• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.3
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• pp.33-40
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• 2001

In order to investigate the operating characteristics of a supersonic steam ejector, the axisymmetric, compressible, Reynolds-averaged, Wavier-Stokes computations are performed using a finite volume method. The secondary and back pressures of the ejector system with a second throat are changed to investigate their effects on the suction mass flow. Three operation modes of the steam ejector system, the critical mode, subcritical mode and back flow mode, are discussed to predict the critical suction mass flow. The present computations are validated with some experimental results. The secondary and back pressures of the supersonic steam ejector significantly affect the critical suction mass flow. The present computations predict the experimented critical mass flow with fairly good accuracy A good correlation is obtained for the critical suction mass flow. The present results show that provided the primary nozzle configuration and secondary pressure are blown, we can predict the critical mass flow with good accuracy.

• Proceedings of the KSME Conference
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• 2001.06e
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• pp.349-354
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• 2001

The present study addresses the open end correction associated with the reflection and discharge phenomena of a weak shock wave from an open end of a duct. The open end correction of the weak shock wave is investigated experimentally and by numerical computation. An experiment is made using a simple shock tube with an open end, and computation is performed to simulate the experimental flow field using the unsteady, axisymmetric, compressible, flow governing equations. The results obtained show that an open end correction should be involved for shock wave discharge and reflection problems generated from the exit of the duct with an open end baffle plate. With a baffle plate less than three times the duct diameter, it is found that the open end correction is a function of both the diameter of the baffle plate and normal shock wave magnitude. However, for a baffle plate larger than three times the duct diameter, it is independent of the baffle plate diameter. The present computations predict the results of shock tube experiment with good accuracy. A new empirical equation for prediction of the open end correction is found for the weak shock reflection and discharge phenomena occurring at the open end of the duct with and without a baffle plate.

• Transactions of the Korean Society for Noise and Vibration Engineering
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• v.11 no.7
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• pp.239-246
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• 2001

The present study addresses the open correction associated with the reflection and discharge phenomena of a weak shock wave from an open end of a duct. The open correction of the weak shock wave is investigated experimentally and by numerical computation. An experiment is made using a simple shock tube with an open end. and computaion is performed to simulate the experimental flow field using the unsteady, axisymmetric compressible. flow governing equations. The results obtained show the an open correction should be involved for shock wave discharge and reflection problems generated from the exit of the duct with an open baffle plate. With a baffle plate less than three times the duct diameter, it is found that the open end correction is a function of both the diameter of the baffle plate and normal shock wave magnitude However, for a baffle plate larger than three the duct diameter it is independent of the baffle plate diametre, The present computations predict the results of shock tube experiment with good accuracy. A new empirical equation for prediction of the open correction is found for the weak shock reflection and discharge phenomena occurring at the open of the duct with and without a baffle plate.

• Journal of computational fluids engineering
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• v.2 no.1
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• pp.66-72
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• 1997

Optimum nozzle design exploiting the method of characteristic(M.O.C) has been in application as an efficient design methodology targeting a less weighted and short expansion nozzle. This paper treats the optimum nozzle design and the analysis of the inviscid compressible flow inside. Based on traditional Rao's method, the optimum nozzle design is coded with minor modifications for the identification of the control surface across which the mass flux should be conserved. Internal flow field is simulated numerically by M.O.C and implicit/explicit Taylor-Galerkin finite element method(F.E.M) with the aid of adaptive remeshing to capture the shock wave, hence improve the accuracy. Designed and calculated flow fields due to the separate analyses show that the mass flux predicted by optimum nozzle design with M.O.C is not conserved across the control surface and the sonic line should be located upstream of the nozzle throat. Rao's optimum nozzle design methodology exaggerates the momentum thrust and tends to overemphasize the engine performance loss.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.5
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• pp.1-9
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• 2003

Detailed flow of a shrouded tail rotor in hover is studied by using a compressible inviscid flow solver on unstructured meshes. The numerical method is based on a cell-centered finite-volume discretization and an implicit Gauss-Seidel time integration. Numerical simulation is made for a single blade attached to the center body and guide by the duct by imposing a periodic boundary condition between adjacent rotor blades. The results show that the performance of an isolated rotor without shroud compares well with experiment. In case of a shrouded rotor, correction of the collective pitch angle is made such that the overall performance matches with experiment to account for the uncertainties of the experimental model configuration. Details of the flow field compare well with the experiment confirming the validity of the present method.

