Experimental and numerical investigations of trailing edge injection in a transonic turbine cascade

2019 ◽  
Vol 92 ◽  
pp. 258-268 ◽  
Author(s):  
Jie Gao ◽  
Ming Wei ◽  
Weiliang Fu ◽  
Qun Zheng ◽  
Guoqiang Yue
Keyword(s):  
Turbine Cascade ◽  
Trailing Edge ◽  
Numerical Investigations ◽  
Transonic Turbine

Investigations of Transonic Turbine Cascade With High Stagger and Low Solidity

1979 ◽  
Author(s):  
M. Inoue ◽  
S. Yamaguchi ◽  
M. Kuroumaru
Keyword(s):  
Transonic Flow ◽  
Laval Nozzle ◽  
Experimental Studies ◽  
Flow Characteristics ◽  
Chord Length ◽  
Turbine Cascade ◽  
Trailing Edge ◽  
Time Marching ◽  
Shock System ◽  
Transonic Turbine

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.

scholarly journals Simulation of Trailing Edge Vortex Shedding in a Transonic Turbine Cascade

1996 ◽  
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.

Simulation of Trailing Edge Vortex Shedding in a Transonic Turbine Cascade

1998 ◽  
Vol 120 (1) ◽  
pp. 10-19 ◽  
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.

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

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.

scholarly journals Three-Dimensional Navier–Stokes Analysis and Redesign of an Imbedded Bellmouth Nozzle in a Turbine Cascade Inlet Section

1996 ◽  
Vol 118 (3) ◽  
pp. 529-535 ◽  
Author(s):  
P. W. Giel ◽  
J. R. Sirbaugh ◽  
I. Lopez ◽  
G. J. Van Fossen
Keyword(s):  
Three Dimensional ◽  
Navier Stokes ◽  
Experimental Measurements ◽  
Inlet Section ◽  
Turbine Cascade ◽  
Blade Cascade ◽  
Inlet Geometry ◽  
Computational Fluid Dynamics Cfd ◽  
Transonic Turbine ◽  
Flow Nonuniformity

Experimental measurements in the inlet of a transonic turbine blade cascade showed unacceptable pitchwise flow nonuniformity. A three-dimensional, Navier–Stokes computational fluid dynamics (CFD) analysis of the imbedded bellmouth inlet in the facility was performed to identify and eliminate the source of the flow nonuniformity. The blockage and acceleration effects of the blades were accounted for by specifying a periodic static pressure exit condition interpolated from a separate three-dimensional Navier–Stokes CFD solution of flow around a single blade in an infinite cascade. Calculations of the original inlet geometry showed total pressure loss regions consistent in strength and location to experimental measurements. The results indicate that the distortions were caused by a pair of streamwise vortices that originated as a result of the interaction of the flow with the imbedded bellmouth. Computations were performed for an inlet geometry that eliminated the imbedded bellmouth by bridging the region between it and the upstream wall. This analysis indicated that eliminating the imbedded bellmouth nozzle also eliminates the pair of vortices, resulting in a flow with much greater pitchwise uniformity. Measurements taken with an installed redesigned inlet verify that the flow nonuniformity has indeed been eliminated.

Experimental and Numerical Investigation on the Performance of a Family of Three HP Transonic Turbine Blades

2004 ◽  
Author(s):  
D. Corriveau ◽  
S. A. Sjolander
Keyword(s):  
Flow Velocity ◽  
Turbine Blades ◽  
Suction Side ◽  
Experimental Results ◽  
Base Pressure ◽  
Suction Surface ◽  
Trailing Edge ◽  
Lower Base ◽  
Shock System ◽  
Transonic Turbine

Experimental results concerning the performance of three high-pressure (HP) transonic turbine blades having fore-, aft- and mid-loadings have been presented previously by Corriveau and Sjolander [1]. Results from that study indicated that by shifting the loading towards the rear of the airfoil, improvements in loss performance of the order of 20% could be obtained near the design Mach number. In order to gain a better understanding of the underlying reasons for the improved loss performance of the aft-loaded blade, additional measurements were performed on the three cascades. Furthermore, 2-D numerical simulations of the cascade flow were performed in order to help in the interpretation of the experimental results. Based on the analysis of additional wake traverse data and base pressure measurements combined with the numerical results, it was found that the poorer loss performance of the baseline mid-loaded profile compared to the aft-loaded blade could be traced back to the former’s higher rear suction side curvature. The presence of higher rear suction surface curvature resulted in higher flow velocity in that region. Higher flow velocity at the trailing edge in turn contributed to reducing the base pressure. The lower base pressure at the trailing edge resulted in a stronger trailing edge shock system for the mid-loaded blade. This shock system increased the losses for the mid-loaded baseline profile when compared to the aft-loaded profile.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Hans-Ju¨rgen Rehder
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Coolant Flow ◽  
Coolant Fluid ◽  
Pressure Side ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Number Range ◽  
Mixing Losses

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

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