Dissertations / Theses on the topic 'Hypersonic propulsion and hypersonic aerothermodynamics'
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1
Fico, Vincenzo. "A high-order method for computational hypersonic aerothermodynamics." Thesis, University of Strathclyde, 2011. http://oleg.lib.strath.ac.uk:80/R/?func=dbin-jump-full&object_id=16946.
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Fiala, Abderrahmane. "Aerothermodynamics of turbulent spots and wedges at hypersonic speeds." Thesis, Imperial College London, 2005. http://hdl.handle.net/10044/1/12013.
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Song, Dong Joo. "Hypersonic nonequilibrium flow over an ablating teflon surface." Diss., Virginia Polytechnic Institute and State University, 1986. http://hdl.handle.net/10919/71192.
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A complex chemical system of teflon/air mixture over an axisymmetric decoy at hypersonic reentry flight conditions has been analyzed by using the nonequilibrium viscous shock-layer method. The equilibrium catalytic wall boundary condition was used to obtain the species concentration at the wall. The species conservation equation for binary mixture (air/teflon) was solved to obtain the concentration of freestream air at the wall. Two test cases were chosen to demonstrate the capability of the current code. Due to lack of experimental or theoretical data, the surface measurable quantities from the current code(VSLTEF) were compared with the equivalent air injection and no-mass injection data obtained from VSL7S code. The current code predicts a higher total heat-transfer rate than that predicted by the seven species nonequilibrium air code (VSL7S) with the same injection rate due to the high diffusional heat-transfer rate. The wall pressure was not affected by blowing, while the skin-friction coefficient was decreased (i.e., 43 % reduction for teflon ablation case ; 53 % for nonequilibrium air injection case at 125 kft) when compared with that of no-mass injection case. A shock-layer peak temperature drop ( 1512° R for 125 kft altitude and 848°R for 175 kft altitude) was observed at both cases. The temperature drops were chiefly due to endothermic reactions (dissociation) of the teflon ablation species. Due to large blowing of teflon, the average molecular weight increased substantially and resulted in a reduction of the specific heat ratio γ and an increase in the Prandtl number at the wall. The impurity of sodium was the major source of free electrons near the wall at the end of the vehicle at 125 kft altitude; however, at 175 kft altitude NO⁺ was the major source of free electrons over the entire body. The peak concentration of Na⁺ increased along the body, but that of NO⁺ decreased at both altitudes; While the chemical reaction rate data used is believed to be the best currently available, uncertainties in this data as were cited by Cresswell et al.(1967) may lead to quantitative changes in the above teflon ablation results.Ph. D.
4
Wilson, Althea Grace. "Numerical study of energy utilization in nozzle/plume flow-fields of high-speed air-breathing vehicles." Diss., Rolla, Mo. : Missouri University of Science and Technology, 2008. http://scholarsmine.mst.edu/thesis/pdf/Wilson_09007dcc804d881b.pdf.
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Thesis (M.S.)--Missouri University of Science and Technology, 2008.Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed April 25, 2008) Includes bibliographical references (p. 57).
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Bae, Yoon-Yeong. "Performance of an aero-space plane propulsion nozzle /." Full-text version available from OU Domain via ProQuest Digital Dissertations, 1989.
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DellaFera, Andrew Brian. "Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods." Thesis, Virginia Tech, 2019. http://hdl.handle.net/10919/95512.
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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.
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Halls-Moore, Michael Louis. "Computational modelling of hypersonic propulsion intakes at off-design conditions." Thesis, Imperial College London, 2009. http://hdl.handle.net/10044/1/59000.
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Hypersonic airbreathing propulsion is soon to become a viable means of transportation and hence determination of engine efficiency at varying flow conditions is important. This report contains the summary of literature surveyed on hypersonic airbreathing propulsion and intake design. An overview of ramjet/scramjets and flow factors that affect their efficiency were studied. A 2D planar intake was studied to generate Mach reflections via a shock-expansion interaction. The Mach reflection geometry was predicted with a slipstream profiler and an algebraic model, which was compared to finite volume based CFD code. An axisymmetric intake with a similar configuration to the planar intake was used to generate shock-expansion interactions. A Method of Characteristics code was written to predict the Von Neumann and Detachment Criteria for the transition between regular and Mach reflections. CFD was used to confirm the existence of a shock reflection transition hysteresis in a traverse of this dual solution domain. An increase in the freestream Mach number and initial flow angle for the planar intakes led to a complex subsonic/supersonic flowfield involving multiple shock reflections and interactions, known as a Type 3 Mach reflection. A temporal analysis was carried out to provide insight into the development of the Type 3 case using CFD. The initial flow angle was increased sequentially to assess the affect on the flowfield topology.
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Zoebelein, Till. "Development of an LU-scheme for the solution of hypersonic non-equilibrium flow." Thesis, Georgia Institute of Technology, 1998. http://hdl.handle.net/1853/12509.
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Silton, Sidra Idelle. "Ablation onset in unsteady hypersonic flow about nose-tips with a forward-facing cavity." Access restricted to users with UT Austin EID Full text (PDF) from UMI/Dissertation Abstracts International, 2001. http://wwwlib.umi.com/cr/utexas/fullcit?p3023560.
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Gupta, Anurag. "The artificially blunted leading edge concept for aerothermodynamic performance enhancement." Diss., Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/12442.
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Bhutta, Bilal A. "A new parabolized Navier-Stokes scheme for hypersonic reentry flows." Diss., Virginia Polytechnic Institute and State University, 1985. http://hdl.handle.net/10919/52287.
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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.
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Henderson, Sean James. "Study of the Issues of Computational Aerothermodynamics Using a Riemann Solver." Wright State University / OhioLINK, 2008. http://rave.ohiolink.edu/etdc/view?acc_num=wright1211933128.
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Bradford, John Edward. "A technique for rapid prediction of aftbody nozzle performance for hypersonic launch vehicle design." Diss., Georgia Institute of Technology, 2001. http://hdl.handle.net/1853/12896.
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Contreras, Daimer Mauthsud Leovan Ospina. "Angle of attack impact in the aerothermodynamics of a hypersonic vehicle with surface discontinuity-like a cavity." Instituto Nacional de Pesquisas Espaciais (INPE), 2017. http://urlib.net/sid.inpe.br/mtc-m21b/2017/04.12.00.58.
