• 한국전산유체공학회:학술대회논문집
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• 1999.05a
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• pp.85-89
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• 1999

Numerical Analysis has been done for the supersonic off-design jet flow due to the pressure difference between the jet and the ambient fluid. The difference of pressure generates an oblique shock or an expansion wave at the nozzle exit, The waves reflect repeatedly at the center axis and on the sonic surface in the shear layer, and the pressure difference is resolved across these waves interacted with the turbulence mixing layer. In this paper, the axi-symmetric Navier-Stokes equation has been used with two equation $k-{\varepsilon}$ turbulence closure model. The second order TVD scheme with flux limiters, based on the flux vector split by the smooth eigenvalue split, has been used to capture internal shocks and other discontinuities. The correction term for the compressible flow and the damping function are used in the turbulence model. Numerical calculations have been done to analyze the off-design jet flow due to the pressure difference. The variation of pressure along the flow axis is compared with an experimental result and other numerical result. The characteristics of the interaction between the shock cell and the turbulence mixing layer have been analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.11a
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• pp.11-14
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• 2001

3-D compressible flow analysis was conducted by using mixing plane method for turbine system which is consisted of partial admission nozzle and rotor. Computational results are shown oblique shock wave in blade leading and trailing edge and also shown flow separation along suction surface of blade due to abrupt blade curvature. But computational results are well agree with 1-D calculation results and experimental data.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.04a
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• pp.43-51
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• 1997

Steady and unsteady numerical simulations are conducted for the experiments performed to investigate the ram accelerator flow field by using the expansion tube facility in Stanford University. Navier-Stokes equations for chemically reacting flows are analyzed by fully implicit and time accurate numerical methods with Jachimowski's detailed chemistry model for hydrogen-air combustion involving 9 species and 19 reaction steps. Although the steady state assumption shows a good agreement with the experimental schlieren and OH PLIF images for the case of $2H_2$+$O_2$+$17N_2$, it fails in reproducing the combustion region behind the shock intersection point shown in the case of $2H_2$+$O_2$+$12N_2$, mixture. Therefore, an unsteady numerical simulation is conducted for this case and the result shows all the detailed flow stabilization process. The experimental result is revealed to be an instantaneous result during the flow stabilization process. The combustion behind the shock intersection point is the result of a normal detonation formed by the intersection of strong oblique shocks that exist at early stage of the stabilization process. At final stage, the combustion region behind the shock intersection point disappears and the steady state result is retained. The time required for stabilization of the reacting flow in the model ram accelerator is found to be very long in comparison with the experimental test time.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.405-408
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• 2005

A theoretical analysis is developed for an annular injection type supersonic ejector having a second-throat downstream under the assumption that the Fabri-chocking is placed in mixing chamber. Non-mixing theory is applied to formulate secondary flow pressure in the region between inlet of the mixing chamber and Fabri-chocking. To describe the shock standing at the inlet of the mixing chamber, two dimensional oblique shock relations are used and it is assumed that the shock affects only primary flow at Fabri-chocking plane. In conclusion, it agrees well with experiments in case of small contracting angle of mixing chamber, under 4degrees.

• Journal of computational fluids engineering
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• v.4 no.2
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• pp.57-66
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• 1999

Numerical Analysis has been done for the supersonic off-design jet flow due to the pressure difference between the jet and the ambient fluid. The difference of pressure generates an oblique shock or an expansion wave at the nozzle exit. The waves reflect repeatedly on the center axis and the sonic surface in the shear layer. The pressure difference is resolved across these reflected waves. In this paper, the axi-symmetric Navier-Stokes equation has been used with the κ-ε turbulence model. The second order TVD scheme with flux limiters, based on the flux vector split with the smooth eigenvalue split, has been used to capture internal shocks and other discontinuities. Numerical calculations have been done to analyze the off-design jet flow due to the pressure difference. The variation of pressure along the flow axis is compared with an experimental result and other numerical result. The characteristics of the interaction between the shock cell and the turbulence mixing layer have been analyzed.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.61-66
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• 2005

