• Advances in aircraft and spacecraft science
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• v.10 no.4
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• pp.317-334
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• 2023

This paper presents a basic methodology and a set of numerical tools for the preliminary design of solid-fueled ramjet missiles. An elementary code determines the baseline system configuration comprised of warhead, guidance-control, and propulsion masses and geometries from specific correlations found in the literature. Then, the system is refined with the help of external and internal ballistics codes. Equations of motion are solved for the flight's ascending, cruising, and descending stages and the internal ballistic set of equations designs the ramjet engine based on liquefying fuels. The combined tools sized the booster and the ramjet sustainer engines for a long-range missile, intended to transport 200 kg of payload for more than 300 km range flying near 14,000 m altitude at Mach 3.0. The refined system configuration had 600 mm in diameter and 8,500 mm in length with overall mass of 2,128 kg and 890 kg/m3 density. Ramjet engine propellant mass fraction was estimated as 74%. Increased missile range can be attained with paraffin-polyethylene blend burning at near constant regression rate through primary air mass flow rate control and lateral 2-D air intakes.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.555-558
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• 2010

This paper presents a design of solid rocket motor with high length to diameter applied composite propellant. Solid rocket motor with high L/D ratio can be generated erosive burning and combustion instability on longitudinal mode. Especially, Erosive burning can effectively prolong the initial pressure spike in some star grain motors. That is, the study shows design of grain, internal ballistics and structural analysis in order to perform system requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.18-25
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• 2006

In order to perform system requirements quickly and accurately, an automatic design program of solid rocket motors(SRM) structure designated as the 'ProDes software' has been developed and verified. from given system design criteria and constraints, it has the capabilities to design, analysis, simulation and drawing process to greatly reduce the over 'design cycle time' and manpower of a project. The conception of the program is modular, and calculations are performed step by step allowing parametric design studies and providing final selected design goal. Each configurations of SRM components and joint types composed of various master models is obtained from the data base module of the library. Between the design results of the ProDes software and those of the previous detail design of the established motor showed good agreements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.374-377
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• 2017

A cold-flow test equipment was designed to carry out the performance verification of TDACS. For that purpose, the pressure dynamics in the solid rocket motor combustor and the cold-flow test was modeled, and the response time showing the dynamic characteristics of each model was obtained. In this paper, the system response time of the cold-flow test was designed to be equal to that of the motor, making the dynamic response in cold-flow and hot gas condition to be similar.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.7-10
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• 2006

A quasi-equilibrium model is used in order to aid movement and ballistic analysis for a propellant actuated device(PAD) system. The validity of the model is examined by experiments of a PAD system. The appropriateness of its usage for application was explored by comparing the tendency of experiments and analysis results, and the coefficients of friction and heat loss were obtained. The design method developed will be applied to the design of PAD systems.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.48-55
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• 2014

In present study, design and hot fire tests of the pryostarter are conducted. To prevent the turbopump RPM overshoot, regressive mass flow rate profile is applied. Sudden decrease of the mass flow rate at the end of the propellant burning is realized as well. Firing test results show good agreements with the design requirements. Through the study with ignition substance variations, combustion products and ignition performances are improved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.265-270
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• 2010

This paper presents the rocket motor support structure to position solid rocket engine within a missile fuselage. When the rocket motor is mounted inside a missile fuselage, fuselage structure must be designed to withstand various structural problems resulting from inserted rocket motor such as axial thrust force, shock/vibration, axial deformation of the rocket motor tank in addition to the flight loads. The motor support structure system proposed in this paper proved to be very simple and efficient while satisfying all the design requirements.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.39 no.6
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• pp.631-638
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• 2015

In this study, semiquantitative failure mode, effects, and criticality analysis (FMECA) for the reliability analysis of a solid rocket propulsion system is performed. The semiquantitative FMECA is composed of failure mode and effects analysis (FMEA) and criticality analysis (CA). To perform FMECA, the structure of the solid rocket propulsion system is divided into 43 parts down to the component level, and FMEA is conducted at the design stage considering 137 potential failure modes. CA is then conducted for each failure mode, during which the criticality number is estimated using the failure rate databases. The results demonstrate the relationship between potential failure modes, causes, and effects, and their risk priorities are evaluated qualitatively. Additionally, several failure modes with higher criticality and severity values are selected for high-priority improvement.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.5
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• pp.132-139
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• 2013

Liquid-propellant rocket engines are widely used all over the world, thanks to their high performances thrust, in particular high thrust-to-weight ratio. The sucess rate of the launching of the liquid propulsion is similar to the solid one even though it has more complex mechanical system. In general, liquid propulsion is seemed as a mature technology, the requirements of a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.3
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• pp.52-59
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• 2001

In this study, a technique has been developed to analyze the performance of the launching system using a gas generator of solid propellant. The physical model which described the thermodynamic behaviors of all launching devices from gas generator to canister and the dynamic behavior of missile was established, applying the empirical coefficients in the heat loss model. The processes of combustion, flow, and heat transfer inside the chamber of gas generator and the launching system were simulated by numerical method. The theoretical analysis guided the optimal design of gas generator and system, which made the launching system satisfy the requirements of good performance and high reliability.