Transition in high-speed boundary layers is investigated using direct numerical simulation (DNS). A compressible Navier-Stokes code that is specifically tailored towards accurate and efficient simulations of boundary layer stability and boundary layer transition was developed and thoroughly validated. Particular emphasis was put into the adoption of a high-order accurate spatial discretization including a boundary closure with the same stencil width as the interior scheme. Oblique breakdown has been shown, using both temporal and spatial DNS, to be a viable route to transition for the boundary layer of the sharp 7° cone at Mach 3.5 investigated by Corke 2002. A 'wedge-shaped' transitional regime was observed to be characteristic for this type of breakdown on the cone geometry. Furthermore, it was shown that the dominance of the longitudinal mode in the nonlinear transition regime of oblique breakdown is due to a continuously nonlinear forced transient growth. That is the primary pair of oblique waves permanently 'seeds' disturbances into the longitudinal mode, where these disturbances exhibit non-modal unstable behavior. In addition to the simulations of controlled transition via oblique breakdown, six simulations have been conducted and analyzed where transition is initiated by multiple primary waves. Despite the broader spectrum of primary waves, typical features of oblique breakdown are still apparent in these simulations and therefore, it may be conjectured, that oblique breakdown initiated by one primary pair of waves is a good model for the nonlinear processes in natural transition. Furthermore, hypersonic boundary layer stability and transition for a flared and a straight cone at Mach 6 was investigated. In particular, a comparative investigation between both geometries regarding the K-type breakdown was performed in order to give some indications towards the open question how strong the nonlinear transition processis altered by the cone flare.

Sharp leading edges required for hypersonic vehicles improve the maneuverability as well as reduce aerodynamic drag. However, due to the sharp design, increased surface temperatures require materials that can withstand these extreme conditions. Ultra-high temperature ceramics are a material group being considered for the leading-edge material, specifically ZrB₂/SiC (ZBS) which has a high thermal shock resistance, melting temperature, and thermal conductivity. Studies done by Tan et. al. has shown that adding samarium (Sm) as a dopant to ZBS has an emittance of 0.9 at 1600^oC and develop oxide scales that have excellent ablation performance. However, it remained unknown how the Sm doped oxide scale formed as well as how the emittance and ablation performance are affected by the microstructure. This study investigates the oxide scale development of 3 mol% doped Sm-ZBS billets as well as how differences in microstructure affect the emittance and ablation performance. Samples were prepared via chemical infiltration of samarium nitrate into spray-dried powders of 80 vol.% ZrB₂/20 vol.% SiC; powders were then pressed into billets and pressureless sintered. Samples cut and polished from these billets were then oxidized for 10, 60, or 300 s, respectively, using an oxyacetylene torch. X-ray diffraction was used to determine the sequence of oxidation of Sm-ZBS, beginning with the formation of ZrO₂ and Sm₂O₃. The final oxide scale was determined to be c₁-Sm_0.2Zr_0.8O_1.9, with a melting temperature exceeding 2500^oC. SEM and EDS were also used to investigate the microstructural formation that occurs from the bursting of convection cells. Samples with different microstructures revealed similar topographical microstructures post-ablation due to the sequence of the oxide formation. However, samples with rougher surfaces and higher porosities had a higher concentration of trapped glass in the cross-sectional oxide scale. It was also found that due to differences in heating the sample during emittance testing compared to ablation testing, the oxide developed was identical for all the samples. It was also found that variances in microstructure had no effect on the spectral emittance of Sm-ZBS at ultra-high temperatures. The fabrication of c₁-Sm_0.2Zr_0.8O_1.9 (SZO) as a bulk billet was also investigated to use as a thermal barrier coating (TBC) in replacement of Sm-ZBS.

