• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.3
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• pp.68-75
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• 2017

Dual bell nozzle is a type of altitude compensation nozzle, which is a nozzle that minimize the losses of the specific impulse at the off-design point of a typical bell nozzle. In this paper, numerical computations are performed to understand the transition characteristics of dual bell nozzles with fixed expansion ratios. The major design variables are the length of extension and the angle of inflection. As the length of the extension increased, the transition altitude and transition duration increased and the reduction of the thrust coefficient decreased. As the angle of inflection increased, the transition altitude and transition duration decreased and the reduction of the thrust coefficient increased.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.198-204
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• 2004

Using liquefied natural gas as a fuel, water, natural gas and liquefied natural gas-cooled firing tests were conducted. With the viewpoint of characteristic velocity, and specific impulse, the effect of OF mixture ratio and fuel inlet temperature into a combustion chamber were analyzed. OF mixture ratio and fuel inlet temperature into a combustion chamber have great influence on the performance. Characteristic velocity and theoretical specific impulse attain the maximum value at 0.72~0.75 and 0.75 of OF mixture ratio, respectively. Engine performance has a tendency to increase, proportional to fuel inlet temperature into a combustion chamber affected by the regenerative cooling.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.409-411
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• 2017

미사일 추력기 체계에 적용되는 하이드라진[$N_2H_4$]추진제는 MSDS-OHS 유해성 분류상 급성독성 물질로서 사용이 제한되고 있는 바, 다양한 대체물질이 개발 중이다. 최근 해외에서 안전성과 취급이 우수한 질산 히드록실암모늄[$NH_3OHNO_3$]과 암모늄 디나이트라마이드[$NH_4N(NO_2)_2$] 기반 단일계 액상추진제가 개발중이며, 이 물질들을 이용한 추력기 시스템 적용 시험이 진행되고 있다. 그러나 저온에서의 연로물질 산성화 반응으로 인한 디나이트라마이드[$N(NO_2)_2{^-}$] 물질의 분해는 나이트레이트[$NO_3{^-}$] 이온 생성을 촉진시키며, 부수적으로 발생하는 침전물은 촉매 및 노즐의 막힘 현상을 유발하므로 추력기 성능의 저해요인으로 작용한다. 그러므로 저온분해 방지를 위한 첨가제 조성 개발 및 열분해 특성 연구가 최근의 관심사이다. 본 연구는 합성/정제/추출한 암모늄 디나이트라마이드 산화제를 주요 조성물로 적용하였으며, 염기성 안정화제를 질량비율 4~5% 첨가하여 산성화 반응을 억제시킨 단일계 액상추진제(KMP) 형태로 제조하였다. 합성한 추진제는 시차주사열량계(DSC)를 이용하여 분해온도를 측정하여 열안정성을 평가해보았다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1072-1076
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• 2017

In this paper, a numerical study has been performed to analyze flow characteristics of rocket propulsion. Through the ground spin test, combustion chamber pressure was measured. Based on the experimental results, numerical analysis was conducted under various nozzle pressure ratio conditions such as standard, operating and base pressure conditions. And it was compared with quasi-1D solution and experimental result. In addition, the difference in thrust characteristics according to the spin/non-spin of the flow conditions was confirmed at the same nozzle pressure ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.4
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• pp.12-18
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• 2021

This study aimed at preparing the solid propellants featuring high pressure exponent available for throttleable rocket system development. The combustion properties of solid propellant based on PCP as a prepolymer were investigated with the different nitramine oxidizer, HMX and HNIW. As a main oxidizer, HNIW could deliver higher burning rate, specific impulse and flame temperature than HMX. In addition, the introduction of AP as a co-oxidizer in PCP/Nitramine propellants could enhance burning rate, specific impulse and flame temperature, showing the lower pressure exponent with increasing the content of fine-sized AP, total solids and plasticizer. Moreover, we examined the temperature sensitivity on burning rate of propellants between 150 psia and 2,500 psia.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.8
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• pp.64-71
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• 2004

A computational study on the supersonic flow around the lateral jet controlled missile has been performed. For this purpose a three dimensional Navier-Stokes computer code(AADL3D) has been developed and case studies have been performed by comparing the normal force coefficient and the moment coefficient of a missile body for different jet flow conditions including jet pressure and jet Mach number. The results show different behavior of normal force and moment variation according to jet pressure variation and jet Mach number variation. From the detailed flow field analyses, it is verified that most of the normal force loss and the pitching moment generation are taken place at the low-pressure region behind the jet nozzle. Furthermore, it is shown that the pitching moment can be efficiently reduced by obtaining the lateral thrust through higher jet Mach number rather than through high jet pressure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.10
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• pp.955-965
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• 2010

The MH-75 propeller for the MAV propulsion is designed using a free vortex design method which considers design parameters such as the hub-tip ratio, the twist angle distribution, the maximum camber location and the chord length of the propeller blade. Aerodynamic characteristics of the MH-75 propeller are predicted by changing the flight speed using the frequency domain panel method. And, the thrust characteristics of the MH-75 propeller are measured using the balance system of the subsonic wind tunnel for the validation of numerical results. The performance characteristics of the MH-75 propeller satisfied with design requirements. Numerical results of the MH-75, which are predicted by the frequency domain panel method, are more agree with experimental results compare with XFOIL.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.133-137
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• 2008

Results of combustion tests performed for a regenerative cooling combustor of a 30 tonf-class liquid rocket engine were described. The combustion chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. The combustion chamber is composed of mixing head, baffle injector, and regenerative cooling chamber. The hot firing tests were performed at design and off-design points. The test results show that the combustion characteristic velocity is in the range of 1738${\sim}$1751 m/sec and the specific impulse of the combustion chamber is in the range of 253${\sim}$270 sec. The peak of combustion characteristic velocity and specific impulse for this combustor is shown at mixture ratio of 2.35 and 2.5, respectively.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.10
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• pp.868-874
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• 2015

Thrust variation is an essential parameter in a space mission such as landing on an atmosphereless planet or docking a spacecraft. In order to achieve the thrust variation control, using throttleable injector is a representative and general method. A dual-manifold injector, one of throttleable injectors, was used to control mass flow rate. Five kinds of injectors were designed and investigated in order to compare the spray characteristics of the dual-manifold injector with various tangential entries. Spray angles and patterns were measured to determine external flow characteristics and film thicknesses were measured in order to investigate the internal flow patterns.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.75-82
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• 1999

Theoretical analysis and performance test of Jet vane type Thrust Vector Control(TVC) were conducted using supersonic cold-flow system. The use of TVC Systems an in particular jet vanes, are currently being researched for use in air launch, ship launch, underwater launch and high altitude maneuvering of tactical missiles and rockets. The necessity to generate control forces to rapidly change the course of the missile is frequently required when traditional, exterior aerodynamic surfaces are unable to produce these forces, when the flow over the control surface is insufficient. This situation can occur at launch, or high angles of attack of the control surfaces. Jet vanes peformed well at all altitudes and environmental conditions, and jet vanes are extremely effective at deflection angles up to as high as $30^{\circ}$, make them ideal for the launch and maneuver applications. In this study, performance test of supersonic cold-flow system and visualization of supersonic jet was conducted, and shape and deflection angle effect of two types of jet vanes are investigated.