• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.1
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• pp.61-69
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• 2004

This paper describes numerical efforts to characterize the flame-holding and air-fuel mixing process of model SCRamjet engine combustor, where a hydrogen jet injected into a supersonic cross flow and in a cavity Combustion phenomena in a model SCRamjet engine, which has been experimentally studied at University of Queensland and Australian National University using a free-piston shock tunnel, was observed around separation region of upstream of the normal injector and inside of cavity. The results show that the separation region and cavity generates several recirculation zones, which increase the fuel-air mixing. Self ignition occurs in the separation-freestream and cavity-freestream interface.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.93-96
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• 2009

Advances in high speed flight technologies and engine efficiencies increase heat load on the aircraft. As the temperature of air flow is too high to cool the structure at hypersonic flight speeds, it is necessary to utilize the aircraft fuel as the primary coolant. By undergoing endothermic reaction, such as thermal decomposition or catalytic decomposition, aircraft fuels have heat sink potential. These fuels are called endothermic fuels. The endothermic reaction can be improved by catalysts, but limited by coke deposition. In this study the essential technologies of endothermic fuels are described, and intended to be used for basic research.

• Journal of ILASS-Korea
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• v.28 no.2
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• pp.89-96
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• 2023

This study presents a numerical simulation investigating hydrodynamic characteristics of high-temperature hydrocarbon aviation fuel injected through a plain orifice injector. The analysis encompassed the temperature range up to the critical point, and the obtained results were compared with prior experimental observations. The analysis unveiled that the injector's exit pressure remains equivalent to the ambient pressure when the fuel injection temperature is below the boiling point. However, when the fuel temperature surpasses the boiling point, the exit pressure of the injector transitions to the saturated vapor pressure corresponding to the fuel injection temperature. Consequently, the exit pressure of the injector increases in tandem with the rapid increase of the saturation vapor pressure due to escalating fuel temperatures. This rise in the exit pressure necessitates a proportional increase in fuel injection pressure to ensure a fixed fuel mass flow rate. Furthermore, the investigation revealed that the discharge coefficient obtained by applying the exit pressure instead of the ambient pressure did exhibit no decrease, but rather was maintained at a nearly constant value, comparable to its level below the boiling point.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.1-7
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• 2018

A combustion test for a supersonic combustor was conducted using a direct-connected type supersonic combustor test facility. The facility was verified for the capability of simulating required flow conditions. The test condition was maintained at Mach 2.0, $915^{\circ}C$ and 496 kPa for 15 s. Using gaseous hydrogen as the fuel, the combustor model was also tested for its ignition and flame holding capability at the fuel equivalence ratio of 0.12. Combustion efficiency was 71%, and the supersonic flow regime was obtained at this test condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.115-121
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• 2008

A core engine for pre-cooled turbojet engines is designed and its component performances are examined both by CFD analyses and experiments. The engine is designed for a flight demonstration of precooled turbojet engine cycle. The engine uses gas hydrogen as fuel. The external boundary including measurement devices is set within $23cm{\times}23cm$ of rectangular cross section, in order to install the engine downstream of the air intake. The rotation speed is 80000 rpm at design point. Mixed flow compressor is selected to attain high pressure ratio and small diameter by single stage. Reverse type main combustor is selected to reduce the engine diameter and the rotating shaft length. The temperature at main combustor is determined by the temperature limit of non-cooled turbine. High loading turbine is designed to attain high pressure ratio by single stage. The firing test of the core engine is conducted using components of small pre-cooled turbojet engine. Gas hydrogen is injected into the main burner and hot gas is generated to drive the turbine. Air flow rate of the compressor can be modulated by a variable geometry exhaust nozzle, which is connected downstream of the core engine. As a result, 75% rotation speed is attained without hazardous vibration and heat damage. Aerodynamic performances of both compressor and turbine are obtained and evaluated independently.

