Since the 1970's, NASA has relied on the use of rigid aeroshells and supersonic parachutes to enable robotic mission to Mars. These technologies are constrained by size and deployment condition limitations that limit the payload they can deliver to the surface of Mars. One candidate technology envisioned to replace the supersonic parachute is the supersonic inflatable aerodynamic decelerator (IAD). This dissertation presents an overview of work performed in maturing a particular type of IAD, the tension cone. The tension cone concept consists of a flexible shell of revolution that is shaped so as to remain under tension and resist deformation. Systems analyses that evaluated trajectory impacts of a supersonic IAD demonstrated several key advantages including increases in delivered payload capability of over 40%, significant gains in landing site surface elevation, and the ability to accommodate growth in the entry mass of a spacecraft. A series of supersonic wind tunnel tests conducted at the NASA Glenn and Langley Research Centers tested both rigid and flexible tension cone models. Testing of rigid force and moment models and pressure models demonstrated the new design to have favorable performance including drag coefficients between 1.4 and 1.5 and static stability at angles of attack from 0º to 20º. A separate round of tests conducted on flexible tension cone models showed the system to be free of aeroelastic instability. Deployment tests conducted on an inflatable model demonstrated rapid, stable inflation in a supersonic environment. Structural modifications incorporated on the models were seen to reduce inflation pressure requirements by a factor of nearly two. Through this test program, this new tension cone IAD design was shown to be a credible option for a future flight system. Validation of CFD analyses for predicting aerodynamic IAD performance was also completed and the results are presented. Inviscid CFD analyses are seen to provide drag predictions accurate to within 6%. Viscous analyses performed show excellent agreement with measured pressure distributions and flow field characteristics. Comparisons between laminar and turbulent solutions indicate the likelihood of a turbulent boundary layer at high supersonic Mach numbers and large angles of attack.

