Hypersonic airbreathing propulsion for access to space has several advantages over conventional rockets. However, to achieve the theoretical propulsion efficiencies, many technological challenges, such as fast and efficient mixing of the fuel with air, still need to be resolved. Several scramjet inlet geometries naturally generate vortices that could be used for mixing enhancement. Thorough characterization of these vortices is required for studying their effect on mixing and to determine the optimum way to utilize these naturally occurring vortical flow structures. For this purpose, an algorithm that tracks and extracts data along a vortex from a computational fluid dynamics (CFD) flow field solution was implemented. It performs a search in slices of the computational solution, based on Galilean-invariant criteria for vortices. The algorithm has been used to analyse the vortices generated by compression corner and within a real scramjet inlet. Results are presented showing the insight in vortex formation and evolution provided by the algorithm.
To investigate the potential performance benefits of tailored fuel injection in a scramjet engine, flow through a rectangular-to-elliptical shape-transitioning scramjet was simulated with two different combustor injection geometries: one with uniformly-spaced boundary-layer fuel injection, and the other with oblique-angle porthole injectors targeting the flow structures observed in engine simulations. Both engines utilized inlet injection to promote rapid ignition and burning of the combustor-injected fuel. The simulations were compared with experimental data for validation purposes. The tailored-injection engine fueled to an equivalence ratio of 1.24 was found to have superior mixing and combustion efficiencies, improving upon the 1.33 equivalence ratio symmetric engine by 7% and 2%, respectively. Tailored injection was shown to be a promising method for improving scramjet performance, potentially allowing higher fuel efficiency and a shorter combustor section.
To investigate the potential performance benefits of tailored fuel injection in a scramjet engine, flow through a rectangular-to-elliptical shape-transitioning scramjet was simulated with two different combustor injection geometries: one with uniform film injection, and the other with oblique-angle cross-flow porthole injectors targeting the flow structures observed in engine simulations. Both engines used inlet injection to promote rapid ignition and burning of the combustor-injected fuel. The simulations were compared with experimental data for validation purposes. The tailored-injection engine fuelled to an equivalence ratio of 1.24 realized higher performance than the film-injection configuration, achieving oxygen-based mixing and combustion efficiencies of 98.9% and 84.9%, respectively. Tailored injection shows promise as a method for improving scramjet performance, allowing higher fuel efficiency with simplified fuel injection, and potentially shorter combustor sections.
Rectangular-to-Elliptical Shape Transition (REST) scramjet engines show promise as an access-to-space technology, due to its desirable on- and off-design performance. However, a Mach 12 REST engine will require signicant improvements to its combustion efficiency to be useful as a part of a hybrid launch system. Fuel injection must therefore be tai- lored to the internal flow of the engine; flow which has been virtually unexamined until now. Simulation of the flow through a Mach 12 REST engine was performed, leading to a characterization of a complex three-dimensional shock structure, which in turn drives the generation of swept separated flow regions in the inlet and combustor. With much of the air flowing along the cowl side of the engine, injection in this region near shock or vortical structures may lead to great increases in engine combustion efficiency.
Drag reduction is important to improving the performance of scramjet engines operating at high Mach numbers. One demonstrated method for reducing skin-friction drag on a surface exposed to hypersonic flow is the injection and combustion of hydrogen fuel in the boundary layer. However, there are other fuels of interest in scramjet applications, and the underlying mechanisms that drive the reduction of skin friction are not well understood. An existing analytical model for boundary-layer combustion of hydrogen is rederived for a general fueling condition and then extended to allow investigation of the underlying flow physics in this model. Applying this theory to ethylene fueling indicates that skin-friction reduction through boundary-layer combustion is possible with fuels other than hydrogen. Analysis of the modeled boundary-layer profiles demonstrates that skin-friction reduction is accomplished through several coupled mechanisms: a change in near-wall viscosity, density changes and combustion act to reduce Reynolds stresses, and the low-momentum fuel stream thickens the boundary layer and changes the wall-normal velocity gradient. Finally, the theory is used to estimate the maximum fraction of fuel that should be used for skin-friction reduction in a typical scramjet engine.
To investigate the potential benefits of hydrogen fuel injection in a three-dimensional scramjet inlet, flow through a Mach 12 rectangular-to-elliptical shape transitioning scramjet inlet was simulated with and without hydrogen fuel injection along its body-side compression surface. The observed flowfields showed that, at an equivalence ratio of 0.33, the fuel cannot escape the body-side boundary layer, having little effect on macroscopic inlet flow structures. Interaction with the surrounding boundary-layer turbulence causes robust mixing of the fuel in the inlet. Fuel radicals are produced immediately following injection, particularly where the fuel plumes interact with thin, hot hypersonic boundary layers sweeping inward from the inlet side walls. Combustion does not proceed until the inlet compression process is nearly complete due to low static pressures in the inlet. Once the well-mixed fuel is processed by the final inlet compression shock, ignition and combustion occurs rapidly in the nearly constant cross-section region just upstream of the throat. The net drag increase observed was less than 5%. This modest increase was primarily due to the fuel-injection process, with almost no additional drag due to fuel combustion. Similar flow structures in other inlets suggest they would benefit from inlet fuel injection.
