One-Dimensional modeling of dual mode scramjet and ramjet flowpaths is a useful tool for scramjet conceptual design and wind tunnel testing. In this thesis, modeling tools that enable detailed analysis of the flow physics within the combustor are developed as part of a new one-dimensional MATLAB-based model named VTMODEL. VTMODEL divides a ramjet or scramjet flow path into four major components: inlet, isolator, combustor, and nozzle. The inlet module provides two options for supersonic inlet one-dimensional calculations; a correlation from MIL Spec 5007D, and a kinetic energy efficiency correlation. The kinetic energy efficiency correlation also enables the user to account for inlet heat transfer using a total temperature term in the equation for pressure recovery. The isolator model also provides two options for calculating the pressure rise and the isolator shock train. The first model is a combined Fanno flow and oblique shock system. The second model is a rectangular shock train correlation. The combustor module has two options for the user in regards to combustion calculations. The first option is an equilibrium calculation with a "growing combustion sphere" combustion efficiency model, which can be used with any fuel. The second option is a non-equilibrium reduced-order hydrogen calculation which involves a mixing correlation based on Mach number and distance from the fuel injectors. This model is only usable for analysis of combustion with hydrogen fuel. Using the combustion reaction models, the combustor flow model calculates changes in Mach number and flow properties due to the combustion process and area change, using an influence coefficient method. This method also can take into account heat transfer, change in specific heat ratio, change in enthalpy, and other thermodynamic properties. The thesis provides a description of the flow models that were assembled to create VTMODEL. In calculated examples, flow predictions from VTMODEL were compared with experimental data obtained in the University of Virginia supersonic combustion wind tunnel, and with reported results from the scramjet models SSCREAM and RJPA. Results compared well with the experiment and models, and showed the capabilities provided by VTMODEL. / Master of Science

The overall performance of a scramjet-powered vehicle not only depends on the performance of individual components but also on how the components interact with one another. Because scramjet engines must be integrated into the vehicle design, the optimization of the design of one component may detrimentally offset the performance of other components. This thesis addresses the optimization of the thrust nozzle of a scramjet-powered vehicle and shows how the optimization must include the integration of the nozzle into the overall vehicle design. The basic scramjet vehicle configuration chosen for the study is the somewhat conventional design of a vehicle with an axisymmetric centrebody, where fuel and payload may be stored, and a quasi-axisymmetric cowl. Six internal intake-combustor-nozzle modules are arranged around the centrebody. Paull et al. (1995) tested a configuration of this type (referred to as model 0) and demonstrated that, at some conditions, a net positive thrust could be produced. They suggest that the thrust nozzle of the vehicle has potential for design change that could lead to significantly improved performance. The main aim of the present study is to test this hypothesis. An important decision that was made early in the present project was that the research would focus on unfuelled operation of the vehicle design. This still allowed the influence of integration of the optimized nozzle to be studied but removed the complications introduced by combustion processes. In order to integrate optimization of design of the thrust nozzle, it was necessary to analyze the performance of the complete scramjet vehicle design. A simple analysis methodology that captures the important physical processes occurring in the flow through and around the vehicle is necessary so that rapid calculations can be made in an iterative optimization program. Therefore, the force prediction methodology that was developed used a combination of simple hypersonics theories and inviscid 3-D CFD modelling instead of using a full 3-D Navier-Stokes equation solver to minimize