The process of shock-induced ignition has been investigated both computa- tionally and experimentally, with particular emphasis on the concept of radical farming. The first component of the investigation contained Computational Fluid Dynamic (CFD) calculations of an ignition delay study, a 2D pre-mixed flow over flat plate at a constant angle to the freestream, and through a generic 2D scramjet model. The focal point of the investigation however examined the complex 3D flow through a generic scramjet model. Five experimental test conditions were ex- amined over flow enthalpies from 3.4 MJ/kg to 6.4 MJ/kg. All test conditions simulated flight at 21000 metres ([symbol=almost equal to] 70000 ft), while the equivalent flight Mach number varied from approximately 8.5 at the lowest enthalpy, to approximately Mach 12 at the highest enthalpy condition. The presence of H2 fuel injected in the intake caused a separated region to form on the lower surface of the model at the entrance to the combustor. A fraction of the total mass of fuel was entrained in this separated region, providing long residence times, hence increased time for the chemical reactions that lead to ignition to occur. In addition, extremely high temperatures were found to exist between each fuel jet. Both fuel and air are present in these regions, therefore the chance of ignition in these regions is high. Streamlines passing through the recirculation zone ignited within this zone, while streamlines passing between the fuel jets ignited soon after entry into the combustor. The first instance of a pressure rise from combustion was observed on the centreline of the model where the reflected bow shock around the fuel jets crossed the centreline of the combus- tor. Upstream of this location the static pressure of the flow was too low for the chemical reactions that release heat to occur. The comparison between the experimental and computational results was lim- ited due to inaccuracies in modelling the thermal state of the gas in the CFD calculations. The gas was modelled as being in a state of thermal equilibrium at all times, which incorrectly models the freestream flow from the nozzle of the shock tunnel, and also the flow downstream of oblique shock wave within the scramjet model. As a result combustion occurs sooner in the CFD calculations than in the experimental result.

Hypersonic vehicles, based on scramjet engines, have the potential to deliver inexpensive access to space when compared with rocket propulsion. The technology, however, is in its infancy and there is still much to be learned from fundamental studies.¶ Flows that represent the conditions inside a scramjet engine can be generated in ground tests using a free-piston shock tunnel and a combustor model. These facilities provide a convenient location for fundamental studies and principles learned during ground tests can be applied to the design of a full-scale vehicle.¶ A wide range of diagnostics have been used for studying scramjet flows, including surface measurements and optical visualisation techniques.¶ The aim of this work is to test the effectiveness of tunable diode laser absorption spectroscopy (TDLAS) as a scramjet diagnostic.¶ TDLAS utilises the spectrally narrow emission from a diode laser to probe individual absorption lines of a target species. By varying the diode laser injection current, the laser emission wavelength can be scanned to rapidly obtain a profile of the spectral line. TDLAS has been used previously for gas-dynamic sensing applications and, in the configuration used in this work, is sensitive to temperature and water vapour concentration.¶ The design of the sensor was guided by previous work. It incorporated aspects of designs that were considered to be well suited to the present application. Aspects of the design which were guided by the literature included the laser emission wavelength, the use of fibre optics and the detector used. The laser emission wavelength was near 1390 nm to coincide with relatively strong water vapour transitions. This wavelength allowed the use of telecommunications optical fibre and components for light delivery. Detection used a dual-beam, noise cancelling detector.¶ The sensor was validated before deployment in a low-pressure test cell and a hydrogen–air flame. Temperature and water concentration measurements were verified to within 5% up to 1550 K. Verification accuracy was limited by non-uniformity along the beam path during flame measurements.¶ Measurements were made in a scramjet combustor operating in a flow generated by the T3 shock tunnel at the Australian National University. Within the scramjet combustor, hydrogen was injected into a flame-holding cavity and the sensor was operated downstream in the expanded, supersonic, post-combustion flow. The sensor was operated at a maximum repetition rate of 20 kHz and could resolve variation in temperature and water concentration over the 3ms running time of the facility.¶ Results were repeatable and the measurement uncertainty was smaller than the turbulent fluctuations in the flow. The scramjet was operated at two fuel-lean equivalence ratios and the sensor was able to show differences between the two operating conditions. In addition, vertical traversal of the sensor revealed variation in flow conditions across the scramjet duct.¶ The effectiveness of the diagnostic was tested by comparing results with those from other measurement techniques, in particular pressure and OH fluorescence measurements, as well as comparison with computational simulation.¶ Combustion was noted at both of the tested operating conditions in data from all three measurement techniques.¶ Computation simulation of the scramjet flow significantly under-predicted the water vapour concentration. The discrepancy between experiments and simulation was not apparent in either the pressure measurements or the OH fluorescence, but was clear in the diode laser results.¶ The diode laser sensor, therefore, was able to produce quantitative results which were useful for comparison with a CFD model of the scramjet and were complimentary to information provided by other diagnostics.