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No. of bitstreams: 1 CAMARGO_Janine_2017.pdf: 19878018 bytes, checksum: 517c1918bf6801b10fe87b0d806d8a50 (MD5) Previous issue date: 2017-03-14 / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES) / Cellulose is a biopolymer available in abundance in nature, which presents very interesting properties mainly for the textile, packaging and paper industry. However, due to its strongly hydrophilic character, resulting from the presence of a large number of free hydroxyl groups in its molecule, its use in certain areas becomes limited. In order to promote the modification of the wetting characteristic of the cellulose, and make it superhydrophobic, samples of this material were submitted to plasma processes performed in two steps: ablation and film deposition. Initially, the effect of variation of ablation time on the creation of adequate surface topography was studied. For this, the samples were exposed to the oxygen plasma at a pressure of 13 Pa and power of 150 W, varying the treatment time from 5 to 60 minutes. Then, the treated samples were submitted to the process called PECVD (Plasma Enhanced Chemical Vapor Deposition), in which a film was deposited on the surface from the precursor hexamethyldisiloxane (HMDSO) in the presence of argon, in a ratio of 70 e 30%, respectively. The deposition time was set to 30 minutes and the power applied was 150 W. In a second investigation, the effect of the thickness of the deposited film was studied. In this case, the cellulose samples were previously exposed to the oxygen ablation plasma for 60 minutes and then led to the PECVD process, with the deposition time varying from 5 to 30 minutes. In a third investigation, the duration of the ablation and deposition steps was reduced to 30 and 1 minute, respectively, in order to optimize the methodology. The samples were characterized by Scanning Electron Microscopy (SEM), Energy Dispersive Spectroscopy (EDS), Fourier Transform Infrared Spectroscopy (FTIR), profilometry, static contact angle measurements and hysteresis and sliding angle measurements. The results demonstrated that longer exposure times for the ablation step provide a significant modification in the material, through the creation of nanoscale structures on the surface of the fibers. The deposition step promoted the formation of films with organosilicon character on the cellulose surface, with thicknesses varying between 144 and 910 nm. The combination of the topography created after 30 minutes of ablation with the thin film resulting from the deposition step for 1 minute provided a surface with low receptivity for both polar (? = 150°) and nonpolar (? ~ 120°) liquids, whose wettability property remained stable with treatment aging time. The low values obtained for hysteresis (9°) and sliding angle (7°) confirm the creation of a superhydrophobic roll-off surface. / A celulose é um biopolímero disponível em abundância na natureza e que possui propriedades bastante interessantes principalmente para a indústria têxtil, de embalagens e papel. No entanto, devido ao seu caráter fortemente hidrofílico, proveniente da presença de um grande número de grupos hidroxila livres em sua molécula, sua utilização em determinadas áreas se torna limitada. Com o objetivo de promover a modificação da característica de molhabilidade da celulose, de modo a torná-la superhidrofóbica, amostras deste material foram submetidas a processos a plasma realizados em duas etapas: ablação e deposição de filme. Inicialmente, estudou-se o efeito da variação do tempo de ablação na criação da topografia adequada da superfície. Para isto, as amostras foram expostas ao plasma de oxigênio a uma pressão de 13 Pa e 150 W de potência, variando-se o tempo de tratamento de 5 a 60 minutos. Em seguida, as amostras tratadas foram submetidas ao processo denominado PECVD (Plasma Enhanced Chemical Vapor Deposition), no qual foi depositado um filme sobre a superfície das mesmas, a partir do precursor hexametildisiloxano (HMDSO) na presença de argônio, numa proporção de 70 e 30%, respectivamente. O tempo de deposição foi fixado em 30 minutos e a potência aplicada foi de 150 W. Numa segunda investigação, foi estudado o efeito da espessura do filme depositado. Neste caso, as amostras de celulose foram previamente expostas ao plasma de ablação com oxigênio durante 60 minutos e posteriormente ao processo de PECVD, variando-se o tempo de deposição de 5 a 30 minutos. Numa terceira investigação, o tempo de duração das etapas de ablação e deposição foi reduzido para 30 e 1 minuto, respectivamente, com o intuito de otimizar a metodologia. As amostras foram caracterizadas por Microscopia Eletrônica de Varredura (MEV), Espectroscopia de Energia Dispersiva (EDS), Espectroscopia de Absorção no Infravermelho por Transformada de Fourier (FTIR), perfilometria, medição de ângulo de contato estático e medição de histerese e ângulo de deslizamento. Os resultados demonstraram que maiores tempos de exposição à etapa de ablação proporcionam uma modificação significativa no material, por meio da criação de estruturas em nanoescala na superfície das fibras. A etapa de deposição promoveu a formação de filmes de caráter organosilicone sobre a superfície da celulose, com espessuras variando entre 144 e 910 nm. A combinação entre a topografia criada após 30 minutos de ablação e o filme de menor espessura resultante da etapa de deposição durante 1 minuto, possibilitaram a obtenção de uma superfície de baixa receptividade tanto a líquidos polares (? =150°) quanto apolares (? ~120°), cuja propriedade de molhabilidade se manteve estável com o tempo de envelhecimento. Os baixos valores de histerese (9°) e ângulo de deslizamento (7°) obtidos confirmam a criação de uma superfície superhidrofóbica do tipo “roll-off”.