<p>Hypersonic flight causes ultra-high surface temperatures which are most intense on sharp leading edges. One way of reducing the surface temperature is to apply a high emittance ceramic (HEC) on the leading edge, increasing the radiation component of heat transfer. An ideal HEC must have a high emittance, while also possessing a strong ablation resistance. From a scientific standpoint, it would be helpful if emittance could be tailored at different wavelengths. For example, materials with tailorable emittance could be used to improve the efficiency of engines, thermo-photo voltaic cells, and other applications. The approach used to create a ceramic with tailorable emittance was to use two different rare-earth elements, adding them to an ultra-high temperature ceramic (UHTC) in small quantities. The samarium element was added to increase the emittance of the UHTC over a large wavelength range (visible to near infrared wavelengths, consistent with the temperature range expected for hypersonic flight), and the erbium element was added to decrease the emittance at specific wavelength ranges. The goal of this study was to create an UHTC with tailorable emittance while maintaining the required ablation resistance. Therefore, ZBS billets with five different Sm to Er ratios and with a nominal total amount of 3 mol.% dopant incorporated were prepared by sintering in vacuum to 2000 °C. The ablation resistance was evaluated by using an oxyacetylene torch and observing at exposure times of 60 s and 300 s, whereas the emittance was evaluated at the Air Force Research Lab facilities via a laser heating testing. The results for the ablation testing showed that ZrB₂-SiC (ZBS) billets co-doped with Sm and Er formed a beneficial <i>c₁</i>-(Sm/Er)_0.2Zr_0.8O_1.9 oxide scale as the majority phase, which is more thermally stable than the <i>m</i>-ZrO₂ oxide scale typically formed in oxidized ZBS systems, resulting in a more adherent oxide scale to the unreacted material. The crystalline oxide scale and the amorphous phase were formed by a convection cell mechanism where the <i>c₁</i>-(Sm/Er)_0.2Zr_0.8O_1.9 crystalline islands precipitate, grow, and coalesce. Moreover, differences in surface temperatures between ZBS samples with different dopant ratios suggest differences in spectral absorptance/emittance between each of the five compositions evaluated. Despite that the emittance profiles with varying Sm:Er molar ratios were similar because <i>m</i>-ZrO₂ was formed as the major oxide phase, the emittance study showed that the erbium oxide influences the emittance profile, as can be noted by the maximum and minimum emittance peaks. Furthermore, results showed that the emittance varies as a function of dopant(s) molar ratios and temperature at shorter wavelength ranges. These changes in the emittance are caused by the different Sm and Er concentration on the surface. Future work should be focused on producing the beneficial <i>c₁</i>-(Sm/Er)_0.2Zr_0.9O_1.8phase directly from the manufacturing process, and therefore, maximize the effect of varying the Sm:Er molar ratios to tailor the emittance. Nonetheless, this study represents the first generation and reported emittance data of UHTC doping ZBS systems with both Sm and Er elements. </p>

