• International Journal of Aerospace System Engineering
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• v.2 no.2
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• pp.29-33
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• 2015

The multidisciplinary design optimization (MDO) of a launch vehicle (LV) with a hybrid rocket engine (HRE) was carried out to investigate the ability of an HRE for a single-stage LV. The non-dominated sorting genetic algorithm-II (NSGA-II) was employed to solve two design problems. The design problems were formulated as two-objective cases involving maximization of the downrange distance over the target flight altitude and minimization of the gross weight, for two target altitudes: 50.0 km and 100.0 km. Each objective function was empirically estimated. Several non-dominated solutions were obtained using the NSGA-II for each design problem, and in each case, a trade-off was observed between the two objective functions. The results for the two design problem indicate that economical performance of the LV is limited with the HRE in terms of the maximum downrange distances achievable. The LV geometries determined from the non-dominated solutions were examined.

• Journal of Aerospace System Engineering
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• v.9 no.2
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• pp.34-40
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• 2015

Staged combustion cycle engines are well known to have high combustion efficiencies and specific impulse. In this study, design of mixing head part of combustion chamber for 8tonf class staged combustion cycle rocket engine (ES-08) was performed. Structural stability of the mixing head part of the combustion chamber is very important design factor because it is loaded by high temperature and high pressure of fuel and oxidizer as well as by thrust load simultaneously. Uniformity of flow distributions of the propellants to the injectors is also important factor. First, a basic configuration for the ES-08 mixing head part was designed on the basis of the structural design requirements. And then, the structural analyses were performed on the basic configuration as well as some of reinforced configurations. As the structural analyses results, the most stable configuration was selected for the ES-08 mixing head part. In order to examine the uniformity of the flow distributions of the propellants through the manifold of the mixing head, flow analysis was performed based on the selected configuration. The results of the flow analysis showed that the fuel and the oxidizer were uniformly supplied to the injector.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.374-377
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• 2007

본 보고서는 소형 로켓 제작에 앞서 행해지는 설계 단계를 기술하였다. 각 요구 조건에 맞춰 사용되어질 추력, 추진제 특성, mass distribution, 동체 형상 설계, Motor 설계에 대해 다뤘다. 로켓 설계에 있어서 가장 중요한 추진제 선정과 노즐형상 설계를 중점적으로 다루었으며, 차후 본 과정에 의하여 소형로켓을 제작하겠다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.97-100
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• 2008

A hybrid sounding rocket carrying about 10kg payload reaching up to 15km altitude has been designed. The commercial seamless aluminium tube and liquid ${N_2}O$ without pressurization devices were chosen as rocket motor case and oxidizer supply system respectively. A hybrid rocket engine performing required propulsion impulse is designed with time dependent internal ballistic scheme. Engine performance, aerodynamic characteristics, and trajectory were predicted by a integral technique of internal ballistics and external ballistics. The design results were evaluated by comparison with previous experimental data, technical reports, and literatures.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.5
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• pp.109-115
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• 2004

A 1-D system design program has been developed for the preliminary design of the turbopump system in liquid rocket engines, which use LOx and kerosene as propellants. Gasgenerator cycle and staged combustion cycle were considered as turbopump type liquid rocket engine systems. In the system analysis, mass flow balance, thrust, specific impulse, mixture ratios, turbopump power, and turbine expansion ratio of engine system were analyzed. Results show that most of the parameters agree well with real engine parameters except gasgenerator. Therefore, the l-D system design program developed in this study can be used to derive the preliminary design parameters of a turbopump with any thrust level liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.133-137
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• 2010

A dual pulse solid rocket motor has several advantages compared to the single one. The range and the terminal velocity of the guided missile can be remarkably increased by the application of the pulse separation device(PSD) to the solid rocket motor which resulted in appropriate thrust distribution. In this study, the subscale dual pulse solid rocket motor with the bulkhead type PSD was designed, manufactured, and fire-tested. The bursting pressure, thermal characteristics, and the structural safety of the PSD were obtained by the tests and the results will be applied to the design of full-scale dual pulse rocket motor.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.45-52
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• 2020

In this paper, a subminiature solid rocket motor(SSRM) was designed to develop a miniature smart-bullet and the designed propellant grain was made of thermoplastic propellant for production convenience of inner shape. The internal ballistics analysis and ground test were performed to investigate the performance of SSRM. And a numerical simulation was carried out to obtain basic data on the design of safety distance between the nozzle outlet and a gunner, the temperature distribution of exhaust gas was analyzed by comparing a numerical simulation and the results of IR camera.

• Proceedings of the KSME Conference
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• 2004.11a
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• pp.562-566
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• 2004

Thermal shock is a physical phenomenon that occurs in the condition of the exposure of a rapidly large temperature and pressure change of in the quenching condition of material. The rocket nozzle is exposed to high temperature combustion gas, it may have failure and erosion deformation. So, it is important to select a suitable material having excellent thermal shock properties and evaluate these materials in rocket design. In this study, the temperature gradient and crack initiation of rocket nozzle material is investigated using by FEM under thermal shock condition. This is very important information in the design process of thermal structure.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.10-13
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• 2002

Pogo is the instability resulting from the interaction between rocket structure and propulsion system of liquid propellant rocket. The coupling of structure and propulsion system can lead to severe problem in rocket. For the analysis of pogo, a time-invariant linearized mathematical model is developed for a selected flight time. Propulsion system is modeled using element representations for each components. The constitutive equation of propulsion system is a homogeneous second-order equation form in the Laplace domain. Rocket structure is modeled using FEM. From the results of modal analysis of structure, the behavior of structure can be represented. System equations for coupling structure and propulsion system are composed of all propulsion system equations and vehicle motion equations reacting on the vehicle by each component of propulsion system. The stability is obtained by the eigen solution of system matrix. The optimization of the design variables such as size, place of accumulator for suppressing pogo instability is carried out. This article of study can be used to determine the degree of stability, and guide the design of pogo suppression system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.215-218
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• 2009

A design of a subscale solid propellant rocket motor was conducted to do the similitude experiments for the large scale rocket motor. One of the main factor to subscale was the mach number of the solid propellant flume through a nozzle exit The analysis of the flume flow was done to obtain the mach number for the large and subscale rocket motor. The flume shapes on the non dimensional axises by the nozzle exit diameter was matched each other. The propellant grain of a subscale solid rocket motor was designed by the profile of pressure vs time obtained by the mach number of the flume shape. Some analyses of the theoretical solution were compared with the results of the ground static test.