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.

<div>A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel</div><div>at 6° angle of attack, extending several previous experiments on the growth and breakdown of</div><div>stationary crossflow instabilities in the boundary layer. </div><div><br></div><div>Measurements were made using infrared</div><div>imaging and surface pressure sensors. Detailed measurements of the stationary and traveling</div><div>crossflow vortices, as well as various secondary instability modes, were collected over a large</div><div>region of the cone.</div><div><br></div><div>The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was</div><div>adapted for application to crossflow work. It was demonstrated that the roughness elements were</div><div>the primary factor responsible for the appearance of the specific pattern of stationary streaks</div><div>downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness</div><div>insert was created with a high RMS level of normally-distributed roughness to excite the naturally</div><div>most-amplified stationary mode.</div><div><br></div><div>The nonlinear breakdown mechanism induced by each type of roughness appears to be</div><div>different. When using the discrete RIM roughness, the dominant mechanism seems to be the</div><div>modulated second mode, which is significantly destabilized by the large stationary vortices. This</div><div>is consistent with recent computations. There is no evidence of the presence of traveling crossflow</div><div>when using the RIM roughness, though surface measurements cannot provide a complete picture.</div><div>The modulated second mode shows strong nonlinearity and harmonic development just prior</div><div>to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal</div><div>band, leading to a peak in the heating simultaneously with the peak amplitude of the measured</div><div>secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic</div><div>heating pattern is reminiscent of experiments on a flared cone and earlier computations</div><div>of crossflow on an elliptic cone.</div><div><br></div><div>When using the distributed roughness there are several differences in the nonlinear breakdown</div><div>behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,</div><div>which is expected given computational results. Furthermore, the traveling crossflow waves become</div><div>very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there</div><div>appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability</div><div>under the upwelling of the stationary vortices. The traveling crossflow and the secondary</div><div>instability interact nonlinearly prior to breakdown.</div>

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.) - University of Queensland, 2003. / Includes bibliography.

<div>A 7° half-angle cone with a highly-swept fin was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel at 0.0° angle of attack. Previous measurements of the surface heat transfer using temperature sensitive paint revealed heating streaks on the cone surface related to streamwise vortices generated by the fin shock. High-frequency measurements of the cone-surface pressure fluctuations revealed that transition occurs in the streak region at sufficiently-high freestream unit Reynolds numbers under quiet flow. In this work, high-resolution measurements of the surface heat transfer are obtained using infrared thermography and a polyether-ether-ketone wind-tunnel model. In addition, a novel model design made it possible to measure pressure fluctuations throughout the streak region on the cone surface.</div><div><br></div><div>A slender cone with a sharp nosetip and a fin swept back 75° with a 3.18 mm leading-edge radius served as the primary geometry for this work. Two laminar heating streaks</div><div>were measured on the cone surface. These travel along a line of nearly-constant azimuth. A hot spot develops in the streak farthest from the fin, which then moves upstream with increasing freestream Reynolds number. Downstream of this hot spot, the streaks begin to spread in azimuth. The heat transfer along the outer streak shows a threefold increase near the hot spot before decreasing back to nearly two times the laminar streak heating. The amplitude of the pressure fluctuations increases simultaneously with the heat transfer, reaching a peak of nearly 9% of the Taylor-Maccoll pressure for a 7° straight cone. Power spectral densities calculated from these fluctuations demonstrate spectral broadening, which is indicative of boundary-layer transition. Using surface-pressure-fluctuation and heat-flux measurements, transition onset was estimated to occur at an axial length Reynolds number of 2.2×10⁶. Pressure sensors that were rotated through the streak region showed that multiple instabilities amplify between the heating streaks, upstream of the transition onset location. Downstream of transition onset, the highest-amplitude instabilities are localized to the hot spot in the outer streak. The effect of freestream noise on transition was also investigated with this geometry. Under conventional noise levels, transition onset was estimated to occur at an axial length Reynolds number of 0.93×10⁶, and only one instability was measured in the streak region with a frequency similar to the second-mode instability.</div><div><br></div><div>Four configurations were tested to investigate the effect of fin sweep and nosetip bluntness under quiet flow. Fins with 70° and 75° sweep were each tested with nominally sharp and 1-mm-radius nosetips. Increasing fin sweep was shown to move the heating streaks on the cone closer to the fin and to decrease the peak-to-peak spacing of the streaks. In addition, transition onset occurred at lower freestream unit Reynolds numbers for the 70° sweep case. Increasing nosetip radius had little effect on the heating streaks, other than to delay the transition location. A blunt nosetip was shown to delay transition more for the 75° sweep fin as compared to the 70° fin. Similar instabilities were measured for all four of the configurations in this work. The frequency of the instabilities appears to be correlated with the peak-to-peak distance of the heating streaks, which can be viewed as an indirect measurement of the vortex diameter.</div><div><br></div><div>Lastly, the first quantitative measurements of heat transfer on the fin were made using the infrared thermography apparatus. Peak heating on the fin, not including the leading edge, is lower than peak heating rates on the cone. One broad heating streak was measured close to the corner, and smaller low-heating streaks were measured farther outboard. The heating within the streak closest to the corner was shown to agree well with a fully-laminar computed basic state, indicating that the flow on the fin is laminar up to at least 6.31×10⁶ m⁻¹. Using miniaturized Kulite sensors, pressure fluctuations were measured at twelve locations on the fin surface. No obvious conclusions could be drawn from these Kulite measurements, and there is no clear indication that transition occurs on the fin within the maximum quiet</div><div>freestream conditions.</div>

