This thesis aims to assist the development of a multiblock implicit Navier-Stokes code for hypersonic flow applications. There are mainly three topics, which concern the understanding of basic Riemann solvers, the implementing of implicit zonal method, and grid adaption for viscous flow. Three problems of Riemann solvers are investigated. The post-shock oscillation problem of slowly moving shocks is examined, especially for Roe's Riemann solver, and possible cures are suggested for both first and second order schemes. The carbuncle phenomenon associated with blunt body calculation is cured by a formula based on pressure gradient, which will not degrade the solutions for viscous calculations too much. The grid-dependent characteristic of current upwind schemes is also demonstrated. Several issues associated with implicit zonal methods are discussed. The effects of having different mesh sizes in different zones when shock present are examined with first order explicit scheme and such effects are shown to be unwanted therefore big mesh size change should be avoided. Several implicit schemes are tested for hypersonic flow. The conservative DDADI scheme is found to be the most robust one. A simple and robust implicit zonal method is demonstrated. A proper treatment of the diagonal Jacobian and choosing the updating method are found to be crucial. The final topic concerns the calculation and grid adaption of viscous flow. We study the linear advection-diffusion equation thoroughly. The results are unfortunately not applicable to Navier-Stokes equations directly. Nevertheless a suggestion on the mesh size control for viscous flow is made and demonstrated. An attempt to construct a cell-vertex TVD scheme is described in the appendix.

No description available.

No description available.

A body exposed to hypersonic flow is subjected to extremely high wall heating rates, owing to the conversion of the kinetic energy of the oncoming flow into heat through the formation of shock waves and viscous dissipation in the boundary layer and this is one of the main concerns in the design of any hypersonic vehicle. The conventional way of tackling this problem is to use a blunt fore-body, but it also results in an increase in wave drag and puts the penalty of excessive load on the propulsion system. An alternative approach is to alter the flow field using external means without changing the shape of the body; and several such methods are reported in the literature. The superiority of such methods lie in the fact that the effective shape of the body can be altered to meet the requirements of low wave drag, without having to pay the penalty of an increased wall heat transfer rate. Among these techniques, the use of local energy addition in the freestream to alter the flow field is particularly promising due to the flexibility it offers. By the suitable placement of the energy source relative to the body, this method can be effectively used to reduce the wave drag, to generate control forces and to optimise the performance of inlets. Although substantial number of numerical investigations on this topic is reported in the literature, there is no experimental evidence available, especially under hypersonic flow conditions, to support the feasibility of this concept. The purpose of this thesis is to experimentally investigate the effect of energy deposition on the flow-field of a 120� apex angle blunt cone in a hypersonic shock tunnel. Energy deposition is done using an electric arc discharge generated between two electrodes placed in the free stream and various parameters influencing the effectiveness of this technique are studied. The effect of energy deposition on aerodynamic parameters such as the drag force acting on the model and the wall heat flux has been investigated. In addition, the unsteady flow field is visualised using a standard Z-type schlieren flow visualisation setup. The experimental studies have shown a maximum reduction in drag of 50% and a reduction in stagnation point heating rate of 84% with the deposition of 0.3 kW of energy. The investigations also show that the location of energy deposition has a vital role in determining the flow structure; with no noticeable effects being produced in the flow field when the discharge source is located close to the body (0.416 times body diameter). In addition, the type of the test gas used is also found to have a major influence on the effectiveness of energy deposition, suggesting that thermal effects of energy deposition govern the flow field alteration mechanism. The freestream mass flux is also identified as an important parameter. These findings were also confirmed by surface pressure measurements. The experimental evidence also indicates that relaxation of the internal degrees of freedom play a major role in the determination of the flow structure. For the present experimental conditions, it has been observed that the flow field alteration is a result of the interaction of the heated region behind the energy spot with the blunt body shock wave. In addition to the experimental studies, numerical simulations of the flow field with energy deposition are also carried out and the experimentally measured aerodynamic drag with energy deposition is found to match reasonably well with the computed values.

