Defects formed in solid rocket propellant during manufacturing, transportation, storage, and assembly can lead to alterations in the thrust time profiles and possibly catastrophic failure of the entire rocket. In order to determine the effects of temperature, loading rate, and thickness on this particulate composite, tests were conducted at three temperatures and two loading rates. Both uncracked and edge cracked "biaxial" specimens were produced from solid rocket propellant. The stress relaxation modulus and stress-strain data were obtained from load curves formed during "biaxial" tension tests. Near crack tip displacements and strains were calculated from photographs taken of a surface grating on the pre-cracked specimens during crack propagation. The effect of thickness, temperature, and loading rate on the stress intensity factor was also studied. Finally, by applying continuum theory the displacement singularity was determined at different stages of crack growth. From the stress strain data, it was found that temperature had a greater influence on behavior than loading rate over the ranges studied. The crack growth in the composite material consists of a series of crack opening, crack blunting, and crack growth/resharpening stages which are highly nonlinear. However, the thick specimen at low temperature did not follow this crack growth mechanism. At -65°F the thick specimen developed transverse constraints which caused a brittle fracture to occur when the specimen was loaded. Determination of the displacement singularity order for the sharp cracks was found to be consistent with the theoretical results predicted by Benthem. / Master of Science

Solid propellant research has mainly been directed towards more accurate characterization of the propellant material nature and more reliable structural analysis of the grain. Internal star grain design is among the most popular grain shapes that are used in today's propulsion system. Due to its complex geometry, stress concentrations are inevitably present around the highly curved area. Furthermore, this geometric effect together with various loading conditions throughout the grain's service life actually causes numerous defects inside its body. However, little is known concerning the three-dimensional fracture mechanism of the surface cracks which are the most common defects detected in the real rocket motor grain. After a brief evaluation of the current status of solid propellant research, stress analysis of a star grain model under internal pressure was performed by both photoelastic experiments and finite element calculations. These results illustrated the stress concentration effect around the star finger tip in addition to the global stress distribution across the whole section. Meanwhile, the deformation of the grain's outer surface was also obtained from the finite element results. A series of photoelastic experiments was conducted on cracked specimens with surface flaws emanating both on and off the axis of symmetry starting from the star finger tip. For the symmetric crack problem, cracks with different depths were intensively studied and the three-dimensional stress intensity factor (SIF) distribution was obtained for each test. These experimental data were further used to construct three analytical models, the "equivalent" radius model, the weight function model and the notch-root crack model, to expand the application range of the experimental data base so that a symmetric crack's SIF distribution with an arbitrary depth can be predicted. Moreover, surface cracks initiated off the axis of symmetry were also investigated by considering two off-axis angles. The crack shape and propagation path were achieved through a series of experiments and two methods were developed to effectively predict the possible crack growth path under sufficient pressure. The SIF distribution around the crack border was obtained for different offaxis angles and the factors that might influence the distributions were addressed based on the comparisons between the symmetric and asymmetric cracks, and the asymmetric cracks with different geometries. / Ph. D.