Despite being an active topic of research for over 50 years, scramjet technology has only recently matured to a point where flight tests are being successfully carried out at the lower end of the hypersonic regime. While this progress is encouraging, a renewed interest in low-cost, reliable, and environmentally responsible access to space has identified scramjets capable of accelerating to speeds as high as Mach 12 as desirable. One class of scramjets thought to be capable of hypervelocity performance are those that employ three-dimensional streamtraced compression inlets to efficiently compress captured air. One promising example of this type of scramjet is the Mach 12 Rectangular-to-Elliptical Shape-Transitioning (REST) engine. The aims of this study are to investigate and characterize the flow physics behind the Mach 12 REST engine's current performance, and then attempt to improve its combustion performance by tailoring the engine’s fuel injection to its internal flow field without otherwise modifying the engine geometry. To meet these aims, the engine was studied both numerically and experimentally. The first-ever combusting simulations of a REST scramjet operating at Mach 12 conditions were performed for the Mach 12 REST engine using the CFD research code US3D. The simulations covered a range of conditions, including: unfuelled engine flow, inlet-fuelled flow, and various combined inlet/combustor fuelling configurations. The simulations were found to match well with the experiments they were designed to reproduce and be compared against. A comparison of simulations with experimental in-flow conditions and their equivalent flight conditions on an otherwise identical engine showed that experiments in the T4 Stalker tube reproduce engine pressure and heat flux distributions well. The tunnel condition tends to capture less incoming flow than the engine at flight conditions, which leads to the ground-tested engine over-predicting the engine equivalence ratio. The Mach 12 REST inlet was found to produce a thick bubble-shaped boundary layer along its bodyside compression surface, due to the compression effects of the inlet sidewalls acting on a thick turbulent boundary layer ingested from the vehicle forebody. This thick boundary layer forces the majority of inlet-captured air into a high-density, high Mach number flow region along the engine cowlside wall. The inlet also produces a symmetric pair of high-temperature swept separations that enter the engine isolator along the sidewalls of the engine. When fuel is injected from the bodyside surface of the inlet, it remains trapped inside the thick, turbulent boundary layer, where it becomes well-mixed and begins to burn just upstream of the inlet throat. As much as 50% of this fuel is burned by the time it enters the engine isolator, while its injection and burning increases the inlet's drag by less than 5%. This burning bodyside flow region thermally compresses the remaining air flow within the engine isolator, and provides a source of heat and combustion radicals for the ignition of fuel injected further downstream. Overall combustion efficiency of inlet-injected fuel at the engine exhaust plane was found to be nearly 80% at high equivalence ratios. Flow within the Mach 12 REST combustor is strongly shock-dominated. This is caused by both the cowl closure shock train transmitted from the inlet, and a strong recompression shock generated at the start of the combustor. This recompression shock is generated by the flow passing over a backward step at the entrance to the combustor, and is reinforced on the cowlside of the engine by compression caused by the engine flow path turning to realign with the nominal direction of flight. Fuel injected from the face of the combustor step was combined with inlet injection in an attempt to reduce skin friction drag through boundary layer combustion. This was found to be ineffective: the individual fuel jets never coalesced into a single continuous flow structure, and the fuel layer was quickly disrupted by the pressure gradients induced by the engine turn. The step-injected fuel was quickly ignited by the bodyside combustion region spreading circumferentially around the combustor wall, and its final combustion efficiency was found to be 82.9%. Boundary layer injection from the step was replaced with tailored fuel injection, in which fuel was injected directly into the cowlside core flow, and the swept separation regions along the engine sidewall. This was found to improve combustion efficiency to 84.9%, verifying the validity of the tailoring approach. The tailored-injection Mach 12 REST engine was found to produce the same level of thrust as the boundary layer injection case despite having a lower fuel mass flow rate. The uninstalled specific impulse of the tailored injection engine was 5% higher than that of the step-injection configuration. Improvements to the tailored fuel injection geometry, combined with small modifications to the engine’s geometry and operation may be sufficient to boost thrust and reduce internal viscous drag enough to allow a flight-model scramjet to achieve net thrust.
The determination of aerodynamic coefficients for complex high-speed vehicles still requires experimental measurement. This paper details the development of the free-flight measurement technique within the University of Oxford High Density Tunnel. In particular, a novel image processing technique is developed for calculating the position of the model from high-speed video. This study focuses on the measurement of high Reynolds number experimental aerodynamic data for a subscale model of the Skylon space plane of Reaction Engines. Testing was undertaken at a Mach 7 test condition replicating flight at an altitude of 63.5 km. Results for lift, drag, and pitching moment coefficients were obtained over a range of angles of attack. Lift coefficient was nearly linear over the range of angles of attack tested, and drag coefficient was parabolic in shape, but sensitive to model yaw. The vehicle was also shown to be statically unstable, a common characteristic of canard configuration vehicles.
Analytical solutions to the magnetohydrodynamic shock refraction problem are used to explore the limiting interface angles for which regular refraction will occur, for a particular choice of problem parameters and varying magnetic field strength. Beyond this limit, irregular shock refraction occurs and no analytical solutions exist at present. Thus, the wave structure resulting from irregular magnetohydrodynamic shock refraction is determined via ideal simulations, allowing its properties to be explored.
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