• Proceedings of the KSME Conference
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• 2003.11a
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• pp.496-501
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• 2003

Computational study is performed to understand the fluidic thrust vectoring control of an axisymmetric nozzle, in which secondary gas injection is made in the divergent section of the nozzle. The nozzle has a design Mach number of 2.0, and the operation pressure ratio is varied to obtain the different flow features in the nozzle flow. The injection flow rate is varied by means of the injection port pressure. Test conditions are in the range of the nozzle pressure ratio from 3.0 to 8.26 and the injection pressure ratio from 0 to 1.0. The present computational results show that, for a given nozzle pressure ratio, an increase of the injection pressure ratio produces increased thrust vector angle, but decreases the thrust efficiency.

• Aerospace Engineering and Technology
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• v.6 no.1
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• pp.136-145
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• 2007

A computation was made to predict the movement of the thrust center position due to the rocket nozzle deflection. Three dimensional computations were done for the nozzle deflection angles of 0/1/3 degrees, and the oscillation of aerodynamic coefficients, not observed for the axisymmetric cases, was encountered. The position of the thrust center was found to be at -16 mm and -4 mm for the deflection angles of 1 and 3 degrees, respectively, and it can be concluded that the thrust center movement due to nozzle deflection is negligible. In addition to the computational results, the mechanism of thrust generation in a rocket engine is described with a brief mathematical derivation as it is sometimes mistaken. Also presented are some descriptions on the problem of pressure center definition for symmetric cases such as a rocket external flow problem and the nozzle deflection case.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.4
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• pp.392-398
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• 2009

Theoretical thrust equations for the diverse nozzle expansion condition were derived. By using the obtained thrust equations, parametric studies were carried out to estimate the effect of pressure exponent, minimum operation pressure, ambient pressure and extinguishment pressure on thrust modulation performance in pintle-nozzle solid rocket motors. Analysis results showed that thrust turndown ratio can be easily attained by small nozzle-throat area variation at high pressure exponent, low minimum operation pressure, high ambient pressure and high extinguishment pressure condition. At those conditions, the highest chamber pressure to obtain the intended thrust turndown ratio can be minimized.

• Journal of Industrial Technology
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• v.20 no.B
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• pp.45-53
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• 2000

A numerical study is done on the thrust vector control using gaseous secondary injection in the rocket nozzle. A commercial code, PHOENICS, is used to simulate the rocket nozzle flow. A $45^{\circ}-15^{\circ}$ conical nozzle is adopted to do numerical experiments. The flow in a rocket nozzle is assumed a steady, compressible, viscous flow. The exhaust gas of the rocket motor is used as an injectant to control the thrust vector of rocket at the constant rate of secondary injection flow. The injection location which is on the wall of rocket is chosen as a primary numerical variable. Computational results say that if the injection position is too close to nozzle throat, the reflected shock occurs. On the other hand, the more mass flow rate of injection is needed to get enough side thrust when the injection position is moved too far from the throat.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1108-1111
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• 2017

The nozzle exit shape is scarfed according to the external shape of missile when the nozzle axis should be canted from missile axis due to missile system application. There is inevitable thrust loss for the scarfed nozzle comparing to non-scarfed nozzle. The numerical analysis is necessary to calculate the thrust loss in design process, and ground tests of rocket motor were performed to verify the calculation results. From the comparison of non-scarfed nozzle and scarfed nozzle experiment results, the thrust loss from calculation was about 16.6% and that from experiments was about 15.0%.

• Proceedings of the KSME Conference
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• 2004.04a
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• pp.1637-1642
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• 2004

Experimental and computational studies were performed to investigate the effectiveness of a thrust vectoring method using a counterflow concept. A shadowgraph method was used to visualize the supersonic jet expanded from a two-dimensional convergent-divergent nozzle and deflected by a now suction. The primary nozzle pressure and suction nozzle pressure ratios are varied between 3.0 and 5.0, and between 0.2 and 1.0 respectively. The present experimental and computational results showed that, for a given primary nozzle pressure ratio, a decrease in the suction nozzle pressure ratio produced an increased thrust vector angle, and during the change processes of the suction pressure, a hysteresis effect of the thrust vectoring was found through the wall pressure distributions.

• Journal of computational fluids engineering
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• v.16 no.3
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• pp.1-7
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• 2011

As a preliminary design study to achieve target aerodynamic performance, this work was conducted on an original nozzle with 9 flutes in order to design a fluted nozzle with 12 flutes. The thrust and rolling moment of the nozzle with 12 flutes were analyzed using a CFD code according to the depth and rotation angle of the flutes. Based on this, a fluted nozzle with 12 flutes was optimized to yield the same thrust as that of the original nozzle with 9 flutes. The response surface method was applied for shape optimization of the fluted nozzle and design variables were selected to determine the depth angle and rotation angle of the flutes. An optimized shape that led to a thrust as strong as that of the original nozzle was obtained.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.5
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• pp.27-34
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• 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
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• 2010.11a
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• pp.170-176
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• 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.

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