Rocket plume flowfields have an importance due to their influence on the signature of the rocket and also on the distribution of the plume gases around the vehicle. Little information on the co-flowing situation exists other than a previous study at Oxford. This thesis thus represents a significant database for co-flowing rocket plumes of this form. The work presented deals with two new aspects of co- flowing rocket plumes in that detailed flowfield measurements have been made and plumes from twin nozzle have been investigated for the first time in this thesis. This study on twinjet rocket plumes was carried out using the University of Oxford Gun Tunnel. Twinjet rockets with nozzle exit Mach numbers of 3 and 5 were tested in quiescent and in co-flow at Mach 7 using nitrogen and hydrogen injections. A major feature of the twinjet case was the so-called impingement shock between the flows from the two nozzles. It was discovered that this shock was insensitive to the freestream and scaling parameters are suggested for its geometry. Comparisons with single equivalent thrust nozzles are made at downstream locations and similar Pitot pressure profiles were observed for nitrogen injection in a nitrogen freestream after approximately 3 nozzle diameters downstream. Shear layers were studied and fluctuations in this region were measured by fast-response Pitot pressure and heat transfer probes sampled at 1.1 MHz. The extent of the shear layer was deduced using a new Oxford Total Temperature Probe. With the freestream stagnation temperature at approximately 650 K and injected gas at 350 K, a linear variation for the deduced total temperature across the shear layer was obtained. This was consistent with the Pitot pressure variations across this region. Convective heat transfer coefficient fluctuations and flow total temperature fluctuations across rocket flowiields were obtained using three thin-film heat transfer probes and found to be closely correlated. Experimental results for the twinjet and the single jet were compared with CFD simulations and good overall agreements were achieved. Instrumentation for the hypersonic experiments was investigated and a fast-response (~ 20 kHz) Pitot probe suited for flows heavily contaminated with particulate was developed and tested.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1993. / Vita. Also available via the Internet.

Thesis (Ph.D.) - University of Queensland, 2003. / Includes bibliography.

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

abstract: This thesis examines the modeling, analysis, and control system design issues for scramjet powered hypersonic vehicles. A nonlinear three degrees of freedom longitudinal model which includes aero-propulsion-elasticity effects was used for all analysis. This model is based upon classical compressible flow and Euler-Bernouli structural concepts. Higher fidelity computational fluid dynamics and finite elementmethods are needed formore precise intermediate and final evaluations. The methods presented within this thesis were shown to be useful for guiding initial control relevant design. The model was used to examine the vehicles static and dynamic characteristics over the vehicles trimmable region. The vehicle has significant longitudinal coupling between the fuel equivalency ratio (FER) and the flight path angle (FPA). For control system design, a two-input two-output plant (FER - elevator to speed-FPA) with 11 states (including 3 flexible modes) was used. Velocity, FPA, and pitch were assumed to be available for feedback. Propulsion system design issues were given special consideration. The impact of engine characteristics (design) and plume model on control system design were addressed.Various engine designs were considered for comparison purpose. With accurate plume modeling, effective coupling from the FER to the FPA was increased, which made the peak frequency-dependent (singular value) conditioning of the two-input two-output plant (FER-elevator to speed-FPA) worse. This forced the control designer to trade off desirable (performance/robustness) properties between the plant input and output. For the vehicle under consideration (with a very aggressive engine and significant coupling), it has been observed that a large FPA settling time is needed in order to obtain reasonable (performance/ robustness) properties at the plant input. Ideas for alleviating this fundamental tradeoff were presented. Plume modeling was also found to be particularly significant. Controllers based on plants with insufficient plume fidelity did not work well with the higher fidelity plants. Given the above, the thesismakes significant contributions to control relevant hypersonic vehicle modeling, analysis, and design. / Dissertation/Thesis / M.S. Electrical Engineering 2010

