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.

The measurement of Pitot pressure in high enthalpy expansion tubes

Expansion tubes operating at total flow enthalpies of 100 MJ kg(-1) or more have characteristical test times of 30-50 mus. Under these conditions, the response time of the Pitot pressure measuring device is critical when performing flow calibration studies. The conventional technique of using a commercial pressure transducer protected by shielding has not always proven to be effective, due to the relatively large (and variable) response time caused by the shielding. A device called the stress wave bar gauge has been designed and calibrated and shown to be an effective way to measure the Pitot pressure with a response time of only 2-3 mus.

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.

Heat transfer and flow behind a step in high enthalpy superorbital flow

Heat transfer levels have been investigated behind a rearward-facing step in a superorbital expansion tube. The heat transfer was measured along a flat plate and behind 2 and 3mm steps with the same length to step height ratio. Results were obtained with air as the test gas at speeds of 6.76kms(-1) and 9-60kms(-1) corresponding to stagnation enthalpies of 26MJ/kg and 48MJ/kg respectively. A laminar boundary layer was established on the flat plate and measured heat transfer levels were consistent with classical empirical correlations. In the case of flow behind a step, the measurements showed a gradual rise in heat transfer from the rear of the step to a plateau several step heights downstream for both flow conditions. Reattachment distance was estimated to be approximately 1.6 step heights downstream of the 2mm step at the low enthalpy condition through the use of flow visualisation.

Skin-friction measurements in high-enthalpy hypersonic boundary layers

Skin-friction measurements are reported for high-enthalpy and high-Mach-number laminar, transitional and turbulent boundary layers. The measurements were performed in a free-piston shock tunnel with air-flow Mach number, stagnation enthalpy and Reynolds numbers in the ranges of 4.4-6.7, 3-13 MJ kg(-1) and 0.16 x 10(6)-21 x 10(6), respectively. Wall temperatures were near 300 K and this resulted in ratios of wall enthalpy to flow-stagnation enthalpy in the range of 0.1-0.02. The experiments were performed using rectangular ducts. The measurements were accomplished using a new skin-friction gauge that was developed for impulse facility testing. The gauge was an acceleration compensated piezoelectric transducer and had a lowest natural frequency near 40 kHz. Turbulent skin-friction levels were measured to within a typical uncertainty of +/-7%. The systematic uncertainty in measured skin-friction coefficient was high for the tested laminar conditions; however, to within experimental uncertainty, the skin-friction and heat-transfer measurements were in agreement with the laminar theory of van Driest (1952). For predicting turbulent skin-friction coefficient, it was established that, for the range of Mach numbers and Reynolds numbers of the experiments, with cold walls and boundary layers approaching the turbulent equilibrium state, the Spalding & Chi (1964) method was the most suitable of the theories tested. It was also established that if the heat transfer rate to the wall is to be predicted, then the Spalding & Chi (1964) method should be used in conjunction with a Reynolds analogy factor near unity. If more accurate results are required, then an experimentally observed relationship between the Reynolds analogy factor and the skin-friction coefficient may be applied.

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.

Numerical modelling of radiating superorbital flows

Experimental aerodynamic studies of the flows around new aerocapture spacecraft configurations are presently being done in the superorbital expansion tubes at The University of Queensland. Short duration flows at speeds of 10--13 km/s are produced in the expansion tube facility and are then applied to the model spacecraft. Although high-temperature effects, such as molecular dissociation, have long been a part of the computational modelling of the expansion tube flows for speeds below 10 km/s, radiation may now be a significant mechanism of energy transfer within the shock layer on the model. This paper will study the coupling of radiation energy transport for an optically thin gas to the flow dynamics in order to obtain accurate predictions of thermal loads on the spacecraft. The results show that the effect of radiation on the flowfields of subscale models for expansion tube experiments can be assessed by measurements of total heat transfer and radiative heat transfer.

Axisymmetric shock wave interaction with a cone: a benchmark test

Results of the benchmark test are presented of comparing numerical schemes solving shock wave of M-s = 2.38 in nitrogen and argon interacting with a 43 degrees semi-apex angle cone and corresponding experiments. The benchmark test was announced in Shock Waves Vol. 12, No. 4, in which we tried to clarify the effects of viscosity and heat conductivity on shock reflection in conical flows. This paper summarizes results of ten numerical and two experimental applications. State of the art in studies regarding the shock/cone interaction is clarified.

Coupled thermal, structural and vibrational analysis of a hypersonic engine for flight test

This paper describes a relatively simple and quick method for implementing aerodynamic heating models into a finite element code for non-linear transient thermal-structural and thermal-structural-vibrational analyses of a Mach 10 generic HyShot scramjet engine. The thermal-structural-vibrational response of the engine was studied for the descent trajectory from 60 to 26 km. Aerodynamic heating fluxes, as a function of spatial position and time for varying trajectory points, were implemented in the transient heat analysis. Additionally, the combined effect of varying dynamic pressure and thermal loads with altitude was considered. This aero-thermal-structural analysis capability was used to assess the temperature distribution, engine geometry distortion and yielding of the structural material due to aerodynamic heating during the descent trajectory, and for optimising the wall thickness, nose radius of leading edge, etc. of the engine intake. A structural vibration analysis was also performed following the aero-thermal-structural analysis to determine the changes in natural frequencies of the structural vibration modes that occur at the various temperatures associated with the descent trajectory. This analysis provides a unique and relatively simple design strategy for predicting and mitigating the thermal-structural-vibrational response of hypersonic engines. (C) 2006 Elsevier SAS. All rights reserved.

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.

Modern developments in hypersonic wind tunnels

The development of new methods of producing hypersonic wind-tunnel flows at increasing velocities during the last few decades is reviewed with attention to airbreathing propulsion, hypervelocity aerodynamics and superorbital aerodynamics. The role of chemical reactions in these flows leads to use of a binary scaling simulation parameter, which can be related to the Reynolds number, and which demands that smaller wind tunnels require higher reservoir pressure levels for simulation of flight phenomena. The use of combustion heated vitiated wind tunnels for propulsive research is discussed, as well as the use of reflected shock tunnels for the same purpose. A flight experiment validating shock-tunnel results is described, and relevant developments in shock tunnel instrumentation are outlined. The use of shock tunnels for hypervelocity testing is reviewed, noting the role of driver gas contamination in determining test time, and presenting examples of air dissociation effects on model flows. Extending the hypervelocity testing range into the superorbital regime with useful test times is seen to be possible by use of expansion tube/tunnels with a free piston driver.

Performance of a scramjet combustor with combined normal and tangential fuel injection

The performance of a scramjet combustor with combined normal and tangential injection was experimentally investigated. Experiments were performed on a 500-mm cylindrical scramjet combustor at a freestream Mach number of 4.5, a nozzle supply pressure of 35.8 MPa, and a nozzle supply enthalpy of 5.8 MJ/kg. Hydrogen fuel was injected normally through portholes to promote combustion and tangentially through a slot to reduce viscous drag. A series of fuel injectors were used to vary the proportion of tangential to normal fuel between 45 and 100%. Reductions in the viscous drag of up to 25% were observed with the greatest reductions occurring at the lowest total equivalence ratio tested for each injector. However, the average pressure produced by combustion with combined normal and tangential injection was approximately 50% less than that produced by normal injection alone. An analysis of the change in specific impulse of the scramjet combustor indicated that the best overall performance was produced by 100% normal injection.