Theses supervised by Doç. Dr. Mustafa Arif Karabeyoğlu
13 theses · Koç University
Ignition capability of CO2 in hybrid rockets for Mars ascent vehicles
CO2 is a major combustion product arises from the combination of carbon as the fuel and oxygen as the oxidizer. However,CO2 can burn with metallic powders that makes it interesting compound for the combustion applications. Metals have high reactivity levels compared to carbon thus breaks the carbon oxygen bond.This reaction releases significant amount of energy. Although Metal/CO2 com-bustion has been studied by many researchers, process has not been evaluated for practical rocket applications. Therefore, this thesis aims to understand the ignition characteristics of the CO2 in hybrid rocket motors. Propellant combination of Paraffin/Metal/N2O/CO2 is selected to accomplish the combustion.Preliminary experiments are performed by using lab-scale hybrid motors. In addition, a scale up motor is used to better understand the ignition capability of the CO2. Micron sized aluminum and magnesium is casted within the paraffin wax asthe additive. Motor operating in "blowdown" mode uses N2O/CO2 oxidizer mixture in saturated liquid state. Results indicate that magnesium has better ignition capability with CO2 than aluminium. The maximum successful combustion is achievedup to 75% CO2 by weight percentage in the oxidizer mixture for 60%Mg by mass in the paraffin wax. CO2 slows down the chemical kinetics thus adiabatic flame temperature has significant role on ignitibility limit. Ignitions are quenched at Tflame below 1600 K for scale up motor.Paraffin/Mg/N2O/CO 2propellant combination is a feasible candidate for Mars Ascent Vehicles. In-situ CO2 and Mg usage for propulsion systems significantly reduces mass that is needed to brought from the Earth. Although high level of CO2 seems more efficient for in-situ MAV, the minimum mass is needed to be brought from the Earth is found at 60% CO2 80% Mg propellant combination.
Design & analysis and development of turbine system for hybrid rocket applications
Most hybrid rocket applications utilize high-pressure tanks to transfer their oxidizer into the combustion chamber. These applications are called pressure–fed and are considered the most straightforward technique for pressurizing oxidizer tanks. On the other hand, more complex systems such as liquid rocket engines employ a turbopump assembly to increase the pressure of the oxidizer and feed the combustion chamber of the launch vehicle. Although turbopump assembly adds more weight to the launch vehicle and brings mechanical complexity, the reduction in high pressure and oxidizer tanks' masses increases the launch vehicle's performance. Moreover, turbopumps deliver oxidizer at a constant mass flow rate and pressure. It is another advantage of using turbopumps in hybrid rockets since pressure-fed rockets suffer from "blowdown" phenomenon between high pressure and oxidizer tank. Within the scope of this thesis, a turbine for turbopump assembly is designed. Designed turbine systems for different scales are analysed in CFD software for further improvement. Additionally, hybrid rocket fundamentals and motor design parameters are introduced. Finally, the developed system is tested with water and liquid oxygen to validate meridional and CFD analysis. Performance analysis and obtained results clearly show that the usage of turbopumps in a hybrid rocket system improves the overall performance of hybrid rocket launch vehicles even though additional weights are added to the launch system.
Solenoid vana çoklu-fizik modeli ve optimizasyonu
There are several limitations when designing solenoid systems. With proper search space selection possible designs can be narrowed down lightest, fastest and most power efficient depending on the application. A method was developed starting from computing required coil design to overcome sealing forces, minimizing structural thicknesses to reduce weight and looking into the effect of manufacturing tolerances and clearances as a perturbation matrix. This was achieved by a multi-physics simulation where the magnetic circuit is modeled using magnetic reluctances of parts and clearances. The fluid flow was computed assuming ideal gas, 1D flow and adiabatic process.
