HyShot scramjet testing in the HEG

Gardner, Anthony Donald (2007). HyShot scramjet testing in the HEG PhD Thesis, School of Engineering , University of Queensland.

       
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Author Gardner, Anthony Donald
Thesis Title HyShot scramjet testing in the HEG
School, Centre or Institute School of Engineering
Institution University of Queensland
Publication date 2007
Thesis type PhD Thesis
Supervisor Dr Allan Paull
Abstract/Summary A study of the suitabiiity of ground-based faciiities for simulating the iiow and combustion in scramjets has been carried out. A wind tunnei model was developed for the HEG from the HyShot flight test model, which recreated the internal geometry at a scale of 1 : 1. The model was split into two engines, which were instrumented with pressure transducers and thermocouples. A fuei injection system injected cold gaseous hydrogen into the flow through portholes in the wall of one combustion chamber. The model was tested at various angles of attack and equivalence ratios and testing with a nitrogen freestream was used to quantify the effect of fuel injection without combustion. in this study, ZFD and shock tunnel experiments have been performed, these then being compared with the flight test data from the HyShot I1 flight. Two new conditions for the HEG were designed analytically and refined using the one-dimensional Lagrangian code i i D and experimentally-determined efficiency coefficients, aiiowing the dennition of a new condition with a maximum of two tests in the HEG. A new Mach 7.8 contoured nozzle was calibrated and a new piston with a braking system to prevent reverse movement after diaphragm burst was designed and tested. -1 I ne engine pressures without combustion in the HEG agreed well with theory and CFD. Measurements with thermocouples in the HEG showed boundary layer transition in the combustion chamber between Reynolds numbers of 6 12000 and 782000. ivieasurements of the position of the initiation of combustion in the HEG showed that it was a function of mixing only and was independent of the combustion kinetics. Formation of a stable boundary layer separation on the cowl in the combustion chamber was observed for tests at higher equivalence ratio and angle of attack. The size of the separation was observed to increase with increasing equivalence ratio, angle of attack and wall temperature, and with decreasing freestream pressure. The expansion corner attached to the cowl reduced the size of the subsonic region, allowing separation to occur without choking the engine. The flight data was analysed at 27.5 km and 32.4km altitude, corresponding to the experimental conditions in the HEG. In the HyShot I1 flight, the combustion chamber pressures without combustion were 30% and 40% higher than expected at 32.4 km and 27.5 km altitude respectively. It has been shown that failure of the top and bottom tips of the combustion chamber was not responsible, but that failure of the side tips of the combustion chamber or a hysteresis in the iniet from higher altitudes couid not be ruled out as the source of the pressure discrepancy. The fuel-on flight data at 28 km altitude had a large separated region and the data at and 3.53" angle of attack agreed well with that taken in the HEG at Condition XI and 6.0" angle of attack. Experiments suggested that the fuel-on fiight data at 33 km altitude had only small separated regions, but the elevated pressure in the combustion chamber before combustion, prevented good agreement with the HEG data. -1 I nis work has demonstrated that the supersonic combustion and boundary iayer separation in a mixing-limited scramjet can be accurateiy simulated in a shock tunnel at an enthalpy of 3 MJIkg.

 
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Created: Fri, 21 Nov 2008, 15:22:38 EST