A. Paull

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.

Experiments on supersonic combustion ramjet propulsion in a shock tunnel

Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg-1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg-1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pitot and static pressure measurements showed that the combustion was supersonic. The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.

Test flow disturbances in an expansion tube

The operation of an expansion tube is investigated with particular attention given to the test flow disturbances which have limited their utility in the past. Theoretical bounds for the duration of uniform test flow are first explored using one-dimensional ideal-gas relations, together with shock-tube boundary-layer entrainment effects. It is seen that test flow duration is limited either by the arrival of the downstream edge of the test-gas unsteady expansion or by the arrival of the upstream edge of this expansion after it has been reflected from the driver-test gas interface. These bounds are seen to be in good agreement with measurements made with large driver-gas expansion ratios. For small expansion ratios additional disturbances are observed in the test gas. Similar disturbances are also observed in the driver gas. It is postulated that these disturbances first appear in the driver gas and are transmitted into the test gas before the test gas is expanded. These disturbances remain with the test gas as it is expanded and subsequently produce unsteady conditions at the test section. Theoretical calculations for the range of frequencies which occur in the test gas before the expansion are obtained by modelling the disturbances as acoustic waves. It is shown that only the high-frequency components of the disturbances in the driver gas can penetrate the driver-test gas interface and this provides a mechanism for suppressing disturbances in the test gas. Additional analytical calculations for the shift in frequency produced as an acoustic wave traverses an unsteady expansion are also presented and it is shown that all frequencies of a given acoustic wave mode converge to one frequency. This focusing of frequencies is seen to occur in three different facilities.

The HyShot Scramjet Flight Experiment - Flight Data and CFD Calculations Compared

An analysis has been made using CFD of selected points on the HyShot scramjet flight experiment trajectory. Two-dimensional intake calculations have been done to assess the influence of angle of attack on the performance of the intake and cowl shock/boundary layer bleed duct, and have shown that at low angles of attack the configuration is well behaved, but at high angles of attack the cowl shock separates the boundary layer on the main intake ramp. This sends a separation shock into the combustion chamber. For selected altitudes during the flight for which the maximum angle of attack is a local minimum and for which there is zero yaw, intake and combustor calculations have been made. The calculations consistently predict that supersonic combustion has been obtained. At higher altituudes, reasonable agreement between the measured and predicted pressures is found, although the fuel-off pressures are underpredicted by approximately 15%. At lower altitude, further into the experiment, the flight data is well underpredicted. Further investigations of the flight data are required to assess this, including the possibility of leading edge ablation and/or intake distortion due to the high flight heat loads.

Hypervelocity Skin-Friction Reduction by Boundary-Layer Combustion of Hydrogen

Shock-tunnel measurements of Stanton number and skin-friction coefficient are reported for the injection of hydrogen through a 1.6-mm slot into a turbulent boundary layer in a l-m-long duct. The mainstream Mach number of 4.5, stagnation enthalpy of 7.8 MJ/kg, pressure of 50 kPa, and temperature of 1500 K provided a combination of flow variables that was sufficient to ensure boundary-layer combustion of the hydrogen. The experiments were also simulated by a numerical mode). The experiments and the numerical model indicated that the Stanton number was only slightly affected by boundary-layer combustion. However, the numerical simulation indicated that injection with combustion caused a reduction of approximately 50% in the skin friction coefficient, whereas the experiments yielded an even greater effect, with the reduction in skin-friction coefficient reaching 70-80% of the values of skin friction with no injection. Numerical simulation of a constant pressure flow indicated that boundary-layer combustion caused the skin-friction reduction to persist for at least 5 m downstream.

Comparison of Supersonic Combustion Between Impulse and Vitiation-Heated Facilities

A comparison has been made between supersonic combustion in two commonly used, but fundamentally different, facilities for scramjet research, a vitiation-heated blowdown tunnel and a free-piston shock tunnel. By passing the shock-tunnel freestream flow through a normal shock and then expanding it to Mach 2.5, combustor inlet conditions and geometries were nominally replicated between the two facilities. A constant-area rectangular duct and a diverging duct, both employing central-strut hydrogen injection, were used. Boundary-layer separation and choking in the constant-area duct limited supersonic combustion comparisons up to a fuel equivalence ratio of the order of 0.3. The experimental results also show that the onset of boundary-layer separation occurs at the same combustor pressure loads and that it behaves similarly in the different facilities. With the diverging duct, comparisons were made up to an equivalence ratio of 1.05. Agreement between the results obtained in the two facilities is within experimental error when the different freestream and boundary layers are accounted for.

Numerical Investigation of a Spiked Blunt Nose Cone at Hypersonic Speeds

When flying at hypersonic speeds, it is a fundamental requirement to reduce the high drag resulting from a blunt nose cone in the ascent stage to increase the payload weight on the one hand and decrease the amount of energy needed to overcome the Earths gravity on the other. However, an aerospike can be attached on the front of the nose cone to obtain a high drag and heat load reduction. Different Mach numbers at different altitudes have been chosen to investigate the effect of the aerospike on the nose cones surrounding flowfield. The drag and the heat load reduction is numerically evaluated at Mach numbers of 5.0, 7.0, and 10.0. Different lengths of the aerospike are investigated between 1 and 4 times the diameter of the dome of the nose cone. Additional modifications to the tip of the spike to obtain different bow shocks are examined, including a sharp front, a blunt spike, and an aerodome mounted on the tip of the spike. To solve the very complicated flowfield, the flow solver CFD-FASTRAN is used.

Radical Farming in Scramjets

This paper outlines the theory of radical farming in scramjets and describes the experimental scramjet model that was designed to investigate it. Experiments were conducted at two conditions; a 3MJ/kg condition corresponding to Mach 7.9 flight at an altitude of 24km and a 4MJ/kg condition corresponding to Mach 9.1 flight at an altitude of 32km. The results are presented as pressure distributions on the flowpath wall and specific impulse estimates.

HIFiRE-1 Preliminary Aerothermodynamic Measurements

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory (AFRL) and Australian Defence Science and Technology Organisation (DSTO). HIFiRE flight one flew in March 2010. Principle goals of this flight were to measure hypersonic boundary-layer transition and shock boundary layer interactions in flight. The flight successfully gathered pressure, temperature and heat transfer measurements during ascent and reentry. HIFiRE1 has provided transition measurements suitable for calibrating N-factor prediction methods for flight, and has produced some insight into the structure of the transition front on a cone at angle of attack. Pressure and heat transfer measurements in the shock-boundary-layer interaction were obtained. Preliminary analysis of the shock boundary layer interaction shows intermittent pressure fluctuations qualitatively similar to those measured in wind tunnel experiments. A large amount of data was obtained on the flight, and significant data reduction efforts continue.

Ground Testing of the HyShot Supersonic Combustion Flight Experiment in HEG and Comparison with Flight Data

The first phase of the HyShot supersonic combustion ramjet (scramjet) flight experiment program of The University of Queensland in Australia was designed to provide benchmark data on supersonic combustion for a flight Mach number of approximately M=8. The second flight of the HyShot program, performed on July 30th 2002, was successful and supersonic combustion was observed along the specified trajectory range. The operating range of the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Centre (DLR) was recently extended. The facility now has the capability of testing a complete scramjet engine with internal combustion and external aerodynamics at M=7.8 flight con-ditions in altitudes of about 30 km. A post-flight analysis of the HyShot flight experiment was performed using an operational scramjet wind tunnel model with a geometry which is identical to that of the flight configuration.