Nas últimas décadas, tem havido um crescente interesse pelo desenvolvimento de novos sistemas propulsivos que permitam conciliar baixo custo, reduzido impacto ambiental, menor tempo de desenvolvimento e maior segurança de operação. Dentro deste contexto, este trabalho teve como objetivo o desenvolvimento de um combustível para motor foguete, com baixa toxicidade e elevada densidade de empuxo, à base de etanol e monoetanolamina catalisada com diferentes materiais catalíticos e hipergólico com o peróxido de hidrogênio (H2O2). Primeiramente, foi desenvolvido um sistema para concentração do H2O2. Paralelamente, foram estudados os fatores físicos e químicos que influenciam o tempo de indução do par hipergólico e elaborado um programa experimental para avaliar a velocidade de ignição dos diferentes combustíveis preparados a partir da dissolução de catalisadores em monoetanolamina. Dentre os materiais catalíticos testados o nitrato de cobre foi aquele que apresentou o melhor desempenho. A proporção dos constituintes do combustível, ideal, foi de 61,0% de monoetanolamina e 30,1% de etanol e 8,9% de Cu(NO3)2.3H2O em massa. Finalmente, os resultados analíticos e experimentais geraram informações para a fabricação e testes de um propulsor de 50 N de empuxo teórico, operando com este combustível e com H2O2 90% como oxidante. Este estudo mostrou que a adição de etanol ao sistema reduz, significativamente, o atraso de ignição e aumenta o impulso específico do sistema. O custo destes propelentes é bem inferior àqueles empregados tradicionalmente em propulsão e o desempenho bastante similar, não sendo, entretanto, agressivos ao meio ambiente. / In recent decades, interest in development of new propulsion systems has grown. The new systems reconcile low cost, reduced environmental impact, quick development, and safer operation. The objective of this study was to develop a rocket fuel that was not highly toxic and had high thrust density. The fuel is based on ethanol and monoethanolamine; Di_erent catalysts and hypergolic materials were used with hydrogen peroxide (H2O2). While an H2O2 concentration system system was developed, the physical and chemical factors that inuenced the induction time of hyperbolic pairs were studied and an experimental program was developed that would evaluate the ignition speed of di_erent catalysts that were dissolved in monoethanolamine. Copper nitrate was the best catalyst of those tested. The ideal ratio of fuel components was 61.0% monoethanolamine to 30.1% ethanol to 8.9% Cu(NO3) 2.3H2O by mass. Finally, the experimental and analytical results generated the information needed for manufacture and testing of the thruster. The thruster could theoretically generate 50 N of thrust using the ideal fuel and 90% H2O2 as an oxidant. This study showed that adding ethanol to the system signi_cantly reduced ignition delay and increased the system\'s speci_c thrust. This fuel costs much less that those that are normally used in rockets and the performance if very similar. In addition, it causes less damage to the environment.

No description available.

No description available.

