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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 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.