Quotidiennement, des objets orbitant à proximité de la Terre (ou géocroiseurs) impactent cette dernière. Lorsque la dimension de l'objet atteint une taille critique (autour de 50m de diamètre),les conséquences au sol peuvent devenir dramatiques.De plus, ces objets ont une occurrence d'impact séculaire, donc à l'échelle d'une vie humaine. L'entrée d'un tel objet met en œuvre de nombreux phénomènes, parfois peu ou pas connus de manière précise : AéroThermoDynamique (ATD) de l'écoulement, rayonnement, ablation, fragmentation. La grande variété de conditions d'entrée étudiées nécessite de plus une étude paramétrique approfondie. Notre thèse est que la phase de rentrée et les phénomènes s’y déroulant jouent un rôle fondamental dans la prévision des risques d'impact au sol. Ainsi, nous avons quantifié ces phénomènes afin d'en établir leurs conséquences pendant la rentrée puis au sol : Nombre et tailles des fragments, empreinte au sol, vitesse(s), masse(s) et énergie cinétique finales. Des simulations ATD préliminaires ont permis de voir que l'écoulement post-choc était en équilibre thermochimique et rayonnait de façon importante. De ce fait des calculs de rayonnement au niveau de la ligne d'arrêt pour différents points de vol ont été effectués, en vu de développer une loi analytique permettant d’estimer correctement le flux radiatif pour nos conditions d’entrée. Cette étude a mis en défaut la représentativité des formules analytiques pré-existantes pour les conditions considérées ici. Du fait du flux thermique incident, un géocroiseur perd de la masse par ablation. Deux modélisations de ce phénomène ont été réalisées, afin d'en évaluer l'incidence en terme de pertes de masses et changements de forme, et donc sur la trajectoire. Nous avons également modélisé le phénomène de fragmentation, de l'initiation de la rupture du fait des contraintes mécaniques à la génération de fragments et à leur dynamique d'évolution. Cette étude a montré l'importance de ce phénomène sur la prévision d'impact, en particulier sur le nombre de fragments impactant et leur énergies cinétiques d'impact. De plus, les interactions entre fragments réduisent la dispersion au sol.Enfin des simulations de trajectoires 1D et 3D avec modélisations de l’ ablation et la fragmentation ont été effectuées sur 3 exemples d'entrée. Elles ont mis en évidence l'importance des paramètres d'entrée (vitesse et incidence en particulier) dans l'estimation de l'impact au sol, et démontré l'influence protectrice de l'atmosphère dans l'estimation des conséquences au sol, du fait en particulier du phénomène de fragmentation, et dans une moindre mesure d'ablation. / Near Earth Objects (NEOs) impact Earth everyday. When the objet reaches a critical size (>50m), ground consequences might be dramatic. Moreover, NEOs have a secular occurrence, i.e. at a human scale. A NEO entry object involves various phenomena, poorly or not known: flow AeroThermoDynamics (ATD), radiation, ablation, fragmentation. The variety ofstudied entry conditions implies also an extensive parametric study. My thesis is that the entry and the phenomena that take place in this phase has a crucial role in the prediction of impact consequences. That why I have quantified these phenomena in order to assess their consequences on the ground impact: number and sizgg of the fragments, ground print, velocity, mass and kinetic energy. ATD simulations showed the aftershock flow was in thermochemical equilibrium, and highly radiates. In order to correctly estimate the radiative flux for the entry conditions of a NEO, an analytical law has been developed. During its entry, a NEO loses mass and change its shape because of ablation.To estimate the consequence on the trajectory of the NEO, two models of this phenomenon have been elaborated. Fragmentation has been modelled, from the origin of breakup to the mechanism offragment generation and flight dynamics of these fragments. This study showed the importance of these phenomena on ground consequences prediction, especially on the number of fragments impacting, their kinetic energies, and their positions on ground. Eventually, trajectory simulations (1 D&3 D), ta ken into account these phenomena, have been conducted. They highlighted the importance ofentry speed and slope on ground consequences.These simulations also demonstrated the protective role of the atmosphere on ground consequences, especially because of the fragmentation.

<p>This research focuses on experiments on compression-induced shock wave/boundary-layer interactions conducted in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) at Purdue University. The BAM6QT facilitates a low-freestream-noise hypersonic test environment more similar to that experienced in flight than a conventional wind tunnel. Measurements were captured on two sliced 7° half-angle cones with finite-span compression ramps. On the first, the slice was cut parallel to the axis of the cone to build upon previous measurements in hypersonic flow. While similar geometries have been analyzed for over 30 years in experiment and computation, there are significant gaps in understanding of the underlying mechanisms leading to instability and transition on the ramp. Further, in low-noise Mach 6 flow, the boundary layer separated at the leading edge of the slice, which is unlikely to occur on a real flight vehicle. Thus, on the second model, the slice was cut at a 4° incline to the</p> <p>cone axis to facilitate the growth of an attached laminar boundary layer on the slice. Using this configuration, the ramp-induced boundary-layer thickening initiated between the slice leading edge and the ramp leading edge, allowing the investigation of a ‘naturally’ formed separated region. </p> <p><br></p> <p>Data were captured at angles of attack ranging from 0° to 6°, on compression ramp angles ranging from 10° to 20°, and for freestream Reynolds numbers of 2.5×10^6/m to 12×10^6/m. To analyze the mean-flow behavior of the separation bubble as it changes with the above parametrics, time-averaged schlieren visualization was used to provide off-surface visualization of the flowfield, allowing estimates of reattachment position and separation bubble size. In all cases, reattachment position was shown to move upstream with an increase in angle of attack, an increase in ramp angle, and an increase in Reynolds number. However, on the model with the inclined slice, the Reynolds number impacted reattachment location to a much lesser extent. </p> <p><br></p> <p>Heat transfer measurements on the ramp revealed regions with the most significant aerothermal loading. Streamwise streaks of high heating originating at the ramp edges and centerline were observed to increase in magnitude with an increase in Reynolds number, angle of attack, and ramp angle. On the model with the inclined slice, many streaks of high heating were observed that increased in quantity and magnitude with angle of attack and ramp angle. Root mean squared pressure fluctuations computed from surface pressure measurements were shown to follow similar trends to centerline heat transfer results for both models. Angle of attack, ramp angle, and slice angle are shown to play a dominant role in transition. Finally, the importance of quiet tunnels is made remarkably clear, as the BAM6QT operating in its conventional-noise configuration resulted in drastically different results.</p> <p><br></p> <p>For measurement of shock wave/boundary-layer instabilities, schlieren frames were captured at 100,000 fps to allow measurement of low-to-mid-frequency fluctuations of the recirculation zone edge. Shear layer flapping frequencies were found to occur at around 1100–1200 Hz, which increased with angle of attack to up to 1600 Hz. It is likely that this is an inherent instability in the separation bubble itself, rather than a function of freestream disturbances, and may be indicative of an ‘expansion and relaxation’ effect known as bubble breathing. Additional measurements using low-frequency-capable pressure sensors must be captured to determine whether this breathing effect manifests on the model slice or ramp. </p>

