In this thesis, combustion processes in a laboratory-scale methane based low pressure rocket motor (LPRM) is studied experimentally and numerically. Experiments are conducted to measure flame temperatures and chamber temperature and pressure. Single reaction-four species reacting flow of gaseous methane and gaseous oxygen in the combustion chamber is also simulated numerically using a commercial CFD solver based on 2-D, steady-state, viscous, turbulent and compressible flow assumptions. LPRM geometry is simplified to several configurations, i.e. Channel and Combustion Chamber with Nozzle and FWD. Flow in a Bunsen burner is simulated inside Channel geometry in order to validate the reaction model. Grid independence study is also conducted for reacting as well as non-reacting flows. Numerical model is calibrated based on experimental results. Results of the computational model are found in a good agreement with the experimental data after calibrating specific heats of the products. Parametric study is conducted in order to investigate the effects of different mass flow rates and chamber pressures on flow and combustion characteristics of a LPRM to provide insight to future studies.

A condição ótima de funcionamento de uma tubeira em um motor foguete com escoamento isentrópico, implica que a velocidade na garganta (seção de menor área) seja equivalente à velocidade do som local, condição de Mach 1 e bocal engasgado. Pode-se alcançar essa condição reduzindo a área da seção do escoamento até a área crítica, velocidade sônica. Após a garganta acontece a expansão e se alcança velocidades supersônicas no divergente. Para manter a condição de Mach 1 na garganta em motores foguetes, trabalha-se com pressões superiores à necessária para se engasgar o bocal. Isto ocorre porque tenta-se compensar instabilidades ou variações de volumes produzidos na combustão ou queima. Usando uma pressão de trabalho maior, impõe-se que a condição de Mach 1 fique mantida durante toda a queima do combustível, isso implica em usar tubos mais resistentes à pressão e maior massa do tubo-motor. Observou-se experimentalmente que em algumas situações construtivas se podem modificar a pressão e temperatura necessárias para engasgar o bocal aumentando o comprimento da garganta. O comprimento do estrangulamento pode estabelecer uma condição para formação e evolução da camada limite e esta condição restringir a área nominal, modificando o regime do escoamento. Um equipamento especialmente desenvolvido para esse ensaio compara resultados de cinco modelos de motores, divididos em dois grupos, cada grupo com áreas de entrada, garganta e saída iguais, porém com comprimentos diferentes de garganta. Em análise experimental, observou-se que a pressão de trabalho e a temperatura são influenciadas pelo comprimento da garganta, interferindo na relação entre as pressões internas e de garganta e apresentando condições de engasgamento mensuráveis. Essas medidas foram conduzidas no presente estudo de doutorado. / The optimum operational condition of a rocket motor nozzle with isentropic flow implies that the velocity at the throat (the section with smallest area) is equivalent to the speed of the local sound. This speed is also called Mach 1 and it is said that at this condition the nozzle is choking. One can achieve this condition by reducing the cross-sectional area of the flow to the critical area resulting in a sonic speed. Beyond the nozzle throat, in the divergent section of the motor, flow expansion occurs and reaches supersonic speeds. To maintain the condition of Mach 1 at the throat, higher pressures than the one necessary to choke the nozzle are applied. This practice is done in order to compensate for jitter or variations of volumes produced in the combustion process. Using a higher operating pressure guarantees that a Mach 1 speed is maintained throughout the combustion process. Consequently, due to this higher operating pressure, more resistant tubes are needed to withstand this higher pressure and an increase in the motor weight is inevitable. It was observed experimentally that some constructional modifications of the motor can alter the pressure and temperature required for choking. This was noted with increasing the bottleneck length of the nozzle throat which was able to establish a condition for the formation and evolution of the boundary layer, restricting the nominal area and thus modifying the flow regime. In this study, the results of five engine models are compared using a specially designed equipment. The rockets were divided into two groups, each with equal inlet, throat, and exit areas, but having different throat lengths. In experimental analysis, it was observed that the working pressure and temperature are influenced by the length of the throat, interfering in the relationship between the internal pressures and throat presenting measurable choking conditions which were conducted in this doctorate thesis study.

