• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.57-60
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• 2008

There exist a lot of factors and restrictions for the design of solid rocket motors like burning rate, of solid propellant, demanded thrust, chamber pressure, diameter, length, weight and acceleration. For the optimization of these factors and restrictions, integrated design software for internal/external ballistic analysis was developed and verified by the performance test of solid rocket motors.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.4
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• pp.44-53
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• 2022

The depressurization rate in a thrust variable solid rocket motor is the major factor that has the greatest influence on the thrust termination performance. In this study, the depressurization rates range of model solid rocket motor was identified and major factors affecting the depressurization rate were found. It is important for actual system design to understand the depressurization rate of the system that can satisfy the target performance as well as the extinguishing characteristics of the solid propellant. The methodology for obtaining the depressurization rate model in this study is considered to be applicable to the optimal design of the thrust terminable propulsion system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.167-174
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• 2011

Solid propulsion systems have simple structures compared to other propulsion systems and are suitable to long-term storage. However the systems generally have limits on control of thrust levels. In this paper we suggest control algorithms for combustion chamber pressure of variable thrust solid propulsion systems using special nozzles such as pintle valve. For this we use a simple pressure change model by considering only mass conservation within the combustion chamber, design a classical algorithm and also a nonlinear controller using feedback linearization technique. Derived thrust equation and designe a thrust control model. We design the proportion-integral controller for linearizing about operating point. We also demonstrate the performance of controller model through numerical simulations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.147-150
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• 2005

Space launch vehicles mainly use the liquid-propellant propulsion system which has easy thrust control ability and high specific impulse for that the payload like satellite and spacecraft should be entered into exact orbit. However, the liquid-propellant propulsion system is very difficult to develop because it is more complicate than the solid rocket propulsion system and demands very high technology. In space launch vehicle developing procedure the system design level is very important thing to reduce cost, shorten schedule, and improve the performance. The system design process was introduced for selecting the best liquid-propellant propulsion system on this paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.228-231
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• 2011

Burning rate of a propellent is an essential factor when designing a propulsion system. In order to come up with burning rate, first we need to design and build propellent grain to get neutral pressure curve. Then check the pressure with ground test and calculate the burning rate using burning rate equation. This burning rate is then compared to the burning rate of a propellent which was resulted from making a standardized specimen and combusting it using a strand burner. An accurate burning rate is calculated after comparing those two burning rates. For this study, compact propulsion system was designed, produced, tested and analyzed in order to get burning rates, an essential factor in propulsion system design, in an effective way.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.238-238
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• 2012

Report is devoted to place and role of numerical simulation in design of rocket propulsion systems. In introduction advanced solutions in liquid propellant rocket engines design are presented. Further essence of design process described briefly. The central place of method of solution of direct problem in design process was shown. Numerical simulation for solving direct problem of fluid dynamic was used as the alternative to theoretical and experimental approaches. Main features of numerical models of processes in propulsion systems were observed. Some results of simulation and (or) design of different types of chemical propulsion system were presented also. The combined rocket engine, rocket engine with injection of after-turbine gas into supersonic part of the nozzle, solid propellant engine and hybrid propulsion engine are under consideration.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.601-604
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• 2004

Precise position and attitude control of pico-satellite requires huge number of impulses of the order of 10$^{-6}$ Ns. MEMS solid rocket array is a promising propulsion system but the higher degree of miniaturization causes unreliable operation mainly due to quenching. In order to breakthrough this situation, a novel design of solid micro-rocket is proposed, which generates tiny impulses repetitively from a single rocket not from array. This unique micro-rocket is based on the utilization of quenching, which causes propellant reaction to sustain only in a small area. A test chip of a micro solid propellant tank and micro heater array is fabricated and ignition test is conducted. Obtained results show the feasibility of this concept and future direction of this quenching-based propulsion is discussed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.330-339
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• 2008

Ground-based interceptors(GBI) comprise a major element of the strategic defense against hostile targets like Intercontinental Ballistic Missiles(ICBM) and reentry vehicles(RV) dispersed from them. An optimum design of the subsystems is required to increase the performance and reliability of these GBI. Propulsion subsystem design and optimization is the motivation for this effort. This paper describes an effort in which an entire GBI missile system, including a multi-stage solid rocket booster, is considered simultaneously in a Genetic Algorithm(GA) performance optimization process. Single goal, constrained optimization is performed. For specified payload and miss distance, time of flight, the most important component in the optimization process is the booster, for its takeoff weight, time of flight, or a combination of the two. The GBI is assumed to be a multistage missile that uses target location data provided by two ground based RF radar sensors and two low earth orbit(LEO) IR sensors. 3Dimensional model is developed for a multistage target with a boost phase acceleration profile that depends on total mass, propellant mass and the specific impulse in the gravity field. The monostatic radar cross section (RCS) data of a three stage ICBM is used. For preliminary design, GBI is assumed to have a fixed initial position from the target launch point and zero launch delay. GBI carries the Kill Vehicle(KV) to an optimal position in space to allow it to complete the intercept. The objective is to design and optimize the propulsion system for the GBI that will fulfill mission requirements and objectives. The KV weight and volume requirements are specified in the problem definition before the optimization is computed. We have considered only continuous design variables, while considering discrete variables as input. Though the number of stages should also be one of the design variables, however, in this paper it is fixed as three. The elite solution from GA is passed on to(Sequential Quadratic Programming) SQP as near optimal guess. The SQP then performs local convergence to identify the minimum mass of the GBI. The performance of the three staged GBI is validated using a ballistic missile intercept scenario modeled in Matlab/SIMULINK.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.207-210
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• 2008

We developed an integrated design system for a solid rocket motors. We can do a conceptual design of a solid rocket motor easily and quickly with this system. It consists of four modules, or, size design, structure design, grain design and performance analysis module. Size design module determines the lengths and diameters of some major parts, which results in fixing the whole size of a motor. Structure design module has many master models, which enables a designer can do a conceptual design of almost parts of motor structures. Grain design module can design a solid fuel according to the result of structure design. Finally performance analysis module verifies the proposed design with the output from grain design module.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.4
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• pp.18-25
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• 2011

Solid propulsion systems have simple structures compared to other propulsion systems and are suitable for long-term storage. However the systems generally have limits on control of thrust levels. In this paper we suggest control algorithms for combustion chamber pressure of variable thrust solid propulsion systems using special nozzles such as pintle valve. For the pressure control within the chamber, we use a simple pressure change model by considering only mass conservation within the combustion chamber, design a classical algorithm and also a nonlinear controller using the feedback linearization technique. Also we derive the equation of the thrust for an under-expanded one-dimensional nozzle and then design a proportional-intergral controller after linearizing the thrust model for an operating point. Finally, we demonstrate the performance of the controller through a numerical simulation.