• Proceedings of the Korean Society of Marine Engineers Conference
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• 2001.11a
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• pp.130-134
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• 2001

In the re-entry control system, errors apt to induce because the time derivative of drag acceleration is analytically estimated. Still more, the difficulty of estimation of the exact drag coefficient in hypersonic velocity and the nun-reality of the scale height cause a steady-state drag error. This paper proposes the additional method of the disturbance observer. This reduces the steady-state drag error according to the following series. First, this method estimates a error in drag acceleration time derivative by the analytic calculation and then creates the new drag acceleration time derivative using the estimated error. The performance of the re-entry control system is verified about 32 reference trajectories.

• 한국전산유체공학회:학술대회논문집
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• 1997.10a
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• pp.92-98
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• 1997

A finite-difference method based on characteristic upwind flux difference splitting has been studied on the blowing effect on the wall heat transfer over blunt-bodies. As the blowing rates increased, the wall heat transfer rate decreased and the temperature gradient also decreased compared with no blowing case. The heat trasfer rate at Mach No. 20 was almost twice higher than that of Mach No. 15 at 50km altitude. The surface blowing can be an effective mechanism to reduce the surface heat transfer rate at hypersonic flight condition.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.159-165
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• 2003

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.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.153-157
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• 2003

Supersonic and hypersonic aircrafts must pass wide range of speed to reach high speed region. But for existing engines the most efficient operating speed ranges are decided according to their flying speed, so an engine which mixes several engines like TRJ (Turbo Ramjet) and ARJ (Air Turbo Ramjet) has been planed. This mixed type engine has inefficiency that more than two engines must be installed simultaneously, but the pulse detonation engine (PDE) that uses detonation wave has a strong point that it can operate in all speed range with single engine. This paper deals with the simulation of the pulse detonation engine which uses hydrogen peroxide $(H_2O_2)$ mono propellant. Hydrogen peroxide is low-cost propellant, and it is reacted without oxidizer. Comparison between $H_2-O_2$ mixture with $H_2O_2$ mono propellant about thrust, pressure, temperature and velocity shows that $H_2O_2$ is a very useful propellant.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2011.10a
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• pp.171-177
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• 2011

In this paper, a square simply supported panel flutter have been considered at high supersonic flow by using coupled fluid-structure (FSI) analysis that based on time domain method. The Reynolds-Average Navier Stokes (RANS) equation with Spalart-Allmaras turbulent model were applied for unsteady flow problems of panel flutter. A fully implicit time marching schemed based on the Newmark direct integration method is used for calculating the coupled aeroelastic governing equations of it. In addition, the SOL 145 solver of MSC.NASTRAN was used to investigate flutter velocity based on PK-method of Piston theory. Our numerical results indicated that there is a good agreement result between Piston Theory in MSC.NASTRAN and coupled fluid-structure analysis.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.21 no.2
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• pp.1-6
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• 2013

In high-speed vehicle, selection of fuel, configuration of components and cooling system are required to solve the heating issue by aerodynamic heating and inner combustion process. This subsystem consists of fuel tank, supply pump, various control valve, heat exchanger, including reactor, connecting line, adiabatic structures and insulations. In this paper, applicable fuel property is considered at flight characteristic of hypersonic vehicles. In this regard, current state of fuel/cooling system technology is identified.

• International Journal of Aeronautical and Space Sciences
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• v.6 no.2
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• pp.46-55
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• 2005

A rocket-powered vehicle is designed conceptually which uses an engine running on methane and oxygen and delivering 10 tons of thrust. The aerodynamic coefficients of the vehicle are taken to be those of the Japan's HOPE-X, and the weight of this vehicle is estimated using a method developed by NASA. The resulting vehicle will be about 9 meters long, 5.8 meters in wing span, weigh about 2 tons empty, carry a maximum of 5.6 tons of propellant, and endure a g-load of 4.5. The craft will be able to carry five passengers, in addition to a pilot, and fly for space tourism between a northern and a southern airport with a maximum g-load varying from 3g to 4g depending on the route flown.

• International Journal of Aeronautical and Space Sciences
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• v.12 no.4
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• pp.396-402
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• 2011

Aircraft combat survivability is an essential factor in the design of combat aircrafts that operate in an enemy air defense area. The combat aircrafts will be confronted with anti-aircraft artillery and/or surface-to-air missiles (SAM) from the ground, and their survivability can be divided into two categories: susceptibility and vulnerability. This article studies the prediction of susceptibility in the case of a one-on-one engagement between the combat aircraft and a surface-based threat. The weighted score method is suggested for the prediction of susceptibility parameters, and Monte Carlo simulations are carried out to draw qualitative interpretation of the susceptibility characteristics of combat aircraft systems, such as the F-16 C/D, and the hypersonic aircraft, which is under development in the United States, versus ground threat from the SAM SA-10.

• International Journal of Aeronautical and Space Sciences
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• v.16 no.1
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• pp.64-76
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• 2015

This paper suggests applying pseudospectral model predictive method for exo-atmospheric guidance. The method is a fusion of pseudospectral law and model predictive control, in which a two point boundary value problem is formulated using model predictive approach and solved by applying pseudospectral law. In this work, the method is applied to exo-atmospheric guidance with specific target requirement. The existing exo-atmospheric guidance methods suffice general requirements for guidance, but cannot ensure specific target constraints; whereas, the presented method is able to do so. The proposed guidance law is assessed through simulation of perturbed cases, and the tests suggest that the method is able to operate semi-autonomously under control and thrust vector perturbations.

• 한국전산유체공학회:학술대회논문집
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• 1999.05a
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• pp.73-78
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• 1999

극초음속 유동장의 정확한 해석을 위해 AUSMPW+ 수치기법과 충격파 포착시 생기는 수치오차를 제거하기 위해 충격파 정렬 기법(Shock-Aligned Grid Technique)을 개발하였다. AUSMPW+ 수치기법은 자체 수치점성이 적은 수치기법으로 점성 경계층 계산시 정확한 계산결과를 보여주며 기존의 AUSM 계열이 가지는 문제점인 물성치의 진동 현상을 제거한 수치기법이다. 원통형과 무딘 물체 주위의 극초음속 유동장 해석을 통해 공력이 진동현상 없이 정확하게 계산됨을 확인하였다. 그리고 충격파 정렬 기법의 특성을 파악하기 위해 충격파 반사문제와 충격파-충격파 상호작용 문제를 해석하여 수치오차 없이 충격파를 포착할 수 있음을 보였다. 또한 화학적 평형 비평형 유동 영역까지 충격파 정렬 격자 기법을 확장하였다.