• Journal of the Computational Structural Engineering Institute of Korea
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• v.30 no.6
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• pp.495-500
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• 2017

In the boundary layer of supersonic or hypersonic vehicles, there is the conversion from kinetic energy to thermal energy, called aerodynamic heating. Aerodynamic heating has to be considered to design supersonic vehicles, because it induces severe heat flux to surface. Transient heat transfer analysis with CFD is used to predict thermal response of vehicles, however transient heat transfer analysis needs excessive computing powers. Loosely coupled method is widely used for evaluating thermal response, however it needs to be revised for overestimated heat flux. In this research, quasi-transient method, which is combined loosely coupled method and conjugate heat transfer analysis, is proposed for evaluating thermal response with efficiency and reliability. Defining reference time of splitting flight scenario for transient simulation is important on accuracy of quasi-transient method, however there is no algorithm to determine. Therefore the research suggests the algorithm with various flow conditions to define reference time. Supersonic flow field of blunt body with constant acceleration is calculated to evaluate quasi-transient method. Temperature difference between transient and quasi-transient method is about 11.4%, and calculation time reduces 28 times for using quasi-transient method.

• Advances in aircraft and spacecraft science
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• v.5 no.5
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• pp.581-594
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• 2018

This paper is the follow-on of a previous paper by the author where it was pointed out that the forthcoming, manned exploration missions to Mars, by means of complex geometry spacecraft, involve the study of phenomena like shock wave-boundary layer interaction and shock wave-shock wave interaction also along the entry path in Mars atmosphere. The present paper focuses the chemical effects both in the shock layer and on the surface of a test body along the Mars orbital entry and compares these effects with those along the Earth orbital re-entry. As well known, the Mars atmosphere is almost made up of Carbon dioxide whose dissociation energy is even lower than that of Oxygen. Therefore, although the Mars entry is less energized than the Earth re-entry, one can expect that the effects of chemistry on aerodynamic quantities, both in the shock layer and on a test body surface, are different from those along the Earth re-entry. The study has been carried out computationally by means of a direct simulation Monte Carlo code, simulating the nose of an aero-space-plane and using, as free stream parameters, those along the Mars entry and Earth re-entry trajectories in the altitude interval 60-90 km. At each altitude, three chemical conditions have been considered: 1) gas non reactive and non-catalytic surface, 2) gas reactive and non-catalytic surface, 3) gas reactive and fully-catalytic surface. The results showed that the number of reactions, both in the flow and on the nose surface, is higher for Earth and, correspondingly, also the effects on the aerodynamic quantities.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.6
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• pp.83-90
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• 2016

Ramjet mode combustion test was performed for a dual-mode ramjet engine model. The engine model consists of an air intake, a combustor and a nozzle. The combustor in the model has a large backward-facing step, designed to be used as a part of a rocket-based combined cycle engine. The test was performed at the flight speed of Mach 5 and the altitude of 24 km. Strong combustion was established only when the fuel was injected from both of the bottom-side and cowl-side wall. When the total fuel stoichiometric ratio was 1.0, distributed as 0.5 on the cowl side and 0.5 on the bottom side, the flow became subsonic at some portion in the combustor by thermal choking, i.e., ramjet mode was established for this condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.2
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• pp.10-18
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• 2010

Scramjet engine is one of the core parts of hypersonic vehicle of next generation and being investigated by many countries. Korea Aerospace Research Institute performed a ground test of the model scramjet engine S1 in 2007. And, S2 model which is improved from S1 model in engine startability and thrust was tested with HIEST (High Enthalpy Shock Tunnel) at Kakuda Space Center of JAXA. Design condition of S2 model was Mach 6.7, however, it was tested at Mach 7.7 as an off-design condition test. As a test result, flow separation was found at the inside of the intake, but the engine showed stable combustion pressure distribution. Furthermore, compared to other test models, S2 model showed a good performance value in thrust and specific impulse.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.5
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• pp.65-76
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• 2020

On the initial design process of a scramjet vehicle such as the trajectory prediction, it is inevitable to estimate the aerodynamic performance of a three-dimensional effect. Despite the necessity of intensive computing for the three-dimensional model, it is inefficient in predicting a wide range of aerodynamic performance. In this study, an engineering model for aerodynamic performance was developed based on two-dimensional computational fluid analysis and linearized supersonic inviscid flow theory. Correspondingly, the three-dimension aerodynamic performance relations are presented based on the two-dimensional results. And the additional three-dimensional computation was performed to evaluate the adequacy for the extended relations.