• Journal of the Korean Society of Propulsion Engineers
• /
• v.11 no.4
• /
• pp.59-66
• /
• 2007

In this paper, experimental investigation results of the effect of partial admission ratio on the performance of axial turbine was presented. A supersonic impulse turbine of gas generator cycle liquid rocket engine turbopump was used for the test. for experimental purpose, a nozzle block, in which total 14 number of axi-symmetric convergent-divergent nozzles are arranged circumferentially, was designed and manufactured. Partial admission ratio was controlled by changing the number of active nozzles. High pressure air was used as working medium for the test. The experimental result revealed that the performance of the supersonic impulse turbine does not much affected by the partial admission ratio for supersonic impulse turbine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.43-48
• /
• 2010

Present study was conducted numerical analysis for velocity compound supersonic impulse turbine with the rotor overlaps, and the performance characteristics were analyzed through the numerical results. Tip overlap was more effective than hub overlap through the analysis. a case, overlap applied the hub and tip of the rotor, has the largest improvement for the turbine performance in parametric study cases. In case of overlap for the 2nd stage rotor, however, improvement of the turbine performance was not visibly large. Because, power generated in the 2nd stage was 22~23% of whole generated turbine power.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.11a
• /
• pp.145-148
• /
• 2008

Korea Aeronautics Research Institute (KARI) is developing a turbo pump that has 1-stage impulse turbine and relatively high tip clearance for safety. The objective of this research is to investigate the effect of reaction on tip clearance loss in axial turbines. Both cascades were tested in a subsonic wind tunnel. In each cascade, total pressure was measured for tip clearance ranging from 1% to 20% of chord. In results, increasing tip clearance, total pressure loss in reaction turbines is continually increased but impulse turbines keep almost same level of mass averaged total pressure loss. When tip clearance becomes more than 10% of chord, mass-averaged total pressure loss in impulse turbines is less than in reaction. This means that when tip clearance is more than 10% of chord, impulse turbines have better efficiency than reaction turbines.

• Aerospace Engineering and Technology
• /
• v.9 no.2
• /
• pp.185-192
• /
• 2010

Design of velocity-compounded turbine for 75ton class LRE turbopump application and pressure compounded turbine for 30ton class LRE turbopump has been performed. 1D calculation and CFD analysis were conducted in determining blade and flow passage shape of velocity compounded turbine iteratively. Finally, 23.1% improved specific power and 5% reduced weight turbine to the original design was developed. In case of pressure-compounded supersonic turbine design, rotational speed was increased by 50% and the effect of carryover ratio, 2nd nozzle installation angle, leakage flow of 2nd nozzle, and work sharing factor was studied. Final 1D design resulted 36% increased specific power and 51% reduced weight comparing to the original single-row impulse turbine. It is anticipated that nozzle flow path design will be very important for the accomplishment of expected performance of pressure-compounded turbine and nozzle shape optimization will be conducted through the CFD analysis.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.8
• /
• pp.99-106
• /
• 2003

Geometrical design variables are numerically investigated to improve aerodynamic performance of the supersonic impulse turbine of a turbopump in a liquid rocket engine. Aerodynamic redesign was performed for maximization of the blade power. Four design variables considered are blade angle, blade thickness and radii of upper and lower arc blade with appropriate constraints. A fast Navier-Stokes solver was developed and Chien's k-$\varepsilon$ turbulence modelling was used for turbulence closure. In initial shape, a flow separation was found in the middle of blade chord. However, it disappeared in final shape via its geometrical design variable change. About 3.2 percent of blade power was increased from this research.

• Journal of Aerospace System Engineering
• /
• v.11 no.1
• /
• pp.8-13
• /
• 2017

This study was performed to evaluate the turbine performance of a turbopump in the third stage engine of the Korea Space Launch Vehicle-II. The turbine is a supersonic impulse type with a single rotor. One nozzle is for starting and four remaining nozzles are for steady operation. A similarity test was carried out in the high air test facilities at the Korea Aerospace Research Institute. Test results showed that turbine efficiency changed much more from rotational speed variations than by pressure ratio variations. These results showed characteristics similar to other supersonic impulse turbines.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.35 no.2
• /
• pp.143-148
• /
• 2007

An impulse turbine, which is a main component of a liquid rocket engine, needs to be a small size with light weight and generate large power. Since the impulse turbine is being operated under complicated supersonic conditions, flow analysis and performance prediction largely depend on CFD technique. In order to increase the reliability of the prediction code, however, it often requires an experimental data to compare. In this research a rotating turbine rotor with multiple blades is simulated with a two-dimensional stationary cascade to check the effect of major flow parameters. Mach number is measured at nozzle exit by using a pitot tube and the blade thrust was also measured with a load cell. The measured thrust coefficient and the power are compared well with the designed conditions, which proves the design procedures are properly taken.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.15 no.1
• /
• pp.19-28
• /
• 2011

As a preview study, present research analysed the performance characteristics of a velocity compound supersonic impulse turbine with the rotor overlaps before adapting the overlap has the best turbine performance. This research was conducted for the turbine with square cross-section nozzles instead of axisymmetric nozzles and wrap around nozzles. Through 3-dimensional flow analysis for the turbine by a commercial flow analysis package, tip overlap case was more effective to improve the turbine performance than case hub overlap, and overlap case applied the hub and tip of the rotor had the largest improvement for the turbine performance in the cases. In case of overlap for the 2nd stage rotor, improvement of the turbine performance was not visibly large. Because, generated power in the 2nd stage is 22~23% of whole generated turbine power.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.561-565
• /
• 2009

The performance of supersonic impulse turbine was investigated experimentally. Experiment was performed with the compressed air instead of the high temperature burned gas because of the limitation of test facility and danger. As a result of the experiment with the compressed air, the performance in the real gas(burned gas) was predicted by the similarity method. The nozzle area of prototype turbine was calculated based on the real gas. So, it is difficult to satisfy the similarity conditions completely. Two similarity conditions were set and the design point for real gas was existed between two similarity conditions. And, the new turbine test model with calculated nozzle area based on the compressed air was tested. Therefore, similarity point of the new turbine test model was also existed between above two similarity points. It means that the design point for real gas was similar to the test point with the new turbine model.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.11a
• /
• pp.117-121
• /
• 2008

The effect of tip clearance on the performance of a supersonic impulse turbine was investigated experimentally. Test was performed using high pressure air in wide ranges of pressure ratio and rotational speeds. Test revealed that efficiency gradient of the subject turbine at a reference test point was a very low value of 0.05. Turbine efficiency was varied non-linearly with respect to tip clearance. It has been found that efficiency gradient is proportional to the cube of rotational speed at a fixed turbine pressure ratio. It also has been found that efficiency gradient shows its minimum at a reference test pressure ratio irrespective of rotational speeds.