New flux-split algorithms are developed for high velocity, high-temperature flow situations, when finite-rate chemistry and non-equilibrium thermodynamics greatly affect the physics of the problem. Two flux-vector-split algorithms, of the Steger-Warming and of the Van Leer type, and one flux-difference-split algorithm of the Roe type are established and utilized for the accurate numerical simulation of flows with dissociation, ionization, and combustion phenomena. Several thermodynamic models are used, including a simplified vibrational non-equilibrium model and an equilibrium model based upon refined statistical mechanics properties. The framework provided is flexible enough to accommodate virtually any chemical model and a wide range of non-equilibrium, multi-temperature thermodynamic models. A theoretical study of the main features of flows with free electrons, for conditions that require the use of two translational temperatures in the thermal model, is developed. Interesting and unexpected results are obtained, because acoustic wave speeds of the symmetric form u±α no longer appear. A simple but powerful asymptotic analysis is developed which allows the establishment of the fundamental gas-dynamic properties of flows with multiple translational temperatures. The new algorithms developed demonstrate their accuracy and robustness for challenging flow problems. The influence of several assumptions on the chemical and thermal behavior of the flows is investigated, and a comparison with results obtained using different numerical approaches, in particular spectral methods, is provided, and proves to be favorable to the present techniques. Other calculations in one and two space dimensions indicate large sensitivities with respect to chemical and thermodynamic modeling. The algorithms developed are of sufficient generality to begin to examine these effects in detail. Preliminary numerical simulations are performed using elementary modeling of transport phenomena. / Ph. D.

The use of flux-splitting techniques on the Euler equations is considered for high Mach number, high temperature flows in which the fluid is assumed to be inviscid air in equilibrium. Three different versions of real gas extensions to the Steger-Warming and Van Leer flux-vector splitting, and four different versions of real gas extensions to the Roe flux-difference splitting, are compared with regard to general applicability and ease of implementation in existing perfect gas g algorithms. Test computations are performed for the M = 5, high temperature flow over a 10-degree wedge and the M = 24.5 flow over a blunt body. Although there were minor differences between the computed results for the three types of flux-splitting algorithms considered, little variation is observed between different versions of the same algorithm. / Master of Science

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.

Thesis (Ph. D.)--University of Washington, 1990. / Vita. Includes bibliographical references (leaves [100]-105).

Thesis (Ph. D.)--University of Missouri-Columbia, 1996. / Typescript. Vita. Includes bibliographical references (leaves 254-260). Also available on the Internet.

Thesis (Ph. D.)--University of Missouri-Columbia, 1996. / Typescript. Vita. Includes bibliographical references (leaves 254-260). Also available on the Internet.

Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1993. / Vita. Abstract. Includes bibliographical references (leaves 162-167). Also available via the Internet.

Experiments were performed with two different helium injector models at different injector transverse and yaw angles in order to determine the mixing rate and core penetration of the injectant and the flow field total pressure losses. when gaseous injection occurs into a supersonic freestream. Tested in the Virginia Tech supersonic tunnel. with a freestream Mach number of 3.0 and conditions corresponding to a freestream Reynolds number of 5.0 x 107 1m. was a single. sonic. 5X underexpanded, helium jet at a downstream angle of 30° relative to the freestream. This injector was rotated from 0° to _28° to test the effects of injector yaw. The second model was an array of three supersonic, 5X underexpanded helium injectors with an exit Mach number of 1.7 and a transverse angle of 15°. This model was tested in the NASA Langley Mach 6.0, High Reynolds number tunnel, with freestream conditions corresponding to a Reynolds number of 5.4 x 10⁷ /m. The injector array as tested at yaw angles of 0° and -15°. Surface flow visualization showed that significant flow asymmetries were produced by injector yaw. Nanosecond exposure shadowgraph pictures were taken, showing the gaseous injection plume to be unsteady, and further studies demonstrated this unsteadiness was related to shock waves orthogonal to the injectant bow shock, that were generated at a frequency of 30 kHz. The primary data technique used, was a concentration probe which measured the molar concentration of helium in the flow field. Concentration data and other meanflow data was taken at several downstream axial stations and yielded contours of helium concentration, total pressure, Mach number, velocity, and mass flux, as well as the static properties. From these contour plots, the various mixing rates for each case were determined. The injectant mixing rates, expressed as the maximum concentration decay, and mixing distances were found to be unaffected by injector yaw, in the Mach 3.0 experiments, but were adversely affected by injector yaw in the Mach 6.0 experiments. One promising aspect of injector yaw was the that as the yaw angle was increased, lateral motion of the injectant plume became significant, and the turbulent mixing region area increased by approximately 34%. Comparisons of the 15° transverse angled injection into a Mach 6.0 flow to previous experiments with 15° injection into a Mach 3.0 freestream, demonstrated that there is a significant decrease in initial mixing, at Mach 6.0, resulting in a much longer mixing distance. From a parametric computational study of the Mach 6.0 experiments, the effects of adjacent injectors was found to decrease lateral spreading while increasing the vertical penetration of the injectant plume, and marginally increasing the injectant core decay rate. Matching of the computational results to the experimental results was best achieved when using the Baldwin-Lomax turbulence model without the Degani-Schiff modification. / Ph. D.

There has come about, in recent years, a renewed interest in aerospace vehicles operating in the hypersonic regime. With this interest has come a need to not only reestablish the hypersonic test capability that was available in the 1960s but to enhance this capability to meet the demanding needs of today's proposed vehicles. This will require more capable hypersonic wind tunnels with larger test sections, longer run times, and test gases more closely resembling the fluid to be encountered by the vehicle being developed. This document will review the current hypersonic testing capability, examine the operating characteristics of several hypersonic vehicles to develop a set of hypersonic testing requirements, and develop a preliminary design of a required hypersonic facility that addresses the demonstrated requirements. An order of magnitude cost estimate is also presented. / Master of Science

Air may be considered a mixture of diatomic nitrogen and oxygen in which all internal molecular energies including molecular vibration are considered. This leads to an adequate thermodynamic description of air up to dissociation. The thermodynamic and transport properties of this "vibrationally excited" gas are presented and compared with those of a perfect gas (which does not include vibration), and of a dissociating gas in chemical equilibrium. The effects of the vibrationally excited gas on Viscous-Shock-Layer flows are then analyzed and compared for a 7° tangent sphere-cone at zero and five degs angle of attack and at altitudes between 50 and 200 kft. The nose radius is 0.15 ft and the body is 30 nose radii long. The wall temperature and freestream velocity are constant at 2,000 °K and 25,000 ft/sec, respectively. In general, the vibrationally excited gas results are more accurate than perfect gas, and computationally much faster than equilibrium. The vibrationally excited gas also shows potential for use in the nonequilibrium flow regime where the chemical reaction rates are too high for the "stiff" finite-rate equations. This and other areas for additional research are discussed. / M.S.