Composite propellants are composed of a solid oxidizer that is mixed into a hydrocarbon binder that when polymerized results in a solid mass capable of selfsustained combustion after ignition. Plateau propellants exhibit burning rate curves that do not follow the typical linear relationship between burning rate and pressure when plotted on a log-log scale, and because of this deviation their burning behavior is classified as anomalous burning. It is not unusual for solid-particle additives to be added to propellants in order to enhance burning rate or other properties. However, the effect of nano-size solid additives in these propellants is not fully understood or agreed upon within the research community. The current project set out to explore what possible variables were creating this result and to explore new additives. This thesis contains a literature review chronicling the last half-century of research to better understand the mechanisms that govern anomalous burning and to shed light on current research into plateau and related propellants. In addition to the review, a series of experiments investigating the use of nanoscale TiO2-based additives in AP-HTPB composite propellants was performed. The baseline propellant consisted of either 70% or 80% monomodal AP (223 μm) and 30% or 20% binder composed of IPDI-cured HTPB with Tepanol. Propellants’ burning rates were tested using a strand bomb between 500 and 2500 psi (34.0-170.1 atm). Analysis of the burning rate data shows that the crystal phase and synthesis method of the TiO2 additive are influential to plateau tailoring and to the apparent effectiveness of the additive in altering the burning rate of the composite propellant. Some of the discrepancy in the literature regarding the effectiveness of TiO2 as a tailoring additive may be due to differences in how the additive was produced. Doping the TiO2 with small amounts of metallic elements (Al, Fe, or Gd) showed additional effects on the burning rate that depend on the doping material and the amount of the dopant.

The design, development, integration and testing of a full-featured, Lab-Scale Hybrid Rocket Engine was not only envisioned to be the chosen method of putting student payloads into space, but to be an invaluable teaching resource. The subject of the present thesis is the analysis, design, development, integration and demonstration of a lab-scale hybrid rocket motor. The overarching goal of this project was to establish a working developmental lab model from which further research can be accomplished. The lab model was specifically designed to use a fuel source that could be studied in normal laboratory conditions. As such, the rocket engine was designed to use Hydroxyl Terminated Polybutadiene as the fuel and Liquid Nitrous Oxide as the oxidizer. Developing the rocket engine required the usage of several electronics modules and a software package. The custom-designed electronics modules were a Signal Conditioning & Data Amplification Interface and a Data Acquisition Network. The software package was coded in Visual Basic (VB). A MathCAD regression rate computer model was designed and written to geometrically constrain the engine design. Further, the computer model allowed for the "what-if" situations to be evaluated. Using ProPep, solutions to the Equilibrium Thermodynamics Equations for the fuel and oxidizer mixture were obtained. The resultants were used as initial input to the computer model for predicting the lab-scale rocket's Chamber Pressure, Chamber Temperature, Ratio of Specific Heats and Molecular Weight. Details on the model, the rocket hardware, and the successful test firing are provided.

Hydroxyl terminated polybutadiene (HTPB) based polyurethanes (PUs) are low modulus materials and degrade easily at low temperature. Polycarbodiimide (PCDI) and polysiloxane (PSi) are reactive-type fillers when formed by carbodimidzation and sol-gel process, respectively. During the combustion, PCDI and PSi give off non-toxic, non-corrosive volatile gases, and finally form carbonaceous and siliceous chars. In this study, modified PUs were prepared by incorporating PCDI or PSi into PUs to give high carbon, nitrogen and silicon materials. These modified PUs are kinds of organic-inorganic hybrids with higher modulus and higher thermal stability than HTPB-based PUs. In addition, new silicone based insulation materials were prepared by mixing two silicone rubber materials LSR-2670 and RTV-627 from GE Silicones, in order to improve the heat insulation and to reduce the ablation rate. These inhibitors can keep the rocket motor from the high temperature ablation for a long time, especially castable silicone based heat insulations for the case of the ramjet engines. The mechanical properties at room temperature and the thermal stability of these modified PUs and silicone rubbers were investigated using a tensile tester and a thermogravimetric analyzer (TGA). ATR/FTIR (Attenuated total reflectance / Fourier transform infrared) technique is applied to monitor the synthesis process of PCDI and to examine the change of surface chemistry of insulator before and after thermal degradation via TGA. TGA coupled with FTIR (TGA/FTIR) was used to analyze the kinetics and the mechanism of thermal degradation under nitrogen and/or air. The Friedman and Kissinger methods of analysis were used for calculating the activation energy of degradation from dynamic TGA. The modified PUs (HIPTD-40%Psi¤ÎHIPTD-30%PMPS-PSi) with average activation energy of 88 and 112 kcal/mole (0.5¡Õ£\¡Õ0.9, under N2) and the modified silicone rubber (LR-5%HTB) with activation energy of 46.2~67.0 kcal/mole (0.1¡Õ£\¡Õ0.9, under N2) and 34.0~59.1 kcal/mole (0.1¡Õ£\¡Õ0.9, under air).The maximum degradation temperature (Tmax) and char yield (CY) of thermal degradation were estimated from a series of experiments with heating rates of 1, 3, 5, 10, 20, 30, 40 and 50 ¢J/min, under nitrogen or air. It is apparent that the maximum degradation temperature is dependent on heating rate. By assuming the heating rate for the insulator used in a rocket operating environment is about 5000¢J/min, Tmax calculated for the modified PUs (HIPTD-40%PSi and HIPTD-30%PMPS-PSi under N2) are found as 538 and 562¢J and for the modified silicone rubber (LR-5%HTB under N2 and air) are found as 576 and 562¢J, respectively. CY calculated for the modified silicone rubber (LR-5%HTB under N2 and air) is found as 71.5% and 66.2%. The morphology of modified PUs and silicone rubbers before and after thermal degradation via TGA was observed by optical and scanning electron microscope (SEM).