<p>  </p> <p>High altitude test facilities are needed for ground testing of upper stage rocket engines or small satellite thrusters with high expansion ratio nozzles to ensure full-flowing nozzle conditions. Rocket exhaust diffusers and ejector systems are essential components of these facilities and are frequently used to set desired simulated altitude/low pressure conditions and pump out rocket exhaust products. </p> <p>This dissertation combined experimental and numerical efforts on diffuser-ejector systems. The experimental efforts included the development of a Second Throat Exhaust Diffuser (STED) to aid with the qualification of space thrusters in the Purdue Altitude Chamber Facility. While performing these experiments, we characterized the single and two-stage ejector systems operating in conjunction with the diffuser to obtain and maintain specific simulated altitudes. </p> <p>The concurrent numerical effort focused on validating a Computational Fluid Dynamics (CFD) approach based on Reynolds-averaged Navier–Stokes equations flow simulations. After validating the ejector CFD, we used it to derive a corrective coefficient of a lumped parameter ejector model (LPM) developed previously for the ejectors used in the Purdue Altitude Facility. We created variable coefficient maps for the stages of the two-stage ejector system using the same LPM and the test data from one of our experiments. </p> <p>We designed, manufactured, and then validated a STED for altitude testing of a ~50 lbf hypergolic hybrid motor as a part of a NASA JPL project. The designed STED enabled the operation of the hybrid motor for the full duration of the test firing (about 2 seconds) at a simulated altitude of 102,000 feet, slightly above the targeted altitude of 100,000 feet. We also validated our diffuser CFD approach by creating a simulation using the measured diffuser back pressure and the average motor chamber pressure. </p> <p>We then devised an experiment to investigate several diffuser–ejector system configurations using cold gas thrusters with conical and bell nozzles. The main aim of that experiment was to explore the effects of different thruster nozzle geometries, diffuser geometries, and thruster/ejector operational parameters on the performance of a diffuser–ejector system. For all the configurations tested, we reported on the minimum starting and operating pressure ratios and corresponding correction factors on the normal shock method. The large hysteresis regions obtained mostly with bell nozzles having a high initial expansion angle provided an opportunity to economize the facility resources. In some cases which were later found to violate STED second throat contraction limits, we experienced a choking flow at the second throat. Then, we studied the second throat contraction limits in detail using CFD in addition to the experimental data and explored minimum diffuser second throats enabling diffuser starting and improving aerodynamic efficiency. </p> <p>Finally, we machined a larger scale cold gas thruster with different nozzle geometries (having throat diameters in the range of 0.367 – 0.52 inches) from acrylic rods to study possible flow separation and gas condensation events that could occur during tests in the altitude chamber. The main difference here with the previous experiment was that the diffuser (JPL STED) was fixed, and the two-stage ejector system was used to create the necessary back pressure. With the experiments performed at varying axial gaps between the nozzle exit and diffuser inlet, we were able to investigate the effect of that on the diffuser performance. The experimental data collected in this work and the complementary numerical efforts served to generate the operating envelope of the Purdue Altitude Chamber Facility.  </p>

