• 한국군사과학기술학회지
• /
• 제22권1호
• /
• pp.124-134
• /
• 2019

As the types of North Korea's provocation are diverse and unexpectable in the costal area, ROK navy needs to develop countermeasures, such as costal defence guided rockets. Recently ROK navy developed the PKX-B which is equipped with the new 130 mm guided rocket. The most popular rockets are LOGIR for short range targets, 130 mm guided rocket for middle range targets and Spike-NLOS for long range targets. As various guided rockets are developed, it is required to develop a guided rocket analysis model and it's analysis methodology. In addition, these guided rockets can be installed on any platforms; ground vehicle, aircraft and warship. The paper proposes systematic methodology to estimate the operational effectiveness of costal defence guided rockets. A case study exploiting the ARENA simulation model is explained to demonstrate the implementation of the proposed methodology.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2006년도 제26회 춘계학술대회논문집
• /
• pp.156-160
• /
• 2006

본 연구에서는 액체 추진제 로켓엔진의 연소기에 주로 사용되는 액체-액체 동축 스월형 분사기의 분무특성에 대해 고찰하였다. 액막의 분열에는 선형 안정성 이론[1]을 도입하였고 분열 후 충돌에는 충돌이후 분열이 고려된 Post[2]의 모델을 사용하였으며, solver로는 KIVA[3]를 사용하였다. 이러한 모델을 통해 디젤 엔진에 적합한 고속 분사와 로켓엔진에 적합한 저속 분사를 각각 검증하였고 실험결과와 잘 일치하는 것을 볼 수 있었다.

• 한국추진공학회지
• /
• 제20권1호
• /
• pp.28-36
• /
• 2016

본 논문에서는 핀틀을 이용한 가변추력 고체 추진기관의 추력을 제어하기 위하여 이론적으로 모델을 구하고 압력 제어기를 설계하였다. 고전적인 모델 선형화 및 비례-적분제어를 설계했을 때 실제 모델의 비선형성에 의해 발생하는 제어기 성능 저하를 줄이기 위해 이득 계획 기법을 적용하였다. 시스템의 특징을 고려하여 연소관 내부 체적 변화에 따라 이득을 조절하는 방법과, 연소관 내부 압력에 의해 이득을 조절하는 방법으로 두종류의 이득 계획 제어기를 설계하였다. 각 제어기를 가변 추력기 모델에 적용하여 폐루프 시스템 응답특성을 비교하였으며 가변 추력 추진기관 특성에 따라서 어떤 제어기를 선택하는 것이 유리한지 제안하였다.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2005년도 제24회 춘계학술대회논문집
• /
• pp.57-60
• /
• 2005

A diagnostic investigation is carried out to examine the igniter jet turbulence effects on the internal ballistics of solid rocket motors with divergent port. The numerical studies have been carried out with the help of a two dimensional k-omega turbulence model. It was inferred that increasing the igniter jet turbulence intensity is a possible way to decrease the pressure spike and pressurization rate, marginally during the ignition transient, by altering the location of the secondary ignition in solid rocket motors with non-uniform port.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2008년도 제30회 춘계학술대회논문집
• /
• pp.289-292
• /
• 2008

Highly over-expanded nozzle of the rocket engines will be excited by non-axial forces due to flow separation at sea level operations. Since rocket engines are designed to produce axial thrust to power the vehicle, non-axial static and/or dynamic forces are not desirable. Several engine failures were attributed to the side loads. Present work investigate the unsteady flow in an over-expanded rocket nozzle in order to estimate side load during a shutdown/starting. Numerical computations has been carried out with density based solver on multi-block structured grid. Present solver is explicit in time and unsteady time step is calculated using dual time step approach. AUSMDV is considered as a numerical scheme for the flux calculations. One equation Spalart-Allmaras turbulence model is selected. Results presented here is for two nozzle pressure ratio i.e. 100 and 20. At 100 NPR, restricted shock separation (RSS) pattern is observed while, 20 NPR shows free shock separation (FSS) pattern. Side load is observed during the transition of separation pattern at different NPR.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2009년도 제33회 추계학술대회논문집
• /
• pp.433-438
• /
• 2009

고체 및 액체 추진기괸에 적용하기 위하여 액체 실리콘 함침법(LSI)을 이용한 C/SiC 복합재료를 개발하였다. 연소시험을 통하여 C/SiC 복합재료의 우수한 성능을 확인하였으며 산화에 의한 삭마 모델을 제시 하였다.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2007년도 제29회 추계학술대회논문집
• /
• pp.248-251
• /
• 2007

블로우다운 산화제 공급 방식은 부품 수가 적어 신뢰도가 높고, 별도의 가압 탱크가 존재하지 않아 발사체의 총 중량을 감소시킨다는 측면에서 효과적이지만, 일정한 추력 성능을 보장하지 않는 단점을 내재하고 있다. 따라서 본 연구에서는 각 부품의 수학적 모델링을 통해 블로우다운 산화제 공급방식 하이브리드 로켓의 성능 해석 모델을 제안하여, 고유의 성능 특성에 대하여 고찰하였다.

• 한국항공운항학회지
• /
• 제13권3호
• /
• pp.9-21
• /
• 2005

The stable motion has been judged by mathematical modeling of the conditions that a rocket flies flexibly to take an active part in atmosphere. In this paper, the rocket conditions consist of the air speed, thrust and automatic attitude control. Aerodynamic force, a critical trust and a critical air speed are determined by comparing mathematical knowledges with eigenfrequencies of vibration equation. And then rocket object model is designed. Parameters and eigenfrequencies are used in dimensionless forms for in general applications by eliminating restrictions such as dimension, weight and select of materials.

• 한국군사과학기술학회지
• /
• 제2권2호
• /
• pp.218-225
• /
• 1999

This paper contains the computational simulation of a free rocket with a parachute and the development of a firing table for each range. To obtain the trajectory of the rocket, 6 DOF model of rocket with parachute was generated and the wind tunnel test was done for the input parameters. Good agreement was obtained between the analysis of trajectory and the flight test result. Also the trajectory error analysis was performed by the Monte Carlo simulation. As a result of simulation, the CEP(Circular Error Probability) of the firing table was calculated.

• 한국추진공학회:학술대회논문집
• /
• 한국추진공학회 2008년 영문 학술대회
• /
• pp.260-265
• /
• 2008

A numerical scheme for solid propellant rocket has been studied using preconditioning method to research unsteady combustion processes for the double-base propellant with a converging-diverging nozzle. The Navier-Stokes equation is solved by dualtime stepping method with finite volume method. The turbulence model uses a shear stress transport modeling. The species equation follows up the method of Xinping WI, Mridul Kumar and Kenneth K. Kuo. A preconditioned algorithm is applied to solve incompressible regime inside the combustor and compressible flow at nozzle. Mass flux was evaluated using modified advective upwind splitting method. The simulated result the comparison a fully coupled implicit method and a semi implicit method in terms of accuracy and efficiency. This report shows the result of solid rocket propellant combustion.