Flow behavior near trailing edge of transonic turbine cascade. 1st report Trailing edge normal to blade surface.

In order to clarify the transonic flow characteristics of a turbine cascade with high stagger, low solidity and small deflections, experimental studies were carried out by shortening the chord length of a “Laval-nozzle shaped” blade with thick trailing edge. The behavior of the shock system depends on the amount of overlap between the blades. The relations between the behavior and the performances were discussed in detail. The results may be applied to more standard sections. Lastly, validity of an appropriate time marching analysis for the highly staggered cascade was investigated by comparing with the experiment.

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Experimental Investigation of an Innovative Cooling System (ICS) for High Temperature Transonic Turbine Stages

Volume 5: Turbo Expo 2002, Parts A and B ◽

10.1115/gt2002-30341 ◽

2002 ◽

Cited By ~ 3

Author(s):

E. Go¨ttlich ◽

H. Lang ◽

W. Sanz ◽

J. Woisetschla¨ger

Keyword(s):

Cooling System ◽

Flow Behavior ◽

Leading Edge ◽

Optical Methods ◽

Blade Surface ◽

Camera System ◽

Flow Measurements ◽

Cooling Effectiveness ◽

Ir Camera ◽

Transonic Turbine

Gas turbine design technology requires the development of transonic turbine stages capable of carrying high stage load and of handling hot gas temperatures at turbine inlet. A reliable cooling system is necessary to cope with the shock system in the main flow field especially in the leading edge region of the rotor blades. These requirements are fulfilled by the Innovative Cooling System (ICS) developed at the Institute for Thermal Turbomachinery and Machine Dynamics, Graz University of Technology. The ICS is also able to cover large areas of the blade surface with an effective cooling film and to reduce the metal temperature. In this paper the authors present results on the flow measurements giving the aerodynamic behavior of these cooling films and on the investigation of their cooling effectiveness. The measurements were done on an industrial turbine blade in a linear cascade arrangement. In addition to conventional measurement methods optical methods (Schlieren visualization, Laser Doppler Velocimetry) were employed to investigate and visualize the transonic flow through the linear blade cascade. An infrared (IR) camera system was used to determine the effectiveness of this newly designed cooling system by measuring the temperature distribution on the blade surface. Experimental results concerning aerodynamic flow behavior and cooling effectiveness are presented.

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Aerodynamic Performance of a Transonic Turbine Guide Vane With Trailing Edge Coolant Ejection: Part I—Experimental Approach

Journal of Turbomachinery ◽

10.1115/1.2836698 ◽

1996 ◽

Vol 118 (3) ◽

pp. 519-528 ◽

Cited By ~ 21

Author(s):

C. Kapteijn ◽

J. Amecke ◽

V. Michelassi

Keyword(s):

Gas Turbines ◽

Experimental Approach ◽

Guide Vane ◽

Flow Behavior ◽

Aerodynamic Performance ◽

Inlet Guide Vane ◽

Trailing Edge ◽

Cooling Flow ◽

Turbine Guide Vane ◽

Transonic Turbine

Inlet guide vanes (IGV) of high-temperature gas turbines require an effective trailing edge cooling. But this cooling significantly influences the aerodynamic performance caused by the unavoidable thickening of the trailing edge and the interference of the cooling flow with the main flow. As part of a comprehensive research program, an inlet guide vane was designed and manufactured with two different trailing edge shapes. The results from the cascade tests show that the flow behavior upstream of the trailing edge remains unchanged. The homogeneous values downstream show higher turning and higher losses for the cut-back blade, especially in the supersonic range. Additional tests were conducted with carbon dioxide ejection, in order to analyze the mixing process downstream of the cascade.

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Simulation of Trailing Edge Vortex Shedding in a Transonic Turbine Cascade

Volume 1: Turbomachinery ◽

10.1115/96-gt-483 ◽

1996 ◽

Cited By ~ 1

Author(s):

Tom C. Currie ◽

William E. Carscallen

Keyword(s):

Vortex Shedding ◽

Navier Stokes ◽

Numerical Dissipation ◽

Turbine Cascade ◽

Trailing Edge ◽

Edge Vortex ◽

Transonic Turbine ◽

Exit Mach Number ◽

Good Agreement

Mid-span losses in the NRC transonic turbine cascade peak at an exit Mach number (M2) of ∼1.0 and then decrease by ∼40% as M2 is increased to the design value of 1.16. Since recent experimental results suggest that the decrease may be related to a reduction in the intensity of trailing edge vortex shedding, both steady and unsteady quasi-3D Navier-Stokes simulations have been performed with a highly refined (unstructured) grid to determine the role of shedding. Predicted shedding frequencies are in good agreement with experiment, indicating the blade boundary layers and trailing edge separated free shear layers have been modelled satisfactorily, but the agreement for base pressures is relatively poor, probably due largely to false entropy created downstream of the trailing edge by numerical dissipation. The results emphasize the importance of accounting for the effect of vortex shedding on base pressure and loss.