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The study described in this dissertation was undertaken with the purpose to investigate the impact of discontinuities present on the surface of hypersonic space vehicles. With this perspective in mind, computational simulations of a non-reacting rarefied hypersonic flow over a flat plate with a cavity have been performed by using the Direct Simulation Monte Carlo method. Simulations provided a comprehensive description about the nature of the flowfield structure and the aerodynamic surface properties on the cavity resulting from changes in the length-to-depth (L/H) ratio and changes in the angle of attack of the oncoming flow. A detailed description of the flowfield properties (velocity, density, pressure and temperature) and aerodynamics surface properties (number flux, heat transfer, pressure and skin friction) were obtained by a numerical method that properly account for non-equilibrium effects in the transition flow regime. Results for a cavity defined by L/H ratio of 1, 2, 3 and 4, and flow with angle of attack of 10, 15 and 20 degrees, were compared to those of a flat plate without a cavity with zero-degree angle of incidence and with a flat plate at incidence. The analysis showed that the flow topology inside the cavity, composed by recirculation regions, depended on the L/H ratio as well as on the angle of attack, for the conditions investigated. For L/H < 3 a single vortex core was formed, and filled entirely the cavity. In contrast, for L/H of 3 and 4, two vortices were formed inside the cavity, at the vicinity of the backward and forward faces. The analysis also showed that, for the L/H = 4 case, the flow topology inside the cavity corresponds to that of a ${''}$closed cavity${''}$ in the continuum flow regime for 10-degree angle of incidence, and similar to an open cavity for the others angles of attack investigated. In addition, it was found that the maximum values for the heat transfer, pressure and skin friction coefficients inside the cavity took place on the cavity forward face. It was also found that, maximum values for heat transfer coefficient inside the cavities increased with increasing the angle of attack $\alpha$. However, it was observed that these maximum values are smaller than those observed in a flat-plate without a cavity for the corresponding angle of attack. Consequently, in terms of pressure, the presence of the cavity on the vehicle surface can not be ignored in the vehicle design.O estudo descrito nesta dissertação foi realizado com o propósito de investigar o impacto de descontinuidades presentes na superfície de veículos espaciais hipersônicos. Em busca deste propósito, simulações computacionais de um escoamento hipersônico rarefeito não-reativo sobre uma cavidade foram realizadas usando-se o método Direct Simulation Monte Carlo. As simulações forneceram informações detalhadas sobre a natureza da estrutura do escoamento, propriedades primárias e propriedades aerodinâmicas, em função de mudanças na razão comprimento-profundidade (L/H) da cavidade, e mudanças no ângulo de ataque do escoamento incidindo sobre a cavidade. Uma descrição detalhada, das propriedades primárias (velocidade, massa específica, pressão e temperatura) e das quantidades aerodinâmica na superfície (transferência de calor, pressão e atrito), foi obtida por um método numérico que leva em conta adequadamente os efeitos de não-equilíbrio no regime de transição. Os resultados, para cavidades definidas por L/H de 1, 2, 3 e 4, com ângulos de ataque do escoamento de 10, 15 e 20 graus, foram comparados com os de uma placa plana sem/com a presença de cavidade sem/com incidência. A análise mostrou que a topologia do escoamento dentro da cavidade, composta por regiões de recirculação,dependeu da razão L/H bem como do ângulo de ataque do escoamento, para as condições investigadas. Para L/H < 3, observou-se a formação de um único vórtice ocupando inteiramente a cavidade. Para cavidade com L/H =3 e 4, dois vórtices foram formados dentro da cavidade, nas vizinhanças das faces a montante e a jusante da cavidade. A análise também mostrou que, para uma cavidade com L/H = 4 e 10 graus de incidência, a estrutura do escoamento dentro da cavidade correspondeu aquela de uma cavidade fechada , conforme definido para um escoamento no regime do contínuo. Por outro lado, para L/H = 4 e maiores ângulos de incidência, a estrutura do escoamento correspondeu aquela de uma cavidade aberta , para os ângulos de ataque investigados. Outrossim, verificou-se que os valores máximos para os coeficientes de transferência de calor, pressão e coeficiente de atrito ocorreram na superfície a montante do escoamento dentro da cavidade. Verificou-se também que, os valores máximos para o coeficiente de transferência de calor dentro da cavidade aumentaram com o aumento do ângulo de ataque $\alpha$. Todavia, esses valores máximos foram menores do que aqueles observados sobre uma placa plana sem cavidade com incidência. Como resultado, em termos de pressão, a presença da cavidade sobre a superfície do veículo não pode ser ignorada no projeto do veículo.
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Dreyer, Emily Rose. "Assessment of Reduced Fidelity Modeling of a Maneuvering Hypersonic Vehicle." The Ohio State University, 2021. http://rave.ohiolink.edu/etdc/view?acc_num=osu1610018486409227.
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Zhang, Huaibao. "HIGH TEMPERATURE FLOW SOLVER FOR AEROTHERMODYNAMICS PROBLEMS." UKnowledge, 2015. https://uknowledge.uky.edu/me_etds/64.
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A weakly ionized hypersonic flow solver for the simulation of reentry flow is firstly developed at the University of Kentucky. This code is the fluid dynamics module of known as Kentucky Aerothermodynamics and Thermal Response System (KATS). The solver uses a second-order finite volume approach to solve the laminar Navier– Stokes equations, species mass conservation and energy balance equations for flow in chemical and thermal non-equilibrium state, and a fully implicit first-order backward Euler method for the time integration. The hypersonic flow solver is then extended to account for very low Mach number flow using the preconditioning and switch of the convective flux scheme to AUSM family. Additionally, a multi-species preconditioner is developed.
The following part of this work involves the coupling of a free flow and a porous medium flow. A new set of equation system for both free flows and porous media flows is constructed, which includes a Darcy–Brinkmann equation for momentum, mass conservation, and energy balance equation. The volume-average technique is used to evaluate the physical properties in the governing equations. Instead of imposing interface boundary conditions, this work aims to couple the free/porous problem through flux balance, therefore, flow behaviors at the interface are satisfied implicitly.
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Torres, Luis Carlos Roldan. "Angle of attack effect in the aerothermodynamics of a hypersonic vehicle with a surface discontinuity of gap type." Instituto Nacional de Pesquisas Espaciais (INPE), 2017. http://urlib.net/sid.inpe.br/mtc-m21b/2017/05.23.23.55.
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The development of hypersonic vehicles has become a topic of interest in recent years, since has made it possible to reach inaccessible places such as orbital flights. The construction of these vehicles must be made with specials materials, and must have an efficient aerodynamic shape to withstand high speeds, high temperatures and significant pressure changes. The study described in this dissertation was undertaken with the objective to investigate the impact of discontinuities present on the surface of hypersonic space vehicles. In pursuit of this goal, computational simulations of a low-density hypersonic flow over a flat plate with a gap has been performed by using the Direct Simulation Monte Carlo method. The simulations provided information about the nature of the flowfield structure and the aerodynamic surface properties on the gap resulting from variations in the length-to-depth (L/H) ratio and variations in the angle of attack. A description of the flowfield properties, such as velocity, density, pressure and temperature, and aerodynamics surface quantities, such as, number flux, heat transfer, pressure and skin friction, were obtained by a numerical method that properly account for non-equilibrium effects in the transition flow regime. Results for a gap defined by L/H ratio of 1, 1/2, 1/3 and 1/4, and flow with angle of attack of 10, 15 and 20 degrees, were compared to those of a flat plate without a gap with zero-degree angle of incidence. The analysis showed that the flow topology inside the gap with incidence is slightly different from that for zero-degree angle of incidence for the L/H ratio investigated. It was found that the maximum values for the heat transfer, pressure and skin friction coefficients inside the gap took place on the gap forward face. It was also found that, maximum values for heat transfer coefficient inside the gaps increased with increasing the angle of attack $\alpha$. Nevertheless, it was observed that these maximum values are smaller than those observed in a flat-plate without a gap for the corresponding angle of attack. As a result, in terms of pressure, the presence of the gap on the vehicle surface can not be ignored in the vehicle design.O desenvolvimento de veículos hipersônicos tem se tornado um tema de interesse nos últimos anos, considerando-se a possibilidade de se chegar com tais veículos a locais até então inacessíveis como os voos orbitais. A construção desses veículos exige materiais especiais e deve apresentar uma forma aerodinâmica eficiente para resistir altas velocidades além de temperaturas elevadas e mudanças de pressão significativas. O estudo descrito nesta dissertação foi realizado com o objetivo de investigar o impacto de descontinuidades presentes na superfície de veículos espaciais hipersônicos. Em busca deste objetivo, simulações computacionais de um escoamento hipersônico rarefeito sobre uma placa plana, foi realizada usando-se o método Direct Simulation Monte Carlo. As simulações forneceram informações sobre a natureza da estrutura do escoamento, propriedades primarias e propriedades aerodinâmicas, devido a variações na razão comprimento-profundidade (L/H), e variações no ângulo de ataque. Uma descrição das propriedades primarias, tais como velocidade, massa específica, pressão e temperatura, e das quantidades aerodinâmica, tais como transferência de calor, pressão e atrito na superfície, foi obtida por um método numérico que leva em conta os efeitos de não-equilíbrio no regime de transição. Os resultados para um filete definido por uma razão L/H de 1, 1/2, 1/3 e 1/4, e com ângulo de ataque do escoamento de 10, 15 e 20 graus, foram comparados com os de uma placa plana sem a presença de um filete. A análise mostrou que a estrutura do escoamento dentro do filete com ângulo de ataque é ligeiramente diferente daquela com zero grau de incidência para cada razão L/H investigada. Verificou-se que os valores máximos para os coeficientes de transferência de calor, pressão e coeficiente de atrito ocorreram na superfície a montante do escoamento dentro do filete. Verificou-se também que, os valores máximos para o coeficiente de transferência de calor dentro do filete aumentaram com o aumento do ângulo de ataque $\alpha$. Como resultado, em termos de pressão, a presença do filete sobre a superfície do veículo não pode ser ignorada no projeto do veículo.
