Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, December 2003. / Thesis advisor(s): Jose O. Sinibaldi, Garth V. Hobson. Includes bibliographical references (p. 69). Also available online.

Thesis (Ph. D.)--University of Oklahoma, 1989. / Includes bibliographical references (leaves 54-56).

Prior research in the field of boundary layer turbulence has identified streamwise-elongated large-scale turbulence structures in both low speed compressible and high speed (M=2.0) flow. No experimental work has been done in any flow of M> or =3 in an attempt to identify the presence or quantify the behavior of these structures, nor has any study of favorable pressure gradient or surface roughness element effects on these structures been conducted. This research used high-resolution Particle Imaging Velocimetry in a M = 4.9 blow-down wind tunnel accompanied by a series of data analysis in order to identify the existence of streamwise-elongated large-scale turbulence structures in a hypersonic boundary layer. Furthermore, this study identified physical and statistical behavior which suggests that increasing favorable pressure gradient had a substantial impact on both the structural coherence and relative intensity of these turbulent structures at all boundary layer heights tested. This experiment also identified similar effects on these structures in the lower half of the boundary layer as a result of the introduction of surface roughness elements. Finally, several trends were identified between the averaged turbulence statistics and the behavior of the large-scale streamwise-elongated turbulence structures present in this study.

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.

Thesis (Ph. D.)--University of Washington, 1996. / Vita. Includes bibliographical references (leaves [144]-153).

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Though the field of hypersonic vehicle design is thriving again, few studies to date demonstrate the technology through a mission in which multiple flight conditions and constraints are encountered. This is likely due to the highly integrated and sensitive nature of hypersonic vehicle components. Consequently, a formal Mach 6 through Mach 10 flight envelope is explored which includes cruise, acceleration/climb, deceleration/descend and turn mission segments. An exergy approach to the vehicle synthesis/design, in which trade-offs between dissimilar technologies are observed, is proposed and measured against traditional methods of assessing highly integrated systems. A quasi one-dimensional hypersonic vehicle system simulation program was constructed. Composed of two sub-systems, propulsion and airframe, mechanisms for loss are computed from such irreversible processes as shocks, friction, heat transfer, mixing, and incomplete combustion. The propulsion sub-system consists of inlet, combustor, and nozzle, while the airframe provides trim and force accounting measures. An energy addition mechanism, based on the potential of MHD technology, is utilized to maintain a shock-on-lip inlet operating condition. Thirteen decision variables (seven design and six operational) were chosen to govern the vehicle geometry and performance. A genetic algorithm was used to evaluate the optimal vehicle synthesis/design for three separate objective functions, i.e the optimizations involved the maximization of thrust efficiency, the minimization of fuel mass consumption, and the minimization of exergy destruction plus fuel exergy loss. The principal results found the minimum fuel consumption and minimum exergy destruction measures equivalent, both meeting the constraints of the mission while using 11% less fuel than the thrust efficiency measure. Optimizing the vehicle for the single most constrained mission segment yielded a vehicle capable of flying the entire mission but with fuel consumption and exergy destruction plus fuel loss values greater than the above mentioned integrated vehicle solutions. In essence, the mission-level analysis provided much insight into the dynamics of mission-level hypersonic flight and demonstrated the usefulness of an exergy destruction minimization measure for highly integrated synthesis/design. / Master of Science

Accurate flow path modeling of scramjet engines is a key step in the development of an airframe integrated engine for hypersonic vehicles. A scramjet system model architecture is proposed and implemented using three different engine components: the isolator, combustor, and nozzle. For each component a set of intensive properties are iterated to match prescribed conditions, namely the mass flow. These low-fidelity one-dimensional models of hypersonic propulsion systems are used in tandem with Sandia Labs' Dakota optimization toolbox with the goal of accelerating the design and prototyping process. Simulations were created for the various components of the propulsion system and tied together to provide information for the entire flow-path of the engine given an inlet state. The isolator model incorporated methods to compute the intensive properties such as temperature and pressure of the flow path whether a shock-train exists internally as a dual-mode ramjet or if the engine is operating as a pure scramjet with a shock free isolator. A Fanno flow-like model was implemented to determine the friction losses in the isolator and a relation is iterated upon to determine the strength and length of the shock train. Two combustor models were created, the first of which uses equilibrium chemistry to estimate the state of the flow throughout the combustor and nozzle. Going one step further, the second model uses a set of canonical reactors to capture the non-equilibrium effects that may exist in the combustor/nozzle. The equilibrium combustor model was created to provide faster calculations in early iterations, and the reactor model was created to provide more realistic data despite its longer computational time. The full engine model was then compared and validated with experimental data from a scramjet combustor rig. The model is then paired with an optimization toolbox to yield a preliminary engine design for a provided design space, using a finite element analysis to ensure a feasible design. The implemented finite element analysis uses a coarse mesh with simple geometry to reduce computational time while still yielding sufficiently accurate results. The results of the optimization are then available as the starting point for higher fidelity analyses such as 2-D or 3-D computational fluid dynamics. / Master of Science / Ramjets and scramjets are the key to sustained flight at speeds above five times the speed of sound. These propulsion systems pose a challenging simulation environment due to the wide range of flow seen by the system structure. A scramjet simulation model is formulated using a series of combustion models with the goal of accurately modelling the combustion processes throughout the engine. The combustor model is paired with an isolator model and the engine model is compared against previous studies. A structural analysis model is then paired with the engine simulation, and the combined model is used within an optimizer to find an optimum design.