• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.25-31
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• 2013

A blisk with rotor shroud is usually adopted in LRE turbine to maximize its performance. However it experiences the severe thermal load and resulting damage during engine stating and stop. Shroud splitting is devised to relieve the thermal stress on the turbine rotor. Structural analysis confirmed the reduction of plastic strain at the blade hub and tip. However, split gap at the rotor shroud entails additional tip leakage and results performance degradation. In order to assess the effect of shroud split on the turbine performance, tests have been performed for various settings of shroud split. For the maximum number of shroud splitting, measured efficiency reduction ratio was 2.65% to the value of original shape rotor.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.29-38
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• 2015

In this study, 3-dimensional URANS simulation was performed to analyze the effect of the nozzle-rotor axial gap spacing of a supersonic impulse turbine on turbine performance. The computations were conducted for four different axial gap cases corresponding to about 6%, 10%, 20% and 30% of the blade height, respectively. The results show a good agreement with previous studies and the turbine efficiency decreases drastically in certain range. It is examined that the turbine performance characteristics could change depending on the influence of leading edge shock to the nozzle outlet. It is also found that the entropy rise distributions along the span differ from each other.

• Journal of computational fluids engineering
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• v.10 no.2
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• pp.54-59
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• 2005

For the improvement of aerodynamic performance of the turbine blade in a turbopump for the liquid rocket engine, the optimization of turbine profile shape has been studied. The turbine in a turbopump in this study is a partial admission of impulse type, which has twelve nozzles and supersonic inflow. Due to the separated nozzles and supersonic expansion, the flow field becomes complicate and shows oblique shocks and flow separation. To increase the blade power, redesign ol the blade shape using CFD and optimization methods was attempted. The turbine cascade shape was represented by four design parameters. For optimization, a genetic algorithm based upon non-gradient search hue been selected as an optimizer. As a result, the final blade has about 4 percent more blade power than the initial shape.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.38-44
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• 2011

The performance experiment was tested for the velocity compound turbine of turbopump which was the main part of 75 ton class liquid rocket engine. The seal is installed between the 1st rotor and the reversing vane to reduce the leakage flow. The turbine outlet pressure of the velocity compound turbine by changing the rotating speed was compared with that of baseline turbine with single rotor including the effect on the total performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1038-1044
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• 2017

We are working on improving the performance by applying the three-dimensional design element to the rotor blades of high pressure supersonic impulse turbine that drives turbo pump of liquid rocket engine. In this paper, based on the shape of a rotor blade of a turbopump turbine designed in the past, a three-dimensional shape is applied to a rotor blade through a stacking line change such as sweep and dihedral. After performing the flow analysis, the changes in the turbine performance characteristics for each design element were carefully examined and the results were summarized.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.325-330
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• 2011

A rotor overlap technique was adapted to improve the performance of a axial turbine. The technique secured sufficient flow passage by additional height at the rotor tip and hub. especially in a supersonic turbine, the technique reduced the chance of chocking in the rotor passage, and made to be satisfied the design pressure ratio. However, the technique also made additional losses, like a pumping loss, expansion loss, etc. Therefore, a optimization technique was appled to maximize the improvement of the turbine performance. An approximate optimization method was used for the investigation to secure the computational efficiency. The design variables was shape factors of a rotor overlap. Results indicated that a significant improvement in turbine performance can be achieved through the optimization of the rotor overlap.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.2
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• pp.78-86
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• 2006

This paper presents the performance test result of the 30-ton class turbopump turbine. Test has been conducted using high pressure cold air, The turbine overall performance has been measured for various pressure ratio and rotational speed settings. The nozzle-rotor clearance effect on turbine performance also has been tested for the four kinds of the nozzle-rotor clearance values. We found that turbine efficiency rated 51.1% at its design velocity ratio and pressure ratio of 13.5. We also found that turbine efficiency can be increased by 3.5% for approximately 1mm decrement of the nozzle-rotor clearance from its nominal value.

• The KSFM Journal of Fluid Machinery
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• v.10 no.1 s.40
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• pp.26-33
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• 2007

Four design candidates for a partial admission turbine have been chosen from a preliminary design process. Their performance were estimated through the 3-D numerical analyses using a frozen rotor method. In order to select the optimum design, each flow analysis result was compared with others. Flow characteristics in the passages and some types of losses induced by shocks and wakes were found from calculation results. Based on these calculations, a new rotor blade was redesigned and compared with previous one through flow analysis.

• 유체기계공업학회:학술대회논문집
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• 2004.12a
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• pp.445-451
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• 2004

Four design candidates for a 1.4MW class partial admission turbine have been chosen from a Preliminary design process. Their performance were estimated through the 3-D numerical analyses using a frozen rotor method. In order to select the optimum design, each flow analysis result was compared with others. Flow characteristics in the passages and some types of losses induced by shocks and wakes were found from calculation results. Based on these calculations, a new rotor blade was redesigned and compared with previous one through flow analysis.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.4
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• pp.66-72
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• 2005

In this paper, a small supersonic wind tunnel is designed and built to study the flow characteristics of a supersonic impulse turbine cascade. The flow is visualized by means of a single pass Schlieren system. The supersonic cascade with 2-dimensional supersonic nozzle was tested for various blade leading edge shapes and gaps between the nozzle and cascade. Highly complicated flow patterns including shocks, nozzle-cascade interaction and shock boundary layer interactions are observed.