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Tirtey, Sandy C. "Characterization of a transitional hypersonic boundary layer in wind tunnel and flight conditions." Doctoral thesis, Universite Libre de Bruxelles, 2009. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/210367.
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Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.
Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.
The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods.
Doctorat en Sciences de l'ingénieur
info:eu-repo/semantics/nonPublished
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Rock, Christopher. "Experimental Studies of Injector Array Configurations for Circular Scramjet Combustors." Diss., Virginia Tech, 2010. http://hdl.handle.net/10919/77208.
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A flush-wall injector model and a strut injector model representative of state of the art scramjet engine combustion chambers were experimentally studied in a cold-flow (non-combusting) environment to determine their fuel-air mixing behavior under different operating conditions. The experiments were run at nominal freestream Mach numbers of 2 and 4, which simulates combustor conditions for nominal flight Mach numbers of 5 and 10. The flush-wall injector model consists of sixteen inclined, round, sonic injectors distributed around the wall of a circular duct. The strut injector model has sixteen inclined, round, sonic injectors distributed across four struts within a circular duct. The struts are slender, inclined at a low angle to minimize drag, and have two injectors on each side. The experiments investigated the effects of injectant molecular weight, freestream Mach number, and jet-to-freestream momentum flux ratio on the fuel-air mixing process. Helium, methane, and air injectants were studied to vary the injectant molecular weight over the range of 4-29. All of these experiments were performed to support the needs of an integrated experimental and computational research program, which has the goal of upgrading the turbulence models that are used for Computational Fluid Dynamics predictions of the flow inside a scramjet combustor. The primary goals of this study were to use injector models that represent state of the art scramjet engine combustion chambers to provide validation data to support the development of turbulence model upgrades and to add to the sparse database of mixing results in such configurations. The main experimental results showed that higher molecular weight injectants had approximately the same amount of penetration in the far field as lower molecular weight injectants at the same jet-to-freestream momentum flux ratio. Higher molecular weight injectants also demonstrated a mixing rate that was the same as or slower than lower molecular weight injectants depending on the flow conditions. A comparison of the experimental results for the two different injector models revealed that the flush-wall injector mixed significantly faster than the strut injector in all of the experimental cases.Ph. D.
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Perkins, Hugh Douglas. "Development and Demonstration of a Computational Tool for the Analysis of Particle Vitiation Effects in Hypersonic Propulsion Test Facilities." Case Western Reserve University School of Graduate Studies / OhioLINK, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=case1227553721.
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Lewis, Mark Joel. "The prediction of inlet flow stratification and its influence on the performance of air-breathing hypersonic propulsion systems." Thesis, Massachusetts Institute of Technology, 1988. http://hdl.handle.net/1721.1/14370.
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Tran, Kathleen. "One Dimensional Analysis Program for Scramjet and Ramjet Flowpaths." Thesis, Virginia Tech, 2010. http://hdl.handle.net/10919/30857.
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One-Dimensional modeling of dual mode scramjet and ramjet flowpaths is a
useful tool for scramjet conceptual design and wind tunnel testing. In this thesis,
modeling tools that enable detailed analysis of the flow physics within the combustor are
developed as part of a new one-dimensional MATLAB-based model named VTMODEL.
VTMODEL divides a ramjet or scramjet flow path into four major components: inlet,
isolator, combustor, and nozzle. The inlet module provides two options for supersonic
inlet one-dimensional calculations; a correlation from MIL Spec 5007D, and a kinetic
energy efficiency correlation. The kinetic energy efficiency correlation also enables the
user to account for inlet heat transfer using a total temperature term in the equation for
pressure recovery. The isolator model also provides two options for calculating the
pressure rise and the isolator shock train. The first model is a combined Fanno flow and
oblique shock system. The second model is a rectangular shock train correlation. The
combustor module has two options for the user in regards to combustion calculations.
The first option is an equilibrium calculation with a â growing combustion sphereâ
combustion efficiency model, which can be used with any fuel. The second option is a
non-equilibrium reduced-order hydrogen calculation which involves a mixing correlation
based on Mach number and distance from the fuel injectors. This model is only usable for
analysis of combustion with hydrogen fuel. Using the combustion reaction models, the
combustor flow model calculates changes in Mach number and flow properties due to the
combustion process and area change, using an influence coefficient method. This method
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also can take into account heat transfer, change in specific heat ratio, change in enthalpy,
and other thermodynamic properties.
The thesis provides a description of the flow models that were assembled to create
VTMODEL. In calculated examples, flow predictions from VTMODEL were compared
with experimental data obtained in the University of Virginia supersonic combustion
wind tunnel, and with reported results from the scramjet models SSCREAM and RJPA.
Results compared well with the experiment and models, and showed the capabilities
provided by VTMODEL.Master of Science
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Crowell, Andrew R. "Model Reduction of Computational Aerothermodynamics for Multi-Discipline Analysis in High Speed Flows." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1366204830.
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Bonanos, Aristides Michael. "Scramjet Operability Range Studies of an Integrated Aerodynamic-Ramp-Injector/Plasma-Torch Igniter with Hydrogen and Hydrocarbon Fuels." Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/28847.
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An integrated aerodynamic-ramp-injector/plasma-torch-igniter of original design was tested in a Mâ = 2, unvitiated, heated flow facility arranged as a diverging duct scramjet combustor. The facility operated at a total temperature of 1000 K and total pressure of 330 kPa. Hydrogen (H2), ethylene (C2H4) and methane (CH4) were used as fuels, and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements, and the presence of combustion was determined based on the pressure rises obtained.
Supersonic and dual-mode combustion were achieved with hydrogen and ethylene fuel, whereas very limited heat release was obtained with the methane. Global operability limits were determined to be 0.07 < Ï < 0.31 for hydrogen, and 0.14 < Ï < 0.48 for ethylene. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10º unswept compression ramp injector and similar performance was found with the two arrangements. With hydrogen and ethylene as fuels and the aeroramp/plasma-torch system, the effect of varying the air total temperature was investigated. Supersonic combustion was achieved with temperatures as low as 530K and 680K for the two fuels, respectively. These temperatures are facility/operational limits, not combustion limits.
The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with Ï from a high of about 75% to a low of about 45% at the highest Ï . With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Fuel specific impulse was on average 3000 and 1000 seconds for hydrogen and ethylene respectively, and air specific impulse varied from a low of about 9 sec to a high of about 24 sec (for both fuels) for the To = 1000K test condition.
The GASP RANS code was used to numerically simulate the injection and mixing process of the fuels. The results of this study were very useful in determining the suitability of the selected plasma torch locations. Further, this tool can be used to determine whether combustion is theoretically possible or not.Ph. D.
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Bailet, Gilles. "Radiation and ablation studies for in-flight validation." Thesis, Université Paris-Saclay (ComUE), 2019. http://www.theses.fr/2019SACLC008.