Flow control has always been an intense research subject with the pursuit of favourable control effects like drag reduction, transition delay, and separation prevention. In practice, these flow control effects are achieved using mechanical actuators such as deflectors, vortex generators, transverse jets and so on. However, such mechanical actuators may face the drag penalty and limitation of actuation response time. In recent years, energy deposition has been suggested as a novel flow control technique in high-speed flow with preferable characteristics like non-intrusive, easy arrangement and high actuation frequency. The motivation of this work is to experimentally explore the flow behaviour after the certain amount of energy is deposited in Mach 5 flow. The energy deposition is implemented using a thermal bump (surface energy deposition) and laser beam focusing (volumetric energy deposition).This work starts with the development of a measurement technique of luminescent paint for the present challenging hypersonic testing environment, which is used for the further energy deposition experiment. The successes of the luminescent paint development is demonstrated both on two-dimensional and axisymmetric models. The luminescent paint shows high spatial resolution and the accuracy comparing to the pressure transducer reading. The surface energy deposition is performed using an embedded heating element (thermal bump) on a flat plate. Qualitative and quantitative measurement techniques are utilised to study the modification to the flow structure and the alteration to the distribution of pressure and heat transfer rate after thermal bump is activated. The results reveal the appearance of induced shock wave and suspicious vortices traces due to the activated thermal bump as reported in other literatures. Re-distribution of surface pressure and heat transfer rate are also found.For the volumetric energy deposition, the laser beam is firstly focused in quiescent air in order to understand the induced flow pattern and the impingement to a solid plate. High-speed schlieren photography is utilised to provide an insight to the dynamic evolution of the induced shock wave propagation and plasma kernel development after laser-induced air breakdown. Then, the laser energy deposition is conducted over a flat plate with the presence of Mach 5 flow. The outward motion of the induced shock wave significantly distorts the boundary layer and changes the surface pressure distribution. The results show the different pattern of boundary distortion when laser beam energy is deposited at different positions downstream of the leading edge of flat plate. The entire induced flow pattern is similar to those induced by a pulsed micro-jet. In spite of the laser pulse width of 4 ns, the entire dynamic process lasts about 100 μs.

Abstract The main objective of this thesis was to develop a method than can be used to approximate the pressure forces on air vehicles traveling at hypersonic speed (Mach number > 5). The aerodynamic forces such as lift and drag were calculated from the pressure values on the surface of the airplane. Pitching moment was also tabulated. This work was initiated based on the idea of developing a flow solver proficient and capable of providing aerodynamic data (lift and drag look-up tables) for hypersonic air vehicles that can be fed to a flight simulator (used by the Aviation Systems Department) at the University of Tennessee Space Institute. Several approximation methods are used to solve hypersonic such as shock expansion method. Based on different studies, Computational Fluid Dynamic (CFD) proved to produce very accurate results; however, it is a difficult technique to use. In this thesis work Newtonian Method was adopted as a technique to approximate the aerodynamic forces and hence the performance of hypersonic airplanes, therefore, a computer program (Hyper-N) has been developed for aerodynamic analysis of three dimensional geometries airplane. The program is designed to read in a previously configured list of plates and compute the aerodynamic forces and moments for hypersonic free stream conditions. Programming was completed using MatLab language. The results obtained from the Hyper-N program were for the experimental airplane X-43A which were found to match the results when the shock expansion method is used for the same airplane, [1]. Because of the difficulties involve in using CFD or the complete Navier Stocks equation to obtain the aerodynamic forces on bodies traveling at hypersonic speeds, the Newtonian method is considered to be the most efficient technique to use for preliminary evaluation of the performance of hypersonic airplanes. Modified Newtonian theory and the computational requirement of the code are described. A number of geometric configurations, including the X-43A (experimental hypersonic) airplane, are provided as examples of applications of the Hyper-N program.

