This thesis describes a novel approach to augmenting the combustion performance of hybrid rocket fuels, specifically in terms of regression rate and combustion efficiency. Liquid additives are emulsified into molten paraffin wax using nonionic surfactants and cured to form cylindrical fuel grains. Fuel grains were tested in a lab scale, optically accessible hybrid rocket motor and compared to the performance of neat paraffin fuel grains.

<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>

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.

Historically inducer pumps have been designed with low inlet diffusion that allows for a gradual pressure rise through the machine that has the ability to slowly collapse any cavitation bubbles that may be present. A novel cavitation control device has been developed by researchers at ConceptsNREC that has been shown in previous experimental work to greatly improve the suction performance of a traditionally designed machine. Computational fluid dynamics (CFD) has been employed to understand the effectiveness of the cavitation control device (CCD) at controlling the conditions that lead to cavitation inception and to determine the impact that the CCD has on the flow. Also the upper limit of design incidence ratio where the CCD is no longer able to control the factors that lead to cavitation inception was to be determined through the CFD approach. All machine geometries and test data were provided by researchers at ConceptsNREC. Two cases were selected for validation work and 32 additional designs were employed in a parametric study where the flow coefficient and design incidence ratio were varied over a typical range of interest for a turbopump application. The results of this computational work show that the CCD is able to control the factors that lead to early cavitation inception. The research shows that the addition of the CCD has an overall stabilizing affect on the flow by significantly decreasing the incidence at the leading edge of the blade. It has been determined that the maximum design incidence ratio where the CCD is able to effectively control the factors that lead to cavitation inception is dependent on the flow coefficient and in general the maximum design incidence ratio decreases as the flow coefficient is increased.

An alternative pressure-sensor based method for estimating the acoustic intensity, the phase and amplitude gradient estimation (PAGE) method, is presented. This method is similar to the finite-difference p-p (FD) method, in which the intensity is estimated from pressure measurements made using an array of closely spaced microphones. The PAGE method uses the same hardware as the FD method, but does not suffer from the frequency-dependent bias inherent to the FD method. Detailed derivations of the new method and the traditional FD method are presented. Both methods are then compared using two acoustic fields: a plane wave and a three monopole system. The ability to unwrap the phase component of the PAGE method is discussed, which leads to accurate intensity estimates above previous frequency limits. The uncertainties associated with both methods of estimation are presented. It is shown that the PAGE method provides more accurate intensity estimates over a larger frequency bandwidth. The possibility of using a higher-order least-squares estimation with both methods is briefly demonstrated. A laboratory experiment designed to validate the PAGE method was conducted. The preliminary results from this experiment are presented and compared to analytical predictions. Finally, the application of the PAGE method to a static rocket test firing is presented. The PAGE method is shown to provide accurate intensity estimates at frequencies that are higher than possible with just the FD method.

This research investigated the 3D printing and burning of fuel grains with complex geometry and the development of software capable of modeling and predicting the regression of a cross-section of these complex fuel grains. The software developed did predict the geometry to a fair degree of accuracy, especially when enhanced corner rounding was turned on. The model does have some drawbacks, notably being relatively slow, and does not perfectly predict the regression. If corner rounding is turned off, however, the model does become much faster; although less accurate, this method does still predict a relatively accurate resulting burn geometry, and is fast enough to be used for performance-tuning or genetic algorithms. In addition to the modeling method, preliminary investigations into the burning behavior of fuel grains with a helical flow path were performed. The helix fuel grains have a regression rate of nearly 3 times that of any other fuel grain geometry, primarily due to the enhancement of the friction coefficient between the flow and flow path.

The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.

Aircraft-based space-launch platforms allow operational flexibility and offer the potential for significant propellant savings for small-to-medium orbital payloads. The NASA Armstrong Flight Research Center’s Towed Glider Air-Launch System (TGALS) is a small-scale flight research project investigating the feasibility for a remotely-piloted, towed, glider system to act as a versatile air launch platform for nano-scale satellites. Removing the crew from the launch vehicle means that the system does not have to be human rated, and offers a potential for considerable cost savings. Utah State University is developing a small throttled launch-assist system for the TGALS platform. This "stage zero" design allows the TGALS platform to achieve the required flight path angle for the launch point, a condition that the TGALS cannot achieve without external propulsion. Throttling is required in order to achieve and sustain the proper launch attitude without structurally overloading the airframe. The hybrid rocket system employs gaseous-oxygen and acrylonitrile butadiene styrene (ABS) as propellants. This thesis summarizes the development and testing campaign, and presents results from the clean-sheet design through ground-based static fire testing. Development of the closed-loop throttle control system is presented.