Allan Paull
The University of Queensland, Australia, SoMME, Faculty Member
Research Interests:
HIFiRE 8 is a hypersonic flight test experiment scheduled for launch in late 2018 from the Woomera Test Center in Australia. This project aims to develop a Flight Test Vehicle that will, for the first time, complete 30 seconds of scramjet... more
HIFiRE 8 is a hypersonic flight test experiment scheduled for launch in late 2018 from the Woomera Test Center in Australia. This project aims to develop a Flight Test Vehicle that will, for the first time, complete 30 seconds of scramjet powered hypersonic flight at a Mach Number of 7.0. The engine used for this flight will be a rectangular to elliptic shape transition scramjet. It will be fuelled with gaseous hydrogen. The flight test engine configuration will be derived using scientific and engineering evaluation in the UQ shock tunnel T4 and other potential ground-based facilities. This paper presents current plans for the HIFiRE 8 trajectory, mission events, airframe and engine designs and also includes descriptions of critical subsystems and associated modelling, simulation and analysis activities.
Research Interests:
Research Interests:
Research Interests:
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... more
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 sc...
Research Interests:
... Investigation of an axisymmetric scramjet configuration utilising inlet-injection and radical farming. Hunt, DC, Paull, A., Boyce, RR and Hagenmaier, M. (2009). Investigation of an axisymmetric scramjet configuration utilising... more
... Investigation of an axisymmetric scramjet configuration utilising inlet-injection and radical farming. Hunt, DC, Paull, A., Boyce, RR and Hagenmaier, M. (2009). Investigation of an axisymmetric scramjet configuration utilising inlet-injection and radical farming. ...
Research Interests:
Research Interests:
Measurement of drag have been made in a shock tunnel on a simple integrated vehicle engine combination for hypersonic cruise with hydrogen scramjet propulsion. The test flow Mach number was 6.4, and the velocity was 2.45 kms(exp -1). Zero... more
Measurement of drag have been made in a shock tunnel on a simple integrated vehicle engine combination for hypersonic cruise with hydrogen scramjet propulsion. The test flow Mach number was 6.4, and the velocity was 2.45 kms(exp -1). Zero Drag, which is the necessary condition for cruise, was achieved as the equivalence ratio approached one. It was found that an analysis using established aerodynamic concept was adequate for predicting drag in the case of no combustion. When combustion occurred results of direct connect experiments provided was qualitative guide to the measured levels of drag, and indicated that thrust nozzle combustion was taking place. An heuristic analysis is used to point to the important effect this may have on propulsive lift.
Research Interests:
The results of simultaneous heat transfer and pressure measurements at the walls of three different configurations of a model scramjet are presented. The heat transfer results are compared with results empirically predicted from the... more
The results of simultaneous heat transfer and pressure measurements at the walls of three different configurations of a model scramjet are presented. The heat transfer results are compared with results empirically predicted from the pressure measurements. It is shown that the measured heat transfer rate is comparable with, or lower than, that predicted for a laminar boundary layer. A mathematical model is proposed for the film-cooling effect observed when a hydrogen fuel is injected along a wall of the scramjet. In this mathematical model, the heat transfer rate is shown to be insensitive to the velocity profile in the insulating layer of fuel. The model suggests that the cooling layer is turbulent and that 90 percent of the fuel is mixed with the air.
Research Interests:
ABSTRACT Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology and hypervelocity pulse test facilities are presented. These reports document the tests conducted... more
ABSTRACT Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology and hypervelocity pulse test facilities are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 10 under NASA Grant NAGw-674.
Research Interests:
A series of experiments were initiated to investigate the operation of a two-dimensional, hypersonic, airbreathing engine (scramjet) inclined at angles of attack to the freestream. The experiments were undertaken to obtain data for use in... more
A series of experiments were initiated to investigate the operation of a two-dimensional, hypersonic, airbreathing engine (scramjet) inclined at angles of attack to the freestream. The experiments were undertaken to obtain data for use in the Hyshot flight test program. Experiments on the Hyshot scramjet were under taken in the T4 shock tunnel. Experiments were made at a nominal total enthalpy of 3.0MJkg (exp -1) using a nozzle that produced flows with a Mach number of approximately 6.5. The conditions produced correspond to flight at Mach 7.6 at an altitude range of 35.7-21.4km. A summary of the flow conditions is included. The scramjet was tested at 0, plus 2, plus 4, minus 2 and minus 4 degrees angle of attack. Experiments were also undertaken at 2 and 4 degrees angle of skew.