Historically, hypergolic propellants have utilized fuels based on hydrazine and its<br>derivatives due to their good performance and short ignition delays with the commonly used<br>hypergolic oxidizers. However, these fuels are highly toxic and require special handling<br><div>precautions for their use.</div><div><br></div><div>In recent years, amine-boranes have begun receiving attention as potential alternatives to<br>these more conventional fuels. The simplest of these materials, ammonia borane (AB, NH3BH3)<br>has been shown to be highly hypergolic with white fuming nitric acid (WFNA), with ignition<br>delays as short as 0.6 milliseconds being observed under certain conditions. Additionally,<br>thermochemical equilibrium calculations predict net gains in specific impulse when AB based<br>fuels are used in place of the more conventional hydrazine-based fuels. As such, AB may serve as<br>a relatively less hazardous alternative to the more standard hypergolic fuels.</div><div><br></div><div>Presented in this work are the results of five major research efforts that were undertaken<br>with the objective of developing high performance fuels based on ammonia borane as well as<br>characterizing their combustion behavior. The first of these efforts was intended to better<br>characterize the ignition delay of ammonia borane with WFNA as well as investigate various fuel<br>binders for use with ammonia borane. Through these efforts, it was determined that Sylgard-184<br>silicone elastomer produced properly curing fuel samples. Additionally, a particle size dependency<br>was observed for the neat material, with the finer particles resulting in ignition delays as short as<br>0.6 milliseconds, some of the shortest ever reported for a hypergolic solid fuel with WFNA.</div><div><br></div><div>The objective of the second area of research was intended to adapt and demonstrate a<br>temperature measurement technique known as phosphor thermography for use with burning solid<br>propellants. Using this technique, the surface temperature of burning nitrocellulose (a homogeneous solid propellant) was successfully measured through a propellant flame. During the<br>steady burning period, average surface temperatures of 534 K were measured across the propellant<br>surface. These measured values were in good agreement with surface temperature measurements<br>obtained elsewhere with embedded thermocouples (T = 523 K). While not strictly related to<br>ammonia borane, this work demonstrated the applicability of this technique for use in studying<br>energetic materials, setting the groundwork for future efforts to adapt this technique further to<br>studying the hypergolic ignition of ammonia borane.</div><div><br></div><div>The third research area undertaken was to develop a novel high-speed multi-spectral<br>imaging diagnostic for use in studying the ignition dynamics and flame structure of ammonia<br>borane. Using this technique, the spectral emissions from BO, BO2, HBO2, and the B-H stretch<br>mode of ammonia borane (and its decomposition products) were selectively imaged and new<br>insights offered into the combustion behavior and hypergolic ignition dynamics of ammonia<br>borane. After the fuel and oxidizer came into contact, a gas evolution stage was observed to<br>precede ignition. During this gas evolution stage, emissions from HBO2 were observed, suggesting<br>that the formation of HBO2 at the AB-nitric acid interface may help drive the initial reactant<br>decomposition and thermal runaway that eventually results in ignition. After the nitric acid was<br>consumed/dispersed, the AB samples began burning with the ambient air, forming a quasi-steady<br>state diffusion controlled flame. Emission intensity profiles measured as a function of height above<br>the pellet revealed the BO/BO2-based emissions to be strongest in the flame zone (corresponding<br>to the highest gas temperatures). Within the inner fuel-rich region of the flame, the HBO2 emission<br>intensity peaked closer to the fuel surface after which it unexpectedly began to decrease across the<br>flame zone. This is seemingly in contradiction to the current understanding that HBO2 is a stable product species and may suggest that for this system it is consumed to form BO2 and other boron oxides.</div><div><br></div><div>The fourth area of research undertaken during this broader research effort investigated the<br>use of ammonia borane and other amine borane additives on the ignition delay and predicted<br>performance of novel hypergolic fuels based on tetramethylethylenediamine (TMEDA). Despite<br>these materials being in some cases only sparingly soluble in TMEDA, solutions of ammonia<br>borane, ethylenediamine bisborane, or tetramethylethylenediamine bisborane in TMEDA resulted<br>in reductions of the mean ignition delays of 43-51%. These ignition delay reductions coupled with<br>the significantly reduced toxicity of these fuels compared to the conventional hydrazine-based<br>hypergolic fuels make them promising, safer alternatives to the more standard hypergolic fuels.<br>Attempts were made to improve these ignition delays further by gelling the TMEDA, allowing for<br>amine borane loadings beyond their respective solubility limits. Moving to these higher loadings<br>had mixed results however, with the ignition delays of the AB/EDBB-based fuels increasing<br>significantly with higher AB/EDBB loadings. The ignition delays of the TMEDABB-based fuels<br>on the other hand decreased with increasing TMEDABB loadings, though the shortest were still<br>comparable to those found with the saturated fuel solutions.</div><div><br></div><div>The final research area that was undertaken was focused on scaling up and developing fuel<br>formulations based on ammonia borane for use in a small-scale hypergolic hybrid rocket motor.<br>Characterization of the regression rate behavior of these fuels under motor conditions suggested<br>the fuel mass flow rate was driven primarily by the thermal decomposition of the ammonia borane.<br>This mechanism is fundamentally different from that which governs the regression rate of most<br>conventional solid fuels used in hybrid rockets as well as that of ethylenediamine bisborane, a<br>similar material in the amine borane family of fuels. Understanding this governing mechanism further may allow for its exploitation to enable high, nearly constant fuel mass flow rates<br>independent of oxidizer mass fluxes. If successful, this would enable further optimization of the<br>design for rocket systems utilizing these fuels, resulting in levels of performance that rival that of<br>the more conventional hydrazine-based fuels.<br></div>

need OCR / Master of Science

There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement for existing industry standard fuels for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, the NASA Marshall Space Flight Center, and the NASA Glenn Research Center each either recently concluded or currently has ongoing programs in the synthesis and development of these potential new propellants. In order to perform conceptual designs using these new propellants, most conceptual rocket engine powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. Very little thermophysical property data exists for most of these potential new HEDM propellants. Experimental testing of these properties is both expensive and time consuming and is impractical in a conceptual vehicle design environment. A new technique for determining these thermophysical properties of potential new rocket engine propellants is presented. The technique uses a combination of three different computational methods to determine these properties. Quantum mechanics and molecular dynamics are used to model new propellants at a molecular level in order to calculate density, enthalpy, and entropy. Additivity methods are used to calculate the kinematic viscosity and thermal conductivity of new propellants. This new technique is validated via a series of verification experiments of HEDM compounds. Results are provided for two HEDM propellants: quadricyclane and 2-azido-N, N-dimethylethanamine (DMAZ). In each case, the new technique does a better job than the best current computational methods at accurately matching the experimental data of the HEDM compounds of interest. A case study is provided to help quantify the vehicle level impacts of using HEDM propellants. The case study consists of the National Aeronautics and Space Administrations (NASA) Exploration Systems Architecture Study (ESAS) Lunar Surface Access Module (LSAM). The results of this study show that the use of HEDM propellants instead of hypergolic propellants can lower the gross weight of the LSAM and may be an attractive alternative to the current baseline hypergolic propellant choice.

