• Journal of the Korean Society of Propulsion Engineers
• /
• v.14 no.6
• /
• pp.25-30
• /
• 2010

Dual throat nozzle(DTN) is recently attracting much attention as a new concept of the thrust vectoring technique. This DTN is designed with two throats, an upstream minimum and a downstream minimum at the nozzle exit, with a cavity in between the upstream throat and exit. In the present study, a computational work has been carried out to analyze the performance of a dual throat nozzle(DTN) at various mass flow rate of secondary flow and nozzle pressure ratios(NPR). Two-dimensional, steady, compressible Navier-Stokes equations were solved using a fully implicit finite volume scheme. The present computational results were validated with some experimental data available. Based upon the present results, The control effectiveness of thrust-vector is discussed in terms of the thrust coefficient and the discharge coefficient.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.6 no.2
• /
• pp.1-7
• /
• 2002

In the modern propulsion systems, requited thrust is obtained using a nozzle. Sometimes shock and induced boundary layer separation is generated in an over-expanded convergent-divergent supersonic nozzle. It occurs because the nozzle expansion ratio is too large for a given nozzle pressure ratio (NPR). This phenomenon can be explained that it redefines effective nozzle geometry, shorer nozzle geometry and lower pressure ratio, in a given pressure ratio. Numerical studies were conducted about a fixed geometry 2D nozzle in overexpanded condition and compared with Hunter's experimental result. For the numerical simulation of the supersonic nozzle, Navier-Stokes equations are considered and as a turbulent model, $\kappa$-$\varepsilon$ /$\kappa$-$\omega$ blended SST two equation turbulent model is used. The characteristics of $\lambda$-shape shock systems due to the interaction of shock and boundary layer was investigated in a low NPR. And the result of comparison of thrust value shows that a fixed geometry nozzle can cover required flight mission.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.227-230
• /
• 2008

Igniter system of Solid Rocket Motor is classfied as Forward-end type and After-end type. Forward-end type is used for sustentation of combustion pressure by nozzle plug. But After-end type is used for sustentation of combustion pressure by igniter mount. Igniter Mount is assembled on nozzle throat. Igniter mount holds combustion pressure and igniter release energy. A study has qualificated result of Dynamic behavior for After-end igniter mount of Static Firing Test and Finite element method.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.5
• /
• pp.61-70
• /
• 2017

Robust thrust distribution method of solid DACS is researched. For the case of the system which has higher number of actuation nozzles than the degree of freedom of thrust to be controlled, the robust thrust allocation law which accommodate the abnormal operation is suggested. Assuming the situation that some nozzles are uncontrollable, the error between nozzle throat area command and response can be calculated. The error is used for realtime reshaping of weighting matrix. From the weighting effect, the nozzle which operated abnormally has low responsibility for the command then, the thrust error is reduced. The suggested algorithm is verified by the simulation of abnormal operation condition of DCS and ACS nozzle respectively.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.2
• /
• pp.72-82
• /
• 2017

Nozzleless booster is one of the applicable components for integral rocket ramjet (IRR). In order to predict nozzleless solid booster performance, the simplified theoretical analysis was applied for L/D=5, 6, 7, 9, 11, and 13. Al-HTPB and Zr-HTPB propellant with a high metal content were used to increase the hardness because of the combustion gas flow effect. It was found that the trends between the simplified theoretical analysis and experiments were similar.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.1072-1076
• /
• 2017

In this paper, a numerical study has been performed to analyze flow characteristics of rocket propulsion. Through the ground spin test, combustion chamber pressure was measured. Based on the experimental results, numerical analysis was conducted under various nozzle pressure ratio conditions such as standard, operating and base pressure conditions. And it was compared with quasi-1D solution and experimental result. In addition, the difference in thrust characteristics according to the spin/non-spin of the flow conditions was confirmed at the same nozzle pressure ratio.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.15 no.5
• /
• pp.27-34
• /
• 2011

A numerical analysis was peformed for the supersonic aircraft with variable pitch thrust vector nozzle. Based on the requirement of the mixed turbofan engine of the supersonic aircraft, two dimensional thrust vector nozzle with variable pitch angle was designed. To investigate the effect of the thrust vectoring nozzle, the numerical analysis was conducted by using Fluent under the several pitch deflection angle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.14-17
• /
• 2011

Numerical simulation was carried out for the SNECMA modulatable thrust devices, with four different nozzle contraction angle $45^{\circ}$, $60^{\circ}$, $83^{\circ}$, and $90^{\circ}$, respectively. Results show that $83^{\circ}$ nozzle contraction is better in that it comes up with good thrust level with small 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.170-176
• /
• 2010

A numerical analysis was peformed for the supersonic aircraft with variable pitch thrust vector nozzle. Based on the requirement of the mixed turbofan engine of the supersonic aircraft, two dimensional thrust vector nozzle with variable pitch angle was designed. To investigate the effect of the thrust vectoring nozzle, the numerical analysis was conducted by using Fluent under the several pitch deflection angle.