A condição ótima de funcionamento de uma tubeira em um motor foguete com escoamento isentrópico, implica que a velocidade na garganta (seção de menor área) seja equivalente à velocidade do som local, condição de Mach 1 e bocal engasgado. Pode-se alcançar essa condição reduzindo a área da seção do escoamento até a área crítica, velocidade sônica. Após a garganta acontece a expansão e se alcança velocidades supersônicas no divergente. Para manter a condição de Mach 1 na garganta em motores foguetes, trabalha-se com pressões superiores à necessária para se engasgar o bocal. Isto ocorre porque tenta-se compensar instabilidades ou variações de volumes produzidos na combustão ou queima. Usando uma pressão de trabalho maior, impõe-se que a condição de Mach 1 fique mantida durante toda a queima do combustível, isso implica em usar tubos mais resistentes à pressão e maior massa do tubo-motor. Observou-se experimentalmente que em algumas situações construtivas se podem modificar a pressão e temperatura necessárias para engasgar o bocal aumentando o comprimento da garganta. O comprimento do estrangulamento pode estabelecer uma condição para formação e evolução da camada limite e esta condição restringir a área nominal, modificando o regime do escoamento. Um equipamento especialmente desenvolvido para esse ensaio compara resultados de cinco modelos de motores, divididos em dois grupos, cada grupo com áreas de entrada, garganta e saída iguais, porém com comprimentos diferentes de garganta. Em análise experimental, observou-se que a pressão de trabalho e a temperatura são influenciadas pelo comprimento da garganta, interferindo na relação entre as pressões internas e de garganta e apresentando condições de engasgamento mensuráveis. Essas medidas foram conduzidas no presente estudo de doutorado. / The optimum operational condition of a rocket motor nozzle with isentropic flow implies that the velocity at the throat (the section with smallest area) is equivalent to the speed of the local sound. This speed is also called Mach 1 and it is said that at this condition the nozzle is choking. One can achieve this condition by reducing the cross-sectional area of the flow to the critical area resulting in a sonic speed. Beyond the nozzle throat, in the divergent section of the motor, flow expansion occurs and reaches supersonic speeds. To maintain the condition of Mach 1 at the throat, higher pressures than the one necessary to choke the nozzle are applied. This practice is done in order to compensate for jitter or variations of volumes produced in the combustion process. Using a higher operating pressure guarantees that a Mach 1 speed is maintained throughout the combustion process. Consequently, due to this higher operating pressure, more resistant tubes are needed to withstand this higher pressure and an increase in the motor weight is inevitable. It was observed experimentally that some constructional modifications of the motor can alter the pressure and temperature required for choking. This was noted with increasing the bottleneck length of the nozzle throat which was able to establish a condition for the formation and evolution of the boundary layer, restricting the nominal area and thus modifying the flow regime. In this study, the results of five engine models are compared using a specially designed equipment. The rockets were divided into two groups, each with equal inlet, throat, and exit areas, but having different throat lengths. In experimental analysis, it was observed that the working pressure and temperature are influenced by the length of the throat, interfering in the relationship between the internal pressures and throat presenting measurable choking conditions which were conducted in this doctorate thesis study.

Propõe-se aqui uma metodologia de projeto aero-termo-estrutural de motores foguete a propelente sólido. O projeto de um motor foguete deve ser realizado com o objetivo de cumprir requisitos de uma missão. Para cada veículo espacial, com uma nova missão, um novo motor pode ser projetado, necessitando para isso de uma série de ferramentas robustas, capazes de compreender todas as combinações de esforços existentes no funcionamento de um motor, sob condições de altas pressões e temperaturas. A metodologia aqui proposta é testada e validada em bancada de ensaios desenvolvida para este fim. Os resultados obtidos mostram que a metodologia utilizada se aproxima bastante dos resultados teóricos e pode ser ajustada por coeficientes de eficiência com grande facilidade. / It is proposed here an aero-thermo-structural design methodology for solid propellant rocket motors. The design of a rocket motor must be carried out in order to fulfill requirements of a mission. For each new space vehicle, with a new mission, a new motor can be designed, requiring for it a variety of robust tools, able to comprise all combinations of load existing in the operation of a motor under high pressures and temperatures. The methodology proposed here is tested and validated in bank of tests developed for this purpose. The results show that the methodology is very close to the theoretical results and can be adjusted by coefficients of efficiency with great ease.

