Much Research on composite solid propellants has been performed over the past few decades and much progress has been made, yet many of the fundamental processes are still unknown, and the development of new propellants remains highly empirical. Ways to enhance the performance of solid propellants for rocket and other applications continue to be explored experimentally, including the effects of various additives and the impact of fuel and oxidizer particle sizes on burning behavior. One established method to measure the burning rate of composite propellant mixtures in a controlled laboratory setting is to use a constant-volume pressure vessel, or strand burner. To provide high-pressure burn rate data at pressures up to 360 atm, the authors have installed, characterized and improved a strand burner facility at the University of Central Florida. Details on the facility and its improvements, the measurement procedures, and the data reduction and interpretation are presented. Two common HTPB/AP propellant mixtures were tested in the original strand burner. The resulting burn rates were compared to data from the literature with good agreement, thus validating the facility and related test techniques, the data acquisition, data reduction and interpretation. After more than 380 successful recordings, an upgraded version of the strand burner, was added to the facility. The details of Strand Burner II, its improvements over Strand Burner I, and its characterization study are presented.

A labscale hybrid rocket motor test stand has been developed for research at Cal Poly. The primary focus of research using this rig has been the development of regenerative cooling techniques using nitrous oxide as coolant and oxidizer, as well as validation of technologies relating to the annular aerospike nozzle. In order to prevent undesirable deflection of the cantilevered spike, a structural stiffening web, referred to as “The Spider,” is proposed. The Spider resembles a three-spoked wheel, with the aerospike held by the inner hub and the chamber walls abutting the outer radius. The Spider, placed upstream of the nozzle, is subject to thermal loads due to radiation and convection from the gases, and conduction from the outer annulus, as well as mechanical loads from thermal expansion and gas flow. Simulation tools are developed in three phases to produce an accurate model of the spatio-temporal distribution of these loads. A prototype of the Spider instrumented with thermocouple probes is designed, manufactured, and subjected to a series of hotfire tests. Results from three experimental runs are gathered and compared to simulated results. Good agreement is shown for the most part between the two datasets, with a single noticeable discrepancy for one measured temperature location. The high fidelity in the mean rate of temperature change for all stations indicates that the convective heat load is accurately modeled. The simulation results, confirmed by experiment, indicate that in order for the Spider to survive in the steady-state during an actual burn, an active cooling strategy is necessary. Two actively cooled concept designs are presented and discussed, and future avenues of research are suggested.

Hybrid rocket propulsion technology shows promise for the next generation of sounding rockets and small launch vehicles. This paper seeks to provide details on the process of developing hybrid propulsion systems to the academic and amateur rocket communities to assist in future research and development. Scaling hybrid rocket motors for use in sounding rockets has been a challenge due to the inadequacies in traditional boundary layer analysis. Similarity scaling is an amendment to traditional boundary layer analysis which is helpful in removing some of the past scaling challenges. Maintaining geometric similarity, oxidizer and fuel similarity and mass flow rate to port diameter similarity are the most important scaling parameters. Advances in composite technologies have also increased the performance through weight reduction of sounding rockets through and launch vehicles. Technologies such as Composite Overwrapped Pressure Vessels (COPV) for use as fuel and oxidizer tanks on rockets promise great advantages in flight performance and manufacturing cost. A small scale COPV, carbon fiber ablative nozzle and a N class hybrid rocket motor were developed, manufactured and tested to support the use of these techniques in future sounding rocket development. The COPV exhibited failure within 5% of the predicted pressure and the scale motor testing was useful in identifying a number of improvements needed for future scaling work. The author learned that small scale testing is an essential step in the process of developing hybrid propulsion systems and that ablative nozzle manufacturing techniques are difficult to develop. This project has primarily provided a framework for others to build upon in the quest for a method to easily develop hybrid propulsion systems sounding rockets and launch vehicles.