This study investigates the stress wave force balance technique for the measurement of forces on a fuelled hypersonic flight vehicle in an impulse-type test facility. A three component force balance for the measurement of lift, thrust and pitching moment on a supersonic combustion ramjet engine was designed, built, calibrated and tested. The force balance was designed using finite element analysis and consisted of four stress bars instrumented for the measurement of strain. Relative errors of less than 2% were obtained for the recovered simulated calibration loads, while errors of less than 3% were obtained for lift and thrust components for simulated fuel-on and fuel-off force loading distributions. Tests in a calibration rig showed that the balance was capable of recovering the magnitude of point loads to within 3% and their lines of action to within 1% of the chord of the model. Additional errors result when testing in a wind tunnel. The uncertainties for the experiments with fuel injection are estimated at 9%, 7% and 9% for the coefficients of lift, thrust and pitching moment. The scramjet vehicle was 0.566m long and weighed approximately 6kg. It consisted of an inlet, combustion chamber and thrust surface. Fuel could be injected through a series of injectors located on the scramjet inlet. The scramjet model was set at zero angle of attack. Experiments were performed in the T4 Free Piston Shock Tunnel at a total enthalpy of 3.3MJ/kg, a nozzle supply pressure of 32MPa and a Mach number of 6.6, with equivalence ratios up to 1.4. Fuel-off force coefficients were measured to within 2% of theoretical values based on predictions using CFD and hypersonic theory. The fuel-off centre-of-pressure was measured to within 4% of the predicted value. The force coefficients varied linearly with equivalence ratio. Good comparison of the measured lift and thrust forces with theoretical values was obtained with increasing flow rates of fuel. The lift-to-drag ratio increased from 3.0 at the fuel-off condition to 17.2 at an equivalence ratio of 1.0. Poor agreement between the measured pitching moment and theoretical values was obtained due to difficulties in predicting the pressure distribution with heat addition on the latter parts of the thrust surface. A shift in the centre-of-pressure of approximately 10% of model chord was measured as the equivalence ratio varied from 0.0 to 1.0. For the design tested, the thrust produced was not enough to overcome drag on the vehicle, even at the highest equivalence ratio tested. Tests at higher stagnation enthalpies (up to 4.9MJ/kg) showed the lift and pitching moment coefficients remained constant with an equivalence ratio of 0.8 but the thrust coefficient decreased exponentially with increasing stagnation enthalpies. Good agreement of experimental values of lift and thrust force with predicted values was obtained for equivalence ratios of 0.0 and 0.8. Choking occurred at stagnation enthalpies of less than 3.0MJ/kg and a nozzle supply pressure of 32MPa with fuel injection at an equivalence ratio of approximately 0.8, resulting in a drag force of approximately 2.5 times the fuel-off drag force. Tests at a nozzle supply enthalpy of 3.3MJ/kg and nozzle supply pressures of 32, 26 and 16MPa were performed at equivalence ratios of 0.0 and 0.8. The fuel-off lift coefficient remained constant but the thrust coefficient increased. This is attributed to a reduction in skin friction associated with longer lengths of laminar boundary layers as the Reynolds number was decreased. The measured fuel-off lift and thrust coefficients agreed with the predicted values to within the known test flow and force prediction uncertainties. Combustion did not occur at a nozzle supply pressure of 16MPa. This work has demonstrated that overall scramjet vehicle performance measurements (such as lift-to-drag ratio and shifts in centre-of-pressure) can be made in a free piston shock tunnel.