the calculation time. The pressure forces and viscous forces on the model were calculated separately for each component of the scramjet vehicle designs. The theory of van Driest (1956) for skin friction drag in turbulent boundary layers was found to produce best agreement with measurements. Therefore, it was employed to estimate the turbulent skin friction drag in the present research. To validate the current force prediction methodology, the net axial drag force on an unfuelled quasi-axisymmetric scramjet model derived from the design of Paull et al. (1995) and designed for operation at Mach 6 (model 1) was measured in the T4 Stalker tube at The University of Queensland using a single component Stress Wave Force Balance. (The design of Paull et al. (1995) is referred to here as model 0.) Tests were performed with Mach 6, Mach 8, and Mach 10 nozzles attached to the end of the shock tube. In order to get a wide range of flow conditions the nozzle-supply enthalpy was varied from 3 to 10 MJ/kg and the nozzle-supply pressure from 35 to 45 MPa. A reduction of the drag coefficient of model 1 was observed with decreasing nozzle-supply enthalpy for each of the tunnel nozzles tested. The performance of model 1 was analyzed using the force prediction methodology. Generally, the force prediction results were in good agreement with experimental results. The results indicate that the internal intakes provide 50% of the total drag. The skin friction drag in the combustion chambers and the nozzles account for 30% of the total drag. In order to investigate the influence on the overall performance of the vehicle obtained by improving the nozzle performance, optimization and parametric studies of quasi-axisymmetric scramjet nozzle designs were conducted. The vehicle which was optimized in this study is of a similar configuration to the model used in Paull et al. (1995). The vehicles are optimized for minimum fuel-off net axial drag for a design flight Mach number of 8 using the force prediction methodology and the Nelder and Mead (1965) optimization algorithm. The optimization studies focused on the combustion chamber and the nozzle. Therefore, the shape of the conical forebody and the intake were not changed. The external flow over the cowl was taken into account during the optimization studies. The results showed that a long nozzle with a large external cowl deflection angle, which allowed the nozzle area ratio to be increased, did not give better performance than a short nozzle with a smaller area ratio. This was due to the competing effects of increased external drag on the cowl and increased nozzle thrust as the nozzle area ratio increased. The optimum shape gave limited improvement compared with that of Paull et al. (1995). While fuelled performance of the vehicle was not the focus of the present investigation, a preliminary theoretical study of fuelled operation was performed. A parametric study to vary the nozzle length and external cowl deflection angle was performed for different flight Mach numbers. The results indicate a larger nozzle and higher external cowl deflection angle are appropriate for fuelled cases compared with unfuelled cases. The net axial force on a model with a geometry close to the optimum design (model 2) was measured in the T4 shock tunnel in order to check that the optimization procedure was valid. Model 2 showed generally better performance than other models experimentally. For the Mach 6 nozzle tests, although model 2 has some performance losses due to the spillage of flow around the intakes, model 2 shows approximately a 20% lower drag coefficient than model 1 and shows slightly better performance than model 0. For all test conditions, a break-down of the components of the drag coefficient indicates that the nozzle of model 2 produces approximately three times more thrust than the nozzle of model 0 and approximately twice more than that of model 1. For the Mach 8 nozzle tests, model 2 has approximately a 20% lower drag coefficient than model 1. However, for the Mach 10 nozzle tests, no significant differences between the models were observed in the measurements. Finally, the measurements and optimization study indicate that when model 2 is fuelled, it could be expected to be capable of cruise up to Mach 8 because of its very effective nozzle.

No description available.