<p>This dissertation covers a computational investigation into hypersonic flight vehicle geometric imperfections, with a focus on wing-elevon-cove configurations. The primary region of focus for the overall research was the cove region at the juncture of the main wing element and the elevon. This region is associated with the shock-wave/boundary-layer interaction produced by the control surface deflection. There also exists a centrifugal instability at the cove, due to streamline curvature, which is associated with the production of Görtler vortices. The content includes three projects revolving around hypersonic wing-elevon-cove flows. These flows were computed with improved delayed detached-eddy simulation.</p> <p><br></p> <p>The first project was a computational investigation simulating the NASA experimental study done by W.D. Deveikis and W. Bartlett in 1978. This experiment consisted of hypersonic high Reynolds number wind tunnel tests for a shuttle-type reentry vehicle. The computational aerothermodynamic surface loadings for this project were compared to the experimental published data. Grounded with the agreement with mean surface data, this project expanded on the topics explored in the experimental study to include topics such as flow visualization and statistical analysis. The second and third project are extensions of this work and were done in collaboration with Purdue University and the University of Tennessee Space Institute (UTSI). A swept wing-elevon-cove model was designed by Carson Lay, of Purdue University, and is currently being employed in ongoing experiments in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) and at the Tennessee Aerothermodynamics Laboratory (TALon). A computational investigation on hypersonic high Reynolds number wing-elevon-cove flows was conducted with this model, where both corresponding experimental facility conditions were employed. At this time, the experimental data are limited; however, future experimental and computational collaboration is expected.</p> <p><br></p> <p>The motivation behind this research was to expand the knowledge on hypersonic wing-elevon-cove flows, gap heating, and the low-frequency unsteadiness in shock-wave/boundary-layer interactions. Therefore, the intended goal of this work was to provide an accurate characterization of the three hypersonic wing-elevon-cove flows. This was accomplished by using computational data to produce flowfield visualizations, analyze aerothermodynamic loadings, and conduct statistical flow analyses. The results on the three hypersonic wing-elevon-cove computations are presented, analyzed, and discussed throughout this dissertation.</p>

<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>

High Mach number, low-Reynolds number (high-altitude) reentry flowfield predictions are an important problem area in computational aerothermodynamics. Available numerical tools for handling such flows are very few and significantly limited in their applicability. A new implicit fully-iterative Parabolized Navier-Stokes (PNS) scheme is developed to accurately predict such low-Reynolds number flows. In this new approach the differential equations governing the conservation of mass, momentum and energy, and the algebraic equation of state for a perfect gas are solved simultaneously in a coupled manner. The idea is presented that by treating the governing equations in this manner (rather than eliminating the pressure terms in the governing equations by using appropriate differentiated forms of the equation of state) it may be possible to have an unconditionally time-like numerical scheme. The stability of a simplified version of this new PNS scheme is also studied, and it is demonstrated that these simplified equations are unconditionally time-like in the subsonic as well as the supersonic flow regions. A pseudo-time integration approach is used in addition to a new second-order accurate fully-implicit smoothing, to improve the efficiency of the solution algorithm. The new PNS scheme is used to predict the flowfield around a seven-deg sphere-cone vehicle under high- and low-Reynolds number conditions. Two test case, Case A and Case B, are chosen such that Case A has a large freestream Reynolds number (2.92x10⁵), whereas Case B has a freestream Reynolds number of 1.72x10³, which is smaller than the usual limit of applicability of the non-iterative PNS schemes (Re~10⁴ or larger). Comparisons are made with other available numerical schemes, and the results substantiate the stability, accuracy and efficiency claims of the new Parabolized Navier-Stokes scheme. / Ph. D.

The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body. Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.

Thermal Protection System (TPS) is required to shield an atmospheric entry vehicle against the high surface heating environment experienced during hypersonic flight. There are significant uncertainties in the tools and models currently used for the prediction of entry aeroheating and TPS material thermal response. These uncertainties can be reduced using experimental data. Analysis of TPS ground and flight data has been traditionally performed in a direct fashion. Direct analyses center upon comparison of the computational model predictions to data. Qualitative conclusions about model validity may be drawn based on this comparison and a limited number of model parameters may be iteratively adjusted to obtain a better match between predictions and data. The goal of this thesis is to develop a more rigorous methodology for the estimation of surface heating and TPS material response using inverse estimation theory. Built on theoretical developments made in related fields, this methodology enables the estimation of uncertainties in both the aeroheating environment and material properties from experimental temperature data. Unlike direct methods, the methodology developed here is capable of estimating a large number of independent parameters simultaneously and reconstructing the time-dependent surface heating profile in an automated fashion. This methodology is applied to flight data obtained from thermocouples embedded in the Mars Pathfinder and Mars Science Laboratory entry vehicle heatshields.