Material performance determination for nozzle erosion in hybrid rocket motors
Performance parameters such as thrust coefficient (CF) and specific impulse (Isp) are critical for rocket motor design. An increase in the nozzle throat area causes a change in CF, thrust, and Isp. Therefore, no erosion or minimal erosion is expected. Nozzle erosion in hybrid rocket motors is a prominent factor that should be studied since hybrid rocket motors (HRM) are still developing. This study investigates the erosion rates of different materials at various oxidizer-to fuel ratios (O/F), equivalence ratios, and chamber pressures. Moreover, the effects of throat length on erosion were investigated as well. Performances of different graphite grades, metals, composites, and silicon carbide as nozzle materials were determined. The silicon carbide sintering procedure and the use of silicon carbide in hybrid rocket motors as nozzle material is one of the most important parts of this study due to being a research topic all by itself. There is only a few research on the use of SiC nozzles in HRM nozzles. Static firing test results are given, and erosion rates of nozzle throats are measured for each test. Nozzle throat diameters are measured before and after the tests using both calipers and image processing toolbox in MATLAB®. Results show that material selection and reduced O/F ratios are the most critical factors for lower erosion rates. Fine graphite grades may be good options for rocket nozzles, but silicon carbide is a promising material due to its high strength.
Design, analysis and testing of cryogenic centrifugal pump for hybrid rocket applications
Chemical rockets are one of the biggest and most critical sub-group of rocket systems and undertake the task of being backbone of the space industry. The dangerous working environment of solid fuels and the complexity and expensiveness of liquid propellants have turned the attention to hybrid rocket systems, which are simpler and more cost-effective solution among chemical rocket propulsion types. In hybrid rocket systems, the oxidizer which is in liquid phase must be transferred into the combustion chamber, where the solid fuel is located, with a certain flow rate and pressure. The feed system that will do this process is very crucial in rocket science as it will directly affect the thrust performance of the rocket system. The liquid oxidizer can be transferred to the combustion chamber using either high-pressure gas or a pump. In the high-pressure gas feeding method, the cold inert gas is used to pressurize the liquid oxidizer to the desired pressure and the pressurized oxidizer is pushed into the combustion chamber. Although this method is cheap and simple, it results in a rocket system with a very high structural weight, as the elements of the feeding system are exposed to high pressure. Due to its high structural weight, this feeding system is generally used in small and medium sounding rocket systems and is not suitable for flight for large systems because it increases the structural weight too much. Contrary to this feeding system, feeding the oxidizer with a pump will only leave certain elements under high pressure. Thus, the latter method reduces the structural weight and increases thrust-to-weight ratio of rocket systems. Even for large-scale systems the use of pump as a pressurizing element is necessary to make the rocket suitable for high ΔV missions. Additionally, thrust profile is improved because the blow-down effect is minimized with the pump supplying the oxidizer with a constant flow rate. The first goal of this thesis is to design a small-scale centrifugal pump to pressurize LOX to be used in a hybrid rocket system. This pump system, which is manufactured with appropriate techniques, has been tested with water and LOX and the results are compared with CFD results. After the reliability tests of the system, static hot-firing test is carried out with a small-scale hybrid motor with paraffin fuel and the performance of the hybrid engine system is evaluated. A new pump has been designed to be used in the feeding system of a large-scale hybrid rocket system with the experience gained from the small-scale pump design process. CFD analysis of the large-scale pump along with an appropriately sized hybrid rocket motor performance analysis have been performed. Another hybrid rocket system with a high-pressure gas feeding system with the same total impulse is designed to compare the performances of both pressurization system methods. Therefore, a detailed simulations of entire feeding systems have been made. The simulation results are compared to a pressure-fed hybrid rocket system with the same total impulse, and it is observed that the use of pumps in the feed system significantly reduces the structural weight and improves ΔV performance of the hybrid rocket.
Fuel grain modeling and optimization for hybrid propulsion systems using additive manufacturing
Hybrid propulsion systems have garnered significant interest in recent years due to their potential to offer a safer and cheaper alternative to traditional solid and liquid rocket engines. A key component of these systems is the fuel grain, which plays a vital role in the performance and efficiency of the propulsion system. However, the practical feasibility of hybrid propulsion has been hindered by the low regression rates. This thesis focuses on the improvement of regression rates in hybrid rocket fuel grains through the utilization of additive manufacturing. Acrylonitrile Butadiene Styrene (ABS) exhibits a lower regression rate than conventional fuels like paraffin or HTPB, making it a promising candidate for this study. Additive manufacturing allows the production of complex geometries that were previously unattainable with traditional manufacturing techniques. This study explores the impact of the grid infill structure with different densities on the performance of ABS fuel grains. Single port fuel grains with infill densities of 70\%, 90\%, and 100\% are fabricated using a 3D printer. A small-scale test setup was employed for the static fire tests of the fuel grains with nitrous oxide. The test results show that utilizing a lower infill ratio inside the fuel grain increases the regression rate significantly while ensuring combustion stability and structural integrity.