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Dévoiler les mystères du système solaire pour comprendre les mécanismes de la formation de la Terre, pour rechercher des signes de vie ou pour développer des colonies sur d’autres planètes, dépend de notre capacité à repousser les limites de l'ingénierie et de la science. Pour cela, il est important de développer des technologies de pointe pour permettre aux véhicules spatiaux de survivre la phase d'entrée ou de rentrée atmosphérique. Lors de l’entrée ou de la rentrée, l’engin spatial peut être exposé à flux radiatifs intenses qui ne peuvent pas encore être prédits avec précision, imposant ainsi des marges de sécurité sur la conception des systèmes de protection thermique. Ces incertitudes augmentent lorsque le bouclier thermique est constitué d'un matériau ablatif car sa dégradation introduit de nouvelles espèces chimiques réagissant avec le plasma produit devant le véhicule, ce qui affecte le rayonnement. Le but de cette thèse est d’étudier les flux de chaleur radiatifs sur un véhicule de rentrée de petite taille en présence d’un bouclier ablatif (Thermal Protection System, ou TPS), en utilisant des simulations numériques et des expériences pour développer un instrument de vol qui sera embarqué à bord du CubeSat QARMAN.Une évaluation de la trajectoire de rentrée du véhicule QARMAN (masse : 5 kg) a été réalisée en utilisant un code maison à 6 degrés de liberté. Un ensemble de simulations Monte Carlo ont permis de quantifier les incertitudes et ont montré un maximum de ± 15% écart par rapport à la trajectoire nominale. Les spectres sans ablation ont alors été déterminés en utilisant une approche découplée avec deux codes : Stagline (VKI) et SPECAIR (EM2C, CentraleSupélec). Ces simulations ont été effectuées pour la trajectoire nominale ainsi que pour la gamme des incertitudes. Elles ont permis de mettre en évidence un comportement non-linéaire des caractéristiques spectrales par rapport aux valeurs nominales, avec une augmentation drastique vers la fin de la mission.Les effets de l'ablation ont été étudiés avec une nouvelle technique de mesure développée au cours de cette thèse. Basée sur deux sondes de mesure de rayonnement, l’une refroidie et l’autre recouverte d’un matériau ablatif, cette méthode permet de quantifier l'émission et l'absorption induite par tout type de TPS ayant des interactions gaz-surface avec l'écoulement, dans l’hypothèse que les raies d’émission et d’absorption des espèces ablatives ne soient pas superposées. La méthode a été validée sur un échantillon de graphite TPS. Elle a ensuite été appliquée à la prédiction du rayonnement attendu lors de la mission QARMAN (Cork P50 TPS). Cette étude a également permis de sélectionner un spectromètre d’émission adapté à la mission QARMAN et aux objectifs de la thèse (plage de 350 à 800 nm pour une masse de 68 g).Un instrument de mesure de rayonnement standard a été testé et les limites de cet appareil ont été établies. Deux nouvelles technologies ont été développées et la charge utile (spectromètre d’émission INES) a été construite et intégrée au véhicule QARMAN. Un étalonnage spectral et thermique dédié a également été développé pour maximiser la qualité du retour scientifique en prenant en compte les variations de température dans la baie de charge utile de QARMAN.L’instrument proposé est, à ce jour, la seule charge utile non intrusive capable d’effectuer des mesures radiatives sans limitations liées à la contamination par les poussières et gaz d'ablation. L’instrument peut aussi fournir des mesures de la récession, de la sublimation et du gonflement du TPS avec une précision d'au moins 0,2 mm. Le fonctionnement de l'appareil a été démontré pour une grande variété de conditions de test, y compris différents profils d'enthalpie, mélanges de gaz et matériaux de TPSUnveiling the mysteries of the solar system to understand the mechanisms of Earth’s formation, to search for signs of life, or to develop settlements on other planets, depends on our abilities to push the limits of engineering and science. One of the key aspects of space exploration is the development of advanced technologies to sustain the entry/reentry phase. During entry or reentry, the spacecraft may be exposed to intense radiative fluxes that cannot be accurately predicted yet, thus imposing high safety margins on the design of thermal protection systems. These uncertainties rise when the heat shield is made of an ablative material as its degradation introduces new chemical species reacting with the flow affecting radiation processes. The goal of this thesis is to study the radiative heat fluxes onto a small size reentry vehicle in the presence of an ablative TPS, using numerical simulations and experiments to develop a flight instrument that will be carried onboard the QARMAN CubeSat.An assessment of the reentry trajectory of the 5-kg QARMAN vehicle was performed using a custom 6-degree of freedom code. An extensive set of Monte Carlo simulations allowed to quantify uncertainties and showed a maximum of ±15% deviation from the nominal trajectory. The spectra without ablation were then computed using a decoupled approach with two codes: Stagline (VKI) and SPECAIR (EM2C, CentraleSupélec). These simulations were performed for the nominal trajectory as well as for the range of uncertainties. They showed a nonlinear behavior of the spectral features deviations from nominal with a drastic increase toward the end of the mission.The effects of ablation were studied with a new measurement technique developed during this thesis. Based on two radiation measurement probes, one cooled and the other with an ablative surface, it allows to quantify the emission and absorption induced by any kind of TPS having gas-surface interactions with the flow, provided that the radiative emission or absorption features of the ablative species do not fully overlap. The method was validated on a graphite TPS sample. It was then applied to determine the radiation expected during the QARMAN mission (Cork P50 TPS). This study also allowed to select an emission spectrometer (350-800 nm range for a 68-g mass).A standard radiation instrument was tested and the limits of this device shown. On those lessons learned, two new technologies were developed and an emission spectrometer payload (INES) was built and integrated into the QARMAN reentry CubeSat. A dedicated spectral and thermal calibration was also developed to maximize the quality of the scientific return by tackling the non-standard internal temperature variations of QARMAN’s payload bay.Relying on two inventions made during this study, the apparatus is at the time of writing, the only non-intrusive payload capable of making radiative measurements without limitations due to ablation dust contamination. The instrument can also provide measurements of recession, sublimation and swelling of the TPS with a precision of at least 0.2 mm. Operation of the apparatus was demonstrated for a wide variety of test conditions, including different enthalpy profiles, gas mixtures and TPS materials
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Benyo, Theresa L. "Analytical and Computational Investigations of a Magnetohydrodynamic (MHD) Energy-Bypass System for Supersonic Turbojet Engines to Enable Hypersonic Flight." Kent State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=kent1369153719.
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Morham, Brett G. "Numerical Examination of Flow Field Characteristics and Fabri Choking of 2D Supersonic Ejectors." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/340.
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An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio.
The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered.
For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.
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Ferrier, Loïc. "Analyse aérothermodynamique de l'entrée atmosphérique d'un géocroiseur à occurence séculaire." Thesis, Toulouse, ISAE, 2012. http://www.theses.fr/2012ESAE0017/document.
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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'ablationNear 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
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Retaureau, Ghislain J. "On recessed cavity flame-holders in supersonic cross-flows." Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/43703.
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Flame-holding in a recessed cavity is investigated experimentally in a Mach 2.5 preheated cross-flow for both stable and unstable combustion, with a relatively low preheating. Self-sustained combustion is investigated for stagnation pressures and temperatures reaching 1.4 MPa and 750 K. In particular, cavity blowout is characterized with respect to cavity aspect ratio (L/D =2.84 - 3.84), injection strategy (floor - ramp), aft ramp angle (90 deg - 22.5 deg) and multi-fuel mixture (CH₄-H₂ or CH₄-C₂H₄ blends). The results show that small hydrogen addition to methane leads to significant increase in flame stability, whereas ethylene addition has a more gradual effect. Since the multi-fuels used here are composed of a slow and a fast chemistry fuel, the resulting blowout region has a slow (methane dominant) and a fast (hydrogen or ethylene dominant) branch. Regardless of the fuel composition, the pressure at blowout is close to the non-reacting pressure imposed by the cross-flow, suggesting that combustion becomes potentially unsustainable in the cavity at the sub-atmospheric pressures encountered in these supersonic studies. The effect of preheating is also investigated and results show that the stability domain broadens with increasing stagnation temperature. However, smaller cavities appear less sensitive to the cross-flow preheating, and stable combustion is achieved over a smaller range of fuel flow rate, which may be the result of limited residence and mixing time. The blowout data point obtained at lower fuel flow rate fairly matches the empirical model developed by Rasmussen et al. for floor injection phi = 0.0028 Da^-.8, where phi is the equivalence ratio and Da the Damkohler number. An alternate model is proposed here that takes into account the ignition to scale the blowout data. Since the mass of air entrained into the cavity cannot be accurately estimated and the cavity temperature is only approximated from the wall temperature, the proposed scaling has some uncertainty. Nevertheless the new phi-Da scaling is shown to preserve the subtleties of the blowout trends as seen in the current experimental data.
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Martos, João Felipe de Araújo. "Investigação experimental do veículo hipersônico aeroespacial 14-XB." reponame:Repositório Institucional da UFABC, 2014.
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Orientador: Prof. Dr. João Felipe de Araújo MartosDissertação (mestrado) - Universidade Federal do ABC, Programa de Pós-Graduação em Engenharia Mecânica, 2014.
