Researcher:
Karabeyoğlu, Mustafa Arif

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Mustafa Arif

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Karabeyoğlu

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Karabeyoğlu, Mustafa Arif

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2014 - 2019232020 - 202211

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    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; 114595
    CO2 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.
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    Influence of micro-aluminum addition to paraffin-based fuels on graphite nozzle erosion rate
    (American Institute of Aeronautics and Astronautics, 2020) Karakaş, Hakkı; Kara, Ozan; Kahraman, Büşra; Eren, Büşra Nimet; Özkol, İbrahim; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595
    One of the most used ablative cooling material for the hybrid rocket motors is the carbon graphite and the erosion of this material during the operation of these motors causes the loss of chamber pressure and Isp. Especially, long burning upper stage rocket motors suffer more from the nozzle erosion. One of the ways to combat the nozzle erosion is to change the fuel composition in order to reduce the oxidizing species at the nozzle and create a protective liquid fuel film on the surface of the nozzle. To study this method, we have decided to use a paraffin-based fuel because of its high regression rate and aluminum powder addition for nozzle erosion reduction. Aluminum is added as a fuel ring at the front of the paraffin-based fuel which makes this method to be scalable for bigger hybrid rocket motors. In our experimental tests with gaseous oxygen as an oxidizer, aluminum addition decreased the nozzle erosion rate nearly %45 and increased the nozzle erosion onset by 5 seconds. Both of these combined improve the hybrid rocket motor considerably. This new and innovative method to add an energetic material such as aluminum to the hybrid rocket fuel as a fuel ring offers very low-cost, easy to implement and highly effective way of reducing the nozzle erosion and by doing that improving the performance of the motor to a great extent.
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    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; 114595
    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. 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.
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    Small satellite architecture optimization: electric propulsion moon imaging mission
    (American Institute of Aeronautics and Astronautics Inc, AIAA, 2015) N/A; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Kara, Ozan; Faculty Member; Master Student; Department of Mechanical Engineering; College of Engineering; Graduate School of Sciences and Engineering; 114595; N/A
    This study underlies small satellite architecture optimization by using existing electric propulsion systems for the Moon missions. The estimated objective is panoramic imaging of the Moon accompanied with future in-situ applications. Edelbaum’s low thrust trajectory transfer with optimal control theory is used to calculate the required ΔV. During the journey, 1.5h eclipse duration effects the solar array design. The optimized xenon propellant density and pressure are 1350 kgm3 and 8.3 MPa within 300K. Two types of optimization process revealed based on hexagonal SC architecture. The iterative method with LEO departured ion thruster has 23 mN with minimum 213 kg total mass. Corresponding SC volume is 0.70 m3, propellant mass is 64 kg. This scenario cost $108.5M and takes 980 days. Same thruster level for GEO departure case takes 880 days with 58 kg xenon gas. The total cost reduces $2.5M. For HALL engine design, LEO departure case needs 0.8 m3, 247 kg SC including 82 kg xenon. 77 mN thrust operates 208 days towards the Moon that ends up with $121M total cost. If the SC to be launched from GEO, flight time reduces 45 days by consuming 65 kg propellant. Total SC mass, volume and power values are 230 kg, 0.71 m3 and 1351W which cost $115M. Results are compared with previous Moon or electric propulsion missions such as SMART-1, LADEE, Clementine and Hayabusa. For future applications of small satellites, innovative concepts are envisioned for in-space, Earth-independent exploration and space education. 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. .
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    Performance additives for hybrid rockets
    (Springer-Verlag Berlin, 2017) N/A; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595
    Addition of performance-improving materials in the solid hydrocarbon fuels of hybrid rockets has been studied extensively both in academia and also in the industry. The primary motivation has been to improve the specific impulse, density impulse, and regression rate performance of the propulsion system. Despite the fact that hybrid rockets are particularly suitable for the inclusion of performance additives, which are typically in solid phase, successful implementation has been quite difficult to achieve. In this paper, we evaluate the feasibility of using performance additives with the following primary objectives: (a) develop a comprehensive survey of fuel additives (and the best binders) for hybrid rockets in order to establish the state of the art in the field, (b) rank these additives based on performance and a number of important practical factors, and (c) recommend a subset of promising additives for further evaluation. Even though, this feasibility study primarily makes use of the information in the open literature, new thermochemical calculations has also been conducted in order to establish the theoretical performance of various propellant systems operating at a common reference state (i.e., chamber pressure, nozzle expansion ratio, etc.).
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    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; 114595
    Internal 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). 
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    Hybrid rocket nozzle erosion with microaluminum-added fuel
