In an attempt to generalize previous models of the bidirectional vortex mean flow, a new solution is presented that can cope with arbitrary injections and outlet conditions. In the process, the steady, inviscid and axisymmetric equations of motions are reduced to one partial differential equation for the stream function, known as the Bragg-Hawthorne equation, which is solved exactly. The solution is shown to be highly dependent on the imposed boundary conditions: the mean flow changes according to the manner by which the fluid is injected or extracted from the vortex chamber. From the stream function, the velocity is obtained along with the vorticity and pressure distributions which are carefully derived and analyzed. The results are then compared to several inviscid models found in the literature. After determining an exact inviscid solution to the problem, viscous effects at the core are added to overcome the known singularity that arises at the centerline. The governing equations are hence revisited while keeping the viscous diffusion term in the tangential momentum equation. The core region, where viscous effects lead to the onset of a forced vortex, is rescaled using appropriate transformations. An asymptotic approximation is then applied to linearize and solve the resulting ODE for the tangential vi velocity. The inner viscous solution is then matched to the outer inviscid result using Prandtl’s Matching Principle. Finally, the viscous correction is passed onto the vorticity and pressure formulations.

Rocket propulsion using the hybrid-propellant scheme is a technology that offers much promise in applications where high-performance liquid rocket engines are deemed too complex and solid rocket motors are considered to lack performance or safety. However, despite extensive research, there is still a lack of knowledge in the theoretical aspects of hybrid rocketry, especially in the area of fuel-oxidizer mixing and fuel regression rate. This lack of a good theoretical model makes the implementation of good, practical solutions and mature, well-functioning designs more diffcult. This disadvantages the hybrid rocket engine when compared to liquid rocket engines or solid rocket motors.In this study, a hybrid rocket engine burning polymethyl methacrylate (PMMA) with compressed air has been designed to the point of a preliminary design defnition. PMMA is a transparent material, and this has been utilized to create a transparent-chamber rocket engine where engine processes can be studied with various optical methods withoutinterrupting or disturbing the operation of the engine. The function of hybrid rocket engines, the technological solutions involved in designing working hybrid rocket engines and the constituent parts of hybrid rocket engines have been studied. The nature of the trade-offs between performance and simplicity that occur when designing a rocket engine are also studied, with a focus on maximizing simplicity, safety and minimizing expenses, while still designing an engine that fulfills basic requirements. The results include a design defnition with a preliminary user’s guide, a feasibility study, and a summary of the results of the hybrid rocket performance model that was used to determine appropriate design parameters.

<p>This thesis focuses on the development of reduced pressure testing capabilities at Zucrow Laboratories. A two-stage ejector on loan from NASA Marshall is used in series with a supersonic diffuser to allow for the testing of up to100 lb_f rocket engines at equivalent altitudes of up to 100,000 ft. The objective of this research is to implement a one-dimensional (1-D) model which accurately predicts the performance of the two-stage ejector in real time, informing the maximum thrust and simulated altitude capabilities within the altitude chamber located in room 134A of ZL3 during experimental testing.</p>

Progress in conventional rocket engine technologies, based on constant pressure combustion, has plateaued in the past few decades. Rotating detonation engines (RDEs) are of particular interest to the rocket propulsion community as pressure gain combustion may provide improvements to specific impulse relevant to booster applications. Despite recent significant investment in RDE technologies, little research has been conducted to date into the effect of nozzle design on rocket application RDEs. Proper nozzle design is critical to capturing the thrust potential of the transient pressure ratios produced by the thrust chamber. A computational fluid dynamics study was conducted based on hotfire conditions tested in the Purdue V1.3 RDE campaign. Three geometries were investigated: nozzleless/blunt body, internal-external expansion (IE-) aerospike, and flared aerospike. The computational study found the RDE's dynamic exhaust plume enhances the ejection physics beyond that of a typical high pressure device. For the nozzleless geometry, the base pressure was drawn down below constant pressure estimates, increasing the base drag on the engine. For the aerospike geometries, the occurrance of flow separation on the plug was delayed, which has ramifications on nozzle design for operation at a range of pressure altitudes. The flared aerospike design, which has the ability to achieve much higher area ratios, was shown to have potential performance benefits over the limited IE-aerospike geometry. A new test campaign with the Purdue RDE V1.4 was designed with instrumentation to capture static pressures on the nozzleless and aerospike surfaces. These results were used to validate the results from the computational study. The computational and experimental studies were used to identify new flow physics associated with a rocket RDE important to future nozzle design work. Future computational work is necessary to explore the effect of different parameters on the nozzle performance. More testing, including with an altitude simulation chamber, would help quantify the possible benefit of new aerospike nozzle designs, including the flared aerospike geometry.