Experiments to generate multiple shock waves in an axisymmetric model at hypersonic speeds were conducted in a small reflected shock tunnel. Conical surfaces were used to generate shock waves inside a circular duct chosen to be representative of a scramjet combustor. These shock waves impinged on turbulent boundary layers to produce shock wave/boundary layer interactions (SWBLIs). In the process of observing this phenomenon, the commonly used empirical correlations of Korkegi were tested for accuracy, i.e. the combined pressure ratio across these shocks can be measured and compared to that predicted by these correlations. Korkegi correlates only with Mach number, and is independent of Reynolds number and on how the pressure is applied. A major contribution of this work is to examine how the details of the compression process effect separation. In this study, the history of applying the compression was varied. An analytical method was developed for theoretically estimating the onset of incipient separation using an integrated computation of the momentum flux contained in the boundary layer. By including the summed (negative) contribution of wall shear stress on the integrated momentum flux, the upstream history of the boundary layer was considered. The overall result has a form similar to the Korkegi correlations, plus an additional correction term relating to momentum loss through wall shear stress. The correction term was determined to be a second order effect, which explains why the Reynolds number independent Korkegi correlations work so well over such a large range of conditions. A hypersonic flow test condition conducive to the generation of high Reynolds number flows and turbulent boundary layer production was developed in a small reflected shock tunnel. The experimentally measured flow parameters were matched by numerical simulation using a number of in-house codes at The University of Queensland. This has allowed the unmeasured parameters which are numerically derived to be stated with greater confidence. An internal centre-body with a conical forebody was used to generate conditions of incipient separation. This provided benchmark data for comparison with subsequent experiments with multiple compressions. A semi-vertex angle of 15o was selected based on Large Eddy Simulation (LES) numerical results once the experimental and numerical static wall pressure and heat flux were matched. A two-cone experimental model, which provided for adjustment of the axial separation between the two shock systems, was tested at the same flow conditions as used in the single-cone experiments. A technique of incrementally moving the instrumentation (relative to the centre-body) and repeating the same condition to achieve high resolution in pressure and heat flux distributions with a limited number of transducers was successful. The results verified that it was possible to subject a hypersonic turbulent boundary layer to two quantified compression-expansion systems with an adjustable axial separation between them and capture the first reflected shock in a “shock trap” to remove it's influence from the second SWBLI. The data from this initial two-cone model provided non-separated pressure and heat flux data which was used as a reference to help interpret data from separated flows. The commercially available Reynolds Averaged Navier-Stokes (RANS) numerical code, CFD-Fastran, was used to help design an experimental model which produces boundary layer separation. Algebraic and two-equation turbulence models were applied to a modified two-cone model to show greater pressure rises which would produce boundary layer separation. A modified two-cone model was tested and demonstrated boundary layer separation. Three configurations with varying axial separation between SWBLIs were tested which all produced separation. The configuration that produced the largest pressure ratio and largest separation region at the second SWBLI may represent a geometry whereby the distance from the hollow cylinder inlet and the second cone may represent a critical value. The amount of viscous interaction, generated from the leading edge of the shock trap, and the proximity of the two interactions may be coupled to produce higher than expected values. It is postulated that the boundary layer momentum recovery for the configuration where the second SWBLI was furthest downstream (30 mm configuration), prevented severe separation from occurring. An in-house RANS code, elmer3, was used to simulate the flow of the modified two-cone model. An algebraic turbulence model was applied to this model and comparisons of experimentally measured static wall pressure and heat flux have given good agreement. The wall shear stress was investigated to provide further information concerning the position and size of flow reversal regions. The use of the numerical codes utilised in this study has reinforced their effectiveness for model design and comparison of experimental results.

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.

Tunable diode laser absorption spectroscopy using an external cavity diode laser operating in the infra-red has been developed to monitor CO2 in the freestream of the Longshot hypervelocity facility at the Von Karman Institute for Fluid Dynamics. The Longshot facility offers a unique European facility for ground testing and numerical validation applications, however, some of the traditional data rebuilding aspects are in question. A non-intrusive absorption sensor could significantly aid in improving the knowledge of freestream static values thereby improving the models used in data rebuilding and numerical simulation. The design of such a sensor also expands the spectroscopic capabilities of the Von Karman Institute. The absorption sensor is designed around the single P12 (00001)-(30013) rovibrational transition near 1.6µm (6218.09cm-1 specifically) which yields relatively weak direct absorption levels at about 3.5% per meter for typical Longshot freestream conditions. However, when handled carefully, adequate signal-to-noise can be acquired to exploit significant flow information. By being able to operate in this range, total sensor cost can be easily an a factor of two or more cheaper than sensors designed for the deeper infrared. All sensor elements were mounted to a compact portable optics bench utilizing single-mode optical fibers to allow for quick installation at different facilities by eliminating tedious optical realigning. Scans at 600Hz were performed over 20ms of the 40ms test time to extract core static temperature, pressure and velocity. These results are compared with the current state of the Longshot data rebuild method. The non-uniform flow properties of the shear layer and test cabin rested gas accumulation was of an initial concern. The temperature and density gradients along with significant radial velocity components could result in DLAS temperature, pressure and velocity that are significantly different than that of the target freestream inviscid core values. Fortunately, with the proper selection of the P12 rotational number, this effect could be more or less ignored as the higher temperature and lower density gas of this region is relatively transparent. Ultimately, acquired temperature and density were moderately accurate when compared to Longshot rebuilt results owing primarily to the baseline extraction which poses issues for such low absorption signals. However, the extracted velocity data are quite accurate. This is a definite puls for the sensor as the freestream enthalpy of cold hypersonic facilities is dictated primarily by the kinetic energy contribution. Being able to compare velocity gives insight to the level of vibration non-equilibrium in the flow. The velocity of the DLAS and the Longshot rebuild are quite close. This adds more weight to the argument that vibrational excitation is very low (if present at all) in the free stream and that the van de derWaals equation of state usage and constant specific heat assumption might be an adequate model for the data rebuild after all.