A novel system design for lunar orbital transfer vehicle with hybrid propulsion technology
The need for lunar missions has significantly increased in recent times, primarily due to the Moon being an accessible resource within reach. The Moon attracts attention with its significant elements, and most notably, its water resources. Additionally, it can serve as a crucial launch site for interplanetary transfers. The Moon's low gravity and lack of atmosphere make launches from its surface much more efficient. To realize these advancements, it is necessary to have a presence of numerous vehicles in lunar orbit and on the surface to deepen research. The main obstacles to the sustainability of lunar exploration are accessibility and cost. The payload capacities of current launchers for delivering payloads to the lunar orbit are quite limited. Achieving high-capacity access to the lunar transfer orbit requires additional stages. Therefore, orbital transfer vehicles, which have been implemented in Low Earth Orbit applications so far, can be utilized. Although these vehicles address the accessibility issue to some extent, they fail to provide a solution to the cost problem due to their propulsion systems. Propulsion systems used in these vehicles are expensive due to the fuel they consume and their development cost and handling issues. The main objective of this thesis is to examine a system that will pave the way for future lunar exploration missions. This thesis presents a comprehensive literature review on historical lunar missions and current orbit transfer vehicles. Within the scope of the thesis, an orbit propagator specifically developed for lunar transfer mission planning is introduced. The orbit propagator incorporates a genetic algorithm for mission optimization. Subsequently, tools for hybrid propulsion system sizing and orbit transfer vehicle performance analysis are presented. Finally, a conceptual orbital transfer vehicle is proposed using the developed tools within the thesis.
Novel manufacturing method for large scale and energetic material containing paraffin- based hybrid rocket fuel grains
The propulsion system designs are tailored according to the mission's requirements. Launching and sounding rockets demand high thrust, a condition feasible through the interaction of propellant and oxidizer at high flow rates. In addition to cost, safety, and operational ease considerations, high regression rate and competitive specific impulse make paraffin-based hybrid fuels a robust candidate for both launching and sounding rockets. Centrifugal and direct casting are the standard methods to manufacture paraffin-based fuel grains. The fuel melt's composition may influence the preference for a specific production method. For instance, metallic additives in the demanded fuel formulation make the centrifugal casting method unsuitable due to phase separation between polymeric melt and metallic powder induced by density differences. It makes the direct casting method the only option. However, the direct casting method introduces deviations in the fuel grain's dimensional tolerances. It leads to a substantial reduction in both mechanical properties and density due to exposed micro defects while the melt solidifies. This thesis aims to develop a novel fuel grain manufacturing method with distinctive advantages over classical centrifuge and direct fuel casting techniques. Additionally, it seeks to determine production parameters affecting mechanical properties, identify micro-mechanisms impacting mechanical characteristics, and elucidate specific scenarios wherein the new production method offers favorable outcomes. The fuel formulation is prepared and obtained in a liquid state in the advanced method. It is subsequently transformed into micron-sized particles by spraying into a reactor filled with inert gas through a nozzle. The resulting particles are then placed into molds designed to achieve the desired fuel grain shape and compressed using a uniaxial hydraulic system to obtain a homogeneous solid fuel grain. Within this framework, the thesis study involved the production of over 500 kg of hybrid rocket fuel powder, evaluating more than eight distinct mold designs, and the generation of over 400 samples for the investigation of mechanical and combustion characteristics. Tensile test specimens were extracted from the produced fuel cores, and their mechanical properties and fracture surfaces were examined using an electron microscope. Finally, fuel cores prepared from different formulations tested with appropriate hybrid rocket motors. The test results demonstrate that the new fuel production method facilitates the homogeneous incorporation of metallic additives characterized by a high-density difference into the fuel melt, enabling the utilization of additives requiring processing at low temperatures as fuel. Furthermore, a new additive is proposed for paraffin-based hybrid rocket fuels. The proposed additive, blowing agent, enhance the fuel regression rate up to %178 at 60 g/cm2s oxidizer flux compared to the base fuel formula's regression rate.