An experimental study of an aerodynamic ramp (aeroramp) injector was conducted at Virginia Tech. The aeroramp consisted of an array of two rows with two columns of flush-wall holes that induce vorticity and enhance mixing. For comparison, a single-hole circular injector with the same area angled downstream at 30 degrees was also examined. Test conditions involved sonic injection of helium heated to 313 K, to safely simulate hydrogen into a Mach 4 air cross-stream with average Reynolds number 5.77 e+7 per meter at a jet to freestream momentum flux ratio of 2.1. Sampling probe measurements were utilized to determine the local helium concentration. Pitot and cone-static pressure probes and a diffuser thermocouple probe were employed to document the flow. The main results of this work was that the mixing efficiency value of this aeroramp design which was optimized at Mach 2.4 for hydrocarbon fuel was only slightly higher than that of the single-hole injector at these flow conditions and the mass-averaged total pressure loss parameter showed that the aero-ramp and single-hole injectors had the same overall losses. The natural extension of the investigation was then to look in detail at two major physical phenomena that occurs in a complex injector design such the Aeroramp: the jet-shock interaction and the interaction of the vortical structures produced by the jets injection into a supersonic cross flow. Experimental studies were performed to investigate the effects of impinging shocks on injection of heated helium into a Mach 4 crossflow. It was found that the addition of a shock behind gaseous injection into a Mach 4 crossflow enhances mixing only if the shock is closer to the injection point where the counter-rotating vortex pair (always associated with transverse injection in a crossflow) is not yet formed, and the deposition of baroclinic generated of vorticity is the highest. The final investigation concerned with the interaction of the usual vortex structure produced by jet injection into a supersonic crossflow and an additional axial vortex typical of those that might be produced by the inlet of a scramjet or the forebody of a vehicle to be controlled by jet interaction phenomena. The additional axial vortices were generated by a strut-mounted, diamond cross-section wing mounted upstream of the injection location. The wing was designed to produce a tip vortex of a strength comparable to that of one of the typical counter-rotating vortex pair (CVP) found in the plume of a jet in a crossflow. The profound interaction of supersonic vortices supported by a quantitative description and characterization of the flowfield has been demonstrated. / Ph. D.

A newly designed liquid fuel (kerosene) aeroramp injector/plasma igniter was tested in cold flow using the Virginia Tech supersonic wind tunnel at Mach 2.4. The liquid fuel (kerosene) injector is flush wall mounted and consists of a 2 hole aeroramp array of impinging jets that are oriented in a manner to improve mixing and atomization of the liquid jets. The two jets are angled downstream at 40 degrees and have a toe-in angle of 60 degrees. The plasma torch used nitrogen and air as feedstocks and was placed downstream of the injector as an ignition aid. First, schlieren and shadowgraph photographs were taken of the injector flow to study the behavior of the jets, shape of the plume, and penetration of the liquid jet. The liquid fuel aeroramp was found to have better penetration than a single, round jet at 40 degrees. However, the liquid fuel aeroramp does not penetrate as well as an upstream/downstream impinging jet in a plane aligned with the flow. Next, the Sauter mean droplet diameter distribution was measured downstream of the injector. The droplet diameter was found to vary from 21 to 37 microns and the atomization of the injector does not appear to improve beyond 90 effective jet diameters from the liquid fuel aeroramp. These results were then used to decide on an initial location for the plasma torch. The combined liquid injector/plasma torch system was tested in an unheated (300 K) Mach 2.4 flow with a total pressure of 345 kPa. The liquid fuel (kerosene) volumetric flow rate was varied from 0.66 lpm to 1.22 lpm for the combined liquid injector/plasma torch system. During this testing the plasma torch was operated from 1000 to 5000 watts with 25 slpm of nitrogen and air as feedstocks. The interaction between the spray plume and the plasma torch was observed with direct photographs, videos, and photographs through an OH filter. It is difficult to say that any combustion is present from these photographs. Of course, it would be surprising if much combustion did occur under these cold-flow, low-pressure conditions. Differences between the interaction of the spray plume and the plasma torch with nitrogen and air as feedstocks were documented. According to the OH wavelength filtered photographs the liquid fuel flow rate does appear to have an effect on the height and width of the bright plume. As the liquid fuel flow rate increases the bright plume increases in height by 30% and increases in width slightly (2%). While, a decrease in liquid fuel flow rate resulted