A parallel implicit adaptive-mesh-refinement scheme is proposed for the solution of the Navier-Stokes equations as applied to two-dimensional steady-state hypersonic laminar flows in conjunction with an equilibrium high-temperature equation of state. A finite-volume discretization is applied to the governing equations. Limited piecewise-linear solution reconstruction and Riemann solvers (Roe and HLLE, both modified for a general equation of state) are used to evaluate the inviscid fluxes. The gradients in the viscous fluxes are calculated using diamond-path reconstruction. The system of non-linear algebraic equations resulting from the finite-volume discretization are solved using an inexact Newton method with GMRES to solve the update step of the Newton method. GMRES is preconditioned with Schwarz preconditioning with local block-fill incomplete lower-upper factorization. Multigrid and pseudo-transient continuation are used for startup. Numerical results, including flows at Mach numbers of 7.0, are discussed and demonstrate the validity and efficiency of the scheme.

A parallel implicit adaptive-mesh-refinement scheme is proposed for the solution of the Navier-Stokes equations as applied to two-dimensional steady-state hypersonic laminar flows in conjunction with an equilibrium high-temperature equation of state. A finite-volume discretization is applied to the governing equations. Limited piecewise-linear solution reconstruction and Riemann solvers (Roe and HLLE, both modified for a general equation of state) are used to evaluate the inviscid fluxes. The gradients in the viscous fluxes are calculated using diamond-path reconstruction. The system of non-linear algebraic equations resulting from the finite-volume discretization are solved using an inexact Newton method with GMRES to solve the update step of the Newton method. GMRES is preconditioned with Schwarz preconditioning with local block-fill incomplete lower-upper factorization. Multigrid and pseudo-transient continuation are used for startup. Numerical results, including flows at Mach numbers of 7.0, are discussed and demonstrate the validity and efficiency of the scheme.

Ballute hypersonic aerodynamic decelerators have been considered for aerocapture since the early 1980's. Recent technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. Several technology hurdles have been identified, including the effects of coupled fluid structure interaction on ballute performance and survivability. To date, no aeroelastic solutions of thin-film ballutes in an environment relevant to aerocapture have been published. In this investigation, an aeroelastic solution methodology is presented along with the analysis codes selected for each discipline. Variable-fidelity aerodynamic tools are used due to the long run times for computational fluid dynamics or direct simulation Monte Carlo analyses. The improved serial staggered method is used to couple the disciplinary analyses in a time-accurate manner, and direct node-matching is used for data transfer. In addition, an engineering approximation has been developed as an addition to modified Newtonian analysis to include the first-order effects of damping due to the fluid, providing a rapid dynamic aeroelastic analysis suitable for conceptual design. Static aeroelastic solutions of a clamped ballute on a Titan aerocapture trajectory are presented using non-linear analysis in a representative environment on a flexible structure. Grid convergence is demonstrated for both structural and aerodynamic models used in this analysis. Static deformed shape, drag and stress level are predicted at multiple points along the representative Titan aerocapture trajectory. Results are presented for verification and validation cases of the structural dynamics and simplified aerodynamics tools. Solutions match experiment and other validated codes well. Contributions of this research include the development of a tool for aeroelastic analysis of thin-film ballutes which is used to compute the first high-fidelity aeroelastic solutions of thin-film ballutes using inviscid perfect-gas aerodynamics. Additionally, an aerodynamics tool that implements an engineering estimate of hypersonic aerodynamics with a moving boundary condition is developed and used to determine the flutter point of a thin-film ballute on a Titan aerocapture trajectory.

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.