The ablation performance of phenolic and carbon matrix composites under the free jet conditions of hybrid propellant rocket
Ablative composite materials are commonly preferred in rocket nozzle regions where high temperature durability and shock resistance are required. In hybrid rocket motor applications, these materials are exposed directly to the hot gas products formed by the combustion reaction of fuel and oxidizer. Accordingly, ablation testing under the free-jet of lab-scale hybrid rocket motor would be flight-like simulation for the application of these materials in hybrid propellant rockets. As observed in the literature; Polymer Matrix Composites (PMCs), reinforced with silica and carbon fibers, can withstand short-term exposure to hot gases of combustion, whereas Ceramic Matrix Composites (CMCs), as Carbon/Carbon (C/C), exhibit exceptional durability without significant recession in similar conditions. In this work, chopped and 2D fabric reinforced phenolic matrix PMCs, and C/Cs were tested under the hot exhaust of paraffin fuel and gaseous oxidizer-based lab-scale hybrid rocket for 7 sec. In order to observe the oxygen-to-fuel ratio (O/F) effect on the ablation performances, each sample was tested in both oxygen-rich and fuel-rich firing conditions. Surface temperatures of samples were measured by two-colour IR pyrometer, also ablation rates were determined from the mass loss of samples. Test results show that phenolic matrix and C/C composites are significantly ablated by fiber-matrix debonding. Heat treatment of C/C samples in 1500 °C increased the ablation resistance by enhancing fiber-matrix bonding against hot gas exposure. For both phenolic matrix and C/C composite samples, chopped fiber reinforced samples propose better ablation resistance than 2D reinforcement. The variety of samples provides clear understanding about the ablation mechanisms of composite materials, specifically in the hybrid rocket applications.
Pyrophoric ignition system development for hybrid rocket motors in space applications
Hybrid motors are gaining attention in space applications due to their advantageous features like inherent safety and controllability, standing out against conventional solid and liquid engines. However, achieving rapid and reliable ignition remains a significant challenge for hybrid motors, particularly for their use in space systems. Among the various ignition methods, pyrophorics show promising characteristics. In this thesis, ignition delay characteristics of a pyrophoric liquid triethylaluminum (TEA) with nitrous oxide (N2O), a commonly employed self-pressurizing oxidizer, has been investigated for impinging jets, and the feasibility for an ignition system has been evaluated. A series of experiments were conducted to determine ignition delay values and ignition boundaries for two different oxidizer-to-fuel (O/F) ratios by changing the impingement angle. Results have demonstrated relatively favourable TEA/N2O ignition delays across varying impingement angles. This research also comprehensively evaluates experimental outcomes, post-test equipment conditions along with the theoretical aspects to outline the design criteria for a comprehensive ignition system.