in an increase in height by 9% and an increase in width by 10%. Thus, as the liquid fuel flow rate varies the width and height of the bright plume appear to always increase. This can be explained by noticing that the shape of the bright plume changes as the liquid fuel flow rate varies and perhaps anode erosion during testing also plays a part in this variation of the bright plume. From the OH wavelength filtered photographs it was also shown that the bright plume appears to decrease in width by 9% and increase in height by 22% when the plasma torch is set at a lower power setting. When air is used as the torch feedstock, instead of nitrogen, the penetration of the bright plume can increase by as much as 19% in width and 17% in height. It was also found that the height and width of the bright plume decreased slightly (2%) as the fuel flow rate increased when using air as the torch feedstock. Testing in a hot-flow facility is planned. / Master of Science

An experimental investigation of a flush-wall, diamond-shaped injector was conducted in the Virginia Tech supersonic wind tunnel. The diamond injector was elongated in the streamwise direction and is aimed downstream angled up at 60° from the wall. Test conditions involved sonic injection of helium heated to approximately 313 K into a nominal Mach 4.0 crossstream airflow. These conditions are typical of a scramjet engine for a Mach 10 flight, and heated helium was used to safely simulate hydrogen fuel. The injector was tested at two different injectant conditions. First, it was investigated at a baseline mass flow rate of 3.4 g/s corresponding to an effective radius of 3.54 mm and a jet-to-freestream momentum flux ratio of 1.04. Second, a lower mass flow rate of 1.5 g/s corresponding to an effective ratio of 2.35 mm and a jet-to-freestream momentum flux ratio of 0.49 was studied. The diamond injector was tested both aligned with the freestream and at a 15° yaw angle for the baseline mass flow rate and aligned with the freestream at the lower mass flow rate. For comparison, round injectors angled up at 30° from the wall were also examined at both flow rates. A smaller round injector was used at the lower mass flow rate such that the jet-to-freestream momentum flux ratio was 1.75 for both cases. A concentration sampling probe and gas analyzer were used to determine the local helium concentration, while Pitot, cone-static and total temperature probes were used to determine the flow properties. The results of the investigation can be summarized as follows. For the baseline case, the aligned diamond injector penetrated 44% higher into the crossflow than did the round injector. The addition of yaw angle increased the crossflow penetration to 53% higher than the round injector. The aligned diamond injector produced a 34% wider jet than the round injector, while the addition of yaw angle somewhat reduced this widening effect to 26% wider than the round injector. The aligned and yawed diamond injectors exhibited 10% and 15% lower mixing efficiency than the round injector, respectively. The total pressure loss parameter of the aligned diamond was 22% lower than the round injector, while the addition of yaw angle improved the total pressure loss parameter to 34% lower than the round injector. For the lower mass flow (and momentum flux ratio) case, the diamond injector demonstrated 52% higher penetration and a 39% wider plume than the round injector. The mixing efficiency was nearly identical between the two injectors with just a 4% lower mixing efficiency for the diamond injector. The total pressure loss parameter of the diamond injector was 32% lower than round injector. These results confirm the conclusions of earlier, lower free stream Mach number and higher molecular weight injectant, studies that a slender diamond injector provides significant benefits for crossflow penetration and lower total pressure losses. / Master of Science

The design of supersonic flow paths for scramjet engines and high Mach number shock tunnel nozzles is complicated by high temperature flow effects and multidimensional inviscid/ viscous flow interactions. Due to these complications, design in the past has been enabled by making flow modelling simplifications that detract from the accuracy of the flow analysis. A relatively new approach to designing aerodynamic bodies, which automates design and does not require as many simplifying assumptions to be effective, is the coupling of a computational flow solver to an optimization algorithm. In this study, a new three-dimensional space-marching computational flow solver is developed and coupled to a gradient-search optimization algorithm. This new design tool is then used for the design optimization of an axisymmetric scramjet flow path and two high Mach number shock tunnel nozzles. The flow solver used in the design tool is an explicit, upwind, space-marching, finite-volume solver for integrating the three-dimensional parabolized Navier-Stokes equations. It is developed with an emphasis on simplicity and efficiency. Cross-stream fluxes are calculated using Toro's efficient upwind, linearized, approximate Riemann solver in flow regions of