O veículo hipersônico aeroespacial brasileiro 14-X B (VHA 14-X B) é um demonstrador tecnológico do sistema de propulsão hipersônica aspirada com base em combustão supersônica (scramjet) projetado para voar na atmosfera da Terra a 30 quilômetros de altitude e número de Mach 7. O VHA 14-X B encontra-se em desenvolvimento no Laboratório de Aerotermodinâmica e Hipersônica Professor Henry T. Nagamatsu do Instituto de Estudos Avançados (IEAv). Uma das metodologias mais importantes na análise e desenvolvimento de um veículo aeroespacial hipersônico são os túneis de vento hipersônicos pulsados, os quais são instalações experimentais em terra capazes de simular as condições de voo, fornecendo dados experimentais para o projeto dos veículos aeroespaciais hipersônicos. Neste trabalho, foi utilizado o túnel de vento hipersônico pulsado T3 que possui 61 cm de diâmetro no bocal de saída e é operado com base na técnica de onda de choque refletida. O T3 foi financiado pela Fundação de Amparo a Pesquisa de São Paulo (FAPESP) e foi projetado para Pesquisa e Desenvolvimento na área de combustão supersônica. Trata-se de um tubo de choque equipado com um bocal convergente-divergente utilizado para produzir elevado número Mach e escoamentos de alta entalpia na seção de teste próximos aos encontrados durante o voo de um veículo aeroespacial na atmosfera da Terra a hipervelocidade. O modelo de 1 metro de comprimento em aço inoxidável do 14-X B foi instrumentado, com vinte e oito transdutores de pressão piezelétricos PCB nas superfícies de compressão, câmara de combustão e de expansão. Utilizando o túnel T3 no modo de operação de equilíbrio de interface para atingir escoamento livre com número de Mach entre 7 e 8 foi realizada a investigação experimental. Medidas da pressão estática no intradorso do modelo 14-X B, bem como fotografias schlieren, feitas a partir do bordo de ataque do modelo forneceram dados experimentais, que foram comparados com as análises teórica-analítica e simulações computacionais de dinâmica de fluidos, ambos utilizadas no projeto do modelo do VHA 14-X B.
The Brazilian VHA 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) intended to fly into the Earth¿s atmosphere at 30 km altitude and Mach number 7. The 14-X B was designed o the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics of the Institute for Advanced Studies (IEAv). Hypersonic wind tunnels are one of the most important ground-based experimental facilities intended to simulate the flight conditions providing experimental data to design hypersonic aerospace vehicles and engines. The IEAv 0.60-m nozzle exit diameter Hypersonic Reflected Shock Tunnel named T3 and funded by São Paulo Research Foundation (FAPESP), was designed as a research & development facility for basic investigations in supersonic combustion. The T3 Hypersonic Shock Tunnel is a shock tube equipped with a convergent-divergent nozzle to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of a aerospace vehicle into the Earth's atmosphere at hypersonic flight speeds. A 1-m long stainless steel VHA 14-X B model was instrumented with twenty-eight piezoelectric pressure transducers along its compression surface, combustion chamber and nozzle. It was experimentally investigated on the equilibrium interface operational mode of the T3 Hypersonic Shock Tunnel, yielding a freestream Mach number from 7 to 8. Static pressure measurements at the lower surface of the VHA 14-X B as well as high speed schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp provide experimental data that was compared to the analytic theoretical analysis and computational fluid dynamics simulation, both applied to the design of the VHA 14-X B.
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Poulain, François. "Commande d'un véhicule hypersonique à propulsion aérobie : modélisation et synthèse." Phd thesis, Ecole Nationale Supérieure des Mines de Paris, 2012. http://pastel.archives-ouvertes.fr/pastel-00744985.
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La propulsion aérobie à grande vitesse est depuis longtemps identifiée comme l'un des prochains sauts technologiques à franchir dans le domaine des lanceurs spatiaux. Cependant, les véhicules hypersoniques (HSV) fonctionnant dans des domaines de vitesse extrêmement élevées, de nombreuses contraintes et incertitudes entravent les garanties des propriétés des contrôleurs. L'objet de cette thèse est d'étudier la synthèse de commande d'un tel véhicule.Pour commencer, il s'agit de définir un modèle représentatif d'un HSV exploitable pour la commande. Dans ce travail, nous construisons deux modèles de HSV. Un pour la simulation en boucle fermée, et le second afin de poser précisément le problème de commande.Nous proposons ensuite une synthèse de commande de la dynamique longitudinale dans le plan vertical de symétrie. Celle-ci est robuste aux incertitudes de modélisation, tolérante à des saturations, et n'excite pas les dynamiques rapides négligées. Ses propriétés sont évaluées sur différents cas de simulation. Puis, une extension est proposée afin de résoudre le problème de commande simultanée des dynamiques longitudinale et latérale, sous les mêmes contraintes.Ce résultat est obtenu par une assignation de fonction de Lyapunov, suite à une étude des dynamiques longitudinale et latérale. Par ailleurs, pour traiter les erreurs de poursuite dues aux incertitudes de modélisation, nous nous intéressons au problème de régulation asymptotique robuste par retour d'état. Nous montrons que cette régulation peut être accomplie en stabilisant le système augmenté d'un intégrateur de la sortie. Ceci constitue une extension de la structure de contrôle proportionnel-intégral au cas des systèmes non linéaires.
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Yentsch, Robert J. "Three-Dimensional Shock-Boundary Layer Interactions in Simulations of HIFiRE-1 and HIFiRE-2." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1384195671.
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Hima, Bindu V. "Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration." Thesis, 2009. http://hdl.handle.net/2005/1120.
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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.
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Harrland, Alan. "Hypersonic inlet for a laser powered propulsion system." Thesis, 2012. http://hdl.handle.net/2440/79072.
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The idea of laser powered lightcraft was first conceptualised in the early 1970's as a means of launching small scale satellite payloads into orbit at a much lower cost in comparison to conventional techniques. Propulsion in the lightcraft is produced via laser induced detonation of the incoming air stream, which results in the energy source for propulsion being decoupled from the vehicle. In air breathing mode the lightcraft carries no onboard fuel or oxidiser, allowing theoretically infinite specific impulses to be achieved. Recently interest has been renewed in this innovative
technology through cross-continent and industry research programs aimed at making laser propulsion a reality. In a ground launched satellite, the vehicle must travel through the atmosphere at speeds greatly in excess of the speed of sound in order to achieve the required orbital velocities. Supersonic, and in particular hypersonic, flight regimes exhibit complicated physics that render traditional subsonic inlet design techniques inadequate. The laser induced detonation propulsion system requires a suitable engine configuration that offers good performance over all flight speeds and angles of attack to ensure the required thrust is maintained throughout the mission. Currently a hypersonic inlet has not been developed for the laser powered lightcraft vehicle. Stream traced hypersonic inlets have demonstrated the required performance in conventional hydrocarbon fuelled scramjet engines. This design technique is applied to the laser powered lightcraft vehicle, with its performance evaluated against the traditional lightcraft inlet design. Four different hypersonic lightcraft inlets have been produced employing both the stream traced inlet design methodology, and traditional axi-symmetric inlet techniques. This thesis outlines the inlet design methodologies employed, with a detailed analysis of the performance of the lightcraft inlet at angles of attack and off-design conditions. Fully three-dimensional turbulent computational fluid dynamics simulations have been performed on a variety of inlet configurations. The performance of the lightcraft inlets have been evaluated at differing angles of attack. An idealised laser detonation simulation has also been performed to verify that the lightcraft inlet does not unstart during the laser powered propulsion cycle.Thesis (M.Phil.) -- University of Adelaide, School of Mathematical Sciences, 2012
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Nagashetty, K. "Experimental Investigations on Hypersonic Waverider." Thesis, 2014. http://hdl.handle.net/2005/3195.
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In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6.
The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.
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Saravanan, S. "Experimental Investigation Of The Effect Of Nose Cavity On The Aerothermodynamics Of The Missile Shaped Bodies Flying At Hypersonic Mach Numbers." Thesis, 2007. http://hdl.handle.net/2005/694.
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Hypersonic vehicles are exposed to severe heating loads during their flight in the atmosphere. In order to minimize the heating problem, a variety of cooling techniques are presently available for hypersonic blunt bodies. Introduction of a forward-facing cavity in the nose tip of a blunt body configuration of hypersonic vehicle is one of the most simple and attractive methods of reducing the convective heating rates on such a vehicle. In addition to aerodynamic heating, the overall drag force experienced by vehicles flying at hypersonic speeds is predominate due to formation of strong shock waves in the flow. Hence, the effective management of heat transfer rate and aerodynamic drag is a primary element to the success of any hypersonic vehicle design. So, precise information on both aerodynamic forces and heat transfer rates are essential in deciding the performance of the vehicle. In order to address the issue of both forces and heat transfer rates, right kind of measurement techniques must be incorporated in the ground-based testing facilities for such type of body configurations. Impulse facilities are the only devices that can simulate high altitude flight conditions. Uncertainties in test flow conditions of impulse facilities are some of the critical issues that essentially affect the final experimental results. Hence, more reliable and carefully designed experimental techniques/methodologies are needed in impulse facilities for generating experimental data, especially at hypersonic Mach numbers.