<p dir="ltr">Hypersonic flight is the key to unlocking a nation’s strategic advantage in this century’s military theater. Military powerhouses such as the United States, Russia, India, China, Australia, and the EU publicly possess hypersonic weapons capabilities. Such technology enables intercontinental travel orders of magnitude faster than conventional flights. A trip halfway across the world would take not twenty hours, but two. However, the level of thermal and chemical load the aircraft and these electronic equipment experience while at such high speeds cause them to fail. Thus, ceramic window materials are used to act as a barrier between the hypersonic flight environment and this sensitive electronic equipment. Such materials need to be both mechanically robust, but transparent within the relevant infrared ranges used for target detection. Single-crystal sapphire (alumina) is an infrared window material readily available, plentiful, and easy to microstructurally control and manufacture, but not optimal. Its transparency range is limited to the optical and near-infrared, while it exhibits poor mechanical and dielectric strength. Polycrystalline alumina (PCA) has recently been shown to possess more favorable infrared window characteristics as opposed to its single-crystal counterpart. This is achieved by processing using a platelet powder morphology in a single processing step – hot-pressing. Full densification (> 99.5%) of PCA samples was achieved, demonstrating maximum of 84% optical transparency, but accompanied by grain growth (60+µm), resulting in lower mechanical strength. This research thus works on a two-fold approach to minimizing the grain growth of PCA. Optical tests demonstrated favorable results for lowering isothermal temperatures to reduce grain growth. Weibull values of m = 28.8 and m = 9.7 from 4 point-flexure tests were obtained (ASTM 1161a). Thermal loading via ablation testing compared PCA samples to industry alternatives (single-crystal sapphire) and (equiaxed alumina). Ablation tests revealed the benefit of polycrystalline alumina over sapphire. The benefit of lower isothermal sintering temperatures for reduced grain growth resulted in higher peak load before failure, resulting in greater characteristic strength and minimal transmission lost during a minute of oxyacetylene heat flux exposure. Finally, additional work was done on nanoceramic MgO-Y₂O₃, in a ceramic-processing method like that of PCA. These findings will also be discussed.</p>

<div>High speed flow problems of practical interest require a solution of nonequilibrium aerothermochemistry to accurately predict important flow phenomena including surface heat transfer and stresses. As a majority of these flow problems are in the continuum regime, Computational Fluid Dynamics (CFD) is a useful tool for flow modeling. This work presents the development of a nonequilibrium add-on solver to ANSYS Fluent utilizing user-defined-functions to model salient aspects of nonequilibrium flow in air. The developed solver was verified for several benchmark nonequilibrium flow problems and compared with the available experimental data and other nonequilibrium flow simulations. <br></div><div><br></div><div>The rate of dissociation behind a strong shock in thermochemical nonequilibrium depends on the vibrational excitation of molecules. The Macheret-Fridman (MF) classical impulsive model provides analytical expressions for nonequilibrium dissociation rates. The original form of the model was limited to the dissociation of homonuclear molecules. In this work, a general form of the MF model has been derived and present macroscopic rates applicable for modeling dissociation in CFD. Additionally, some improvements to the prediction of mean energy removed in dissociation in the MF-CFD model has been proposed based on the comparisons with available QCT data. In general, the results from the MF-CFD model upon investigating numerous nonequilibrium flows are promising and the model shows a possibility of becoming the standard tool for investigating nonequilibrium flows in CFD.</div><div><br></div><div>The aerodynamic deorbit experiment (ADE) CubeSat has dragsail to accompany accelerated deorbiting of a CubeSat post-mission. A good estimation of the aerothermal load on a reentry CubeSat is paramount to ensure a predictable reentry. This study investigates the aerothermal load on an ADE CubeSat using the direct simulation Monte Carlo (DSMC) methods and Navier-Stokes-Fourier continuum based methods with slip boundary conditions. The aerothermal load on an ADE CubeSat at 90 km altitude from the DSMC and continuum methods were consistent with each other. The continuum breakdown at a higher altitude of 95 km resulted in a strong disagreement between the continuum and DSMC solutions. Overall, the continuum methods could offer a considerable computational cost saving to the DSMC methods in predicting aerothermal load on an ADE CubeSat at low altitudes.<br> </div>