This study investigates the stress wave force balance technique for the measurement of forces on a fuelled hypersonic flight vehicle in an impulse-type test facility. A three component force balance for the measurement of lift, thrust and pitching moment on a supersonic combustion ramjet engine was designed, built, calibrated and tested. The force balance was designed using finite element analysis and consisted of four stress bars instrumented for the measurement of strain. Relative errors of less than 2% were obtained for the recovered simulated calibration loads, while errors of less than 3% were obtained for lift and thrust components for simulated fuel-on and fuel-off force loading distributions. Tests in a calibration rig showed that the balance was capable of recovering the magnitude of point loads to within 3% and their lines of action to within 1% of the chord of the model. Additional errors result when testing in a wind tunnel. The uncertainties for the experiments with fuel injection are estimated at 9%, 7% and 9% for the coefficients of lift, thrust and pitching moment. The scramjet vehicle was 0.566m long and weighed approximately 6kg. It consisted of an inlet, combustion chamber and thrust surface. Fuel could be injected through a series of injectors located on the scramjet inlet. The scramjet model was set at zero angle of attack. Experiments were performed in the T4 Free Piston Shock Tunnel at a total enthalpy of 3.3MJ/kg, a nozzle supply pressure of 32MPa and a Mach number of 6.6, with equivalence ratios up to 1.4. Fuel-off force coefficients were measured to within 2% of theoretical values based on predictions using CFD and hypersonic theory. The fuel-off centre-of-pressure was measured to within 4% of the predicted value. The force coefficients varied linearly with equivalence ratio. Good comparison of the measured lift and thrust forces with theoretical values was obtained with increasing flow rates of fuel. The lift-to-drag ratio increased from 3.0 at the fuel-off condition to 17.2 at an equivalence ratio of 1.0. Poor agreement between the measured pitching moment and theoretical values was obtained due to difficulties in predicting the pressure distribution with heat addition on the latter parts of the thrust surface. A shift in the centre-of-pressure of approximately 10% of model chord was measured as the equivalence ratio varied from 0.0 to 1.0. For the design tested, the thrust produced was not enough to overcome drag on the vehicle, even at the highest equivalence ratio tested. Tests at higher stagnation enthalpies (up to 4.9MJ/kg) showed the lift and pitching moment coefficients remained constant with an equivalence ratio of 0.8 but the thrust coefficient decreased exponentially with increasing stagnation enthalpies. Good agreement of experimental values of lift and thrust force with predicted values was obtained for equivalence ratios of 0.0 and 0.8. Choking occurred at stagnation enthalpies of less than 3.0MJ/kg and a nozzle supply pressure of 32MPa with fuel injection at an equivalence ratio of approximately 0.8, resulting in a drag force of approximately 2.5 times the fuel-off drag force. Tests at a nozzle supply enthalpy of 3.3MJ/kg and nozzle supply pressures of 32, 26 and 16MPa were performed at equivalence ratios of 0.0 and 0.8. The fuel-off lift coefficient remained constant but the thrust coefficient increased. This is attributed to a reduction in skin friction associated with longer lengths of laminar boundary layers as the Reynolds number was decreased. The measured fuel-off lift and thrust coefficients agreed with the predicted values to within the known test flow and force prediction uncertainties. Combustion did not occur at a nozzle supply pressure of 16MPa. This work has demonstrated that overall scramjet vehicle performance measurements (such as lift-to-drag ratio and shifts in centre-of-pressure) can be made in a free piston shock tunnel.

