Three-component force measurements on a large scramjet in a shock tunnel

A stress-wave force balance for measurement of thrust, lift, and pitching moment on a large scramjet model (40 kg in mass, 1.165 in in length) in a reflected shock tunnel has been designed, calibrated, and tested. Transient finite element analysis was used to model the performance of the balance. This modeling indicates that good decoupling of signals and low sensitivity of the balance to the distribution of. the load can be achieved with a three-bar balance. The balance was constructed and calibrated by applying a series of point loads to the model. A good comparison between finite element analysis and experimental results was obtained with finite element analysis aiding in the interpretation of some experimental results. Force measurements were made in a shock tunnel both with and without fuel injection, and measurements were compared with predictions using simple models of the scramjet and combustion. Results indicate that the balance is capable of resolving lift, thrust, and pitching moments with and without combustion. However vibrations associated with tunnel operation interfered with the signals indicating the importance of vibration isolation for accurate measurements.

Scramjet lift, thrust and pitching-moment characteristics measured in a shock tunnel

Lift, pitching moment, and thrust/drag on a supersonic combustion ramjet were measured in the T4 free-piston shock tunnel using a three-component stress-wave force balance. The scramjet model was 0.567 m long and weighed approximately 6 kg. Combustion occurred at a nozzle-supply enthalpy of 3.3 MJ/kg and nozzle-supply pressure of 32 MPa at Mach 6.6 for equivalence ratios up to 1.4. The force coefficients varied approximately linearly with equivalence ratio. The location of the center of pressure changed by 10% of the chord of the model over the range of equivalence ratios tested. Lift and pitching-moment coefficients remained constant when the nozzle-supply enthalpy was increased to 4.9 MJ/kg at an equivalence ratio of 0.8, but the thrust coefficient decreased rapidly. When the nozzle-supply pressure was reduced at a nozzle-supply enthalpy of 3.3 MJ/kg and an equivalence ratio of 0.8, the combustion-generated increment of lift and thrust was maintained at 26 MPa, but disappeared at 16 MPa. Measured lift and thrust forces agreed well with calculations made using a simplified force prediction model, but the measured pitching moment substantially exceeded predictions. Choking occurred at nozzle-supply enthalpies of less than 3.0 MJ/kg with an equivalence ratio of 0.8. The tests failed to yield a positive thrust because of the skin-friction drag that accounted for up to 50% of the fuel-off drag.

Simulation of a complete reflected shock tunnel showing a vortex mechanism for flow contamination

Simulations of a complete reflected shock tunnel facility have been performed with the aim of providing a better understanding of the flow through these facilities. In particular, the analysis is focused on the premature contamination of the test flow with the driver gas. The axisymmetric simulations model the full geometry of the shock tunnel and incorporate an iris-based model of the primary diaphragm rupture mechanics, an ideal secondary diaphragm and account for turbulence in the shock tube boundary layer with the Baldwin-Lomax eddy viscosity model. Two operating conditions were examined: one resulting in an over-tailored mode of operation and the other resulting in approximately tailored operation. The accuracy of the simulations is assessed through comparison with experimental measurements of static pressure, pitot pressure and stagnation temperature. It is shown that the widely-accepted driver gas contamination mechanism in which driver gas 'jets' along the walls through action of the bifurcated foot of the reflected shock, does not directly transport the driver gas to the nozzle at these conditions. Instead, driver gas laden vortices are generated by the bifurcated reflected shock. These vortices prevent jetting of the driver gas along the walls and convect driver gas away from the shock tube wall and downstream into the nozzle. Additional vorticity generated by the interaction of the reflected shock and the contact surface enhances the process in the over-tailored case. However, the basic mechanism appears to operate in a similar way for both the over-tailored and the approximately tailored conditions.

End-to-end modelling of the HEG shock tunnel

We demonstrate an end-to-end computational model of the HEG shock tunnel as a way to extract more precise test flow conditions and as a way of getting predictions of new operating conditions. For a selection of established operating conditions, the L1d program was used to simulate the one-dimensional gas-dynamic processes within the whole of the facility. The program reproduces the compression tube performance reliably and, with the inclusion of a loss factor near the upstream-end of the compression tube, it provides a good estimate of the equilibrium pressure in the shock-reflection region over the set of six standard operating conditions for HEG.

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.

Effects of hydrogen-air non-equilibrium chemistry on the performance of a model scramjet thrust nozzle

Two aspects of hydrogen-air non-equilibrium chemistry related to scramjets are nozzle freezing and a process called 'kinetic afterburning' which involves continuation of combustion after expansion in the nozzle. These effects were investigated numerically and experimentally with a model scramjet combustion chamber and thrust nozzle combination. The overall model length was 0.5m, while precombustion Mach numbers of 3.1 +/- 0.3 and precombustion temperatures ranging from 740K to 1,400K were involved. Nozzle freezing was investigated at precombustion pressures of 190kPa and higher, and it was found that the nozzle thrusts were within 6% of values obtained from finite rate numerical calculations, which were within 7% of equilibrium calculations. When precombustion pressures of 70kPa or less were used, kinetic afterburning was found to be partly responsible for thrust production, in both the numerical calculations and the experiments. Kinetic afterburning offers a means of extending the operating Mach number range of a fixed geometry scramjet.

Dynamic force balances for short-duration hypersonic testing facilities

Two force balance techniques for use in hypersonic impulse facilities are compared by measuring the drag force on a 30° semi-apex-angle blunt cone model in a hypersonic shock tunnel at a free stream Mach number of 5.75. An accelerometer-based balance and a stress-wave force balance were tested simultaneously on the same model to measure the drag force. It was found that drag force measurements could be made using both techniques in a flow with a 450-μ s test period. The measured drag forces compared well with the theoretical values estimated using Newtonian theory.

Flight data analysis of the HyShot 2 scramjet flight experiment

The development of scramjet propulsion for alternative launch and payload delivery capabilities has been composed largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, a ballistic reentry vehicle experiment called HyShot was devised to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.34 and the other unfueled. Of the two flights conducted, HyShot 1 failed to reach the desired altitude due to booster failure, whereas HyShot 2 successfully accomplished both the desired trajectory and satisfactory scramjet operation. Postflight data analysis of HyShot 2 confirmed the presence of supersonic combustion during the approximately 3 s test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some preflight shock tunnel tests was observed.