach 4.8. Experimental fluorescence images are presented in streamwise and spanwise planes. Nomenclature cp Specific heat capacity at constant pressure, J.kg −1 cp,∞ Specific heat capacity of pure freestream, J.kg −1 c p,f uel Specific heat capacity of pure fuel stream, J.kg −1 f J Boltzmann fraction of the absorbing state g Spectral overlap integral, 1/cm −1 ga Absorption line shape, 1/cm −1 g l Spectral profile of laser, 1/cm −1 k Boltzmann's constant, J.kg −1 mNO Molecular mass of NO, kg mp Molecular mass of perturbing species, kg