• Aerospace Engineering and Technology
• /
• v.8 no.1
• /
• pp.144-153
• /
• 2009

To study the plume effects in the rarefied region, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the rarefied region numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the rarefied flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow region, etc, can be investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.37 no.9
• /
• pp.906-915
• /
• 2009

To study the plume effects in the vacuum area, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the vacuum area numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the vacuum flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow area, etc, can be investigated.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.4
• /
• pp.1-8
• /
• 2014

Overexpanded flow of an axisymmetric thruster nozzle is numerically simulated to investigate effects of nozzle pressure ratio (NPR) on the shock structure and thrust performance. The Reynolds-averaged Navier-Stokes equations with k-${\omega}$ SST turbulence model are solved utilizing FLUENT solver. As the NPR is raised, thrust performance monotonically increases with the shock structure and flow-separation point being pushed toward the nozzle exit. It is also discussed that the flow structure at nozzle-exit plane which is immediately affected by a position of nozzle-interior shocks and expansion waves, has strong influence upon the thrust performance of thruster nozzle.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.33 no.5
• /
• pp.72-78
• /
• 2005

We conducted the performance test of micronozzle having nozzle throat diameter of 1.0, 0.5, 0.25 mm in an ambient pressure. We used N2 gas as a cold gas propellant. We varied chamber pressure from 2 to 20 bar and measured the thrust and mass flow rate. Through the test, we concluded that viscous losses were increased with decreasing chamber pressure. We found that micronozzle performance was higher than orifice performance through thrust comparison.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2012.05a
• /
• pp.392-396
• /
• 2012

Experimental and computational study for steady-state pintle driven nozzle throat flow are carried out by changing four pintle shape. Results show that thruster performance is influenced by pintle shape greatly. This attributes to the distorted throat area and chamber pressure change as the pintle shape and its penetration.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.04a
• /
• pp.189-192
• /
• 2011

In solid rockets, a pintle thruster is a modulatable thrust device which controlls nozzle throat area. In this study, effect of bore on aerodynamic loads in a SNECMA modulatable thruster was carried out. Existence of bore resulted in reduced aerodynamic load.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.483-486
• /
• 2009

An investigation of the thrust vectoring nozzle which can be applied to the supersonic variable exhaust system was performed. The maximum mach number of the model aircraft is 1.8 and mission radius is about 400Nm. The cycle analysis are performed at each operating regime of the aircraft and the specifications of the thrust vectoring nozzle were developed. Based upon the requirement of the thrust vectoring nozzle, two dimensional thrust vectoring nozzle were designed and flow analysis was conducted by deflection of the pitch and yaw angle.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.13 no.6
• /
• pp.17-23
• /
• 2009

The survivability of the temporary nozzle extension for an accurate thrust measurement in a full-scaled combustor has been investigated through thermal analyses. The effects of nozzle extension materials and the thickness of thermal barrier coating (TBC) have been elucidated. It has been found that thermal survivability cannot be guaranteed without TBC. The maximum temperature of the nozzle extension decreased with increasing TBC thickness. For hot firing tests, the TBC is thought to be indispensable to the thermo-structural survivability of the nozzle extension made of steel.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 1998.04a
• /
• pp.31-31
• /
• 1998

위성체의 보조추진시스템은 임구궤도까지의 궤도진입 및 임무궤도상에서의 속도 또는 자세제어에 필요한 임펄스를 제공한다. 단일하이드라진 추력기는 하이드라진(H$_2$H$_4$)과 자발적 촉매(Shell 405)의 발열 및 흡열 열분해 반응에 의해 발생하는 질소($N_2$), 수소(H$_2$), 암모니아(NH$_3$), 혼합가스를 노즐을 통해 방출하므로써 요구되는 impulse를 얻는다. 단일하이드라진 추력기 설계는 주입기, 촉매대, 노즐과 기타 설계 형태에 따른 다지관, 링, 스크린, 지지판 등의 부수적인 부품으로 구성된다. 추력기 제작 과정은 크게 piece-parts 기계가공, HEA(Head End Assembly)와 TCA(Thrust Chamber Assembly)로 구성되고 각 세부공정마다 전수시험 및 검사를 가진다. 연소시험설비는 최소 모사진 공 수준이 고도 100,000 ft(8.4 torr)를 만족시킬 수 있는 진공설비, 시험제어부, 성능변수 측정 및 처리부, 추진제 가압 공급부, 기타 환경 안전 및 부대 설비로 구성된다. 추력기 연소성능시험 절차는 추진제 충전 및 오염 여부 표본 검사, 가압 및 공급 라인 이상여부 확인, 추력기 장착, 추진제 가압 및 공급, 시험장치 보정, 진공 모사 및 연소성능시험, data 처리 등으로 구성된다.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.43 no.4
• /
• pp.304-310
• /
• 2015

Pintle thrusters use pintle stroke to change nozzle throat area, and this controls thrust. Using MATLAB, one-dimensional simulation has been investigated and the results are compared to those of cold flow tests and computational fluid dynamics for the pintle thruster of Chungnam National University. The prediction based on one-dimensional flow theory shows good agreement with measurements for chamber pressure, but deviates for thrust, partly because of nozzle wall separation. Computational results show that nozzle wall separation occurs at an early stage of nozzle expansion, near the design nozzle throat, for the course of pintle strokes. Empirical thrust prediction incorporates nozzle wall separation, and thus 1-D simulation using empirical thrust prediction showed good results for an early stage of pintle stroke.