<div>Thermoacoustic instabilities in rocket engines have been studied for decades and models have been attempted, however, the heat release fluctuations and overall response</div><div>is still poorly understood. To understand the heat release mechanism in a rocket combustion chamber the effect of hydrodynamics and chemical kinetics on the mode/s of combustion need to be studied. Using prior simulations of the CVRC, an initial design for a new model rocket combustor was proposed. The new design improved on past experiments by having better control of all important boundary conditions; facilitate higher fidelity pressure and optical measurements with emphasis on quantifying the results and using them to validate simulation models of the design; and allow good control over the characteristic parameters of the injection mechanics. A prior simulation was done on the proposed design to allow fine tuning of the</div><div>design elements. Three distinct modes of self-excited instability were observed in the experiment, two of which transitioned between one another with a sweep in oxidizer</div><div>temperature. A number of configurations and operating conditions were tested, but the primary focus was on three oxidizer rich cases, at different oxidizer temperatures. The two extreme cases were compared to the simulations conducted. At low oxidizer temperatures there was good agreement, where at high oxidizer temperatures there</div><div>was a fairly good agreement in the type of mechanics observed, but there were a few discrepancies. The vortex shedding off of the fuel collar was captured using chemiluminescence measurements and compared quite well with the simulations. It was found that the fuel collar vortex shedding did not directly contribute to the generation of</div><div>instabilities.</div>

Combustion devices are prone to combustion instabilities. They arise from a constructive coupling between the unsteady heat release rate of the flame and the resonant acoustic modes of the entire system. The occurence of such instabilities can pose a threat to both performance and integrity of combustion systems. Although these phenomena have been known for more than a century, avoiding their appearance in industrial engines is still challenging. The objective of this thesis is threefold: (1) study the dynamics of the resonant acoustic modes, (2) investigate the flame response of a liquid rocket engine under unstable conditions using Large Eddy Simulation(LES) and (3) derive, use and study Time Domain Impedance Boundary Conditions (TDIBCs), i.e. boundary conditions modeling complex acoustic impedances.

Made available in DSpace on 2014-06-11T19:28:33Z (GMT). No. of bitstreams: 0 Previous issue date: 2011-08-22Bitstream added on 2014-06-13T20:58:28Z : No. of bitstreams: 1 melo_hht_me_guara.pdf: 1759227 bytes, checksum: 3dfd437259b41c05e4944f9e56da28d5 (MD5) / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES) / A utilização de imagens digitais para extrair informações de objetos tem sido uma solução amplamente empregada em pesquisas científicas e em processos industriais. A contínua redução nos preços de equipamentos, a facilidade do uso de softwares e a simples integração com recursos de informática tem feito que muitos processos migrem para esta solução mais ágil, confiável e econômica. A indústria aeroespacial, que possui uma cadeia de produção não contínua e exige a avaliação de todos os seus componentes para obtenção de um nível de confiança elevado, encontra no emprego do processamento digital de imagens uma solução versátil e eficaz para análise das características de cada componente. Neste trabalho é apresentado um programa, desenvolvido em LabVIEW™, para medição dos sprays cônicos de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens para sua análise. São apresentadas também as metodologias até então utilizadas para efetuar este tipo de medida. Os sprays dos jatos são desenvolvidos na saída do injetor, são exibidos visualmente durante o teste hidráulico a frio e tem influência direta no desempenho do motor foguete. A utilização desta nova ferramenta permitiu a realização desta medida de forma automática, com o fornecimento da incerteza de medição em níveis de confiança pré-estabelecido e mostrou-se ser mais exata e precisa que as metodologias anteriores / The usage of digital images to extract information from objects has been a solution widely used in scientific research and in industrial processes. The continued reduction in prices of equipment, the facility of software manipulation and the simple integration with computing resources has done many processes to migrate to this more flexible, reliable and economical solution. The aerospace industry, which has a chain of production that is not continuous and requires the evaluation of all its components to obtain a high confidence level, finds in the usage of digital image processing a versatile and effective solution for analysis of the characteristics of each component . This paper presents a program developed in LabVIEW™, to measure the rocket engine conic spray jet by using a digital image processing system for analysis. It is also presented the methodologies previously used to perform this type of measurement. The jet sprays are developed at the exit of the injector, are displayed visually during the cold hydraulic test and it has directly influences on the performance of the rocket engines. The usage of this new tool allowed us to make the measurement automatically with the supply of uncertainty together with a pre-established confidence level and it proved to be more accurate and precise than previous methodologies