In view of the above, an experimental program has been initiated to develop novel techniques of measuring both the aerodynamic forces and surface heat transfer rates. In the present investigation, both aerodynamic forces and surface heat transfer rates are measured over the test models at hypersonic Mach numbers in IISc hypersonic shock tunnel HST-2, having an effective test time of 800 s. The aerodynamic coefficients are measured with a miniature type accelerometer based balance system where as platinum thin film sensors are used to measure the convective heat transfer rates over the surface of the test model.
An internally mountable accelerometer based balance system (three and six-component) is used for the measurement of aerodynamic forces and moment coefficients acting on the different test models (i.e., blunt cone with after body, blunt cone with after
body and frustum, blunt cone with after body-frustum-triangular fins and sharp cone with after body-frustum-triangular fins), flying at free stream Mach numbers of 5.75 and 8 in hypersonic shock tunnel. The main principle of this design is that the model along with the internally mounted accelerometer balance system are supported by rubber bushes and there-by ensuring unrestrained free floating conditions of the model in the test section during the flow duration. In order to get a better performance from the accelerometer balance system, the location of accelerometers plays a vital role during the initial design of the balance. Hence, axi-symmetric finite element modeling (FEM) of the integrated model-balance system for the missile shaped model has been carried out at 0° angle of attack in a flow Mach number of 8. The drag force of a model was determined using commercial package of MSC/NASTRAN and MSC/PATRAN. For test flow duration of 800 s, the neoprene type rubber with Young’s modulus of 3 MPa and material combinations (aluminum and stainless steel material used as the model and balance) were chosen. The simulated drag acceleration (finite element) from the drag accelerometer is compared with recorded acceleration-time history from the accelerometer during the shock tunnel testing. The agreement between the acceleration-time history from finite-element simulation and measured response from the accelerometer is very good within the test flow domain. In order to verify the performance of the balance, tests were carried out on similar standard AGARD model configurations (blunt cone with cylinder and blunt cone with cylinder-frustum) and the results indicated that the measured values match very well with the AGARD model data and theoretically estimated values. Modified Newtonian theory is used to calculate the aerodynamic force coefficient analytically for various angles of attack.
Convective surface heat transfer rate measurements are carried out by using vacuum sputtered platinum thin film sensors deposited on ceramic substrate (Macor) inserts which in turn are embedded on the metallic missile shaped body. Investigations are carried out on a model with and without fin configurations in HST-2 at flow Mach number of 5.75 and 8 with a stagnation enthalpy of 2 MJ/kg for zero degree angle of attack. The measured heating rates for the missile shaped body (i.e., with fin configuration) are lower than the predicted stagnation heating rates (Fay-Riddell expression) and the maximum difference is about 8%. These differences may be due to the theoretical values of velocity gradient used in the empirical relation. The experimentally measured values are expressed in terms of normalized heat transfer rates, Stanton numbers and correlated Stanton numbers, compared with the numerically estimated results. From the results, it is inferred that the location of maximum heating occurs at stagnation point which corresponds to zero velocity gradient. The heat-transfer ratio (q1/Qo)remains same in the stagnation zone of the model when the Mach number is increased from 5.75 to 8. At the corners of the blunt cone, the heat transfer rate doesn’t increase (or) fluctuate and the effects are negligible at two different Mach numbers (5.75 and 8). On the basis of equivalent total enthalpy, the heat-transfer rate with fin configuration (i.e., at junction of cylinder and fins) is slightly higher than that of the missile model without fin.
Attempts have also been made to evaluate the feasibility of using forward facing cavity as probable technique to reduce the heat transfer rate and to study its effect on aerodynamic coefficients on a 41° apex angle missile shaped body, in hypersonic shock tunnel at a free stream Mach number of 8. The forward-facing circular cavities with two different diameters of 6 and 12 mm are chosen for the present investigations. Experiments are carried out at zero degree angle of attack for heat transfer measurements. About 10-25 % reduction in heat transfer rates is observed with cavity at gauge locations close to stagnation region, whereas the reduction in surface heat transfer rate is between 10-15 % for all other gauge locations (which is slightly downstream of the cavity) compared with the model without cavity.
In order to understand the influence of forward facing cavities on force coefficients, measurement of aerodynamic forces and moment coefficients are also carried out on a missile shaped body at angles of attack. The same six component balance is also being used for subsequent investigation of force measurement on a missile shaped body with forward facing cavity. Overall drag reductions of up to 5 % is obtained for a cavity of 6 mm diameter, where as, for the 12 mm cavity an increase in aerodynamic drag is observed (up to about 10%). The addition of cavity resulted in a slight increase in the missile L/D ratio and did not significantly affect the missile lateral components. In summary, the designed balances are found to be suitable for force measurements on different test models in flows of duration less than a millisecond.
In order to compliment the experimental results, axi-symmetric, Navier-Stokes CFD computations for the above-defined models are carried out for various angles of attack using a commercial package CFX-Ansys 5.7. The experimental free stream conditions obtained from the shock tunnel are used for the boundary conditions in the CFD simulation. The fundamental aerodynamic coefficients and heat transfer rates of experimental results are shown to be in good agreement with the predicted CFD.
In order to have a feeling of the shock structure over test models, flow visualization experiments have been carried out by using the Schlieren technique at flow Mach numbers of 5.75 and 8. The visualized shock wave pattern around the test model consists of a strong bow shock which is spherical in shape and symmetrical over the forebody of the cone. Experimentally measured shock stand-off distance compare well with the computed value as well as the theoretically estimated value using Van Dyke’s theory. These flow visualization experiments have given a factual proof to the quality of flow in the tunnel test section.
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"Parametric Analysis of a Hypersonic Inlet using Computational Fluid Dynamics." Master's thesis, 2013. http://hdl.handle.net/2286/R.I.20879.
Full textAbstract:
abstract: For CFD validation, hypersonic flow fields are simulated and compared with experimental data specifically designed to recreate conditions found by hypersonic vehicles. Simulated flow fields on a cone-ogive with flare at Mach 7.2 are compared with experimental data from NASA Ames Research Center 3.5" hypersonic wind tunnel. A parametric study of turbulence models is presented and concludes that the k-kl-omega transition and SST transition turbulence model have the best correlation. Downstream of the flare's shockwave, good correlation is found for all boundary layer profiles, with some slight discrepancies of the static temperature near the surface. Simulated flow fields on a blunt cone with flare above Mach 10 are compared with experimental data from CUBRC LENS hypervelocity shock tunnel. Lack of vibrational non-equilibrium calculations causes discrepancies in heat flux near the leading edge. Temperature profiles, where non-equilibrium effects are dominant, are compared with the dissociation of molecules to show the effects of dissociation on static temperature. Following the validation studies is a parametric analysis of a hypersonic inlet from Mach 6 to 20. Compressor performance is investigated for numerous cowl leading edge locations up to speeds of Mach 10. The variable cowl study showed positive trends in compressor performance parameters for a range of Mach numbers that arise from maximizing the intake of compressed flow. An interesting phenomenon due to the change in shock wave formation for different Mach numbers developed inside the cowl that had a negative influence on the total pressure recovery. Investigation of the hypersonic inlet at different altitudes is performed to study the effects of Reynolds number, and consequently, turbulent viscous effects on compressor performance. Turbulent boundary layer separation was noted as the cause for a change in compressor performance parameters due to a change in Reynolds number. This effect would not be noticeable if laminar flow was assumed. Mach numbers up to 20 are investigated to study the effects of vibrational and chemical non-equilibrium on compressor performance. A direct impact on the trends on the kinetic energy efficiency and compressor efficiency was found due to dissociation.Dissertation/Thesis
M.S. Aerospace Engineering 2013
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Silton, Sidra Idelle 1973. "Ablation onset in unsteady hypersonic flow about nose-tips with a forward-facing cavity." 2001. http://hdl.handle.net/2152/10834.
