• 한국전산유체공학회:학술대회논문집
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• 2003.08a
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• pp.182-187
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• 2003

Numerical analysis of three-dimensional vicous flow is used to compute the design speed operating line of a transonic axial-flow compressor. The Navier-Stokes equation was solved by an explicit finite-difference numerical scheme and the Baldwin-Lomax turbulence model was applied. A spatially-varying time-step and an implicit residual smoothing were used to improve convergence. Two-stage axial compressor of a turboshaft engine developed KARI was chosen for the analysis. Numerical results show reasonably good agreements with experimental measurements made by KARI. Numerical solutions indicate that there exist a strong shock-boundary layer interaction and a subsequent large flow separation. It is also observed that the shock is moved ahead of the blade passage at near-stall condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.66-71
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• 2008

Two cases with circumferential grooves were designed for a transonic compressor, and 3-D numerical simulations were conducted for stall mechanism at three representative speeds. A conclusion can be drawn from the comparison between compressors with or without casing treatment that: with the rising of rotation speed, stall margin increases dramatically under the help of casing treatments, and the case with middle grooves has reasonable compromise between stall margin increment and efficiency cutting. At lower speed, the increment reduces, and grooves at the back of blade tip have more influence on stall margin. Further investigation shows there is a transition in mechanism of compressor stall with the decline of rotational speed: at high rotation speed, the expansion of stall margin mainly results from the suppression of tip leakage vortex by casing treatments, yet it benefits more from the depression of boundary layer separation from suction surface of blade tip.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.41 no.1
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• pp.9-19
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• 2017

In this study, stator shroud injection in a single-stage transonic axial compressor is proposed. A parametric study of the effect of stator shroud injection on aerodynamic performances was conducted using the three-dimensional Reynolds-averaged Navier-Stokes equations. The curvature, length, width, and circumferential angle of the stator shroud injector and the air injection mass flow rate were selected as the test parameters. The results of the parametric study show that the aerodynamic performances of the single-stage transonic axial compressor were improved by stator shroud injection. The aerodynamic performances were the most sensitive to the injection mass flow rate. Further, the total pressure ratio and adiabatic efficiency were the maximum when the ratio of circumferential angle was 10％.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.8-15
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• 2001

In the continuing quest for increased turbomachinery efficiency, the part played by blade profile shape remains crucial. The application of a heated thin metallic film with CTA(constant temperature anemometer) to the measurements of the laminar and turbulent boundary layer behavior(shock-boundary layer-interaction) in a transonic wind tunnel. Results of measurements with hot-film sensors on transonic compressor blades are extremely difficult to interpret because of ambiguous probe signals due to the complexity of the local flow pattern. In order to get the explicit information and give the designer to interpret characteristic signals from hot-film probes, a method was developed by comparing the results with other measuring technic results.

• International Journal of Aeronautical and Space Sciences
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• v.16 no.2
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• pp.157-164
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• 2015

The present work performs three-dimensional flow calculations based on Reynolds Averaged Navier-Stokes (RANS) and Delayed Detached Eddy Simulation (DDES) to investigate the flow field of a transonic rotor (NASA Rotor 37) at near-stall condition. It is found that the DES approach is likely to predict well the complex flow characteristics such as secondary vortex or turbulent flow phenomenon than RANS approach, which is useful to describe the flow mechanism of a transonic compressor. Especially, the DES results show improvement of predicting the flow field in the wake region and the model captures reasonably well separated regions compared to the RANS model. Besides, it is discovered that the three-dimensional vortical flows after the vortex breakdown from the rotor tip region are widely distributed and its vortex structures are clearly present. Near the rotor leading edge, a part of the tip leakage flows in DES solution spill over into next passage of the blade owing to the separation vortex flow and the backflow is clearly seen around the trailing edge of rotor tip. Furthermore, the DES solution shows strong turbulent eddies especially in the rotor hub, rotor tip section and the downstream of rotor trailing edge compared to the RANS solution.

