Thesis (M.S.)--University of Missouri--Rolla, 2007. / Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed March 25, 2008) Includes bibliographical references (p. 95-97).

Thesis (Ph. D.)--University of Washington, 1993. / Vita. Includes bibliographical references (leaves [97]-101).

Thesis (Ph. D.)--Rutgers University, 2010. / "Graduate Program in Mechanical and Aerospace Engineering." Includes bibliographical references (p. 86-90).

Thesis (Ph. D.)--University of Washington, 1995. / Vita. Includes bibliographical references (leaves [76]-82).

Thesis (Ph.D.) - University of Queensland, 2004. / Includes bibliographical references.

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.

A series of experiments were carried out in Purdue University’s Boeing/AFOSR Mach6 Quiet Tunnel, to understand the amplification of streamwise vortices across a separated region in a quiet flow regime. Streamwise vortices were induced on the upstream end of an axisymmetric model consisting of a 7-degree half-angle cone, a cylinder, and a 10-degree flare. The instabilities were seeded using a pre-existing set of roughness inserts, with small, discrete roughness elements. The elements varied in spacing, height, and number of elements. The model was aligned to near 0.0 degree angle of attack. <div><br></div><div>The streamwise, Gortler-like instabilities travelled across the separated region onto the flare, where they were measured with pressure transducers and infrared thermography. The amplification of the instabilities was measured at a variety of Reynolds numbers, under both quiet and conventional noise flow. The results were compared to those of a smooth insert. Heat transfer results showed a streaking pattern, with a peak in heating visible in the streak. Heat flux increased linearly with Reynolds number. If transition was induced, the heat flux would begin to decrease. Power spectral density measurements of the pressure fluctuations indicated that the region within the streak contained two notable instabilities, one between 70 and 150 kHz, and one between 200 and 250 kHz. Transition was only measured in the spectral content in the region on the flare where a ”filling in” of streaks was visible in heat transfer results. Heat flux increased in an nonlinear manner with increasing roughness height. </div><div><br></div><div>The streak positioning and peak heat flux showed a high sensitivity to small, uncontrollable changes in run conditions throughout. Heat transfer results were largely repeatable for small angles of attack, less than 0.1 degrees. The streaks shifted slightly in width and position for angles of attack near 0.1 degrees. Small changes in the streak positioning and heat transfer magnitude were seen in repeatability runs; this is mostly attributable to small changes in initial run conditions. </div>

Hypersonic propulsion has become an increasingly important research field over the past fifty years, and subsequent interest in propulsion systems utilizing supersonic combustion has emerged. Air-breathing engines are desirable for such applications as hypersonic flight vehicles would not need to carry an oxidizer. Therefore, hypersonic air-breathing propulsion systems require an inlet with high mass capture and compressive efficiency. The present work seeks to outline the development and validation of a novel design tool for producing air inlet designs for hypersonic vehicles at variable flight conditions. A Busemann inlet was chosen for its high compressive efficiency, geometric flexibility, and existing experimental validation. The design tool uses the Taylor-Maccoll equation to generate a streamline through a conical flow field. A streamline tracing technique is used to produce three-dimensional inlet surfaces with various capture areas. Additionally, a surface morphing process is implemented to combine inlet profiles for improved engine compatibility. The inlet morphing process allowed for the creation of inlets with offset exit profiles. These offset profiles were evaluated at off-design Mach numbers using Star-CCM+ to quantify efficiency metrics and characterize starting phenomena.

Detonation-based combustion systems are desired for propulsion and power systems due to their ability to provide high thermal efficiency and enable supersonic flight. Detonation combustion in hypersonic flows has traditionally been realized using an oblique detonation wave. However, oblique detonation realization and stabilization in combustion systems is challenging. This communication presents an alternative realization of a detonation mode of combustion through a reacting Mach stem. The detonation is experimentally realized in a hypersonic reacting facility, which is optimized for Mach 5 flow at the combustor inlet and includes a 2D-wedge to stabilize hypersonic reactions at high-enthalpy flow conditions. The Mach stem detonation is analyzed with simultaneous 30 kHz schlieren and chemiluminescence imaging, which reveals the coupling between the Mach stem and the reaction. Further confirmation is provided by comparing the Mach number of the reacting Mach stem with the Chapman-Jouguet (CJ) detonation Mach number. It is found that the Mach number of the reacting Mach stem reaches 94% of the CJ detonation Mach number, confirming that the reacting Mach stem realizes a detonation mode of combustion.

No description available.