This thesis examines the impact of the physical contamination on optical surfaces of spacecraft by an ABS/GOX thruster. Plume contamination presents a significant operational hazard for spacecraft solar arrays and thermal control surfaces can lead to decreased power production and increased spacecraft temperatures. Historically, due to the lack of a reliable, on-demand, and multiple-use ignition methodology, hybrid rockets have never been previously considered for in-space propulsion. Recent advancements in hybrid rocket technologies, have made hybrid systems feasible for in space propulsion. However, prior to this study no research had ever been performed with regard to plume contamination effects due to hybrid rockets. This paper presents the results from a set of preliminary plume contamination measurements on a prototype small spacecraft hybrid rocket system, collected under both ambient and vacuum chamber conditions.

Three-dimensional, additive printing has emerged as an exciting new technology for the design and manufacture of small spacecraft systems. Using 3-D printed thermoplastic materials, hybrid rocket fuel grains can be printed with nearly any cross-sectional shape, and embedded cavities are easily achieved. Applying this technology to print fuel materials directly into a CubeSat frame results in an efficient, cost-effective alternative to existing CubeSat propulsion systems. Different 3-D printed materials and geometries were evaluated for their performance as propellants and as structural elements. Prototype "thrust columns" with embedded fuel ports were printed from a combination of acrylonitrile utadiene styrene (ABS) and VeroClear, a photopolymer substitute for acrylic. Gaseous oxygen was used as the oxidizer for hot-fire testing of prototype thrusters in ambient and vacuum conditions. Hot-fire testing in ambient and vacuum conditions on nine test articles with a combined total of 25 s burn time demonstrated performance repeatability. Vacuum specific impulse was measured at over 167 s and maximum thrust of individual thrust columns at 9.5 N. The expected ΔV to be provided by the four thrust columns of the consumable structure is approximately 37 m/s. With further development and testing, it is expected that the consumable structure has the potential to provide a much-needed propulsive solution within the CubeSat community with further applications for other small satellites.

Hypergolic hybrid motors are advantageous for rocket propulsion due to their simplicity, reliability, low weight, and safety. Many hypergolic hybrid fuels with promising theoretical performance are not practical due to their sensitivity to temperature or moisture. Ammonia borane (AB) has been proposed and studied as a potential hypergolic hybrid fuel that provides both excellent performance and storability. This study investigates the effect of droplet impact velocity, pellet composition, and storage humidity on ignition delay of AB with white fuming nitric acid as the oxidizer. Most ignition delays measured were under 50 ms with many under 10 ms and some even under 2 ms, which is extremely short for hybrid systems. Higher droplet velocities led to slightly shorter ignition delays, and exposing samples to humidity slightly increased ignition delay. An AB pellet composition of at least 20% epoxy binder was found to minimize ignition delay. The epoxy facilitates ignition by absorbing or adhering the oxidizer and slowing the reaction with the fuel, preventing oxidizer expulsion and holding it close to the fuel. These results emphasize the importance of binder properties in hypergolic hybrids. Pellets varying in composition and storage method were extinguished and reignited with the oxidizer to demonstrate reignition capability.

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.

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.