This note reports tests in a shock tunnel in which a fully integrated scramjet configuration produced net thrust. The experiments not only showed that impulse facilities can be used for assessing thrust performance, but also were a... more
This note reports tests in a shock tunnel in which a fully integrated scramjet configuration produced net thrust. The experiments not only showed that impulse facilities can be used for assessing thrust performance, but also were a demonstration of the application of a new technique to the measurement of thrust on scramjet configurations in shock tunnels. These two developments are of significance because scramjets are expected to operate at speeds well in excess of 2 km/sec, and shock tunnels offer a means of generating high Mach number flows at such speeds.
A free-piston shock tunnel has been used to obtain test data on a scramjet combustion chamber with sidewall injection. The results obtained indicate that combustion was strongly influenced by a region of fuel whose temperature was held... more
A free-piston shock tunnel has been used to obtain test data on a scramjet combustion chamber with sidewall injection. The results obtained indicate that combustion was strongly influenced by a region of fuel whose temperature was held below its ignition temperature by wall-cooling effects; this increased the fraction of unburned fuel and resulted in a significant loss of specific impulse. Aerodynamic heating would keep the walls above hydrogen ignition temperature in an actual scramjet powerplant, however. Maximum specific impulse was obtained with a combination of parallel and transverse injection in a long combustion chamber, followed by a dual stage expansion.
Research Interests:
The Center for Hypersonic Education and Research at the University of Maryland and the Centre for Hypersonics in the Department of Mechanical Engineering at the University of Queensland in Brisbane Australia have embarked on a... more
The Center for Hypersonic Education and Research at the University of Maryland and the Centre for Hypersonics in the Department of Mechanical Engineering at the University of Queensland in Brisbane Australia have embarked on a collaborative research program to investigate inversely designed minimum-drag bodies. So-called hypersonic waverider shapes, and low-drag waverider-derived star bodies, have been designed and explored analytically and computationally at the University of Maryland, with a new technique based on power-law flowfields. A quasi-axisymmetric star body was constructed and tested at the University of Queensland in tunnel T4 and compared to an equivalent cone. Experimental results demonstrated performance that exactly matched predictions of analysis. The collaborative interaction and cooperation represented a perfect match of the two research groups strengths and interests, provided an important learning experience for a graduate student, and resulted in some meaningful scientific results.
Research Interests:
Research Interests:
In the current HyShot scramjet flight test the scramjet stays attached to the rocket. In a new HyShot scramjet flight test the scramjet will separate from the rocket to show the ability of producing net thrust. Due to the separation the... more
In the current HyShot scramjet flight test the scramjet stays attached to the rocket. In a new HyShot scramjet flight test the scramjet will separate from the rocket to show the ability of producing net thrust. Due to the separation the scramjet vehicle has to maintain its own flight stability, which in the previous flight test has been provided by the rocket. Previous to the scramjet test the scramjet vehicle will perform a spinning and coning movement due to the spinning of the busting rocket and a turning manoeuvre outside the atmosphere. This movement combined with the aerodynamic forces acting on the vehicle is described in a flight dynamic model. Based on this model an attitude control system has been developed, which will be validated with the HyShot Stability Demonstrator (USD). The new designed USD flight test uses a Zuni rocket motor to bust the USD to Mach 3. At burnout of the rocket motor the USD separates and demonstrates the flight stability. The USD is designed with similar flight dynamic behaviour as the HyShot IX scramjet vehicle. Moreover, the developed subsystems of the USD to maintain the flight test (attitude measurement system, parachute deployment system, low-cost board computer, etc.) will be qualified for the use in the next HyShot scramjet flight tests.
Research Interests:
Research Interests:
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).... more
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 five flew in April 2012. Principle goals of this flight were to measure hypersonic boundary-layer transition on a three-dimensional body. The second stage booster on this flight failed to ignite, so the experiment reached a maximum Mach number of only 3. Nevertheless, supersonic pressure and temperature data were obtained under laminar and turbulent flow, and flight systems were validated. HIFiRE-5 was the first HIFiRE flight to use both the Inertial Sciences Digital Miniature Attitude Reference System (DMARS-R) IMU and Ashtech DG14 Global Positioning System receiver. Results show that a tripped transition occurred on the test article leading edge, but the rest of the configuration showed no gross effects of tripping, with a transition pattern consistent with prior wind tunnel measurements and CFD.
