• Journal of the Korean Society for Aviation and Aeronautics
• /
• v.14 no.2
• /
• pp.27-32
• /
• 2006

The performance-design algorithm of a liquid ramjet engine was studied, which showed the comparable results with CFD calculation except the shock-boundary layer interaction region. In addition to the description of the design algorithm, several important design parameters, such as equivalence ratio, combustion efficiency, air capturing factor, and flight Mach number, are evaluated as predicting one of performance factors, Isp, of a typical ramjet engine, so that the flight envelope might not be determined with narrow perspective performance-operation-area in off-design regime.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.205-210
• /
• 2004

The pulse detonation engine (PDE) has recently expected as a new aerospace propulsion system. The PDE system has high thermal efficiency because of its constant-volume combustion and its simple tube structure. We measured thrust of single-tube pulse detonation rocket (PDR) by two methods using the PDR-Engineering Model (full scale model) for ground testing. The first involved measuring the displacement of the PDR-EM by laser displacement meter, and the second involved measuring the time-averaged thrust by combining a load cell and a spring-damper system. From these two measurements, we obtained 130.1 N of time-averaged thrust, which corresponds to 321.2 sec of effective specific impulse (ISP). As well, we measured the heat flux in the wall of PDE tubes. The heat flux was approximately 400 ㎾/$m^2$. We constructed the PDR-Flight Mode] (PDR-FM). In the vertical flight test in a laboratory, the PDR-FM was flying and keeping its altitude almost constant during 0.3 sec.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2012.05a
• /
• pp.46-51
• /
• 2012

This study aims to develop the performance analysis program on the staged combustion cycle of the liquid rocket engine using liquid oxygen(LOx) as oxidizer, liquid hydrogen(LH2) and RP-1 as fuel. The developed analysis program can obtain the propellant mass flow rate, the specific impulse, and representative design values of engine components for the required thrust satisfying pressure, mass flow, and energy balance conditions. The analysis results show that the the specific impulses (Isp) compared to those of the real engines have been less than 1%. With additional constraints, the program will be improved for the system optimization.