• 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 the Korean Society of Propulsion Engineers
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• v.20 no.5
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• pp.90-98
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• 2016

A high temperature heater is required to supply high temperature air to the hypersonic propulsion system in order to simulate high velocity flying condition during the ground test. Various high temperature heaters were reviewed, categorized, and analyzed in this paper. Heaters were categorized in 4 groups; in-stream combustion heater, electric arc heater, storage heater, and heat exchange heater. Each group has its own advantages and disadvantages, so the heater should be selected considering its purpose.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.289-292
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• 2007

Performance of a dual mode ramjet engine based on the sensitivity analysis of design parameters (the gap between cowl and inlet spike and combustor length) was analyzed from the view points of aerodynamics and thermodynamics. A dual mode engine performing from supersonic to hypersonic (Mach no. 2 to 6) was designed in a proposed flight envelop. The design method and result were comparable to the results of the previous study, Hyperion RLV, and the CFD calculation.

• Journal of computational fluids engineering
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• v.19 no.2
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• pp.49-57
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• 2014

In the present study, 2-D gas-kinetic flow solver on unstructured meshes was developed for flows from continuum to transitional regimes. The gas-kinetic BGK scheme is based on numerical solutions of the BGK simplification of the Boltzmann transport equation. In the initial reconstruction, the unstructured version of the linear interpolation is applied to compute left and right states along a cell interface. In the gas evolution step, the numerical fluxes are computed from the evaluation of the time-dependent gas distribution function around a cell interface. Two-dimensional compressible flow calculations were performed to verify the accuracy and robustness of the current gas-kinetic approach. Gas-kinetic BGK scheme was successfully applied to two-dimensional steady and unsteady flow simulations with strong contact discontinuities. Exemplary hypersonic viscous simulations have been conducted to analyze the performances of the gas-kinetic scheme. The computed results show fair agreement with other standard particle-based approaches for both continuum part and transitional part.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.5
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• pp.93-99
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• 2003

The determination of thrust is essential in design and evaluation of a hypersonic airbreathing propulsion device. Conventional methods to determine the thrust is using thrust stand or force measurement system. However, these conventional methos are not applicable to the case where thrusts stands are impractical, such as free jet testing of engines, and model combustor. For this reason, the thrust determination method from measured pitot pressure is considered and validated. Validation of thrust determination from pitot pressures can be achieved by comparing the actual thrust from thrust stand. For validation purpose, a small-scale supersonic wind tunnel is installed on the thrust stand. Thrusts are measured while pressures are measured simulaneously. Then, the thrust from pitot pressure measurements are compared with the measured thrust and theoretical thrusts.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.15-25
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• 2004

This paper outlines a new convergence acceleration de-signed to solve scramjet flowfields with zones of re-circulation. Named the “marching-window”, the algorithm consists of performing pseudo-time iterations on a minimal width subdomain composed of a sequence of cross-stream planes of nodes. The upstream boundary of the subdomain is positioned such that all nodes upstream exhibit a residual smaller than the user-specified convergence threshold. The advancement of the downstream boundary follows the advancement of the upstream boundary, except in zones of significant streamwise ellipticity where a streamwise ellipticity sensor ensures its continuous progress. Compared to the standard pseudo-time marching approach, the march-ing-window is here seen to decrease the work required for convergence by up to 24 times for supersonic flows with little streamwise ellipticity and by up to 8 times for supersonic flows with large streamwise separated regions. The memory requirements are observed to be reduced sixfold by not allocating memory to the nodes not included in the computational subdomain. The marching-window satisfies the same convergence criterion as the standard pseudo-time stepping methods, hence resulting in the same converged solution within the tolerance of the user-specified convergence threshold. The extension of the marching-window to the weakly-ionized Navier-Stokes equations is also discussed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.42-47
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• 2004

Air-fuel mixing and flame-holding are two important factors that have to be considered in the design of an injection system. Different injection strategies have been proposed with particular concern for rapid air-fuel mixing and flame-holding. Two representative injection techniques can be applied in a supersonic combustor. One of the simplest approaches is a transverse(normal) injection. The cavity flame holder, an integrated fuel injection/flame-holding approach, has been proposed as a new concept for flame holding and air-fuel mixing in a supersonic combustor. This paper describes numerical efforts to characterize the flame-holding and air-fuel mixing process of a model scramjet engine combustor, where hydrogen is injected into a supersonic cross flow and a cavity. The 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, were observed around the separation region of the transverse injector upstream and the inside cavity. The results show that this flow separation generates recirculation regions which increase air-fuel mixing. Self-ignition occurs in the separation-freestream and cavity-fteestream interfaces.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.34-40
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• 2004

Two different type scramjet models with side-wall compression and top-wall compression inlets have been tested in HPTF (Hypersonic Propulsion Test Facility) under the experimental conditions of Mach number 5.8, total temperature 1700K, total pressure 4.5㎫ and mass flow rate 3.5kg/s. The liquid kerosene was used as main fuel for the scramjets. In order to get fast ignition in the combustor, a small amount of hydrogen was used as a pilot. A strut with alternative tail was employed for increasing the compression ratio and for mixing enhancement in the side-wall compression case. Recessed cavities were used as a flameholder for combustion stability. The combustion efficiency was estimated by one dimensional theory. The uniformity of the facility nozzle flow was verified by a scanning pitot rake. The experimental results showed that the kerosene fuel was successfully ignited and stable combustion was achieved for both scramjet models. However the thrusts were still less than the model drags due to the low combustion efficiencies.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.310-319
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• 2017

This paper introduces the method how to measure the thrust in impulse facility. In a Facility having such a short duration time of steady flow, there's no time to reach a steady state of the forces acting on model so that the test model vibrates until the end of the flow. The forces exerted on an engine exist with vibration so that the usual force balance can not be used. SWFB(Stress Wave Force Balance) technique is utilized in a shock tunnel to get the thrust. As an example, a model force balance has been calculated its strain against impulse force by using FEM(Finite Element Method). A transfer function between the impulse force and strain has been obtained by the way of de-convolution.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.47-52
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• 2004

In the design of scramjet intake for hypersonic flight, a variety of aerothermodynamics phenomena are encountered. These phenomena include blunt leading - edge effects, boundary layer development issues, transition, inviscid / viscous coupling, shock - shock interactions, shock / boundary - layer interactions, and flow profile effects. For intakes that are designed to operate within a narrow Mach number / altitude envelope, an understanding of a few of these phenomena might be required. In this work several predominant flowfield phenomena (viscous phenomena, boundary - layer separation, and combustor entrance profile) are discussed to investigate the performance of the intake at the altitude and angle of attack extremes of the HyShot flight experiment.