Research Interests:
Abstract The Hypersonic Collaborative Australia/United States Experiment (HyCAUSE) is a DARPA-sponsored, scramjet technology program involving researchers in Australia and the United States. The program, which began in April 2004,... more
Abstract The Hypersonic Collaborative Australia/United States Experiment (HyCAUSE) is a DARPA-sponsored, scramjet technology program involving researchers in Australia and the United States. The program, which began in April 2004, consisted of ground tests and ...
Research Interests:
A new skin friction gauge has been designed for use in impulsive facilities. The gauge was tested in the T4 free piston shock tunnel, at the University of Queensland, using a 1.5 m long plate that formed one of the inner walls of a... more
A new skin friction gauge has been designed for use in impulsive facilities. The gauge was tested in the T4 free piston shock tunnel, at the University of Queensland, using a 1.5 m long plate that formed one of the inner walls of a rectangular duct. The test gas was fair and the test section free stream flow had a
Research Interests:
External combustion experiments were conducted in a shock tunnel at a freestream Mach number of 7.6. Hydrogen fuel was injected near the leading edge of a simple wedge configuration at conditions conducive to auto-ignition of the fuel-air... more
External combustion experiments were conducted in a shock tunnel at a freestream Mach number of 7.6. Hydrogen fuel was injected near the leading edge of a simple wedge configuration at conditions conducive to auto-ignition of the fuel-air mixture. Pressure and heat flux signals were recorded along the intake and expansion surfaces, and were used in combination with holographic plates to explore the usefulness of intake combustion. The leading edge shock was found to displace and deflect from the intake surface, when mass and heat addition occurred. This alleviated the pressure rise due to burning, indicating a trend towards constant pressure burning on the intake. A pressure rise was still measured on the thrust surface. Interaction of the fuel jet with the leading edge shock was observed, indicating vigorous mixing and good penetration of the fuel. The results indicate that external combustion on an intake, or simply mass addition, could be applied to hypersonic flows to generate additional thrust, lift or a change in pitching moment.
Research Interests:
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... more
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 ...
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... more
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.
Computations are reported of the combusting flows within a hydrogen fuelled model supersonic combustion ramjet (scramjet) operating in a shock tunnel flow. The work is in support of future launches in the HyShot free flight scramjet... more
Computations are reported of the combusting flows within a hydrogen fuelled model supersonic combustion ramjet (scramjet) operating in a shock tunnel flow. The work is in support of future launches in the HyShot free flight scramjet program. The scramjet simulates all features of a free flight scramjet including inlet compression ramps, combustion chamber and thrust surfaces. The device employs ramp injection and its ignition is thought to be shock induced. The purpose of the investigation is to enhance understanding of the detailed processes which operate within the device. The computational results are compared with pressure measurements for cases of no fuel injection and fuel injection into nitrogen and air coflows. Significant combustion pressure rise has been observed in the latter experiment. The comparison shows that the computations are physically sound and quite accurately predict many features of the flows. However, they fail to predict significant combustion in the fuel + air flow. Detailed examination is made of the computed solution from which it is inferred that the temperatures within the flow are too low for combustion. Therefore, unless the real flows are quite unlike the prediction, combustion would not be expected within them. No convincing explanation for this disparity between experiment and theory has yet been found and investigation is continuing.
Pressure measurements were taken along the intake, combustion chamber and thrust surface of a scram-jet model tested in a free-piston driven shock tunnel. The scramjet model incorporated a single rectangular cross-section combustion... more
Pressure measurements were taken along the intake, combustion chamber and thrust surface of a scram-jet model tested in a free-piston driven shock tunnel. The scramjet model incorporated a single rectangular cross-section combustion chamber of constant cross-sectional area. The combustion chamber height was varied from 20 to 32mm. The nominal Mach number was 6.5 and the tests were grouped into two flow conditions with stagnation enthalpies of approximately 3 and 4MJ/kg. Hydrogen fuel was injected on the intake and the equivalence ratio was varied from zero up to a maximum of one. Thrust was calculated by integrating the pressure distribution on the thrust surface over its area. Computational estimations of the internal drag on the model were subtracted from the calculated thrust to provide an indication of net thrust and the specific impulse of the engine. Computational thrust predictions generally agreed with the data; and turbulent boundary layer separation correlations were shown to provide reasonable estimates for the equivalence ratio limit at which choking occurs. Postive net specific impulses were indicated at the higher equivalence ratios for most of the test conditions, with maximum values approaching 500 s.