A greenhouse study was conducted to determine the effects of deposition from static rocket test fires on corn and alfalfa. Seeds were germinated in a wide concentration range of depositional material, called test fire soil (TFS). Additionally, the impact of chloride and aluminum, two major components of test fire soil, on germination was also evaluated. Furthermore, plants were grown in packed columns and exposed to test fire soil, either in the root zone or on foliage. Tissue was weighed and analyzed to compare biomass production and plant composition. Corn and alfalfa exposed to test fire soil in the root zone produced less biomass than controls, but foliar treatment had no effect on biomass production. No kernels were produced by corn exposed to test fire soil in the root zone. Leaves of plants exposed to test fire soil in the root zone accumulated more metals and nutrients than controls, whereas plant tissue treated with test fire soil on the leaves contained only elevated levels of aluminum, although levels were still within reasonable concentrations for plants. Germination of seeds was not affected below 1% test fire soil in soil; however higher concentrations of test fire soil decreased percent germination. Addition of chloride to soil also inhibits germination, but addition of aluminum has no effect on germination percentage. Corn germination was restored in test fire soil leached with 200 mm artificial rainwater. The results of this research contribute information regarding the potential impact of test fire soil from static test fires on crop production. Test fire soil inhibits germination and growth if deposited in the root zone, and even foliar application alters tissue composition. However, plant composition is not altered significantly in terms of feed criteria, and germination can be restored by irrigating the TFS. The effects of test fire soil are attributed to high levels of chloride that induce salt stress. Crop damage may be avoided by conducting static test fires after crops are harvested or providing extra irrigation to soil impacted with the TFS.

The ability of a propulsion system is very crucial for the capability of a missile or a rocket system. Unlike liquid propellant rocket motors, the only control mechanism of the thrust value is the propellant geometry in solid propellant rocket motors. When the operation of solid propellant rocket motor has started, it cannot be stopped anymore. For this main reason the advance of dual pulse motor technology has started. The aim of this study is to investigate the geometrical effects of design parameters on the flow behavior of a dual pulse solid propellant rocket motor by using commercial Computational Fluid Dynamics (CFD) methods. For the CFD analysis, a generic dual pulse rocket motor model is constituted. Within this model, initially four different geometry alternatives of Pulse Separation Device (PSD) are analyzed. To begin PSD analyses, mesh sensitivity analyses are performed on one PSD geometry alternative. By defined grid size, the analyses of PSD geometry alternatives are performed. Computed results were compared in terms of flow behavior (flow streamlines, velocity distribution, turbulent kinetic energy&hellip / etc.) with each other. With the selected PSD geometry alternative the effects of L/D ratio (Length/Diameter ratio) of first pulse chamber, Achamb/APSD ratio (Chamber area/PSD opening area) and APSD/Ath ratio (PSD opening area/Throat area) on the flow behavior is investigated. Flow analyses are performed by simulating the unsteady flow of second pulse operation. With the performed analyses, it is aimed to identify generic geometric definitions for a dual pulse rocket motor.