<p> Hypersonic boundary-layer transition greatly affects aerodynamic heating, skin friction, aircraft stability and other characteristics on flight vehicles. Understanding the factors leading to laminar-turbulent transition is pivotal in hypersonic aircraft design. Various instabilities and modes may facilitate transition at hypersonic speeds including first and second-mode waves, Görtler vortices, and cross-flow which may be stationary or traveling. The research presented here will focus on investigating traveling cross-flow instabilities on a 7° half-angle cone at 6° angle of attack. The experiments were conducted in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) at Purdue University. The low freestream noise of the quiet tunnel facility made it ideal for studying boundary layer transition due to its more, ”flight like” environment when compared to traditional tunnel environments. Previous experiments by Ryan Henderson, Chris Ward, and Joshua Edelman focused on studying the cross-flow instability on right circular cones at angle of attack (AoA) in the BAM6QT. From these experiments it was decided that a means for taking off-surface pressure measurements on a cone was needed. This work sets out to create a micro-pitot traverse system capable of doing such. The system is able to measure pressure fluctuations within the boundary layer of cone models at precise axial, azimuthal and wall-normal locations. The design for the traverse was based off a traverse used at Notre Dame which was designed by David Cavalieri in his PhD dissertation for Illinois Institute of Technology. Micro-pitot probes created using hypodermic tubing and Kulite sensors were created to attach to the end of the traverse and take pressure measurements. The micro-pitot probes were placed such that they formed two distinct spatial pairs capable of measuring both the phase speed and propagation angle of traveling cross-flow instabilities using the difference in time of arrival of the traveling instability between the sensor pairs. The micro-pitot probes developed were made from telescoped hypodermic tubes housing Kulite XCE-061-15A sensors. The telescoped tubing assembly caused attenuation at higher frequencies affecting the micro-pitot probes ability to measure pressure fluctuations at higher frequencies. It was necessary to increase the dynamic performance of the micro-pitot probes in order to capture the cross-flow instability. To accomplish this a custom built frequency 17 compensator was designed to correct for this attenuation. The process for designing the compensator utilized a Mach 4 supersonic jet system (SSJ) to estimate a transfer function model for the tubing assembly. This was done by comparing the spectral content of an untubed Kulite sensor and a micro-pitot sensor in the SSJ. The transfer function model was then used to develop the compensator improving measurements made with the micro-pitot up to 50 kHz. The micro-pitot traverse system was then used in a series of tests in the BAM6QT to validate its ability to function as designed. The traverse needed to provide a rigid platform for the micro-pitot probes during tunnel operation. The deflection of the pitot head was recorded using a shadowgraph system. This allowed real time measurements for the deflection of the pitot head during tunnel operation to be taken. These measurements were compared to theoretical calculations to ensure deflections were within acceptable limits. Also, of key importance was the survivability of the traverse system after repeated runs in the BAM6QT. This focused on the ability of the traverse to continue providing movement in all three-directions and its ability to resist wear in the tunnel environment. The only cause for concern noted over the course of three tunnel entries centered around the motor used for wall-normal movement. This motor suffered repeated damage impairing the traverses ability to function as intended. Observations regarding this issue and solutions implemented to mitigate the impact of this damage are discussed. Finally, the micro-pitot was combined with the traverse system and used in conjunction with surface mounted sensors on a axisymmetric cone to measure traveling cross-flow instabilities. Damage to Kulites needed for the micro-pitot prohibited three sensors from being used in the tunnel. For this reason only propagation angles and phase speed calculations for traveling cross-flow waves were calculated using the surface mounted sensors. However, one micro-pitot sensor was used to measure spectral content near the surface mounted sensors. The spectral content of the micro-pitot was compared to the surface mounted sensors in order to validate that the micro-pitot could measure the desired instability once more are acquired </p>