Research Interests:
The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program. The primary experiment for flight one, launched in March 2010, was to measure boundary-layer transition in hypersonic... more
The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program. The primary experiment for flight one, launched in March 2010, was to measure boundary-layer transition in hypersonic flight on a nonablating, 7 deg half-angle axisymmetric cone with a small bluntness of 2.5 mm radius. The flight gathered pressure, temperature, and heat transfer measurements during ascent and reentry. Although the vehicle reentered the atmosphere at a higher-than-intended angle of attack, the ascent portion of the flight provided smooth-body boundary-layer transition data at freestream Mach numbers greater than 5, where transition was presumed to be dominated by second-mode instability. The angle of attack during this portion of the flight was less than 1 deg. The end of turbulent-to-laminar transition occurred at Reynolds numbers between 10.3×106 and 12.2×106, based on x-location and freestream conditions. Transition was correlated with second-mode N-factors of approximately...
Research Interests:
... by Judy Odam, Allan Paull ... HIFiRE 0 was a low cost payload that provided valuable data regarding the performance of on-board systems such as horizon sensors, mag-netometers, accelerometers, cold gas thrusters, launch lug, flight... more
... by Judy Odam, Allan Paull ... HIFiRE 0 was a low cost payload that provided valuable data regarding the performance of on-board systems such as horizon sensors, mag-netometers, accelerometers, cold gas thrusters, launch lug, flight computers and crucially, the flight software. ...
Research Interests:
ABSTRACT 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... more
ABSTRACT 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 has now the capability of testing a complete scramjet engine with internal combustion and external aerodynamics at M=7.8 flight conditions 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.
Research Interests:
The HIFiRE (Hypersonic International Flight Research and Experimentation) program is a collaboration between the Australian Defence Science and Technology Organisation (DSTO) and the US Air Force Research Laboratory (AFRL) to develop and... more
The HIFiRE (Hypersonic International Flight Research and Experimentation) program is a collaboration between the Australian Defence Science and Technology Organisation (DSTO) and the US Air Force Research Laboratory (AFRL) to develop and demonstrate fundamental hypersonic technologies, through computational analysis, ground testing and flight testing. HIFiRE 7 is the first of the free-flyer scramjet flights scheduled for launch in October 2010. It is an uncontrolled ballistic re-entry experiment to measure thrust generated from two REST (Rectangular-to-Elliptical Shape Transition) inlet scramjets mounted back-to-back. The HIFiRE 7 payload will be launched aboard a spin-stabilized two-stage sounding rocket, utilizing an up-and-over ballistic trajectory. During the re-entry phase, the payload will separate from the sustainer and continue as a free-flyer. The flight experiment will take place at Mach 8 between altitudes of ∼34km and 26km. This paper presents an overview the aerodynamic design of the free-flying vehicle from concept to the current configuration, including possible separation strategies. Copyright
Research Interests:
... Miranda, Brandin Northrop, … Ron Mairs, R Unger, The Boeing Company, Huntington Beach, Kei Y Lau, Todd Silvester, Hans Alesi ... to library · Related research 3 readers. The Hypersonic Flight Experiment SHEFEX. Thino Eggers, José MA... more
... Miranda, Brandin Northrop, … Ron Mairs, R Unger, The Boeing Company, Huntington Beach, Kei Y Lau, Todd Silvester, Hans Alesi ... to library · Related research 3 readers. The Hypersonic Flight Experiment SHEFEX. Thino Eggers, José MA Longo, Marcus Hörschgen, Andreas ...
Research Interests:
A device has been produced which can detect the contamination of the test gas by the driver gas in a reflected shock tunnel. This device monitors the static pressure in a converging duct. The duct is designed to choke at a predetermined... more
A device has been produced which can detect the contamination of the test gas by the driver gas in a reflected shock tunnel. This device monitors the static pressure in a converging duct. The duct is designed to choke at a predetermined contamination level due to the change in the specific heat ratio produced by the contaminants. Experimental results are given for a freestream enthalpy of nominally 6 MJ/kg.