This study investigates the stress wave force balance technique for the measurement of forces on a fuelled hypersonic flight vehicle in an impulse-type test facility. A three component force balance for the measurement of lift, thrust and pitching moment on a supersonic combustion ramjet engine was designed, built, calibrated and tested. The force balance was designed using finite element analysis and consisted of four stress bars instrumented for the measurement of strain. Relative errors of less than 2% were obtained for the recovered simulated calibration loads, while errors of less than 3% were obtained for lift and thrust components for simulated fuel-on and fuel-off force loading distributions. Tests in a calibration rig showed that the balance was capable of recovering the magnitude of point loads to within 3% and their lines of action to within 1% of the chord of the model. Additional errors result when testing in a wind tunnel. The uncertainties for the experiments with fuel injection are estimated at 9%, 7% and 9% for the coefficients of lift, thrust and pitching moment. The scramjet vehicle was 0.566m long and weighed approximately 6kg. It consisted of an inlet, combustion chamber and thrust surface. Fuel could be injected through a series of injectors located on the scramjet inlet. The scramjet model was set at zero angle of attack. Experiments were performed in the T4 Free Piston Shock Tunnel at a total enthalpy of 3.3MJ/kg, a nozzle supply pressure of 32MPa and a Mach number of 6.6, with equivalence ratios up to 1.4. Fuel-off force coefficients were measured to within 2% of theoretical values based on predictions using CFD and hypersonic theory. The fuel-off centre-of-pressure was measured to within 4% of the predicted value. The force coefficients varied linearly with equivalence ratio. Good comparison of the measured lift and thrust forces with theoretical values was obtained with increasing flow rates of fuel. The lift-to-drag ratio increased from 3.0 at the fuel-off condition to 17.2 at an equivalence ratio of 1.0. Poor agreement between the measured pitching moment and theoretical values was obtained due to difficulties in predicting the pressure distribution with heat addition on the latter parts of the thrust surface. A shift in the centre-of-pressure of approximately 10% of model chord was measured as the equivalence ratio varied from 0.0 to 1.0. For the design tested, the thrust produced was not enough to overcome drag on the vehicle, even at the highest equivalence ratio tested. Tests at higher stagnation enthalpies (up to 4.9MJ/kg) showed the lift and pitching moment coefficients remained constant with an equivalence ratio of 0.8 but the thrust coefficient decreased exponentially with increasing stagnation enthalpies. Good agreement of experimental values of lift and thrust force with predicted values was obtained for equivalence ratios of 0.0 and 0.8. Choking occurred at stagnation enthalpies of less than 3.0MJ/kg and a nozzle supply pressure of 32MPa with fuel injection at an equivalence ratio of approximately 0.8, resulting in a drag force of approximately 2.5 times the fuel-off drag force. Tests at a nozzle supply enthalpy of 3.3MJ/kg and nozzle supply pressures of 32, 26 and 16MPa were performed at equivalence ratios of 0.0 and 0.8. The fuel-off lift coefficient remained constant but the thrust coefficient increased. This is attributed to a reduction in skin friction associated with longer lengths of laminar boundary layers as the Reynolds number was decreased. The measured fuel-off lift and thrust coefficients agreed with the predicted values to within the known test flow and force prediction uncertainties. Combustion did not occur at a nozzle supply pressure of 16MPa. This work has demonstrated that overall scramjet vehicle performance measurements (such as lift-to-drag ratio and shifts in centre-of-pressure) can be made in a free piston shock tunnel.

The purpose of this research was to determine if it is feasible to apply Lorentz mixing to supersonic diffusion flames, such as those found in SCRAMjet engines. The combustion rate in supersonic diffusion flames is limited by the rate at which air and fuel mix. Lorentz mixing increases turbulence within a flow, which increases the rate at which species mix and thus increases the rate of combustion. In order to determine the feasibility of Lorentz mixing for this application, a two-dimensional model of supersonic reacting flow with the application of a Lorentz force has been examined numerically. The flow model includes the complete Navier-Stokes equations, the ideal gas law, and terms to account for diffusion of chemical species, heat release due to chemical reaction, change in species density due to chemical reaction, and the Lorentz forces applied during Lorentz mixing. In addition, the Baldwin-Lomax turbulence model is used to approximate turbulent transport properties. A FORTRAN program using the MacCormack method, a commonly used computational fluid dynamics algorithm, was used to solve the governing equations. The accuracy of the program was verified by using the program to model flows with known solutions. Results were obtained for flows with Lorentz forces applied over a series of power levels and frequencies. The results show significant increases in the rate of combustion when Lorentz mixing is applied. The amount of power required to drive Lorentz mixing is small relative to the rate at which energy is released in the chemical reaction. An optimum frequency at which to apply Lorentz mixing was also found for the flow being considered. The results of the current study show that Lorentz mixing looks promising for increasing combustion rates in supersonic reacting flows, and that future study is warranted. In particular, researchers attempting to improve combustion in SCRAMjet engines may want to consider Lorentz mixing as a way to improve combustion. / Graduation date: 1997