• Journal of the Korean Society of Visualization
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• v.9 no.2
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• pp.61-65
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• 2011

Several previous works on rocket nozzle flows have revealed the existence of the transition from FSS to RSS and the occurrence of asymmetric flow associated with the boundary layer separation, which can cause excessive side-loads of the propulsion system. Thus, it is of practical importance to investigate the asymmetric flow behaviors of the propulsion nozzle and to develop its control method. In the present study, the asymmetric flow control method using a cavity system was applied to supersonic nozzle flow. Time-dependent asymmetric flow was experimentally investigated with the rate of change of the nozzle pressure ratio. The results obtained showed that the cavity system installed on nozzle wall would be helpful in fixing the unsteady motions of the boundary layer separation, consequently reducing the possibility of the occurrence of the asymmetric flow.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.51-58
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• 2001

Supersonic, axisymmetric, jets issuing from several kinds of dual, coaxial, nozzles were experimentally investigated. Four different kinds of coaxial, dual nozzles were employed to characterize the major. features of the supersonic, coaxial, dual jets. Two convergent-divergent supersonic nozzles with different impinging angle on the jet axis of were designed to have the Mach number 2.0 and used to compare the coaxial jet flows with those discharging from two sonic nozzles. The primary pressure ratio was changed in the range from 4.0 to 10.0 and the assistant jet ratio from 1.0 to 4.0. The results obtained show that the assistant jets from the annular. nozzle affect the coaxial jet flows and an increase of both the primary jet pressure ratio and assistant jet pressure ratio lead to a longer supersonic length of the dual, coaxial jet.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.35-45
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• 2005

Numerical study is carried out to investigate the effects of chemistry and thermal radiation on the rocket plume flow field at various altitudes. Navier-Stokes equations for compressible flows were solved by a fully-implicit TVD code based on the finite volume method. An infinitely fast chemistry module for hydrocarbon mixture with detailed thermo-chemical properties and a thermal radiation module for optically thick media were incorporated with the fluid dynamics code. The plume flow fields of a kerosene-fueled rocket flying at Mach number zero at sea-level, 1.16 at altitude of 5.06 km and 2.90 at 17.34 km were numerically analyzed. Results showed the plume structures at different altitude conditions with the effects of chemistry and radiation. It is understood that the excess temperature by the chemical reactions in the exhaust gas may not be ignored in the view point of propulsion performance and thermal protection of the rocket base, especially at higher altitude conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.3
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• pp.11-18
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• 2000

When supersonic underexpanded jets are exhausted from the nozzle, complex shock cell configurations such as barrel shock, expansion fan, Mach disc, and exhaust-gas jet boundary are appeared repetitively. The shock cell is smeared by turbulence dissipation and disappeared in long distance from the nozzle. When underexpanded jet is suddenly impinged on a flat plate, it forms very complex flow structure. In this paper, we solve compressible Wavier-Stokes equation adapting finite volume method to obtain jet impingement flow structure and compare calculated data with experimental ones. It is shown that numerical simulation data are in good agreement with experimental one in a short distance between nozzle exit and flat plate and little influence of underexpanded ratio is appeared in jet impingement now distribution.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.6
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• pp.477-483
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• 2016

Velocity profiles of laminar, transition and turbulent boundary layers were investigated by using Particle Image Velocimetry(PIV) measurements on the flat plate at Mach 2.96. The Schlieren visualization and PIV measurements are also used to confirm whether the oblique shock wave generated from the leading edge affects the flow field over the flat plate. The laminar velocity profile measured from the experiment was well matched with the compressible Blasius solution. The velocity profile of the transition boundary layer was well correlated with the theoretical turbulent velocity profile from near the wall and the transition began from Re = $1.41{\times}106$. For the turbulent boundary layer, considering compressibility effects, the Van Driest-transformed velocity satisfies the incompressible log-law. It is found that the log region is extended farther in the wall-normal direction compared to the log region in incompressible boundary layer.