A numerical study of the flow field generated by injection through diamondshaped orifices into a high-speed flow is presented in this document. Jet interaction flows have a wide range of applications in the field of engineering. These applications include the use of jets for fuel injection in scramjets, for reaction control of high-speed aerodynamic bodies and as cooling jets for skins of high-speed vehicles. A necessary requirement in the use of transverse jets for these and other applications is a thorough understanding of the physics of the interaction between the jet and freestream. This interaction generates numerous flow structures that include multiple shocks, vortices, recirculation regions and shear layers. This study involves diamond-shaped orifices that have the advantage of generating weaker or attached interaction shocks as compared to circular injectors. These injectors also negate the effects due to the recirculation region that is formed upstream of the injector. This study was undertaken in order to gain further understanding of the flow features generated by diamond-shaped injectors in a high-speed flow. Numerical simulations were performed using two different levels of turbulence models. Reynolds™ Averaged Navier-Stokes (RANS) simulations were performed using the GASP flow solver while Detached-Eddy Simulation (DES) runs were performed using the Cobalt flow solver. A total of fifteen diamond injector simulations were performed using the RANS model for a 15 half-angle diamond injector. The fifteen simulations spanned over five different injection angles and three jet total pressures. In addition to these, two circular injector simulations were also performed. In addition, low pressure normal injection through diamond and circular orifices simulations were performed using DES. Results obtained from CFD were compared to available experimental data. The resulting flow structure and the turbulent properties of the flow were examined in detail. The normal injection case through the diamond-shaped orifice at the lowest jet total pressure was defined as the baseline case and is presented in detail. In order to study the effect of different components of the vorticity transport equation, an in-house code was used post-process the results from the RANS runs.

In the pursuit of developing hypersonic cruise vehicles, unstart is a major roadblock to achieving stable flight. Unstart occurs when a sudden instability in the combustor of a vehicle’s propulsion system creates an instantaneous pressure rise that initiates a shock. This shock travels upstream out of the inlet of the vehicle, until it is ejected from the inlet and creates a standing shockwave that chokes the flow entering the vehicle, thereby greatly reducing its propulsive capability. In severe cases, this can lead to the loss of the vehicle. This thesis presents the results of a computational study of the dynamics of unstart near Mach 5 and presents some possible precursor signals that may indicate its presence in flight. Using SU2, an open-source CFD code developed at Stanford University, the Unsteady Reynolds-Averaged Navier-Stokes equations are used to develop a model for flow in a scramjet inlet-isolator geometry, both in the fully started state and during unstart. The results of these calculations were compared against experimental data collected by J. Wagner, at the University of Texas, Austin. In the present computations, unstart was initiated through the use of an artificial body force, which mimicked a moveable flap used in the experiments. Once the results of the code were validated against these experiments, a selection of parametric studies were conducted to determine how the design of the inlet-isolator by Wagner affected the flow, and thus how generalizable the results can be. In addition, precursor signals indicative of unstart were identified for further study and examined in the different parametric studies. It was found that a thick boundary layer is conducive to a stronger precursor signal and a slower unstart. In addition, an aspect ratio closer to 1:1 promotes flow mixing and reduces the unstart speed and strength. Moreover, an aspect ratio in this range reduces the precursor signal strength but, if a thick boundary layer is present, will smear the signal out over a larger area, potentially making it easier to detect. <br>

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Shock boundary layer interactions (SBLI) is an undesirable event occurring in high-speed air-breathing propulsion system that stimulates boundary layer separation due to adverse pressure gradients and consequently lead to ow distortion and pressure loss in the intake section. Therefore it is essential to apply ow control mechanisms to prevent this phenomenon. This study involves a novel ow control device called micro-ramp, which is a part of the micro-vortex generator family that has shown great potential in solving the adverse phenomenon. The term micro refers to the height of the device, which is smaller than the boundary layer thickness, δ. It is important to highlight the two main novelties of this investigation. Firstly, it is the first micro-ramp study conducted in the hypersonic ow regime (Mach 5) since most of the previous micro-ramp studies were only performed in subsonic, transonic and supersonic flows. Another novelty is the various experimental techniques that were used in single study for example schlieren photography, oil-dot and oil- ow visualisation and conventional pressure transducers. In addition, advanced ow diagnostic tools such as infrared thermography, pressure sensitive paints (PSP) and particle image velocimetry (PIV) were also employed. T