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Simulation of Trailing Edge Vortex Shedding in a Transonic Turbine Cascade

Journal of Turbomachinery ◽

10.1115/1.2841371 ◽

1998 ◽

Vol 120 (1) ◽

pp. 10-19 ◽

Cited By ~ 11

Author(s):

T. C. Currie ◽

W. E. Carscallen

Keyword(s):

Vortex Shedding ◽

Three Dimensional ◽

Navier Stokes ◽

Numerical Dissipation ◽

Turbine Cascade ◽

Trailing Edge ◽

Edge Vortex ◽

Transonic Turbine ◽

Exit Mach Number

Midspan losses in the NRC transonic turbine cascade peak at an exit Mach number (M2) of ~1.0 and then decrease by ~40 percent as M2 is increased to the design value of 1.16. Since recent experimental results suggest that the decrease may be related to a reduction in the intensity of trailing edge vortex shedding, both steady and unsteady quasi-three-dimensional Navier–Stokes simulations have been performed with a highly refined (unstructured) grid to determine the role of shedding. Predicted shedding frequencies are in good agreement with experiment, indicating the blade boundary layers and trailing edge separated free shear layers have been modeled satisfactorily, but the agreement for base pressures is relatively poor, probably due largely to false entropy created downstream of the trailing edge by numerical dissipation. The results nonetheless emphasize the importance of accounting for the effect of vortex shedding on base pressure and loss.

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Heat Transfer Investigation on a Transonic Turbine Cascade with Pulsating Trailing Edge Cooling

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit ◽

10.2514/6.2012-3880 ◽

2012 ◽

Cited By ~ 1

Author(s):

Bayindir Saracoglu ◽

Guillermo Paniagua

Keyword(s):

Heat Transfer ◽

Turbine Cascade ◽

Trailing Edge ◽

Transonic Turbine

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Unsteady Shock Waves Vortex Shedding Entanglement in a Transonic Turbine Cascade

10.21236/ada587666 ◽

2013 ◽

Cited By ~ 1

Author(s):

Guillermo Paniagua

Keyword(s):

Shock Waves ◽

Vortex Shedding ◽

Turbine Cascade ◽

Transonic Turbine

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Impact of Different Film-Cooling Modes at Trailing Edge on the Aerodynamic Performance of Heavy Duty Gas Turbine Cascade

Applied Mechanics and Materials ◽

10.4028/www.scientific.net/amm.84-85.259 ◽

2011 ◽

Vol 84-85 ◽

pp. 259-263

Author(s):

Xun Liu ◽

Song Tao Wang ◽

Xun Zhou ◽

Guo Tai Feng

Keyword(s):

Gas Turbine ◽

Film Cooling ◽

Large Mass ◽

Turbine Cascade ◽

Heavy Duty ◽

Trailing Edge ◽

Central Difference ◽

Adiabatic Effectiveness ◽

Heavy Duty Gas Turbine ◽

The Impact

In this paper, the trailing edge film cooling flow field of a heavy duty gas turbine cascade has been studied by central difference scheme and multi-block grid technique. The research is based on the three-dimensional N-S equation solver. By way of analysis of the temperature field, the distribution of profile pressure, and the distribution of film-cooling adiabatic effectiveness in the region of trailing edge with different cool air injection mass and different angles, it is found that the impact on the film-cooling adiabatic effectiveness is slightly by changing the injection mass. The distribution of profile pressure dropped intensely at the pressure side near the injection holes line with the large mass cooling air. The cooling effect is good in the region of trailing edge while the injection air is along the direction of stream.

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The Aerothermal Performance of a Cooled Winglet Tip in a High Pressure Turbine Cascade

Journal of Turbomachinery ◽

10.1115/1.4006611 ◽

2013 ◽

Vol 135 (3) ◽

Cited By ~ 10

Author(s):

Chao Zhou ◽

Howard Hodson ◽

Ian Tibbott ◽

Mark Stokes

Keyword(s):

Mass Flow ◽

Heat Load ◽

Blade Surface ◽

Flow Ratio ◽

Turbine Cascade ◽

Tip Leakage Flow ◽

High Pressure Turbine ◽

Cooling Effectiveness ◽

Tip Leakage ◽

Coolant Injection

The aerothermal performance of a winglet tip with cooling holes on the tip and on the blade surface near the tip is reported in this paper. The investigation was based on a high pressure turbine cascade. Experimental and numerical methods were used. The effects of the coolant mass flow rate are also studied. Because the coolant injection partially blocks the tip leakage flow, more passage flow is turned by the blade. As a result, the coolant injection on the winglet tip reduces the deviation of the flow downstream of the cascade due to the tip leakage flow. However, the tip leakage loss increases slightly with the coolant mass flow ratio. Both the computational fluid dynamics tools and experiments using the Amonia–Diazo technique were used to determine the cooling effectiveness. On the blade pressure side surface, low cooling effectiveness appears around the holes due to the lack of the coolant from the cooling hole or the lift-off of the coolant from the blade surface when the coolant mass flow is high. The cooling effectiveness on the winglet tip is a combined effect of the coolant ejected from all the holes. On the top of the winglet tip, the average cooling effectiveness increases and the heat load decreases with increasing coolant mass flow. Due to its large area, the cooled winglet tip has a higher heat load than an uncooled flat tip at engine representative coolant mass flow ratio. Nevertheless, the heat flux rate per unit area of the winglet is much lower than that of an uncooled flat tip. The cycle analysis is carried out and the effects of relative tip-to-casing endwall motion are address.

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