A theoretical analysis is developed for an annular-injection-type supersonic ejector having a second-throat downstream the ejector under the assumption that the Fabri choking is placed in mixing chamber. Non mixing theory is applied to formulate secondary flow pressure in the region between inlet of the mixing chamber and Fabri choking. To describe the shock standing at the inlet of the mixing chamber, two dimensional oblique shock relations are used and it is assumed that the shock affects only primary flow at Fabri choking plane. Physical constraint, which is that primary flow pressure and secondary flow pressure are same at Fabri choking plane, is added. In conclusion, it agrees well with experiments in case of small contracting angle of mixing chamber, under 4degrees.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.4
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• pp.281-286
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• 2007

The Navier-Stokes equations are numerically solved for a two-dimensional small nozzle with the area ratio of 1.8 between the throat and the exit. The shock structures are verified inside the nozzle and near the exit varying with the pressure ratio and the length of the diverging part, respectively. Especially the irregular patterns in the pressure distribution near the throat are analyzed based on the geometric characteristics. It is found that there are similar phenomena in the shock wave structure between the pressure ratio and the length changes. Also there exists a normal shock just between two different oblique shocks crossing each other in special cases.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.10
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• pp.934-939
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• 2007

Thrust vector control(TVC) is the method which generates the side force and moment by controlling the exhausting gas directly from the supersonic nozzle to change the trajectory of a missile quickly. In this paper, performance study on the tapered ramp tabs asymmetrically installed in the supersonic nozzle exhaust for the thurst vector control has been carried out using the supersonic cold flow system. To study the shock wave structure and location of the oblique shock wave produced by the ramp tab, the flow field visualization using the schlieren system is conducted. This paper provides the thrust spoilage, three directional forces and moments and distribution of surface pressure on the region enclosed by the tapered ramp tabs.

• Journal of Mechanical Science and Technology
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• v.17 no.12
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• pp.2107-2115
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• 2003

A supersonic dual coaxial jet has been employed popularly for various industrial purposes, such as gasdynamic laser, supersonic ejector, noise control and enhancement of mixing. Detailed characteristics of supersonic dual coaxial jets issuing from an inner supersonic nozzle and outer sonic nozzles with various ejection angles are experimentally investigated. Three important parameters, such as pressure ratios of the inner and outer nozzles, and outer nozzle ejection angle, are chosen for a better understanding of jet structures in the present study. The results obtained from the present experimental study show that the Mach disk diameter becomes smaller, and the Mach disk moves toward the nozzle exit, and the length of the first shock cell decreases with the pressure ratio of the outer nozzle. It was also found that the highly underexpanded outer jet produces a new oblique shock wave, which makes jet structure much more complicated. On the other hand the outer jet ejection angle affects the structure of the inner jet structure less than the pressure ratio of the outer nozzle, relatively.

• Journal of computational fluids engineering
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• v.18 no.3
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• pp.42-50
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• 2013

The supersonic airplane with flapped biplane, Busemann biplane equipped flap, is superior to drag and noise reduction due to wave cancelation effect between upper and lower airfoils. In this study, it is numerically calculated and analyzed the lift, drag and lift to drag ratio of flapped biplane with respect to various the length and angle of the flap. Euler solver of EDISON CFD, web based computational fluid dynamic solver for the purpose of education, is employed. Depending on the length of the flap, lift and drag increase linearly, and there exists the optimum flap angle which maximize the lift-to-drag ratio at the freestream mach 2.0 on-design condition. The predictable relational expression is driven as liner equation. As a results of comparison with drag of flapped biplane, Busemann biplane, and diamond airfoil with the same lift, the drag of flapped biplane is 88.76% lower than that of the Busemann biplane and 70.67% lower than that of the diamond airfoil. In addition, the change of pressure is compared to confirm the noise reduction effect of flapped biplane at h/c=5 of lower airfoil. The shock strength of flapped biplane is smaller than that of other airfoils.