• 유체기계공업학회:학술대회논문집
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• 1998.12a
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• pp.217-222
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• 1998

CSCM upwind flux difference splitting compressible Navier-Stokes method has been used to predict the transonic flows in centrifugal compressor diffuser. The modified cyclic TDMA and the mass flux boundary conditions were used as boundary conditions of the diffuser analysis. With the mass flux boundary condition and the $130{\times}80{\times}40$ grid, the compressible upwind Navier-Stokes method predicted the transonic diffuser flowfield successfully. Plow changes in the impeller exit region due to the strong interaction between impeller exit and vaned diffuser, broad flow separation on the suction surface near hub and shroud was observed from the results of the mass flow rates 6.0 and 6.2kg/s at 27000 rpm. The static pressure increased and the total pressure decreased through the flow passage of the channel diffuser, which were predicted better from the three-dimensional analysis than from the two-dimensional analysis due to the strong effect of the three-dimensional flow. The mass averaged loss coefficients and pressure coefficients were also studied.

• The KSFM Journal of Fluid Machinery
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• v.18 no.5
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• pp.5-12
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• 2015

This paper presents a numerical investigation of the effect of the dihedral stator on the loss in a transonic axial compressor. Four stator geometries with different stacking line variables are tested in the flow simulations over the whole operating range. It is found that a large shroud loss at the rotor outlet and the subsequent shroud corner separation in the stator passage occur at low mass flow rate. The hub dihedral stator and bowed blade generate unexpected hub-corner-separation, thereby causing a large total pressure loss over the entire operating range. However, the corresponding blockage forces the high momentum flow near the hub to divert toward the upper part of the passage suppressing the negative axial velocity region. The dihedral stator increases deflection angle and secondary vorticity near the endwall where the dihedral is applied. As a result, the endwall loss which is related to the endwall relative velocity decreases.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.10
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• pp.1384-1391
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• 2001

Three-dimensional flow analysis is implemented to investigate the flow through transonic axial-flow compressor rotor(NASA R67) and to evaluate the performances of Abid's low-Reynolds-number k-$\omega$ and Baldwin-Lomax turbulence models. A finite volume method is used fur spatial discretization. The equations are solved implicitly in time by the use of approximate factorization. The upwind difference scheme is used for inviscid terms and viscous terms are approximated with central difference. The flux-difference-splitting method of Roe is used to obtain fluxes at the cell faces. Numerical analysis is performed near peak efficiency and near stall. The results are compared with the experimental data for NASA R67 rotor. Blade-to-Blade Mach number distributions are compared to confirm the accuracy of the code. From the results, it is concluded that Abid'k-$\omega$ model is better for the calculation of flow rate and efficiency than Baldwin-Lomax model. But, the predictions for Mach number and shock structure are almost the same.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.202-212
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• 2008

A three-dimensional computation was conducted to understand effects of the low Reynolds number on the loss characteristics in a transonic axial compressor, Rotor67. As a gas turbine becomes smaller in size and it is operated at high altitude, the operating condition frequently lies at low Reynolds number. It is generally known that wall boundary layers are thickened and a large separation occurs on the blade surface in axial turbomachinery as the Reynolds number decreases. In this study, it was found that the large viscosity did not affect on the bow shock at the leading edge but significantly did on the location and the intensity of the passage shock. The passage shock moved upstream towards leading edge and its intensity decreased at the low Reynolds number. This change had large effects on the performance as well as the internal flows such as the pressure distribution on the blade surface, tip leakage flow and separation. The total pressure rise and the adiabatic efficiency decreased about 3% individually at the same normalized mass flow rate at the low Reynolds number. In order to analyze this performance drop caused by the low Reynolds number, the total pressure loss was scrutinized through major loss categories such as profile loss, tip leakage loss, endwall loss and shock loss.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.30 no.7 s.250
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• pp.622-629
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• 2006

The shape optimization of blade sweep in a transonic axial compressor rotor of NASA Rotor 37 has been performed using response surface method and the three-dimensional Wavier-Stokes analysis. Two shape variables of the rotor blade, which are used to define the rotor sweep, are introduced to increase the adiabatic efficiency of the compressor. Throughout the optimization, optimal shape having a backward sweep is obtained. Adiabatic efficiency, which is the objective function of the present optimization, is successfully increased. Separation line due to the interference between a shock and surface boundary layer on the blade suction surface is moved downstream for the optimized blade compared to the reference one. The increase in adiabatic efficiency for the optimized blade is caused by suppression of the separation due to a shock on the blade suction surface.