• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.513-520
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• 2004

A steady-state/transient performance simulation model was newly developed for the propulsion system of the CRW (Canard Rotor Wing) type UAV (Unmanned Aerial Vehicle) during flight mode transition. The CRW type UAV has a new concept RPV (Remotely Piloted Vehicle) which can fly at two flight modes such as the take-off/landing and low speed forward flight mode using the rotary wing driven by engine bypass exhaust gas and the high speed forward flight mode using the stopped wing and main engine thrust. The propulsion system of the CRW type UAV consists of the main engine system and the duct system. The flight vehicle may generally select a proper type and specific engine with acceptable thrust level to meet the flight mission in the propulsion system design phase. In this study, a turbojet engine with one spool was selected by decision of the vehicle system designer, and the duct system is composed of main duct, rotor duct, master valve, rotor tip-jet nozzles, and variable area main nozzle. In order to establish the safe flight mode transition region of the propulsion system, steady-state and transient performance simulation should be needed. Using this simulation model, the optimal fuel flow schedules were obtained to keep the proper surge margin and the turbine inlet temperature limitation through steady-state and transient performance estimation. Furthermore, these analysis results will be used to the control optimization of the propulsion system, later. In the transient performance model, ICV (Inter-Component Volume) model was used. The performance analysis using the developed models was performed at various flight conditions and fuel flow schedules, and these results could set the safe flight mode transition region to satisfy the turbine inlet temperature overshoot limitation as well as the compressor surge margin. Because the engine performance simulation results without the duct system were well agreed with the engine manufacturer's data and the analysis results using a commercial program, it was confirmed that the validity of the proposed performance model was verified. However, the propulsion system performance model including the duct system will be compared with experimental measuring data, later.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.703-712
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• 2004

A variant of the magnetoplasma jet engine (magjet) is here proposed for airbreathing flight in the hypersonic regime. As shown in Figure 1, the engine consists of two distinct ducts: the high-speed duct, in which power is added electromagnetically to the incoming air by a momentum addition device, and the fuel cell duct in which the flow stagnation temperature is reduced by extracting energy through the use of a magnetoplas-madynamic (MPD) generator. The power generated is then used to accelerate the flow exiting the fuel cells with a fraction bypassed to the high-speed duct. The analysis is performed using a quasi one-dimensional model neglecting the Hall and ion slip effects, and fix-ing the fuel cell efficiency to 0.6. Results obtained show that the specific impulse of the magjet is at least equal to and up to 3 times the one of a turbojet, ram-jet, or scramjet in their respective flight Mach number range. Should the air stagnation temperature in the fuel cell compartment not exceed 5 times the incoming air static temperature, the maximal flight Mach number possible would vary between 6.5 and 15 for a magnitude of the ratio between the Joule heating and the work interaction in the MPD generator varied between 0.25 and 0.01, respectively. Increasing the mass flow rate ratio between the high speed and fuel cell ducts from 0.2 to 20 increases the engine efficiency by as much as 3 times in the lower supersonic range, while resulting in a less than 10% increase for a flight Mach number exceeding 8.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.713-721
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• 2004

Variable air intake and variable exhaust nozzle of hypersonic engines are designed and tested in this study. Dimensions for variable geometry air intake, ram combustor and variable geometry exhaust nozzle are defined based on the requirements of a pre-cooled turbojet engine. Hypersonic Ramjet Engine is designed as a scaled test bed for each component. Actuation forces of moving parts for variable intake and variable nozzle are reduced by balancing the other force in the opposite direction. A demonstrator engine which includes variable intake and variable nozzle is designed and the components are fabricated. Composite material with silicone carbide is applied for high temperature parts under oxidation environment such as leading edge of the variable intake and combustor liner. Internal cooling structure is adopted for both moving and static parts of the variable nozzle. Pressure recovery and mass capture ratio of the variable intake at Mach 5 is obtained by a hypersonic wind tunnel test. Flow characteristics of the variable nozzle are obtained by a low temperature flow test. Wall temperature and heat flux of the nozzle at Mach 3 is obtained by a firing test. As results, the intake and the nozzle are proved to be used at designed pressure and temperature environment.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.1
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• pp.13-20
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• 1998

A steady-state and dynamic simulation program for a small multi-purpose turboshaft engine with the free power turbine was developed. In order to reduce developing cost, time and risk, a turbojet engine whose performance was well-known was used for the gas generator, and life time was improved by replacing turbine material and by using Larson-Miller curves. The component characteristic of the power turbine was derived from scaling the gas generator turbine. Equilibrium equations of mass flow rate and work were used for the steady-state performance analysis, and the Constant Flow Method(CMF) was used for the dynamic performance simulation. The step fuel scheduling was carried out for acceleration in the dynamic simulation. Through this simulation, it was found that the overshoot of the turbine inlet temperature exceeded over the compressor turbine limit temperature.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.5
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• pp.27-35
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• 2021

A technical review of Soviet/Russian supersonic anti-ship cruise missiles is presented. The supersonic anti-ship cruise missiles is one of the weapons for asymmetric power. The supersonic speed of the missiles is very useful both for attacking a time critical target and for improving target-penetration characteristics of the missile. The survivability of the missiles has also been increased by the improved concept of operation. Supersonic cruise missiles is greatly affected by the evolution of propulsion technology. Early supersonic cruise missiles adopt turbojet engines and rocket motors. The use of the integrated rocket-ramjet engine reduced the size of the supersonic missile, so today's supersonic cruise missiles are suitable to be deployed in various platforms. Nowadays, export versions of the missiles are actively being developed.