The Catalyst Development for Space Propulsion Applications
Monopropellant rocket systems work by generating a combination of high-pressure hot gases through an exothermic decomposition reaction of a propellant compound. Afterward, the products are accelerated through a converging–diverging nozzle to provide the desired thrust. These systems can generate thrust in the range of 0.1 to 500 N, with a moderate specific impulse of up to 250 s. In small missiles and satellite engines that require low thrust, monopropellant engines are preferred. This thermal decomposition process has a large activation energy, meaning that the compound does not decompose spontaneously at ambient temperatures and must be suitably heated. A special catalyst can lower the activation threshold to speed up the reaction. As a monopropellant, hydrazine (N2H4) is most often used because the successful development of the Shell S405 catalyst provides high efficiency, with a specific impulse (Isp) value of 237 s. However, its carcinogenic and toxic nature causes transportation and handling problems. Therefore, green alternatives are receiving increasing attention nowadays. Their lower toxicity and safety precautions result in lower costs for manufacturing, handling, and storage. One of the most promising choices among the green monopropellants is highly concentrated hydrogen peroxide (H2O2). Although highly concentrated hydrogen peroxide has a lower Isp value (179s) than hydrazine, H2O2 is ecologically friendly because its decomposition products include oxygen and steam. The development of an appropriate catalyst and a reliable system for green monopropellants are key aspects of catalytic decomposition in thruster systems. The main aim of this thesis is to demonstrate the utilization of an effective and reliable monopropellant thruster. As a chemical approach, the importance of the support and active material of the catalyst is studied, and their properties are analyzed with characterization techniques and reaction kinetic experiments. Then, they are used in the monopropellant thruster as an engineering approach; the catalyst bed is analyzed based on temperature and pressure changes over time during decomposition in the thruster system. In this thesis, the aluminum oxide-supported manganese oxide catalyst is synthesized and tested. A microcrystalline cellulose-templated alumina catalyst support was prepared, and a parametric study was conducted to determine the optimum preparation conditions. The chosen micro cellulose-templated aluminum oxide was 15 wt% of micro cellulose to alumina at a calcination temperature of 900 °C. Macropores were formed on the MnOx/alumina catalyst surface. Thanks to these pores, hot gas was allowed to discharge from the pores without breaking the catalyst. In addition, by increasing the calcination temperature, catalysts with a longer life have been designed for thruster operations. Therefore, the unstable performance of thrusters caused by cracking of the pellet catalyst can be solved by adjusting the porosity of the catalyst. The primary goal of previous studies was to explore methods to improve the mechanical stability of the catalyst within the thruster, which resulted in increased strength but only moderate catalytic activity. As a secondary aim, efforts have now shifted toward enhancing the catalyst's reaction activity while maintaining its improved mechanical properties. The research continued by improving the loading of active material through double impregnation techniques and intermediate heating at 325 °C. According to the BET analysis, high-concentration precursor solutions cause more material to accumulate on the support, leading to pore blockage and a decrease in specific surface area. This observation was also correlated with the diminution of visible pores in SEM images. Increasing the molarity of the manganese precursor and adjusting the synthesis process resulted in significant improvements in catalyst performance, as evidenced by higher MnOx loading and enhanced apparent reaction rates in kinetic and thruster tests at first however, after a while, all catalysts show similar performance. Beyond catalyst synthesis, the study explored various factors affecting thruster performance, including catalyst mass, particle size, the concentration of hydrogen peroxide and the amount of stabilizer in hydrogen peroxide and radiation in space environments. The decomposition of H2O2 using a manganese oxide catalyst exhibits first-order kinetics; however, external mass transfer limitations affect the observed reaction rate. Testing various catalyst amounts (250 mg to 1500 mg) demonstrated a proportional increase in reaction rate, emphasizing the significance of catalyst concentration. The particle size of the catalyst was also crucial in the catalytic bed of the thruster because size can also cause flow instabilities during thruster tests. According to the test results, larger particles (700 µm) resulted in higher pressure drops compared to smaller particles (300 µm), which provided more stable performance with lower pressure loss and greater thrust, indicating their superior effectiveness in thruster applications. Moreover, the presence of stabilizers, such as sodium pyrophosphate, was found to negatively impact catalyst efficiency by inhibiting H2O2 decomposition. Tests with lower-phosphate peroxide grades showed improved catalytic activity, emphasizing the importance of propellant purity in thruster performance. Also, from 87.5wt to 78wt% hydrogen peroxide concentrations are examined, and the catalysts are shown to breakdown them efficiently. Although lower decomposition