Full text
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(8032571), Varun Viswanathan. "Hypersonic Stationary Crossflow Waves: Receptivity to Roughness." Thesis, 2019.
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Experiments were performed on a sharp-nosed 7° half-angle cone at a 6° angle of attack in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) to study the stationary crossflow instability and its receptivity to small surface roughness. Heat transfer measurements were obtained using temperature sensitive paint (TSP) and Schmidt Boelter (SB) heat transfer gauges. Great care was taken to obtain repeatable, quantitative measurements from TSP.
Consecutive runs were performed at a 0° angle of attack, and the heat transfer measured by the SB was found to drop as the initial model temperature increased, while other initial conditions such as stagnation pressure were held constant. This agreed with calculations done using a similarity solution. It was found that repeatable measurements at a 6° angle of attack could be made if the initial model temperature was controlled and the patch location that was used to calibrate the TSP was picked in a reasonable and consistent manner.
The Rod Insertion Method (RIM) roughness, which was used to excite the stationary crossflow instability, was found to be responsible for the appearance of the streaks that were analyzed. The signal-to-noise ratio in the TSP was too low to properly measure the streaks directly downstream of the roughness insert. The heat transfer along the streak experienced linear growth, peaked, and then slightly decayed. It is possible this peak was saturation. The general trend was that the growth of the streaks moved farther upstream as the roughness element height increased, which agreed with past computations and low speed experiments. The growth of the streak also moved farther upstream as the freestream Reynolds number increased. The amplitude of the streaks was calculated by non-dimensionalizing the heat transfer using the laminar theoretical mean-flow solution for a 7° half-angle cone at a 6° angle of attack. The relationship between the amplitude and the non-dimensional roughness height was approximately linear in the growth region of the streaks.
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(6624017), Joshua B. Edelman. "Nonlinear Growth and Breakdown of the Hypersonic Crossflow Instability." Thesis, 2019.
Find full textAbstract:
A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel
at 6° angle of attack, extending several previous experiments on the growth and breakdown of
stationary crossflow instabilities in the boundary layer.
Measurements were made using infrared
imaging and surface pressure sensors. Detailed measurements of the stationary and traveling
crossflow vortices, as well as various secondary instability modes, were collected over a large
region of the cone.
The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was
adapted for application to crossflow work. It was demonstrated that the roughness elements were
the primary factor responsible for the appearance of the specific pattern of stationary streaks
downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness
insert was created with a high RMS level of normally-distributed roughness to excite the naturally
most-amplified stationary mode.
The nonlinear breakdown mechanism induced by each type of roughness appears to be
different. When using the discrete RIM roughness, the dominant mechanism seems to be the
modulated second mode, which is significantly destabilized by the large stationary vortices. This
is consistent with recent computations. There is no evidence of the presence of traveling crossflow
when using the RIM roughness, though surface measurements cannot provide a complete picture.
The modulated second mode shows strong nonlinearity and harmonic development just prior
to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal
band, leading to a peak in the heating simultaneously with the peak amplitude of the measured
secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic
heating pattern is reminiscent of experiments on a flared cone and earlier computations
of crossflow on an elliptic cone.
When using the distributed roughness there are several differences in the nonlinear breakdown
behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,
which is expected given computational results. Furthermore, the traveling crossflow waves become
very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there
appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability
under the upwelling of the stationary vortices. The traveling crossflow and the secondary
instability interact nonlinearly prior to breakdown.
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(6623855), Mark Wason. "CALIBRATION OF HIGH-FREQUENCY PRESSURE SENSORS USING LOW-PRESSURE SHOCK WAVES." Thesis, 2019.
Find full textAbstract:
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.
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.
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.
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.
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(7399604), Phillip Portoni. "Using Suction for Laminar Flow Control in Hypersonic Quiet Wind Tunnels: A Feasibility Study." Thesis, 2019.
Find full textAbstract:
To reduce the risk of using suction in a hypersonic quiet-tunnel nozzle design, this project tested micro-perforated suction sections to remove the boundary layer on an axisymmetric model in the Boeing/AFOSR Mach-6 Quiet Tunnel. The model was a cone-flare geometry tested at 0° angle of attack. The turn from the 7° half-angle cone to the flare was designed to prevent flow separation. The flare was designed to amplify the Görtler instability.
Five suction sections were designed with different perforation patterns and porosities. Four were successfully manufactured, but only the first of the four sections has been tested so far. The first suction section has pores drilled along straight lines with a nominal 5% porosity.
Measurements were made with temperature-sensitive paint and oil-flow visualization on a non-perforated blank to measure the baseline development of Görtler vortices on the flare. Although the signal-to-noise ratio of the measurement techniques were insufficient to measure the vortices, it was confirmed that the boundary layer is laminar for the entire model. Measurements with suction also did not show the Görtler vortices.
Surface pressure fluctuations were measured on the flare. Apparent second-mode waves were detected. The suction measurements showed a slight increase in second-mode peak frequency over the baseline results, as expected.
Concerns had been raised about acoustic noise that might be radiated from the suction section. Thus, fluctuations above the suction section were measured using a pitot probe and using focused-laser differential interferometry. The measurements during suction showed no noticeable increase in fluctuations compared to the baseline results.
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"Modeling, Analysis, and Control of a Hypersonic Vehicle With Significant Aero-Thermo-Elastic-Propulsion Interactions, and Propulsive Uncertainty." Master's thesis, 2010. http://hdl.handle.net/2286/R.I.8592.
Full textAbstract:
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
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Chan, Jonathan. "Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11." Thesis, 2010. http://hdl.handle.net/1807/25446.
Full textAbstract:
This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-ω turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.
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(8793053), Gregory R. McKiernan. "Instability and Transition on a Sliced Cone with a Finite-Span Compression Ramp at Mach 6." Thesis, 2020.
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Initial experiments on separated shock/boundary-layer interactions were carried out within the Boeing/AFOSR Mach-6 Quiet Tunnel. Measurements were made of hypersonic laminar-turbulent transition within the separation above a compression corner. This wind tunnel features freestream fluctuations that are similar to those in
flight. The present work focuses on the role of traveling instabilities within the shear layer above the separation bubble.
A 7 degree half-angle cone with a slice and a finite-span compression ramp was designed and tested. Due to a lack of space for post-reattachment sensors, early designs of this
generic geometry did not allow for measurement of a post-reattachment boundary layer. Oil flow and heat transfer measurements showed that by lengthening the ramp, the post-reattachment boundary layer could be measured. A parametric study was completed to determine that a 20 degree ramp angle caused reattachment at 45% of the
total ramp length and provided the best flow field for boundary-layer transition measurements.
Surface pressure fluctuation measurements showed post-reattachment wave packets and turbulent spots. The presence of wave packets suggests that a shear-layer
instability might be present. Pressure fluctuation magnitudes showed a consistent transition Reynolds numbers of 900000, based on freestream conditions and distance
from the nosetip. Pressure fluctuations grew exponentially from less than 1% to roughly 10% of tangent-wedge surface pressure during transition.
A high-voltage pulsed plasma perturber was used to introduce controlled disturbances into the boundary layer. The concept was demonstrated on a straight 7 degree half-angle circular cone. The perturbations successfully excited the second-mode instability at naturally unstable frequencies. The maximum second-mode amplitudes prior to transition were measured to be about 10% of the mean surface static pressure.
The plasma perturber was then used to disturb the boundary layer just upstream of the separation bubble on the cone with the slice and ramp. A traveling instability was measured post-reattachment but the transition location did not change for any tested condition. It appears that the excited shear-layer instability was not the dominant mechanism of transition.
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(9189470), Abhinand Ayyaswamy. "Computational Modeling of Hypersonic Turbulent Boundary Layers By Using Machine Learning." Thesis, 2020.
