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Publication Metadata only An experimental study on predicting the mass flow rate of self-pressurizing oxidizers through injectors(American Institute of Aeronautics and Astronautics Inc, AIAA, 2021) Kara, Ozan; Department of Mechanical Engineering; N/A; Karabeyoğlu, Mustafa Arif; Karpat, Miray; Faculty Member; Master Student; Department of Mechanical Engineering; College of Engineering; Graduate School of Sciences and Engineering; 114595; N/ASelf-pressurizing propellants are recently gaining attention, specifically in hybrid rocket propulsion systems. Use of self-pressurizing propellants reduces system complexity and overall weight due to their high vapor pressure. N2O has been used widely as an oxidizer since it has a vapor pressure of approximately 5 MPa (730 psi) at room temperature. However, because they operate near or at saturation line, their flow exhibits two-phase behaviour. Therefore, it is difficult to model the feed system and injector flows. A method to predict the two-phase critical mass flow rate has been proposed. In addition, an experimental setup has been designed for validation of the proposed model. Multiple cold flow tests using nitrous oxide has been performed, and data obtained have been compared to that predicted by the two-phase critical flow model. Results have been showed that the proposed model estimates the actual mass flow rate within an error range of 6 to 17%.Publication Metadata only Challenges in the development of large-scale hybrid rockets(Begell House Inc, 2017) N/A; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595Advanced hybrid rockets, which combine fast burning fuels, composite motor construction, and innovative internal ballistic design, have the capability to deliver high performance while retaining the cost, environmental, and simplicity advantages of the classical hybrids. This makes hybrid rocket propulsion a tipping point technology in the sense that a small, short-term investment could have game-changing consequences in the development of green, safe, affordable, and high-performance systems needed for future space missions. In order to demonstrate the advantages of hybrids most effectively, the effort should be concentrated on improving the technology readiness level of the technology for a carefully selected class of missions. That being said, some serious challenges still exist in the development of operational motors, even for applications highly suitable for hybrid propulsion. These challenges, some perceived whereas others are very real, are carefully outlined in this paper. The real-life importance of each challenge is also discussed, along with potential methods to mitigate these issues. The ultimate strategy in the elimination of any practical challenge is that the solution should not compromise the simplicity, cost, and safety advantages of classical hybrid rockets. The solution methodology should be an iterative process that involves a well-balanced combination of theoretical modeling, numerical simulations, and actual motor testing.Publication Metadata only Effect of pressure loss devices on the performance of hybrid rocket systems(International Astronautical Federation (IAF), 2018) N/A; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595Internal ballistic devices that are used to trip the gas flow such as blades, steps, diaphragms or screens are commonly implemented in hybrid rocket motors to improve the mixing in the system. Enhanced mixing typically leads to an improvement in the regression rates and combustion efficiencies. The major issues with the use of these elements are 1) the pressure drop associated with the gas flow over these obstacles, 2) additional weight and 3) additional system complexity and cost. Even though a total pressure drop leads to a reduction in the thrust specific fuel consumption for air breathing propulsion systems, rockets do not suffer a direct hit on their specific impulse. The objective of this paper is to outline a theoretical proof that any pressure drop encountered along the motor axis does not lead to a reduction in the thrust or specific impulse performance of the rocket system. The primary adverse effect of the pressure loss is on the structural mass fraction of the rocket system. Assuming that the pressure at the nozzle entrance does not change, pressure drop requires higher head end pressures leading to increased injector manifold and feed system pressures. Using some example cases, we have estimated the combustion efficiency improvement required to balance the structural mass fraction increase caused by the flow trip devices. Such analysis needs to be conducted to justify the use of pressure drop elements for each design. Even though these arguments were developed for hybrid rockets, the analysis and conclusions are valid for all other chemical rocket types as well (i.e. solids and liquids).Publication Metadata only Electric propulsion optimization of microsatellite moon missions preliminary design application on CubeSats and Turkish small satellite field(International Astronautical Federation (IAF), 2014) N/A; N/A; Department of Mechanical Engineering; Kara, Ozan; Karabeyoğlu, Mustafa Arif; PhD Student; Faculty Member; Department of Mechanical Engineering; Graduate School of Sciences and Engineering; College of Engineering; N/A; 114595Subsystems miniaturization of spacecraft is making scientific microsatellite missions feasible. Specifically interplanetary