For flights beyond Mach 6 ramjets are inefficient engines due to huge total pressure loss in the normal shock systems, combustion conditions that lose a large fraction of the available chemical energy due to dissociation and high structural loads. However if the flow remains supersonic inside the combustion chamber, the above problems could be alleviated and here the concept of SCRAMJET(supersonic combustion ramjet) comes into existence. The scramjets could reduce launching cost of satellites by carrying only fuel and ingesting oxygen from atmospheric air. Further applications could involve defense and transcontinental hypersonic transport. In the current study an effort is made to achieve supersonic combustion in a ground based short duration test facility(combustion driven shock-tunnel), which in addition to flight Mach number can simulate flight Reynolds number as well. In this study a simple method of injection i.e. wall injection of the fuel into the combustion chamber is used. The work starts with threedimensional numerical simulation of a non-reacting gas(air) injection into a hypersonic cross-flow of air to determine the conditions in which air penetrates reasonably well into the cross-flow. Care is taken so that the process does not induce huge pressure loss due to the bow shock which appears in front of the jet column. The code is developed in-house and parallelized using OpenMp model. This is followed by experiments on air injection into a hypersonic cross-flow of air in a conventional shock-tunnel HST2 existing in IISc. The most tricky part is synchronization of injection with start of test-flow in such a short duration(test time 1 millisecond) facility. Next part focuses on numerical simulations to determine the free-stream conditions, mainly the temperature and pressure of air, so that combustion takes place when hydrogen is injected into a supersonic cross-flow of air. The simulations are two-dimensional and includes species conservation equations and source terms due to chemical reactions in addition to the Navier-Stokes equations. This code is also built in-house and parallelized because of more number of operations with the inclusion of species conservation equations and chemical non-equilibrium. However, the predicted conditions were not achievable by HST2 due to low stagnation conditions of HST2. Therefore, a new shock-tunnel which could produce the required conditions is built. The new tunnel is a combustion driven shock-tunnel in which the driver gas is at higher temperature than conventional shock-tunnel. The driver gas is basically a mixture of hydrogen, oxygen and helium at a mole ratio of 2:1:10 initially. The mixture is ignited by spark plugs and the hydrogen and oxygen reacts releasing heat. The heat released raises the temperature of the mixture which is now predominantly helium and small fractions of water vapour and some radicals. The composition of the driver gas and initial pressure are determined through numerical simulations. Experiments follow in the new tunnel on hydrogen injection into a region of supersonic cross-flow between two parallel plates with a wedge attached to the bottom plate. The wedge reduces the hypersonic free-stream to Mach 2. A high-speed camera monitors the flow domain around injection point and sharp rise in luminosity is observed. To ascertain whether the luminosity is due to combustion or not, two more driven gases namely nitrogen and oxygen-rich air are used and the luminosity is compared. In the first case, the free-stream contains no oxygen and luminosity is not observed whereas in the second case higher luminosity than air driver case is visible. Additionally heat-transfer rates are measured at the downstream end of the model and at a height midway between the plates. Similar trend is observed in the relative heat-transfer rates. Wall static pressure at a location downstream of injection port is also measured and compared with numerical simulations. Results of numerical simulations which are carried out at the same conditions as of experiments confirm combustion at supersonic speed. Experiments and numerical simulations show presence of supersonic combustion in the setup. However, further study is necessary to optimize the parameters so that thrust force could be generated efficiently.

In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6. The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.

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.

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.

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.