An Axisymmetric Rocket Ejector Simulation (ARES) was developed to numerically analyze various configurations of an air augmented rocket. Primary and secondary flow field visualizations are presented and performance predictions are tabulated. A parametric study on ejector geometry is obtained following a validation of the flow fields and performance values. The primary flow is calculated using a quasi-2D, irrotational Method of Characteristics and the secondary flow is found using isentropic relations. Primary calculations begin at the throat and extend through the nozzle to the location of the first Mach Disk. Combustion properties are tabulated before analysis to allow for propellant property selection. Secondary flow calculations employ the previously calculated plume boundary and ejector geometry to form an isentropic solution. Primary and secondary flow computations are iterated along the new pressure distributions established by the 1D analysis until a convergence tolerance is met. Thrust augmentation and Specific Impulse values are predicted using a control volume approach. For the validation test cases, the nozzle characteristic net is very similar to that of previous research. Plume characteristics are in good agreement but fluctuate in accuracy due to flow structure formulation. The individual unit processes utilized by the Method of Characteristics are found to vary their outputs by up to 0.025% when compared to existing sources. Rocket thrust and specific impulse are increased by up to 22% for a static system and 15% for an ejector flow at Mach 0.5. Evidence of Fabri conditions were observed in the flow visualization and graphically through the performance predictions. It was determined that the optimum ejector divergence angle for an air augmented rocket greatly depends on the stagnation pressure ratio between the primary and secondary flows.

Cold-flow tests were conducted in mediums of air and water to investigate the internal flow field about the nozzle region of a proposed solid rocket motor (SRM) configuration that would potentially replace the current external boosters on NASA's Space Shuttle. One-eighth-scale clear acrylic models of the proposed submerged aft-dome and aft finned grain elements were constructed to simulate the aft segment of the SRM at ignition and 35 seconds into the firing sequence. Pressure, velocity, and turbulence profiles were obtained during cold air testing, while air bubbles and dye were used for flow visualization during water tunnel testing. The flow visualization experiments indicated the presence of strong inlet vortices, alternating vortex shedding from both grain models' fins, circumferential flow in the aft-dome and around the nozzle, and recirculatory flow in the aft-dome and near an upstream portion of the 35-second grain model. Data acquired during cold air testing showed a turbulent low-velocity flow field in the aft-dome for both grain models. With respect to pressure and mean velocity virtually the entire nozzle/aft-dome region exhibited a minimal sensitivity to nozzle vectoring. / Master of Science

In this thesis a Large Eddy Simulation (LES) of a hybrid rocket engine burning ethylene (C2H4) in nitrous oxide (N2O) is explored. This is done primarily using a solver and solution scheme provided by the Swedish Defence Research Agency (FOI) and an (at this date) unpublished chemistry model. This sheds light on some transiet behaviour of a prior experiment conducted with a model engine that the simulation was based on. Due to time constraints the simulation did not cover the full test of the engine. The results confirm predictions from the experiment that the propellant was fuel rich. Some insight on how oxidizer swirl propagates throughout the engine was discovered as well.

Electrically controllable solid propellants are an area of interest as a viable solution to the lack of throttle-ability in solid propellant rocket motors. Existing studies have focused on propellants compositions using hydroxyl-ammonium nitrate, ammonium nitrate, or lithium perchlorate as oxidizers. Additionally, the thermochemical and electrochemical reaction mechanisms have not yet been fully defined. The research in this thesis explores the nitrate and perchlorate oxidizer families to compare their cation-anion relationships. Using these oxidizers, pseudo electrically controllable solid propellant compositions were created with the addition of multi-wall carbon nanotubes to enhance ohmic heating capabilities. These additives were selected based on theory that with a non-complexing polymer, an oxidizer melt layer is required for ions to dissociate and electrically controlled ignition to occur. Using an applied voltage, ignition delay and current draw experiments were performed to expand on prior findings that ignition delay follows oxidizer melt temperature while mobility is associated with the size of the ionic radii. Additionally, neat oxidizer pellets were electrically decomposed to determine their linear regression rate. These results help to characterize the mechanism of reaction. This advances the knowledge of oxidizers in electrically controllable applications. / Master of Science / Solid propellant rocket motors have been extensively studied and used in both space and military applications because they do not use air as the source of oxygen. Their main limitation is the lack of throttle-ability, or inability to control propellant burning. This is because solid propellants, which are generally composed of an ionic oxidizer salt, a polymer fuel, and additives, are pre-combined and stored within the rocket motor. An emerging viable solution to this limitation is electrically controllable solid propellants. With an applied voltage, the oxidizer is heated and melts, allowing ions to dissociate and current to flow between electrodes. This reaction can then be controlled by turning the power supply on and off. Cations, or ions which have a net positive charge, move to the negatively charged cathode while anions, which have a net negative charge, move to the negatively charged anode. The research in this thesis explores different cation-anion oxidizer pairings using both a propellant composition and as a pure oxidizer under an applied voltage. The results help to characterize the mechanism of reaction of each oxidizer in an electrically controllable context and determine their effectiveness in these propellant applications.