2011 IEEE Congress on Evolutionary Computation (CEC). June 5-8, 2011, Ritz-Carlton, New Orleans, LA, USA

The advantages of Topology Optimization (TO) are realized to a large extent due to the manufacturing freedom that Additive Manufacturing (AM) offer, compared to more conventional manufacturing methods. AM has the advantage of manufacturing shallow and complex structures previously not possible, and consequently opens up a whole new design spectrum. This thesis investigates the possibilities of using Topology Optimization as a tool to find stronger and lighter designs for the inlet turbine manifold in the rocket engine demonstrator Prometheus. The manifold is optimized by giving it more mass, subjecting it to load cases and pushing the topology optimization to make the manifold meet the weight requirement without exceeding the yield strength. Result validation indicates that the pressure and thermal loadings are the most prominent. The current topology optimization tools in ANSYS do not support optimization due to thermal features and thus optimization in the presented work has only been able to consider static structural loads. Nevertheless, it is possible to optimize the manifold due to static structural loads and achieve a manifold which satisfies the weight requirement. However, optimization tools due to thermal loading would be a desirable feature in the future.

Orientador: Fernando de Azevedo Silva / Banca: João Zangrandi Filho / Banca: Silvana Aparecida Barbosa / Resumo: A utilização de imagens digitais para extrair informações de objetos tem sido uma solução amplamente empregada em pesquisas científicas e em processos industriais. A contínua redução nos preços de equipamentos, a facilidade do uso de softwares e a simples integração com recursos de informática tem feito que muitos processos migrem para esta solução mais ágil, confiável e econômica. A indústria aeroespacial, que possui uma cadeia de produção não contínua e exige a avaliação de todos os seus componentes para obtenção de um nível de confiança elevado, encontra no emprego do processamento digital de imagens uma solução versátil e eficaz para análise das características de cada componente. Neste trabalho é apresentado um programa, desenvolvido em LabVIEW™, para medição dos sprays cônicos de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens para sua análise. São apresentadas também as metodologias até então utilizadas para efetuar este tipo de medida. Os sprays dos jatos são desenvolvidos na saída do injetor, são exibidos visualmente durante o teste hidráulico a frio e tem influência direta no desempenho do motor foguete. A utilização desta nova ferramenta permitiu a realização desta medida de forma automática, com o fornecimento da incerteza de medição em níveis de confiança pré-estabelecido e mostrou-se ser mais exata e precisa que as metodologias anteriores / Abstract: The usage of digital images to extract information from objects has been a solution widely used in scientific research and in industrial processes. The continued reduction in prices of equipment, the facility of software manipulation and the simple integration with computing resources has done many processes to migrate to this more flexible, reliable and economical solution. The aerospace industry, which has a chain of production that is not continuous and requires the evaluation of all its components to obtain a high confidence level, finds in the usage of digital image processing a versatile and effective solution for analysis of the characteristics of each component . This paper presents a program developed in LabVIEW™, to measure the rocket engine conic spray jet by using a digital image processing system for analysis. It is also presented the methodologies previously used to perform this type of measurement. The jet sprays are developed at the exit of the injector, are displayed visually during the cold hydraulic test and it has directly influences on the performance of the rocket engines. The usage of this new tool allowed us to make the measurement automatically with the supply of uncertainty together with a pre-established confidence level and it proved to be more accurate and precise than previous methodologies / Mestre