slowly varying data, and an Osher type solver in the remainder of the flow. Vigneron's technique of splitting the streamwise pressure gradient in subsonic regions is used to stabilise the flux calculations. A three-dimensional implementation of an algebraic turbulence model, a finite-rate chemistry model and a thermodynamic equilibrium model are also implemented within the solver. A range of test cases is performed to (1) validate and verify the phenomenological models implemented within the solver, thereby ensuring the simulation results used for design are credible, and (2) demonstrate the speed of the solver. The first application of the new computational design tool is the design of a scramjet flow path, which is optimized for maximum axial thrust at a flight Mach number of 12. The optimization of a scramjet flow path has been examined previously, however, this study differs to others published in that the flow is modelled using a turbulence model and a finite-rate chemical reaction model which add to the fidelity of the simulations. The external shape of the scramjet vehicle is constrained early on in the design process, therefore, the design of the scramjet is restricted to the internal flow path. Because of this constraint, and the large internal surface area of the combustor and the high skin friction iv within the combustor, the net calculated force exerted on the scramjet for both the initial and optimized design is a drag force. The drag force of the initial design, however, is reduced by 60% through optimization. The second application of the design tool is the wall contour of an axisymmetric Mach 7 shock tunnel nozzle, which is computationally optimized for minimum test core flow variation to a level of +/- 0.019 degrees for the flow angularity and +/- 0.26% for the Pitot pressure. The design is verified by constructing a nozzle with the optimized wall contour and conducting experimental Pitot surveys of the nozzle exit flow. The measured standard deviation in core flow Pitot pressure is 1.6%. However, because there is a large amount of experimental noise, it is expected that the actual core flow uniformity may be better than indicated by the raw experimental data. The last application of the computational design tool is a contoured Mach 7 square cross-section shock tunnel nozzle. This is a three-dimensional optimization problem that demonstrates the versatility of the design tool, since the effort required to implement the optimization algorithm is independent of the complexity of the flow-field and flow solver. Optimization results show that the variation in the test core flow properties could only be reduced to a Mach number variation of +/- 7% and flow angle variation of +/- 1.2 degrees ,for a short nozzle suitable for a shock tunnel. The magnitudes of the optimized nozzle exit flow deviations for the short nozzle and two other longer nozzles indicate that generating uniform flow becomes increasingly difficult as the length of square cross-section nozzles is reduced. Overall, the current research shows that coupling a flow solver to an optimization algorithm is an effective and insightful way of designing scramjets and shock tunnel nozzles.

An integrated aerodynamic-ramp-injector/plasma-torch-igniter of original design was tested in a Mâ = 2, unvitiated, heated flow facility arranged as a diverging duct scramjet combustor. The facility operated at a total temperature of 1000 K and total pressure of 330 kPa. Hydrogen (H2), ethylene (C2H4) and methane (CH4) were used as fuels, and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements, and the presence of combustion was determined based on the pressure rises obtained. Supersonic and dual-mode combustion were achieved with hydrogen and ethylene fuel, whereas very limited heat release was obtained with the methane. Global operability limits were determined to be 0.07 < Ï < 0.31 for hydrogen, and 0.14 < Ï < 0.48 for ethylene. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10 unswept compression ramp injector and similar performance was found with the two arrangements. With hydrogen and ethylene as fuels and the aeroramp/plasma-torch system, the effect of varying the air total temperature was investigated. Supersonic combustion was achieved with temperatures as low as 530K and 680K for the two fuels, respectively. These temperatures are facility/operational limits, not combustion limits. The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with Ï from a high of about 75% to a low of about 45% at the highest Ï . With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Fuel specific impulse was on average 3000 and 1000 seconds for hydrogen and ethylene respectively, and air specific impulse varied from a low of about 9 sec to a high of about 24 sec (for both fuels) for the To = 1000K test condition. The GASP RANS code was used to numerically simulate the injection and mixing process of the fuels. The results of this study were very useful in determining the suitability of the selected plasma torch locations. Further, this tool can be used to determine whether combustion is theoretically possible or not. / Ph. D.

No description available.