Design and performance evaluation of lab-scale hybrid thruster using catalytically decomposed hydrogen peroxide oxidizer
- Paper number
IAC-13,C4,P,16.p1,x18577
- Author
Mr. Minwoo Lee, Korea Advanced Institute of Science and Technology (KAIST), Korea, Republic of
- Coauthor
Dr. Eun Sang Jung, Korea Advanced Institute of Science and Technology (KAIST), Korea, Republic of
- Coauthor
Prof. Sejin Kwon, Korea, Republic of
- Year
2013
- Abstract
Hydrogen peroxide has long been used as the rocket propellant and oxidizer. It is safe, easy to obtain, and environment friendly. Hydrogen peroxide is catalytically decomposed into water vapor and oxygen, releasing massive amount of heat. This heat can be used to ignite the solid grain in hybrid rocket. Hybrid rocket has better performance than solid motor, and restartable. Compared to liquid bipropellant rocket, hybrid rocket is easy to construct and cost effective. In this research, design of 80 N scale and 250 N scale up hybrid thruster was conducted and ground experiments were done. 90% hydrogen peroxide and polyethylene was used as oxidizer and solid propellant. Hydrogen peroxide was supplied to solid grain after catalytically decomposed by MnO2/Al2O3 catalyst. As a preliminary research, 80 N scale hydrogen peroxide/polyethylene hybrid rocket was designed and manufactured. NASA CEA code was used to calculate the optimum amount of oxidizer and O/F ratio. Water flow test and monopropellant mode experiment were conducted to observe the pressure drop across the engine components. Hybrid mode combustion test was conducted. Average thrust of 76.0 N and 85.8 N was observed after ignition in paraffin and polyethylene fuel grain. Considering the mass flow rate, pressure drop and thrust, polyethylene is considered as better fuel grain in hybrid rocket using catalytically decomposed hybrid thruster. 250 N hybrid engine was designed based on the result of previous experiments. Thruster using polyethylene fuel grain is designed. Water flow test and hybrid combustion test followed. Carbon phenolic ablative nozzle was adapted to reduce the thermal stress on the thruster. Ablative nozzle was damaged during combustion experiments. Thrust and pressure dropped abruptly after nozzle damage. In two times experiments, 201.5 N and 224.7 N thrust was observed before nozzle damage. This research showed the possibility of development of tropospheric hybrid sounding rocket. Further research can be continued on design and construction of sounding rocket flight model.
- Abstract document
- Manuscript document
(absent)