space exploration mission can be provided by onboard micro propulsion systems. Due to the low requirements, Moon is a feasible destination for a preliminary space mission that universities, companies and governments can perform. This paper addresses the optimization of an electric propulsion system for a potential microsatellite lunar mission. Optimization takes the thrust level as a free variable to find the minimum initial mass along with the associated total burn time. The initial thrust range is chosen between 0.5-6mN. For a given thrust value, corresponding specific impulse, thruster power and thruster mass are determined based on curve fittings. As the next step, the input thrust level is extended up to 42mN to observe optimization over a broader thruster size. A feasible low thrust continuous orbit transfer to the Moon requires a high ΔV which is over 7,000 m/s. Edelbaum's analysis with optimal control theory is utilized to estimate the ΔV value. Edelbaum presents an approach for two non-coplanar circular orbits without any perturbations and shadowing effects. Furthermore, the optimization approach is applied for the preliminary design of a CubeSat Moon Mission. The spacecraft is determined to have an initial mass of 12 kg and requires a total power up to 100W. The particular mission selected for the study starts at 700km LEO and finishes at 200 km LLO. In addition, preliminary mission design presents (1) mass and power budgets, (2) thermal analysis, (3) ADCS selection, (4) structure and array mechanisms, and (5) cost estimation. In this paper, previous and planned small satellite researches in Turkey have also been reviewed. Small satellite projects which are performed by government, universities and industries show that Turkey has a small but growing small satellite activity.Publication Metadata only Experimental study of lunar-based hybrid rocket engine(American Institute of Aeronautics and Astronautics Inc, AIAA, 2021) Yalçıntaş, Ali; Kara, Ozan; Baysal, Mustafa; Department of Mechanical Engineering; N/A; Karabeyoğlu, Mustafa Arif; Yelken, Ümit; Faculty Member; PhD Student; Department of Mechanical Engineering; College of Engineering; Graduate School of Sciences and Engineering; 114595; N/AIn this paper, it was aimed to prepare a metal-based hybrid rocket engine by using elements such as magnesium and aluminum, which are abundant in lunar soil. In thermodynamic performance calculations, a mixture ratio with high specific impulse (Isp) was determined using NASA’s Chemical Equilibrium Analysis (CEA) package program, and a rigid fuel was formed from metal powders by using sodium silicate as the binding component. While determining the mixing ratio of aluminum, magnesium, and sodium silicate, the criterion that was taken into consideration was the temperature values to prevent residue formation at the combustion chamber and nozzle throat. The temperature values above the boiling points of the combustion products were tried to be obtained both in the combustion chamber and at the nozzle throat. Thus it was aimed to make a hybrid rocket engine that could be used for extended runtimes. Experimental studies of this hybrid rocket engine fuel obtained from the elements found in the lunar soil and rocks were carried out.Publication Metadata only Performance analysis of N2O/CO2 oxidizer mixture with paraffin based micro-aluminum and magnesium fuels for mars ascent vehicles(International Astronautical Federation, IAF, 2019) Karakaş, Hakkı; Kahraman, Büşra; Eren, Busra N.; N/A; Department of Mechanical Engineering; Kara, Ozan; Karabeyoğlu, Mustafa Arif; PhD Student; Faculty Member; Department of Mechanical Engineering; Graduate School of Sciences and Engineering; College of Engineering; N/A; 114595The objective of this research is to perform experimental test of a hybrid rocket motor by using N2O/CO2 mixture as the oxidizer and paraffin wax as fuel with the addition of metallic powders such as micron size aluminum and magnesium. A lab scale hybrid motor with 70 mm grain length and 30 mm outer grain diameter is used to measure the key performance parameters such as specific impulse (Isp), regression rate (R-), and combustion efficiency (?comb). The 3-micron sized aluminum powder with %98.9 purity has been used as fuel additive in the paraffin wax. Numerous tests are performed with N2O/CO2 mixtures as the oxidizer. CO2 as a saturated liquid is mixed with liquid N2O between 8 to 30% by mass in order to measure the motor performance characteristics such as average regression rate, r- and combustion efficiency, ?comb. Experiments were performed in the blowdown mode using the self-pressurizing capability of the mixed oxidizer around 45-50 bar. Chamber pressure of the hybrid motor was in the 15-30 bar range. Successful ignition and motor operation of the mixed oxidizer is achieved all the way up to 22% CO2 in the mixture. In addition, 62 micron magnesium powder is also used to evaluate the performance parameters of micron-sized aluminum based hybrid motor. Experiments showed that magnesium based fuels can combust with over 25% CO2 in the oxidizer mixture due to the easy combustibility nature of the Mg. Although Mg indicates lower regression rate characteristics than Al, usually bring higher combustion efficiency out. Therefore, initial conclusion from this study indicate that mixing CO2 with N2O is a viable method of burning CO2 in a practical rocket system for Mars missions. Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved.Publication Metadata only Propulsion system design for Mars ascent vehicles by using the in-situ CO2(American Institute of Aeronautics and Astronautics Inc, AIAA, 2021) Kara, Ozan; Karpat, Miray; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595CO2 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. Although Metal CO2 combustion has been studied by many researchers, process has not been evaluated for practical rocket applications. Therefore, this research aims to present ignition characteristics of the CO2 in actual hybrid rockets. Propellant combination of Para f fin/ Metal CO2 is selected to accomplish the combustion. In addition, combustion feasibility of CO2 is evaluated for Mars Ascent Vehicles. Hybrid propulsion system by using Para f fin/ Metal CO2 propellant combination is totally feasible for Mars Ascent Vehicles. In-situ propellant combination provides significant cost savings as well as ease of manufacturing on Mars. In-situ propellant mass optimization provides that only 20% fuel (paraffin wax) and 40% oxidizer (nitrous oxide) are needed to be brought from the Earth. It means that Para f fin/ Metal CO2 / CO2 propellant combination uses 80% Mg and 60% CO2 for a practical MAV system. Furthermore, 240 kg Mars Sounding Rocket is designed by using Para f fin/ Metal CO2 / CO2 for ballistic hopper missions.Publication Metadata only Systems advantages of electric pump fed upper stage hybrid rocket(International Astronautical Federation (IAF), 2018) N/A; Department of Mechanical Engineering; N/A; Gegeoğlu, Kaan; Karabeyoğlu, Mustafa Arif; Kara, Ozan; Master Student; Faculty Member; PhD Student; Department of Mechanical Engineering; Graduate School of Sciences and Engineering; College of Engineering; Graduate School of Sciences and Engineering; N/A; 114595; N/APropellant feed systems in liquid rocket engines deliver fuel and oxidizer from tanks to the combustion chamber at required flow rate and pressure. In hybrid rockets, only the liquid propellant is fed into the combustion chamber where solid propellant resides. The feed system has also a critical role for performance especially at upper stages of launch vehicles. There are two conventional feed system architectures that are mostly used. One employs a turbo pump while the other uses tank pressure to feed propellants. Briefly, propellant pumps are used in high pressure and high performance applications; however using turbo pumps makes design more complex and heavy. In contrast, the pressure fed system has simpler design although it is limited to low chamber pressures since high-pressure requirements make propellant tanks heavy. In addition there is a constant need for making elements such as tanks, valves, feed lines and pressurization devices lighter, simpler and more efficient. With advancing technologies in electric motor and batteries, electric pump feed system as an alternative to conventional feed system types has started to emerge. Most of the hybrid rocket applications evade turbo pumps since it requires carrying at least one another liquid propellant to drive turbines. Therefore, the performance of electric pump fed system over traditional pressure fed system is analyzed in this research on an upper stage of hybrid rocket engine. For this purpose, an optimization procedure has been applied on both systems and ∆V values as performance parameter are compared. on missions utilizing electric pump fed system, a noticeable improve of ∆V by 6.6% is reported. Optimization showed that longer burn times are also in the favor of electric pump fed systems.Publication Metadata only Understanding regression rate characteristics of CO2 as an oxidizer compound in paraffin based hybrid rocket motors(International Astronautical Federation, IAF, 2020) N/A; Department of Mechanical Engineering; Kara, Ozan; Karabeyoğlu, Mustafa Arif; PhD Student; Faculty Member; Department of Mechanical Engineering; Graduate School of Sciences and Engineering; College of Engineering; N/A; 114595CO2 is considered as a major combustion product in nature. However, CO2 can only burn with metallic powders due to high reactivity levels of metals compared to carbon. Therefore, Metal/CO2 combustion mechanism can be a potential propellant candidate for Mars Ascent Vehicles. In-situ CO2 can be used in Mars atmosphere reduces required Earth-based propellant mass. Propellant combination in this hybrid rocket motor consist of paraffin wax based solid fuel including metallic aluminium and magnesium. CO2 is mixed with N2O in various percentages by mass as the motor oxidizer. The goal of this research is to understand the maximum CO2 percentage by mass can be ignited with selected propellant combination (classical hybrid rocket motor). Maximum CO2 is achieved as 75% in oxidizer mixture by using 60% Magnesium 40% Paraffin solid fuel. Oxidizer flow rate at 75% CO2 is obtained as 60 g/s. Aluminum based solid fuels has maximum CO2 percentage of 50%. Magnesium provides better ignition characteristics with CO2. Therefore, the ignition boundary in mass flow rate vs. CO2 percentage is presented for both Al and Mg fuels. Regression rate characteristics does not change with the addition of CO2 in nitrous. Combustion efficiency of Mg based experiments are higher than Al and above 90%. Ignition results indicate that CO2 percentage can be increased with changing internal ballistics of hybrid motor. Paraffin/Metal/CO2 propellant combination can be a potential candidate for Mars Ascent Vehicles.