<div>Wind-tunnel experiments with a sharp 7-degree half-angle cone and a 33% scale Boundary Layer Transition (BOLT) model were performed in the Boeing/AFOSR Mach 6 Quiet Tunnel to investigate the effects of forward- and backward-facing steps on boundary-layer instability and transition. Each model was modified to include intentional steps just downstream of the nosetip. Experiments were performed at different freestream Reynolds numbers and varying step sizes. Infrared thermography was used to calculate surface heat transfer, and high-frequency pressure sensors were used to measure pressure fluctuations. A replica measurement technique was used to accurately measure step heights on the BOLT flight vehicle and the wind tunnel model.</div><div><br></div><div>A 7-degree half-angle cone was tested at 0-degree and 6-degree angles of attack. Step heights ranged from 0.610 mm to 1.219 mm. At a 0-degree angle of attack, no significant increases in heat transfer were observed with any of the forward- or backward-facing steps. However, a 250 kHz instability was measured with the forward-facing steps. Growth of the instability was similar to a second-mode. At a 6-degree angle of attack, an increase in heat transfer was observed on the windward ray with the forward-facing steps. Sharp increases in heating rates and increased pressure fluctuations were indications of boundary-layer transition. Elevated heating rates and pressure fluctuations were not measured with the backward-facing steps.</div><div><br></div><div>The BOLT model was tested at 0-degree, 2-degree, and 4-degree angles of attack and 2-degree and 4-degree yaw angles. Step heights ranged from 0.076 mm to 1.016 mm. At a 0-degree angle of attack and 0-degree yaw angle, thin wedges of heating were observed with the backward-facing steps. Instabilities were measured near these wedges of heating and are thought to be caused by a secondary instability. The effects of the steps were magnified on the windward side of the BOLT model at angles of attack. Wedges of heating were wider and more intense. At higher angles of attack, the onset of heating was further upstream. Sensors near and directly underneath the wedges of heating measured pressure fluctuations that were indicative of a turbulent flow. Wedges of heating were also observed at a 4-degree yaw angle, but only with the 1.016 mm backward-facing step.</div>

<div>Many important measurements of low-amplitude instabilities related to hypersonic laminar-turbulent boundary-layer transition have been successfully performed with 1-MHz PCB132 pressure sensors. However, there is large uncertainty in measurements made with PCB132 sensors due to their poorly understood response at high frequency. The current work continues efforts to better characterize the PCB132 sensor with a low-pressure shock tube, using the pressure change across the incident shock as an approximate step input. </div><div> </div><div> New vacuum-control valves provide precise control of pre-run pressures in the shock tube, generally to within 1\% of the desired pressure. Measurements of the static-pressure step across the shock made with Kulite sensors showed high consistency for similar pre-run pressures. Skewing of the incident shock was measured by PCB132 sensors, and was found to be negligible across a range of pressure ratios and static-pressure steps. Incident-shock speed decreases along the shock tube, as expected. Vibrational effects on the PCB132 sensor response are significantly lower in the final section of the driven tube.</div><div> </div><div> Approximate frequency responses were computed from pitot-mode responses. The frequency-response amplitude varied by a factor of 5 between 200--1000 kHz due to significant resonance peaks. Measurements with blinded PCB132 sensors indicate that the resonances in the frequency response are not due to vibration. </div><div> </div><div> Using the approximate frequency response measured with the shock tube to correct the spectra of wind-tunnel data produced inconclusive results. Correcting pitot-mode PCB132 wind-tunnel data removed a possible resonance peak near 700 kHz, but did not agree with the spectrum of a reference sensor in the range of 11--100 kHz. </div>

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