A numerical analysis was conducted to study the effects of geometrically scaling scramjet inlet-combustor isolators. Three-dimensional fully viscous numerical simulation of the flow inside constant area rectangular ducts, with a downstream back pressure condition, was analyzed using the SolidWorks Flow Simulation software. The baseline, or 1X, isolator configuration has a 1” x 2.67” cross section and 20” length. This baseline configuration was scaled up based on the 1X configuration mass flow to 10X and 100X configurations, with ten and one hundred times the mass flow rate, respectively. The isolator aspect ratio of 2.67 was held constant for all configurations. To provide for code validation, the Flow Simulation program was first used to analyze a converging-diverging channel and a wind tunnel nozzle. The channel case was compared with analytical theory and showed good agreement. The nozzle case was compared with AFRL experimental data and showed good agreement with the entrance and exit conditions (Pi0= 40 psia, Ti0= 530ºR, Pe= 18.86 psia, Te= 456ºR, respectively). While the boundary layer thickness remained constant, the boundary layer thickness with respect to the isolator height decreased as the scale increased. For all the isolator simulations, a shock train was expected to form inside the duct. However, the flow simulation failed to generate this flow pattern, due to improper sizing of the isolator and combustor for a 3-D model or having a low pressure ratio of 2.38. Instead, a single normal shock wave was established at the same relative location within the length of each duct, approximately 80% of the duct length from the isolator entrance. The shape of the shock changed as the scale increased from a normal shock wave, to a bifurcated shock wave, and to a normal shock train, respectively for the 1X, 10X, and 100X models.

A scramjet inlet was designed for use on a small scale, Mach six, ethylene-fuelled vehicle. The inlet used strut-based cantilevered fuel injectors and a well-defined mixing duct to mix fuel prior to the combustor. Designed using theoretical and numerical analyses, the resulting inlet configuration consisted of a single body shock inlet with vertical fuel injector struts and four cantilevered injectors per strut side. This inlet was 80 cm long and 42 cm high. Numerical analysis of the vehicle was conducted with computational fluid dynamics by solving the Favre-Averaged Navier-Stokes equations; turbulence was simulated using the Wilcox k-ω model. Multispecies simulations in two and three dimensions were used to evaluate the design. Analysis of the simulated flow features, thrust potential and mixing efficiency demonstrated favourable vehicle performance. In particular, the inlet allowed for complete combustion when lean equivalence ratios of less than 0.7 were used.

A scramjet inlet was designed for use on a small scale, Mach six, ethylene-fuelled vehicle. The inlet used strut-based cantilevered fuel injectors and a well-defined mixing duct to mix fuel prior to the combustor. Designed using theoretical and numerical analyses, the resulting inlet configuration consisted of a single body shock inlet with vertical fuel injector struts and four cantilevered injectors per strut side. This inlet was 80 cm long and 42 cm high. Numerical analysis of the vehicle was conducted with computational fluid dynamics by solving the Favre-Averaged Navier-Stokes equations; turbulence was simulated using the Wilcox k-ω model. Multispecies simulations in two and three dimensions were used to evaluate the design. Analysis of the simulated flow features, thrust potential and mixing efficiency demonstrated favourable vehicle performance. In particular, the inlet allowed for complete combustion when lean equivalence ratios of less than 0.7 were used.

A numerical analysis was conducted to study the effects of geometrically scaling scramjet inlet-combustor isolators. Three-dimensional fully viscous numerical simulation of the flow inside constant area rectangular ducts, with a downstream back pressure condition, was analyzed using the SolidWorks Flow Simulation software. The baseline, or 1X, isolator configuration has a 1” x 2.67” cross section and 20” length. This baseline configuration was scaled up based on the 1X configuration mass flow to 10X and 100X configurations, with ten and one hundred times the mass flow rate, respectively. The isolator aspect ratio of 2.67 was held constant for all configurations. To provide for code validation, the Flow Simulation program was first used to analyze a converging-diverging channel and a wind tunnel nozzle. The channel case was compared with analytical theory and showed good agreement. The nozzle case was compared with AFRL experimental data and showed good agreement with the entrance and exit conditions (Pi0= 40 psia, Ti0= 530ºR, Pe= 18.86 psia, Te= 456ºR, respectively). While the boundary layer thickness remained constant, the boundary layer thickness with respect to the isolator height decreased as the scale increased. For all the isolator simulations, a shock train was expected to form inside the duct. However, the flow simulation failed to generate this flow pattern, due to improper sizing of the isolator and combustor for a 3-D model or having a low pressure ratio of 2.38. Instead, a single normal shock wave was established at the same relative location within the length of each duct, approximately 80% of the duct length from the isolator entrance. The shape of the shock changed as the scale increased from a normal shock wave, to a bifurcated shock wave, and to a normal shock train, respectively for the 1X, 10X, and 100X models.