<p dir="ltr">Prediction of boundary-layer transition is a critical element of hypersonic vehicle design</p><p dir="ltr">due to the impact transition has on boundary-layer separation, heat transfer, and aerodynamic</p><p dir="ltr">control. Transition is affected by many factors including surface roughness. The</p><p dir="ltr">roughness on a hypersonic vehicle can cause a boundary-layer to become turbulent. However,</p><p dir="ltr">there is a limited understanding of the impacts that roughness can have, and the conditions</p><p dir="ltr">under which it is important.</p><p dir="ltr">The rocket-sled track at Holloman Air Force Base was selected as a ground-test facility</p><p dir="ltr">for transition measurements. The present work is about understanding the mechanism of</p><p dir="ltr">transition on blunt ogives or blunt cones with moderate nose radii, as it appears that nosetip</p><p dir="ltr">roughness affects boundary-layer transition on the afterbody for moderate nose radii. A</p><p dir="ltr">single test-track shot is to be executed for a blunt ogive to determine if the test track can</p><p dir="ltr">make useful measurements of boundary-layer transition.</p><p dir="ltr">Initially, the present research used a boundary-layer solver to estimate target roughnesses</p><p dir="ltr">that would be applied to the nose tip. Preliminary analysis was conducted on test cases for</p><p dir="ltr">sharp cones and blunt cones. However, due to time constraints, the target roughnesses were</p><p dir="ltr">then estimated with a higher fidelity code by Brad Wheaton of JHU APL. Two separate</p><p dir="ltr">roughness targets were established for the upper and lower sides of the hemispherical nosetip.</p><p dir="ltr">The focus of this work then shifted to measurements of the roughness that was applied</p><p dir="ltr">by others to the hemisphere nose tip for a blunt ogive. Utilizing the Zygo ZeGage 3D optical</p><p dir="ltr">profiler, roughness scans were collected both directly under the profiler head and indirectly</p><p dir="ltr">using rubber molds. Profilometer measurements were also recorded. Twelve iterations were</p><p dir="ltr">completed to allow the polisher to develop appropriate procedures for applying the roughness,</p><p dir="ltr">given the material and curvature. The first five iterations involved roughness applied to</p><p dir="ltr">cylindrical-shaped test areas. After achieving the target roughnesses on these test areas,</p><p dir="ltr">the hemispherical ends of test specimens were then polished and measured until both the</p><p dir="ltr">rough and smooth halves met the roughness target. During this time, the three roughness measurement</p><p dir="ltr">techniques were refined until good agreement was reached between them. When the roughness-application and </p><p dir="ltr">roughness-measurement techniques were sufficiently mature,</p><p dir="ltr">the actual blunt-ogive nose tip was then polished until the roughness target was achieved.</p>

<p dir="ltr">Experiments were performed in the Boeing/AFOSR Mach-6 Quiet Tunnel to study the effect of flight-derived discrete roughness elements repeated in an axisymmetric pattern near the nose of a sharp 7° cone. The aim of the roughness array was to simulate natural vehicle roughness and attempt to introduce a deterministic roughness pattern with the ability to cancel out the instabilities caused by the natural roughness. The cone was pitched at a 6° of attack to determine the three-dimensional flow field effects of the roughness elements. Tests were also ran at 0° of attack for comparison. Quiet flow testing included the designed-for freestream unit Reynolds number of 10.8x10⁶, and Reynolds numbers above and below. In noisy flow, comparable Reynolds numbers were also tested at to isolate the effects of noise in a conventional flow wind tunnel.</p><p dir="ltr">Infrared thermography and surface pressure sensors were used to document the behavior of the boundary layer. It was found that the roughness pattern was in general unsuccessful in controlling the added boundary layer instabilities as intended at 6° of attack, but it did create different instability amplitudes and heating patterns. Additionally, it was determined to reduce Mack's second-mode instability amplitudes at 0° of attack.</p><p dir="ltr">Additionally, work was done to document and characterize the roughness patterns found on samples of hypersonic glide vehicles PRIME (SV-5D or X-23) and ASSET (ASV-3). These samples were taken in the form of molded impressions of the surface which were able to be analyzed with an optical profilometer and considered for future experimental distributed roughness studies.</p>