Find full textAbstract:
A key component of research in the aerospace industry constitutes hypersonic flights (M>5) which includes the design of commercial high-speed aircrafts and development of rockets. Computational analysis becomes more important due to the difficulty in performing experiments and reliability of its results at these harsh operating conditions. There is an increasing demand from the industry for the accurate prediction of wall-shear and heat transfer with a low computational cost. Direct Numerical Simulations (DNS) create the standard for accuracy, but its practical usage is difficult and limited because of its high cost of computation. The usage of Reynold's Averaged Navier Stokes (RANS) simulations provide an affordable gateway for industry to capitalize its lower computational time for practical applications. However, the presence of existing RANS turbulence closure models and associated wall functions result in poor prediction of wall fluxes and inaccurate solutions in comparison with high fidelity DNS data. In recent years, machine learning emerged as a new approach for physical modeling. This thesis explores the potential of employing Machine Learning (ML) to improve the predictions of wall fluxes for hypersonic turbulent boundary layers. Fine-grid RANS simulations are used as training data to construct a suitable machine learning model to improve the solutions and predictions of wall quantities for coarser meshes. This strategy eliminates the usage of wall models and extends the range of applicability of grid sizes without a significant drop in accuracy of solutions. Random forest methodology coupled with a bagged aggregation algorithm helps in modeling a correction factor for the velocity gradient at the first grid points. The training data set for the ML model extracted from fine-grid RANS, includes neighbor cell information to address the memory effect of turbulence, and an optimal set of parameters to model the gradient correction factor. The successful demonstration of accurate predictions of wall-shear for coarse grids using this methodology, provides the confidence to build machine learning models to use DNS or high-fidelity modeling results as training data for reduced-order turbulence model development. This paves the way to integrate machine learning with RANS to produce accurate solutions with significantly lesser computational costs for hypersonic boundary layer problems.
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(8292123), Julien Keith Louis Brillon. "Modeling Thermochemical Nonequilibrium Processes and Flow Field Simulations of Spark-Induced Plasma." Thesis, 2020.
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This study is comprised of two separate parts: (1) modeling thermochemical nonequilibrium processes, and (2) flow field simulations of spark-induced plasma. In the first part, the methodology and literature for modeling thermochemical nonequilibrium processes in partially ionized air is presented and implemented in a zero-dimensional solver, termed as NEQZD. The solver was verified for a purely reacting flow case as well as two thermochemical nonequilibrium flow cases. A three-temperature electron-electronic model for thermochemical nonequilibrium partially ionizing air mixture was implemented and demonstrated the ability to capture additional physics compared to the legacy two-temperature model through the inclusion of electronic energy nonequilibrium. In the second part of this work, full scale axisymmetric simulations of the flow field produced by the abrupt heat release of spark-induced plasma were presented and analyzed for two electrode configurations. The heat release was modeled based on data from experiments and assumed that all electrical power supplied to the electrodes is converted to thermal energy. It was found that steeper electrode walls lead to a greater region of hot gas, a stronger shock front, and slightly larger vortices.
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"Modeling, Analysis, and Control of a Hypersonic Vehicle with Significant Aero-Thermo-Elastic-Propulsion Interactions: Elastic, Thermal and Mass Uncertainty." Master's thesis, 2011. http://hdl.handle.net/2286/R.I.8873.
Full textAbstract:
abstract: This thesis examines themodeling, 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 analyses. This model is based upon classical compressible flow and Euler-Bernouli structural concepts. Higher fidelity computational fluid dynamics and finite element methods are needed for more 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 vehicle's static and dynamic characteristics over the vehicle's 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. Aerodynamic heat modeling and design for the assumed TPS was incorporated to original Bolender's model to study the change in static and dynamic properties. De-centralized control stability, feasibility and limitations issues were dealt with the change in TPS elasticity, mass and physical dimension. The impact of elasticity due to TPS mass, TPS physical dimension as well as prolonged heating was also analyzed to understand performance limitations of de-centralized control designed for nominal model.Dissertation/Thesis
M.S. Electrical Engineering 2011
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(8960978), Lalit Rajendran. "DEVELOPMENT OF IMAGE-BASED DENSITY DIAGNOSTICS WITH BACKGROUND-ORIENTED SCHLIEREN AND APPLICATION TO PLASMA INDUCED FLOW." Thesis, 2021.
Find full textAbstract:
There is growing interest in the use of nanosecond surface dielectric barrier discharge (ns-SDBD) actuators for high-speed (supersonic/hypersonic) flow control. A plasma discharge is created using a nanosecond-duration pulse of several kilovolts, and leads to a rapid heat release and a complex three-dimensional flow field. Past work has been limited to qualitative visualizations such as schlieren imaging, and detailed measurements of the induced flow are required to develop a mechanistic model of the actuator performance.
Background-Oriented Schlieren (BOS) is a quantitative variant of schlieren imaging and measures density gradients in a flow field by tracking the apparent distortion of a target dot pattern. The distortion is estimated by cross-correlation, and the density gradients can be integrated spatially to obtain the density field. Owing to the simple setup and ease of use, BOS has been applied widely, and is becoming the preferred density measurement technique. However, there are several unaddressed limitations with potential for improvement, especially for application to complex flow fields such as those induced by plasma actuators.
This thesis presents a series of developments aimed at improving the various aspects of the BOS measurement chain to provide an overall improvement in the accuracy, precision, spatial resolution and dynamic range. A brief summary of the contributions are:
1) a synthetic image generation methodology to perform error and uncertainty analysis for PIV/BOS experiments,
2) an uncertainty quantification methodology to report local, instantaneous, a-posteriori uncertainty bounds on the density field, by propagating displacement uncertainties through the measurement chain,
3) an improved displacement uncertainty estimation method using a meta-uncertainty framework whereby uncertainties estimated by different methods are combined based on the sensitivities to image perturbations,
4) the development of a Weighted Least Squares-based density integration methodology to reduce the sensitivity of the density estimation procedure to measurement noise.
5) a tracking-based processing algorithm to improve the accuracy, precision and spatial resolution of the measurements,
6) a theoretical model of the measurement process to demonstrate the effect of density gradients on the position uncertainty, and an uncertainty quantification methodology for tracking-based BOS,
Then the improvements to BOS are applied to perform a detailed characterization of the flow induced by a filamentary surface plasma discharge to develop a reduced-order model for the length and time scales of the induced flow. The measurements show that the induced flow consists of a hot gas kernel filled with vorticity in a vortex ring that expands and cools over time. A reduced-order model is developed to describe the induced flow and applying the model to the experimental data reveals that the vortex ring's properties govern the time scale associated with the kernel dynamics. The model predictions for the actuator-induced flow length and time scales can guide the choice of filament spacing and pulse frequencies for practical multi-pulse ns-SDBD configurations.
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Perez, Jaime Enrique. "EVALUATION OF GEOMETRIC SCALE EFFECTS FOR SCRAMJET ISOLATORS." 2010. http://trace.tennessee.edu/utk_gradthes/739.
Full textAbstract:
A numerical analysis was conducted to study the effects of geometrically scaling scramjet inlet-combustor isolators. Three-dimensional fully viscous numerical simulation of the flow inside constant area rectangular ducts, with a downstream back pressure condition, was analyzed using the SolidWorks Flow Simulation software. The baseline, or 1X, isolator configuration has a 1” x 2.67” cross section and 20” length. This baseline configuration was scaled up based on the 1X configuration mass flow to 10X and 100X configurations, with ten and one hundred times the mass flow rate, respectively. The isolator aspect ratio of 2.67 was held constant for all configurations. To provide for code validation, the Flow Simulation program was first used to analyze a converging-diverging channel and a wind tunnel nozzle. The channel case was compared with analytical theory and showed good agreement. The nozzle case was compared with AFRL experimental data and showed good agreement with the entrance and exit conditions (Pi0= 40 psia, Ti0= 530ºR, Pe= 18.86 psia, Te= 456ºR, respectively). While the boundary layer thickness remained constant, the boundary layer thickness with respect to the isolator height decreased as the scale increased. For all the isolator simulations, a shock train was expected to form inside the duct. However, the flow simulation failed to generate this flow pattern, due to improper sizing of the isolator and combustor for a 3-D model or having a low pressure ratio of 2.38. Instead, a single normal shock wave was established at the same relative location within the length of each duct, approximately 80% of the duct length from the isolator entrance. The shape of the shock changed as the scale increased from a normal shock wave, to a bifurcated shock wave, and to a normal shock train, respectively for the 1X, 10X, and 100X models.