• Journal of the Korea Institute of Military Science and Technology
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• v.11 no.2
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• pp.152-160
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• 2008

In the present paper, a transonic mixed-flow compressor that has relatively lower frontal area than that of centrifugal compressors is discussed, and aerodynamic design as well as performance prediction are performed. Main design constraints are compressor exit Mach number of 0.3 and flow angle of 30degrees at the design point, and maximum overall compressor diameter of 177mm, that is 7.0inch. The mass flow rate of design point and pressure ratio are 1.05kg/s and 5.2:1, respectively. The aerodynamic design results show that the transonic compressor designed with forward-swept inducer and curved diffuser can have the target performance with efficiency of 75% within the given constraints. And the compressor exit flow characteristics are discussed here.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.4
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• pp.56-62
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• 2008

This study presents the aerodynamic design and numerical analysis results on a transonic centrifugal compressor which is used for gas turbine systems. Mean-line analysis and quasi-3D analysis are used for the aerodynamic design, and Reynolds-averaged Navier-Stokes analysis is applied to flow analysis of the compressor. The aerodynamic parameters for a transonic compressor, such as pressure coefficient, swirl parameter, blade loading, are discussed, and flow characteristics in the impeller and diffuser are discussed.

• 유체기계공업학회:학술대회논문집
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• 2003.12a
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• pp.639-645
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• 2003

Design optimization of a transonic compressor rotor (NASA rotor 37) using response surface method and three-dimensional Navier-Stokes analysis has been carried out in this work. Baldwin-Lomax turbulence model was used in the flow analysis. Two design variables were selected to optimize the stacking line of the blade, and mass flow was used as a design variable, as well, to obtain new design point at peak efficiency. Data points for response evaluations were selected by D-optimal design, and linear programming method was used for the optimization on the response surface. As a main result of the optimization, adiabatic efficiency was successfully improved, and new design mass flow that is appropriate to an improved blade was obtained. Also, it is found that the design process provides reliable design of a turbomachinery blade with reasonable computing time.

• The KSFM Journal of Fluid Machinery
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• v.15 no.6
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• pp.77-84
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• 2012

The aerodynamic performance of a single-stage transonic axial compressor was experimentally evaluated by measuring pressure and temperature distribution at the inlet and outlet of the compressor. The compressor was developed by Korea Aerospace Research Institute through multidisciplinary design optimization (MDO) method, especially integrating aerodynamic performance and structural stability. The test results show that the pressure ratio is 1.65 and the efficiency is 85.8 % at design point, where the corrected speed is 22,000 rpm and the corrected mass flow rate is 15.4 kg/s, and it has a good agreement with the design target and computational results. The distribution of pressure ratio is very steep at design speed, compared with the trend of other subsonic compressors. Also the static pressure distribution on the stator casing shows that the blade loading is gradually increasing through the stage as designed.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.32 no.4
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• pp.317-326
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• 2008

The main purpose of the present study is to perform shape optimizations of transonic compressor blade in order to enhance its performance. In this study, the Latin hypercube sampling of design of experiments and the weighted average surrogate model with the help of a gradient based optimization algorithm are used within design space by the lower and upper limits of each design variable and for finding optimum designs, respectively. 3-D Reynolds-averaged Navier-Stokes solver is used to evaluate the objective functions of adiabatic efficiency and pressure ratio. Six variables from lean and airfoil thickness profile are selected as design variables. The results show that the adiabatic efficiency is enhanced by 1.43% by efficiency optimization while the pressure ratio is increased very small, and pressure ratio is increased by 0.24% by pressure ratio optimization.

• The KSFM Journal of Fluid Machinery
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• v.14 no.6
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• pp.96-101
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• 2011

Korea Aerospace Research Institute is performing 3 stage transonic axial compressor development program. This paper introduces design step of the compressor, the performance test results and its analysis. In the fore part of the paper, aerodynamic process of the 3 stage axial compressor is presented. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3, while the maximum relative Mach number in the rotor stage is between 1.3~1.4 to increase the compressor flow coefficient. The final design has been confirmed by iterating three dimensional CFD calculations to verify design target and some design intentions. In the latter part of the paper, its performance test processes and results are presented. The performance test result shows that the overall compressor performance targets; pressure ratio and efficiency are well achieved. The stator static pressure distributions show that the blade loading is gradually increasing from the downstream of the compressor.

• Proceedings of the KSME Conference
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• 2005.11a
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• pp.74.2-74.2
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• 2005

• The KSFM Journal of Fluid Machinery
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• v.15 no.2
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• pp.12-19
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• 2012

This study presents unsteady and unstable characteristics of three stage transonic axial compressor, developed by Korea Aerospace Research Institute. As approaching to the unstable operating region at the 103% design speed of the compressor, a modal type stall precursor appears in front of highly loaded 3rd rotor row at first, and it propagates to the upstream. On the contrary, actual stall cell initiates from the stall precursor in front of the 1st rotor row, and it propagates to the downstream of the compressor. After the stall region reached the 3rd stage and stall cell rotates circumferentially about 360 deg, it develops to one dimensional compressor surge mode. It shows a mild surge behaviour with 3~4 Hz frequency. From the test data, it can be suggested that there is a priority to give an optimum blade loading distributions to construct a multi stage transonic axial compressor stages either to secure more stable compressor operating ranges, or to maximize the compressor efficiency.

• The KSFM Journal of Fluid Machinery
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• v.9 no.3 s.36
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• pp.7-13
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• 2006

Shape optimization of a rotor blade in a single-stage transonic axial compressor has been performed using a response surface method and three-dimensional Navier-Stokes analysis. Two shape variables of the rotor blade, which are used to define a blade skew, are introduced to increase an adiabatic efficiency. Throughout the shape optimization of a rotor blade, the adiabatic efficiency is increased to about 2.2 percent compared to that of the reference shape of the stator. The increase in efficiency for the optimal shape of the rotor is due to the pressure enhancement, which is mainly caused by moving the separation position on the suction surface of rotor blade to the downstream direction.

• 유체기계공업학회:학술대회논문집
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• 2002.12a
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• pp.244-249
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• 2002

It is known that tip clearance flows reduce the pressure rin, flow range and efficiency of the turbomachinery. So, the clear understanding about flow fields in the tip region is needed to efficiently design the turbomachinery. The Navier-Stokes code with the proper treatment of the boundary conditions has been developed to analyze the three-dimensional steady viscous flow fields in the transonic rotating blades and a numerical study has been conducted to investigate the detail flow physics in the tip region of transonic rotor, NASA Rotor 67. The computational results in the tip region of transonic rotors show the leakage vortices, leakage flow from pressure side to suction side and their interaction with a shock Depending on the operating conditions, the position of shock-wave on the blade surface are v8y different close to the blade tip of the transonic compressor rotor. The shock-wave position dose to the blade tip had the dose relationship with the starting position of leakage vortex and the direction of leakage flow.