Effects of the O/F Ratio on the Performance of a Low Thrust LO X /Methane Rocket Engine with an ElecPump-fed Cycle
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ORIGINAL PAPER
Effects of the O/F Ratio on the Performance of a Low Thrust LOX /Methane Rocket Engine with an ElecPump-fed Cycle Byungil Yu1,3 · Hyun-Duck Kwak2 · Hong Jip Kim3 Received: 26 March 2020 / Revised: 6 August 2020 / Accepted: 2 September 2020 © The Korean Society for Aeronautical & Space Sciences 2020
Abstract Recently, electrically driven pump-fed cycle rocket engines (i.e., ElecPump engines) have received considerable attention. ElecPump engines are superior to pressure-fed engines and will likely replace gas generator cycle engines under a 100 kN thrust level. This study focuses on LOX /methane ElecPump engines with a low-thrust level and estimates the nominal operation point of the engine. The objective variable was set to the maximum velocity increment of the stage. The mixture ratio and combustion pressure were explored as the design variables. Thus, the optimal mixture ratio was proposed as the design variable, considering not only the engine weight but also the structural weight. Keywords Liquid rocket engine · Liquid oxygen · Liquid methane · Specific impulse · Electric propulsion
List of Symbols A CEA CF c∗ E GG g0 Isp Ivac MR m O/F P p PFS
B 1
Area Chemical equilibrium with applications Thrust coefficient Characteristic velocity Electric energy Gas generator Standard acceleration of gravity at sea level Specific impulse Vacuum specific impulse Mass ratio (m 1 /m 0 ) Mass Oxidizer-to-fuel mass flow ratio Electric power Pressure Propellant feed system
Hong Jip Kim [email protected] Engine Test and Evaluation Team, Korea Aerospace Research Institute (KARI), 169-84 Gwahak-ro, Yuseong-gu, Daejeon 34133, Republic of Korea
2
Turbopump Team, Korea Aerospace Research Institute (KARI), 169-84 Gwahak-ro, Yuseong-gu, Daejeon 34133, Republic of Korea
3
School of Mechanical Engineering, Chungnam National University, 99 Daehak-ro(St), Yuseong-gu, Daejeon 34134, Republic of Korea
RP-1 r T tb V v v δE,bat δinv δmotor δP,bat δpu δs δtp η ρ τ
Rocket propellant-1 or refined petroleum-1 Tank radius Absolute temperature Burning time Volume Velocity Velocity increment Battery energy density Inverter power density Motor power density Battery power density Pump power density Mass per unit surface area Turbopump power density Efficiency Density Thrust
Subscripts bat bp cc E f FT in
Battery Battery pack Combustion chamber Energy Fuel (methane) Fuel tank Inlet, input
123
International Journal of Aeronautical and Space Sciences
inv mot o OT out P p pu s tp vac 0 1
Inverter Motor Oxidizer (LOX ) Oxidizer tank Outlet, output Power Propellant Pump Surface Turbopump Vacuum Initial condition Final condition
1 Introduction Recently, several studies have evaluated engine cycles that supply propellants to rocket engines and pumps powered by electric motors (hereinafter referred to as ElecPump) [1–11]. In the late 1980s, Johnson and Bigert developed the electric propellant pump system for the satellite apogee engine under contract from the European Space Agency [1]. Because of the technology development of e
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