To be competitive in the market, it is very important to design cost effective and reliable products. For this purpose, it is necessary to consider reliability as an integral part of the design procedure. Therefore, reliability which is a design parameter that affects cost and safety of a system should be taken into consideration in early phases since it is very difficult to change design at the later phases. Reliability of a rocket motor can be evaluated by reliability testing but these tests are very expensive and difficult since the tests are destructive and test sample size is determined by the binomial law. Because of the difficulties in reliability testing, in early design phases reliability can be evaluated by using reliability prediction results. This thesis report includes application of probabilistic approach for a solid rocket motor design to evaluate its reliability in preliminary design phase. In this study, it is aimed to assess the solid rocket motor ballistic performance reliability and casing structural reliability, determine important parameters affective on the solid rocket motor reliability and find a new design point to improve the reliability. Variations in dimensions and material properties are considered as the sources of failures and the limit states for acceleration, total impulse and maximum stress in the casing are approximated with response surface method by considering these variations. With the response surface functions, Monte Carlo simulation is used to assess failure probability and distributions of the rocket motor performance. Besides the assessment of the reliability, capability of the response surface functions to estimate the rocket motor performance and effects of the input parameters on the rocket motor performance and performance variation are also examined. By considering the effect of the input parameters, a new design point is proposed to decrease the total probability of failure.

Design process of a solid rocket motor with the objective of meeting certain mission requirements can be specified as a search for a best set of design parameters within the overall design constraints. In order to ensure that the best possible design amongst all achievable designs is being achieved, optimization is required during the design process. In this thesis, an optimization tool for internal ballistic design of solid rocket motors was developed. A direct search method Complex algorithm is used in this study. The optimization algorithm changes the grain geometric parameters and nozzle throat diameter within the specified bounds, finally achieving the optimum results. Optimization tool developed in this study involves geometric modeling of the propellant grain, burnback analysis, a 0-dimensional ballistic performance prediction analysis of rocket motor and the mathematical optimization algorithm. The code developed is verified against pretested rocket motor performance.

In this study, a detailed service life assessment procedure for solid propellant rockets under random environmental temperature and transportation loads is introduced. During storage and deployment of rocket motors, uncontrolled thermal environments and random vibratory loads due to transportation induce random stresses and strains in the propellant which provoke mechanical damage. In addition, structural capability degrades due to environmental conditions and induced stresses and strains as well as material capability parameters have inherent uncertainties. In this proposed probabilistic service life prediction, uncertainties along with degradation mechanisms are taken into consideration. Vibration loads are accounted by utilizing acceleration spectral density values which are induced during various deployment scenarios of ground, air and sea transportation. Furthermore, thermal loads are represented with a mathematical model being a harmonic function of time. Throughout the finite element analyses, a linear viscoelastic material model is to be used for the propellant. Change in the structural capability of the propellant with time is calculated using Laheru&#039 / s cumulative damage model. Moreover, to include aging effect of the propellant, Layton model is used. To determine the effects of induced stress and strains under variations and uncertainties in the random loads and material constants, mathematical surrogate models are constructed using response surface method. Limit state functions are utilized to predict failure modes of the solid rocket motor. First order reliability method is used to calculate reliability and probability of failure of the propellant grain. With the proposed methodology, instantaneous reliability of the propellant grain is determined within a confidence interval.

Propõe-se aqui uma metodologia de projeto aero-termo-estrutural de motores foguete a propelente sólido. O projeto de um motor foguete deve ser realizado com o objetivo de cumprir requisitos de uma missão. Para cada veículo espacial, com uma nova missão, um novo motor pode ser projetado, necessitando para isso de uma série de ferramentas robustas, capazes de compreender todas as combinações de esforços existentes no funcionamento de um motor, sob condições de altas pressões e temperaturas. A metodologia aqui proposta é testada e validada em bancada de ensaios desenvolvida para este fim. Os resultados obtidos mostram que a metodologia utilizada se aproxima bastante dos resultados teóricos e pode ser ajustada por coeficientes de eficiência com grande facilidade. / It is proposed here an aero-thermo-structural design methodology for solid propellant rocket motors. The design of a rocket motor must be carried out in order to fulfill requirements of a mission. For each new space vehicle, with a new mission, a new motor can be designed, requiring for it a variety of robust tools, able to comprise all combinations of load existing in the operation of a motor under high pressures and temperatures. The methodology proposed here is tested and validated in bank of tests developed for this purpose. The results show that the methodology is very close to the theoretical results and can be adjusted by coefficients of efficiency with great ease.