The manufacturing process for solventless extruded double base propellants involves a number of rolling and reworking stages. Throughout these processes a decrease in weight average molecular weight was observed, this was attributed to denitration. Differential scanning calorimetery data indicated that the reworking stages of extruded double base propellant manufacture were crucial to the homogenisation of the propellant mixture. To determine the homogeneity of the final extruded product, a sample was analysed across its diameter. No variations in stabiliser concentration, molecular weight, or Vickers hardness were detected. An accelerated thermal ageing trial simulating up to 8 years of ageing at 25°C was carried out to evaluate the storage characteristics. Reductions in stabiliser concentration, number average molecular weight, weight average molecular weight and polydispersity compared with un-aged samples were observed. The glass transition temperature measured using differential scanning calorimetery decreased by ~3°C. The decrease was attributed to the initial denitration reducing the energy of bond rotation and shortening the polymer chains, both factors reducing the energy required for movement. Modulus values determined from dynamic mechanical analysis temperature scanning experiments, did not detect significant variation between un-aged and aged samples. Though it was considered that variations would be likely if a more extensive ageing program was completed. In order to evaluate propellant behaviour at very high and low frequencies, time temperature superposition (TTS) and creep testing were carried out. The TTS technique superpositioned data well, allowing future investigation of high frequency propellant properties. Creep testing was considered to be an appropriate approach, though the equipment available was not optimised for such testing. This thesis is concerned with understanding how propellants are manufactured from nitrocellulose, nitroglycerine and other constituents. It is also about how the propellants decompose during long periods of time in storage, and how these changes can be measured using thermal and mechanical methods. It is about how the physical, chemical and thermal properties of the propellant composition change throughout the manufacture. This is relevant as it could be used to develop more efficient manufacturing processes, allow operators to adjust processes to tailor product properties or be used to re-design manufacturing to compensate for a different starting material. The thesis also considers how and why the properties of the product change over the course of years of storage. A specific focus on whether changes in mechanical and thermal properties occur, and if so how they can be detected.

In 2004, President Bush addressed the nation and presented NASA's new vision for space exploration. This vision included the completion of the International Space Station, the retirement of the Space Shuttle, the development of a new crew exploration vehicle, and the return of humans to the moon by 2020. NASA's Exploration Systems Architecture Study (ESAS) produced a transportation architecture for returning humans to the moon affordably and safely. This architecture requires the development of two new Shuttle-derived launch vehicles, an in-space transportation vehicle, a lunar descent and landing vehicle, and a crew exploration vehicle for human transportation. The development of an in-space propellant transfer capability could greatly improve the performance, cost, mission success, and mission extensibility of the overall lunar architecture, providing a more optimal solution for future exploration missions. The work done in this thesis will analyze how this new capability could affect the current NASA lunar architecture, and will outline the value proposition of propellant refueling to NASA. A value proposition for propellant refueling will be provided to establish why an architecture that utilizes propellant refueling is better equipped to meet the goals of the Vision for Space Exploration than the current baseline design. The primary goal addressed in this research is the development of a sustainable and affordable exploration program. The value proposition will outline various refueling strategies that can be used to improve each of the architecture Figures of Merit. These include a decrease in the Life Cycle Cost of both the lunar and Mars exploration campaigns, the ability to more than double the mission payload that can be delivered to the lunar surface during cargo missions, improving the probability of successfully completing each lunar mission, decreasing the uncertainty, and therefore risk, experienced during the development process, and improving the extensibility of the exploration architecture by utilizing a greater portion of the lunar program for future crewed mission. The ability to improve these Figures of Merit provides NASA with a more valuable architecture because NASA is able to achieve a greater return on its large initial investment.