    (Amer Inst Aeronautics Astronautics) Karakas, Hakki; Ozkol, Ibrahim; N/A; Department of Mechanical Engineering; Kahraman, Büşra; Karabeyoğlu, Mustafa Arif; Master Student; Faculty Member; Department of Mechanical Engineering; Graduate School of Sciences and Engineering; College of Engineering; N/A; 114595
    Because of its relatively low cost and wide availability, carbon graphite is one of the most widely used ablative nozzle materials in hybrid rocket propulsion. The erosion characteristics of the nozzle material have paramount importance because it directly influences the I-sp performance. This is especially the case for upper-stage or in-space rocket motors operating with very long burn times. In this study, the effect of aluminum-added fuel on the graphite nozzle erosion is studied. In the experimental studies, a high regression rate paraffin-based fuel is loaded with micrometer-size aluminum powder for nozzle erosion reduction. In our approach, aluminum is added at high concentrations as a fuel ring in front of the main paraffin-based fuel which contains no aluminum. Based on the motor tests conducted with gaseous oxygen as an oxidizer, it is shown that aluminum addition decreased the nozzle erosion rate up to 45% and increased the nozzle erosion onset time by 1 to 3s. The new method of introducing an energetic powder in a fuel ring positioned at the fore end of the motor offers an easy and scalable way of reducing the nozzle erosion and improving the I-sp performance of the rocket motor.
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    Assessment of using electric pumps on hybrid rockets
    (American Institute of Aeronautics and Astronautics Inc, AIAA, 2019) Gegeoğlu, Kaan; Kahraman, Mehmet; Üçler, Çağlar; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595
    In small hybrid rockets, liquid oxidizer is fed into the combustion chamber by pressurizing the oxidizer tank to achieve the desired injector pressure. However, larger applications employ propellant pumps to keep the propellant tank mass at reasonable levels. Conventionally, pumps are implemented on rockets by means of turbopumps. However, turbopumps are complex systems with serious development time and costs. With the current advancements in batteries and electrical machinery, pumps driven by electric motors become viable actors. In the case of hybrid rockets, operation of a single liquid propellant with a single pump has significant advantages compared to their use in the bipropellant liquid engines. In this paper, an example case of an upper stage hybrid rocket system is introduced and a propellant pump is designed using CFD modeling. The advantage of using electric motor driven pumps has been established via weight analysis and a comparison between alternative feed systems. Turbopump system is not considered due to its complexity as an alternative system. Therefore, comparison is limited to the cold gas pressurization feed system. It has been shown that a significant advantage in total feed system weight is achieved not only with a high-head pump but also with a low-head pump with assistive pressurization. A 69.9 % mass reduction for double-stage pump system and a 45.2 % mass reduction for the single-stage pump system was calculated compared to the reference pressure-fed system.
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    Testing N2O/CO2 oxidizer mixtures with paraffin based μaluminum fuels for Mars Ascent Vehicles
    (American Institute of Aeronautics and Astronautics Inc, AIAA, 2019) Karakaş, Hakkı; Department of Mechanical Engineering; N/A; Karabeyoğlu, Mustafa Arif; Kara, Ozan; Faculty Member; PhD Student; Department of Mechanical Engineering; College of Engineering; Graduate School of Sciences and Engineering; 114595; N/A
    The 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. The impact of 3-micron size aluminum on key performance parameters such as specific impulse (Isp), regression rate (r), and combustion efficiency (ηcomb) has been studied experimentally using a lab scale hybrid motor with 70 mm grain length and 30 mm outer grain diameter. Thermochemical performance of the various propellants studied in this paper are evaluated using NASA’s Chemical Equilibrium Analysis (CEA) software. Numerous tests are performed with N2 O/CO2 mixtures as the oxidizer. CO2 as a saturated liquid is mixed with liquid N2 O between 8 to 30% by mass in order to measure the motor performance characteristics such as average regression rate, r and combustion efficiency, ηcomb. 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. Initial conclusion from this study indicate that mixing CO2 with N2 O is a viable method of burning CO2 in a practical rocket system for Mars missions.
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    Regression rate enhancement of hybrid rockets by introducing novel distributed tube injector
    (Amer Inst Aeronautics Astronautics, 2022) Kahraman, Mehmet; Ozkol, Ibrahim; Department of Mechanical Engineering; Karabeyoğlu, Mustafa Arif; Faculty Member; Department of Mechanical Engineering; College of Engineering; 114595
    Low fuel regression rate is one of the major drawbacks of hybrid rocket motors. In this study, a novel injector concept is proposed to provide a substantial enhancement in the fuel regression rate. A tubular injector, Distributed Tube Injector (DTI), is inserted in the center of the cylindrical fuel port in order to tailor the oxidizer flow introduced into the combustion chamber with the desired combination of radial, tangential, or axial components of velocity. This concept has been tested using a 500 N thrust class hybrid rocket motor, which uses a paraffin-based fuel and supercharged N2O (L) as the oxidizer. As a result of 30 hot firings conducted in the test program, it is determined that the DTI configuration provides regression rates up to 3.9 times higher than the regression rates obtained using fore-end injector commonly employed in conventional hybrid rockets. Using the motor test data, a comprehensive nondimensional and scalable regression rate relation has been established. This nondimensional regression rate equation can be used to design the internal ballistic configuration of hybrid rocket motors using the novel injector.