Research Interests:
Abstract.A device has been developed to detect the arrival of the driver gas in a shock tunnel. The detector is small enough to be used in conjunction with other experiments. It works by choking a duct when the specific heat ratio is... more
Abstract.A device has been developed to detect the arrival of the driver gas in a shock tunnel. The detector is small enough to be used in conjunction with other experiments. It works by choking a duct when the specific heat ratio is increased past a critical value. Times are given for the onset of a 7.5% contamination level in flows with freestream enthalpies ranging from 3–9 MJ/kg. These results are compared with and are shown to be in agreement with measurements made with a mass spectrometer. Results displaying the rate at which the test gas is contaminated are also given.
Research Interests:
Research Interests:
Research Interests:
... Miranda, Brandin Northrop, … Ron Mairs, R Unger, The Boeing Company, Huntington Beach, Kei Y Lau, Todd Silvester, Hans Alesi ... to library · Related research 3 readers. The Hypersonic Flight Experiment SHEFEX. Thino Eggers, José MA... more
... Miranda, Brandin Northrop, … Ron Mairs, R Unger, The Boeing Company, Huntington Beach, Kei Y Lau, Todd Silvester, Hans Alesi ... to library · Related research 3 readers. The Hypersonic Flight Experiment SHEFEX. Thino Eggers, José MA Longo, Marcus Hörschgen, Andreas ...
Research Interests:
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... more
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 Earth's 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 cone's 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.
Research Interests:
The performance of a power law star body, designed for low drag, is calculated numerically and verified experimentally. The methodology of the shape design is presented along with a comparison to other forebody shapes and an evaluation of... more
The performance of a power law star body, designed for low drag, is calculated numerically and verified experimentally. The methodology of the shape design is presented along with a comparison to other forebody shapes and an evaluation of scaling effects. A star body is then constructed and tested in the reflected shock tunnel, T4, at the University of Queensland. Various test conditions were used to determine both the on design and off design performance of the star body. These results are then compared to tests done on a cone of equivalent volume and length. The resulting star body shape is shown to have 20% less drag than the equivalent cone at a Mach number of 6.4 and a Re = 1.5 x 10.
Research Interests:
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... more
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.
Research Interests:
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... more
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 35.8MPa and a nozzle supply enthalpy of 5.8MJ/kg. Hydrogen fuel was injected normally to promote combustion and tangentially to reduce viscous drag. A series of fuel injectors were used to vary the proportion of tangential to normal fuel, with between 45 and 100 % of the fuel injected tangentially. 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 indicated that the best overall performance was produced by 100% normal injection.
Research Interests:
Despite the very short flow duration, it is possible to conduct direct-connect combustor experiments in impulse facilities by using an unstarted converging-diverging nozzle located in the freestream flow. Such an unstarted inlet contains... more
Despite the very short flow duration, it is possible to conduct direct-connect combustor experiments in impulse facilities by using an unstarted converging-diverging nozzle located in the freestream flow. Such an unstarted inlet contains a very large subsonic region, which would be expected to require considerable time to establish and reach steady flow. Pressure measurements indicate that steady flow was achieved after approximately 1.5 ms of freestream flow, which at low tunnel enthalpies leaves sufficient time for meaningful combustor measurements to be made before driver gas arrival.
Research Interests:
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... more
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.
Research Interests:
Shock-tunnel measurements are reported of skin friction with supersonic hydrogen-air combustion in a constant area duct. A floating-element skin-friction gauge was used, in which the shear force was applied directly to a piezoceramic... more
Shock-tunnel measurements are reported of skin friction with supersonic hydrogen-air combustion in a constant area duct. A floating-element skin-friction gauge was used, in which the shear force was applied directly to a piezoceramic measuring element. The experiments were conducted at stagnation enthalpies of 5.7 and 6.8 MJ kg(-1), a precombustion Mach number of similar to 4.5, and with a maximum duct Reynolds number of 1.3 x 10(7). The measurements showed that, although supersonic combustion caused the skin friction to fluctuate with time, it did not affect the mean value of the skin friction coefficient, and this mean value could be predicted using existing turbulent: boundary-layer theory. Measurements of heat transfer also established that Reynolds analogy could be used in both the fuel-off and fuel-on flows.