Since the end of World War II the world has seen several nations expand into the space age. Also after the Second World War, the Cold War began and many nations found themselves allying themselves with either the hegemony of the West or the Communists. Space was no exception in this dilemma, as weaker nations began to develop their own indigenous space programs and had technological diffusion from one of the hegemonies. Japan and India are two democracies that both sought support for their indigenous space programs from the west, particularly from the U.S. These two nations emerged from poverty and a broken infrastructure during the 1950s and have grown over the last sixty years into two of the most advanced space-faring nations in the world. These two nations have overcome several external and internal factors ranging from Communist expansion to bureaucratic strife. Japan and India have been and remain the two leading democratic nations in Asia that have risen to the rank of space power.

This work presents a preliminary design of a liquid rocket engine test platform to support research into liquid propulsion systems and rocket engine components, including injectors, ignition systems, combustion chambers and engine cooling systems. The liquid propellants, specifically liquid oxygen and ethanol, are pressure-fed using gaseous nitrogen. The test platform supports engine thrust values up to 1 kN, as well as varying oxidizer/fuel ratios up to 4.0 and varying ethanol concentrations between 70 and 100%. The test platform will integrate with a mobile control centre, which was designed concurrently, and provides remote control of the test procedures and data acquisition of all relevant pressure, temperature, mass flow and thrust data. The propellant feed assembly can support both cold and hot fire testing campaigns and is equipped with numerous safety features including inert gas purge lines, emergency drain lines and emergency shut-down and de-pressurization procedures.

Solid rocket motors consist of a thin metallic or composite shell filled with a soft rubbery propellant. Such motors are vulnerable and prone to buckling due to sudden external pressures produced by nearby detonation. The stability conditions of rocket motors subjected toaxisymmetric, external pressure loading are examined. The outer cases of motors are considered as isotropic (metallic) or anisotropic (composite), thin and high-strength shells, which are the main structures of interest in the stability analyses. The inner, low-strength elastic cores are modeled as linear and nonlinear elastic foundations. A general, refined, Sanders' nonlinear shell theory, which accounts for geometric nonlinearity in the form of von Karman type of nonlinear strain-displacement relations, is used to model thin-walled, laminated,composite cylindrical shells. The first order shear deformable concept is adopted in the analyses to include the transverse shear flexibility of composites. A winkler-type of linear and nonlinear elastic foundation is applied to model the internal foundations. Pasternak-foundation constants are also chosen tomodify the proposed elastic foundation model for the purpose of shear interactions. A set of displacement-based finite element codes have been formulated to determine critical buckling loads and mode shapes. The effect of initial imperfections on the structural responses are also incorporated in the formulations. A variety of numerical examples are investigated to demonstrate the validity and efficiency of the purposed theory under various boundary condiitions and loading cases. First, linear eigenvalue analysis is used to examine approximate buckling loads and buckling modes as well as symmetric conditions. An iterative solution procedure, either Newton-Raphson or Riks-Wempner method is employed to trace the nonlinear equilibrium paths for the cases of stress, buckling and post-buckling analyses. Both ring and shell-type models are applied for the structural analyses with different internal elastic foundations and initial imperfections. / Ph. D.