In the aerospace propulsion sector safety is an important driver to costs, vehicle design and mission capabilities. Hybrid rockets are considered some of the safest forms of chemical propulsion. That factor alone makes hybrid rocket propulsion systems desirable options. Hybrid systems often benefit from multiple additional advantages over conventional solid and liquid propellant systems, including: minimal environmental impact, higher density impulses, start-stop-restart capabilities, simplistic random throttle control, low development costs, and basic transportation and storage requirements. A major issue that continues to impact the effective use of hybrid systems, is that classical hybrid rocket fuels suffer in low regression rates. If fuel regression rates can be improved upon without diminishing any of the other beneficial factors to a hybrid rocket motor then a far greater market for such systems can be generated. In this work, additively manufactured polypropylene solid fuel grains were infused with gels as a means of significantly altering the fuel burning rates in a lab scale hybrid rocket motor. Gels based on Jet-A were created using both particulate (fumed silica, micro aluminum, nano aluminum) and polymeric (paraffin wax) gellants. The particle structure of the aluminum powders was characterized by means of microscopic imaging, particle size measurement, and thermal mass response analysis. The rheological behavior of the gels was characterized in order to determine the relationship between melt layer viscosity, viscoelastic properties, and combustion performance. High speed color video recording was used on select grains for spatially and temporally resolved three-color camera pyrometry analysis. The process showed promise in determining aluminized gel burn time across an entire rocket firing. The performance of the gel infused grains was compared to a traditional center perforated fuel grain, under similar flows of gaseous oxygen. Rocket motors fired with gel infused grains exhibited pressure increases of greater than 40%. Gel infused fuel grains demonstrated regression rate enhancements up to 90% higher than the baseline. The estimated gel regression rates were over 500% higher than the host polypropylene fuel. When the O/F was maintained near stoichiometric or lean conditions, c∗ efficiencies of the gel infused grains were similar to that of the baseline indicating thorough combustion of the gels. At low oxygen mass flows, the effects of gel infusion are not as significant, which is consistent with the liquefying fuel entrainment concept. / Master of Science / In the field of air and space flight, safety is an important driver to costs, vehicle design and mission capabilities. Hybrid rockets are considered some of the safest forms of vehicle lift systems compared to similar forms. That factor alone makes hybrid rockets desirable options. Hybrid systems often benefit from multiple additional advantages over similar systems often used, including: minimal environmental impact, greater force for a given time and volume of fuel, start-stop-restart capabilities, simplistic random motor control, low development costs, and basic transportation and storage requirements. A major issue that continues to impact the effective use of hybrid systems is that classical hybrid rocket fuels suffer in low burn rates. If fuel burn rates can be improved upon without diminishing any of the other beneficial factors to a hybrid rocket motor then a far greater market for such systems can be generated. In this work, specially manufactured solid fuel grains were combined with gels as a means of significantly altering the fuel burning rates in a small scale test setup. Gels based on a type of jet fuel were created using multiple gel forming and modifying materials. The structure of two types of small scale aluminum powders was characterized by means of microscopic imaging, particle size measurement, and weight response to thermal changes. Properties specific to the gels were characterized in order to determine performance relationships to individual material properties. High speed color video recording was used on select grains for space and time resolved three-color camera temperature analysis. The process showed promise in determining aluminized gel burn time across an entire rocket firing. The performance of the gel modified grains was compared to a traditional fuel grain design, under similar flows of gaseous oxygen. Rocket motors fired with gel modified grains exhibited pressure increases of greater than 40%. Gel modified fuel grains demonstrated burn rate enhancements up to 90% higher than the traditional fuel grain design. The estimated gel burn rates were over 500% higher than the host polypropylene fuel. When ideal conditions were maintained, fuel burn efficiencies of the gel modified grains were similar to that of the traditional fuel grain design indicating ideal burning of the gels. At low oxygen flow rates, the effects of gel addition are not as significant, which is consistent with an expectant similar concept.

This thesis set out to find a blend of thermoplastics that had better combustion properties than the current ABS (acrylonitrile butadiene styrene) plastic or “Lego TM plastic” used by Utah State University. The current work is in an effort to eliminate toxic propellants from small space applications. High and low density polyethylene plastics were used because they are common plastic waste items. In this way rocket fuel can be made from these items to reduce the waste found in landfills. Three plastics were considered for replacement and as mixture components with the ABS plastic, namely low and high density polyethylene, and high impact polystyrene. These plastics failed to have superior combustion properties when used in rockets designed to achieve 12 pounds or less of thrust compared to the current ABS plastic.

2011 IEEE Congress on Evolutionary Computation (CEC). June 5-8, 2011, Ritz-Carlton, New Orleans, LA, USA

The staged combustion aft-injected hybrid (SCAIH) rocket motor is a promising design for the future of hybrid rocket propulsion. Advances in computational fluid dynamics and scientific computing have made computational modelling an effective tool in design and development. The focus of this thesis is the numerical modelling of the SCAIH rocket motor in a turbulent combustion, high-speed, reactive flow accounting for solid soot transport and radiative heat transfer. The SCAIH motor has a shear coaxial injector with liquid oxygen injected centrally at sub-critical conditions: 150K, 150m/s (Mach≈0.9), and a gas-generator gas-solid mixture of one-third carbon soot by mass injected in the annual opening at 1175K, and 460m/s (Mach≈0.6). Flow conditions in the near injector region and the flame anchoring mechanism are of particular interest. Overall, the flow is shown to exhibit instabilities and the flame is shown to anchor directly on the injector faceplate with temperatures in excess of 2700K.