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Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

This document details original numerical studies performed by the author pertaining to solid propellant combustion. Detailed kinetic mechanisms have been utilized to model the combustion of the pseudo-propellants RDX/GAP and AP/HTPB. A particle packing model and a diffusion flame model have been utilized to develop a burning rate and an aluminum agglomeration model. The numerical model for RDX/GAP combustion utilizes a "universal" gas-phase kinetic mechanism previously applied to combustion models of several monopropellants and pseudo-propellants. The kinetic mechanism consists of 83 species and 530 reactions. Numerical results using this mechanism provide excellent agreement with RDX and GAP burning rate data, and agree qualitatively with RDX/GAP pseudo-propellant data. The numerical model for AP/HTPB combustion utilizes the same universal mechanism, with chlorine reactions added for modeling AP combustion. Including chlorine, there are 106 species and 611 reactions. Global condensed-phase reactions have been developed for six AP percentages between 59% and 80% AP. The AP/HTPB model accurately predicts burning rates, as well as temperature and species profiles. The numerical burning rate model utilizes a three-dimensional particle-packing model to generate cylindrical particle packs. Particle-size distributions have been modeled using a three-parameter lognormal distribution function. Pressure-dependent homogenization has been used to capture pressure effects and reduce cpu time. A "characteristic" burning path is found through each particle pack. Numerical results showed that different path-finding approaches work better depending on the propellant formulation and combustion conditions. Proposed future work and modifications to the present model are suggested. The numerical agglomeration model utilizes the same particle packing model and particle-size distribution function as in the burning rate model. Three preliminary models have been developed examining the ideas of pockets, separation distance, and aluminum ignition. Preliminary model results indicate the importance of predicting aluminum particle ignition. In the final model, the surface is regressed numerically through each particle pack. At each surface location, calculations are performed to determine whether aluminum particles combine and/or ignite. Ignition criteria have been developed from the results of the diffusion flame model and an analysis of particle-pack cross-sections. Numerical results show qualitative agreement with each experimentally observed trend. Proposed future work and modifications to the present model are suggested.

<p>In an AP-HTPB propellant microstructure, the local strain rate depends on the AP crystal size and the material, while the local temperature rate depends on the impact velocity, AP crystal size, and the material.  Larger AP crystals lead to higher local strain rates and higher local temperature rates, which means hot spots are more likely to occur in AP-HTPB propellants with more large AP crystals.</p>

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Hydroxyl-terminated polybutadiene (HTPB) is a common solid rocket fuel that is used in solid-fueled ramjets (SFRJs) for hypersonic propulsion. Renewed interest in hypersonic systems has developed a need to simulate HTPB combustion using computational fluid dynamics (CFD). However, the current chemical kinetic mechanisms lack the experimental data necessary to accurately simulate SFRJ conditions. Shock tube experiments at the University of Central Florida (UCF) were carried out to study the interaction of HTPB pyrolysis products (ethylene, propene, 1,3-butadiene, benzene, and toluene) with air. Experiments were conducted at 5 ± 0.5 atm and 1100 – 1800 K. Mixtures ranged from fuel lean to fuel rich (Φ = 0.5, 1.0, 4.0, 4.76) and contained individual fuel components or a fuel blend of the HTPB pyrolysis products. Species time-history of carbon monoxide (CO) and time-resolved temperature were measured with a quantum cascade laser (QCL) in the mid-infrared (MIR) region and compared against chemical kinetics models: AramcoMech3.0, HyChem, and NUIG 1.3. Additional OH* measurements were taken with an emission detector to capture ignition delay time (IDT). The data is used to develop an improved chemical kinetic mechanism for simulating HTPB combustion in air at SFRJ conditions.