Measurement of the aerodynamic performance of a scramjet-powered vehicle in a shock tunnel

A force balance system for measuring lift, thrust and pitching moment has been used to measure the performance of fueled scramjet-powered vehicle in the T4 Shock Tunnel at The University of Queensland. Detailed measurements have been made of the effects of different fuel flow rates corresponding to equivalence ratios between 0.0 and 1.5. For proposed scramjet-powered vehicles, the fore-body of the vehicle acts as part of the inlet to the engine and the aft-body acts as the thrust surface for the engine. This type of engine-integrated design leads to a strong coupling between the performance of the engine and the lift and trim characteristics of the vehicle. The measurements show that the lift force increased by approximately 50% and centre-of-pressure changed by approximately 10% of the chord of the vehicle when the equivalence ratio varied from 0.0 to 1.0. The results demonstrate the importance of engine performance to the overall aerodynamic characteristics of engine-integrated scramjet vehicles and that such characteristics can be measured in a shock tunnel.

Investigation of test time at undertailored conditions in a shock tunnel.

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Design, modelling and analysis of a six component force balance for hypervelocity wind tunnel testing

A combination of modelling and analysis techniques was used to design a six component force balance. The balance was designed specifically for the measurement of impulsive aerodynamic forces and moments characteristic of hypervelocity shock tunnel testing using the stress wave force measurement technique. Aerodynamic modelling was used to estimate the magnitude and distribution of forces and finite element modelling to determine the mechanical response of proposed balance designs. Simulation of balance performance was based on aerodynamic loads and mechanical responses using convolution techniques. Deconvolution was then used to assess balance performance and to guide further design modifications leading to the final balance design. (C) 2001 Elsevier Science Ltd. All rights reserved.

The BCD transient recorder revisited: firmware for the ISA-to-USB bridging card and software for the graphical user interface

This report describes recent updates to the custom-built data-acquisition hardware operated by the Center for Hypersonics. In 2006, an ISA-to-USB bridging card was developed as part of Luke Hillyard's final-year thesis. This card allows the hardware to be connected to any recent personal computers via a (USB or RS232) serial port and it provides a number of simple text-based commands for control of the hardware. A graphical user interface program was also updated to help the experimenter manage the data acquisition functions. Sampled data is stored in text files that have been compressed with the gzip for mat. To simplify the later archiving or transport of the data, all files specific to a shot are stored in a single directory. This includes a text file for the run description, the signal configuration file and the individual sampled-data files, one for each signal that was recorded.

Fuel stagnation temperature effects on mixing with supersonic combustion flows

An experimental study of the effect of fuel stagnation temperature on mixing in a supersonic hydrogen-air flame is described, The combustor consisted of a constant-area rectangular duct with a centrally located fuel-injection strut that spanned the width. A high-enthalpy stream of air was supplied by a free-piston shock tunnel, and heated hydrogen fuel, supplied by a gun-tunnel, was injected into the freestream as a coflowing planar jet. The freestream total enthalpies were 5.6, 6.5, and 9 MJ/kg, and fuel stagnation temperatures were 300, 450, and 700 K, Raising the fuel stagnation temperature increased the fuel velocity to be near that of the airstream and resulted in a decrease in the mixing rate, Even as the fuel and air velocities became equal, significant mixing still occurred because of a large difference in density, Increasing the freestream enthalpy reduced the difference between the initial air temperature and the adiabatic flame temperature, which in turn reduced the heat addition, and subsequently, the amount of pressure rise in the duct.

Skin-friction measurements in a supersonic combustor with crossflow fuel injection

Shock-tunnel experiments have been performed to measure the effect on skin-friction drag in a supersonic combustor of flow disturbances induced by hydrogen fuel injection transverse to the airstream. Constant-area, circular cross section combustors of lengths varying up to 0.52 m were employed. The experiments were done at a stagnation enthalpy of 7.2 MJ . kg(-1) and a Mach number of 4.3, with a boundary layer that was turbulent downstream of the 0.14-m station in the combustors. Combustor skin-friction drag was measured by a method based on the stress wave force balance, the method being validated by agreement between fuel-off skin-friction drag measurements and predictions using existing skin-friction theories. When fuel was injected, it was found that the drag remained at fuel-off values. Thus, the streamwise vortices and other flow disturbances induced by the fuel injection, mixing, and combustion, which are expected to be present in a scramjet combustor, did not influence the skin-friction drag of the combustors.

Transducer for direct measurement of skin friction in hypervelocity impulse facilities

An acceleration compensated transducer was developed to enable the direct measurement of skin friction in hypervelocity impulse facilities. The gauge incorporated a measurement and acceleration element that employed direct shear of a piezoelectric ceramic. The design integrated techniques to maximize rise time and shear response while minimizing the affects of acceleration, pressure, heat transfer, and electrical interference. The arrangement resulted in a transducer natural frequency near 40 kHz. The transducer was calibrated for shear and acceleration in separate bench tests and was calibrated for pressure within an impulse facility. Uncertainty analyses identified only small experimental errors in the shear and acceleration calibration techniques. Although significant errors were revealed in the method of pressure calibration, total skin-friction measurement errors as low as +/-7-12% were established. The transducer was successfully utilized in a shock tunnel, and sample measurements are presented for flow conditions that simulate a flight Mach number near 8.

Upstream porthole injection in a 2-D scramjet model

Injection from portholes upstream of the combustion chamber was investigated as a method of delivering fuel into a scramjet. This method reduces the viscous drag on a model by allowing a reduction in the length of the combustion chamber. At experimental enthalpies of 3.0 MJ/kg in the T4 shock tunnel, there was no evidence of combustion in the intake, either by shadowgraph or pressure measurements. Combustion was observed in the combustion chamber. A theoretical extension of these results is made to a hot wall scenario.