temperatures resulted in minor thrust losses, these are unlikely to impair the overall performance of the propulsion system. Experiments also examined the effects of cold starts and preheating on thruster operation. Preheating the catalyst bed to 150 °C was found to provide optimum thrust performance by reducing thrust response times. Subsequently, gamma radiation was applied to both hydrogen peroxide and the catalyst to examine the effects of space radiation exposure. Ionizing radiation can cause hydrogen peroxide to decompose into oxygen, reducing its concentration and impacting thruster performance. However, tests on 88% hydrogen peroxide exposed to 158.4 Gray showed no significant change in concentration. While radiation can degrade stabilizers such as sodium pyrophosphate, the changes were minimal. Catalysts were also tested for radiation effects, showing some shifts in manganese oxidation states (increase in Mn²⁺, decrease in Mn⁴⁺); however, the catalyst's structure and functionality remained stable in thruster and kinetic tests, with only a slight reduction in apparent reaction rates. In summary, this thesis presents a comprehensive study of catalyst development for hydrogen peroxide decomposition in monopropellant thrusters. The findings demonstrate significant advancements in catalyst stability, efficiency, and performance, providing valuable insights for future applications in space propulsion systems
Development of a propulsion technology using lunar resources
This dissertation explores the development of a hybrid rocket propulsion technology utilizing lunar resources, specifically magnesium (Mg) and aluminum (Al), for sustainable space exploration. The study addresses the challenges of in-situ resource utilization (ISRU) to minimize dependency on Earth-supplied fuels, focusing on the potential of metals found in lunar soil. Through thermochemical analysis, Mg and Al are evaluated as primary fuels due to their energetic properties and abundance on the Moon. The research involves the design, production, and testing of hybrid rocket engines using metal powders combined with sodium silicate as a binder, creating a robust fuel with favorable combustion properties. Experimental tests were conducted on various fuel compositions to optimize the combustion characteristics and ensure high performance under lunar conditions. A specific focus is placed on the relationship between oxygen-to-fuel (O/F) ratio and the specific impulse (Isp) of the metal-based hybrid engine, achieving significant results with Mg-Al mixtures. The experimental results demonstrate that the optimal fuel composition, consisting of 75\% magnesium and 20\% aluminum, can sustain combustion efficiency with an Isp close to 350 seconds in a vacuum. Additionally, the study addresses the challenges posed by metal oxide formation and its impact on combustion chamber stability. Solutions, including thermal management and fuel structuring, are presented to mitigate the risks associated with high-temperature operations. The results of this research contribute to the broader field of ISRU and have significant implications for future lunar missions, particularly in the context of NASA's Artemis program and other international lunar exploration initiatives. This work presents a scalable and sustainable propulsion system, positioning metal-based hybrid rockets as a viable solution for long-term lunar exploration and colonization.
Yüksek kademe karma roketler için elektrik tahrikli kriyojenik pompa geliştirmesi
The performance of a rocket propulsion system utilizing liquid propellants depends on the propellant feed system (PFS) performance and capability. A PFS must be able to deliver the propellant at the desired flow rate and pressure into the combustion chamber. The most commonly used mechanisms for PFS consist of pressurization by a cold gas or pressurization by a pump. The first method requires pressurizing the entire feed system including propellant tanks to its delivery pressure while the second one pressurizes only a small portion of the feed line but requires a mechanical power source. These two methods have been used widely where pumps are employed on high thrust and power applications while cold gas pressurization is employed on relatively small and less sophisticated systems. The reason for this is that only capable power sources for high-performance pumps were turbines and this approach was not viable for lower thrust level systems due to its complexity and cost. As the space industry leans towards micro-satellites, the development of launch vehicles in this class is gaining more attention. Weight disadvantage of pressure-fed systems, complexity, development cost and time of turbopump systems and emerging power electronics, battery, and electric motor technologies arouse the question of the viability of electric driven pumps on micro-satellite level propulsion systems. This study is aimed to show the advantages of electric pump utilization over cold gas pressurization in hybrid rockets through an electric pump development process. For this, a hydraulic and mechanical design is made from a set of system requirements and the performance of this pump is assessed by means of CFD analysis. The CFD model is validated by pump tests. Then two separate hybrid rocket propulsion systems are designed, one of which employs electric pump and the other is pressure-fed. A system-level comparison of performance on different operational parameters, concluding high fidelity evidence for the thesis argument. Final results show that electric pumps are quite beneficial for upper stage hybrid rocket propulsion systems.