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.

The Trailing Edge Loss of Transonic Turbine Blades

1990 ◽  
Vol 112 (2) ◽  
pp. 277-285 ◽  
Author(s):  
J. D. Denton ◽  
L. Xu
Keyword(s):  
Surface Pressure ◽  
Turbine Blades ◽  
Back Pressure ◽  
Base Pressure ◽  
Mass Energy ◽  
Suction Surface ◽  
Trailing Edge ◽  
Transonic Turbine ◽  
Energy And Momentum ◽  
Unique Relationship

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading, it is shown that there is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading, which operates choked but with subsonic axial velocity, there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two-dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results. The effects of boundary layer thickness and trailing edge coolant ejection are considered briefly. Coolant ejection acts to reduce the mainstream loss. It is shown that suction surface curvature downstream of the throat may be highly beneficial in reducing the loss of blades with thick trailing edges operating at high subsonic or low supersonic outlet Mach numbers.

scholarly journals The Aerodynamics of Trailing-Edge-Cooled Transonic Turbine Blades: Part 1 — Experimental Approach

1997 ◽  
Author(s):  
Mathias Deckers ◽  
John D. Denton
Keyword(s):  
Turbine Blades ◽  
Base Pressure ◽  
Stagnation Temperature ◽  
Trailing Edge ◽  
Static Pressure Distribution ◽  
Discernible Effect ◽  
Two Dimensional Flow ◽  
Shock System ◽  
Flow Through ◽  
Transonic Turbine

The research presented in this part of the paper involved a detailed experimental study of the flow through transonic turbine blading with trailing edge coolant ejection. The tests were carried out on (nearly) flat plate models representing the region of uncovered turning downstream of the throat. The investigation focused on the aerodynamic aspects associated with trailing edge ejection in steady two-dimensional flow over a range of exit Mach numbers, coolant pressure ratios and temperature ratios. The experiments showed that the simple existence of the coolant cavity leads to a substantial change of the flow field in the near wake. Consequently, the slotted unblown base was found to have considerably less loss than the solid one. The effect of coolant ejection is shown to cause a substantial increase in base pressure and reduction in overall loss. The surface static pressure distribution and boundary layers were affected by the coolant in two ways: directly from downstream, via the base pressure, and indirectly through a changed trailing edge shock system. However, the coolant stagnation temperature ratio was found to have no discernible effect on the base pressure and loss.

The Trailing Edge Loss of Transonic Turbine Blades

1989 ◽  
Author(s):  
J. D. Denton ◽  
L. Xu
Keyword(s):  
Surface Pressure ◽  
Turbine Blades ◽  
Back Pressure ◽  
Base Pressure ◽  
Mass Energy ◽  
Suction Surface ◽  
Trailing Edge ◽  
Transonic Turbine ◽  
Energy And Momentum ◽  
Unique Relationship

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading it is shown that there is is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading which operates choked but with subsonic axial velocity there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results.

Investigation on the Reduction of Trailing Edge Shock Losses for a Highly Loaded Transonic Turbine

2016 ◽  
Author(s):  
Wei Zhao ◽  
Weiwei Luo ◽  
Qingjun Zhao ◽  
Jianzhong Xu
Keyword(s):  
Turbine Blades ◽  
Strong Shock ◽  
Suction Side ◽  
Loss Reduction ◽  
Blade Profile ◽  
Trailing Edge ◽  
Pressure Loss Coefficient ◽  
Reflected Shock ◽  
Transonic Turbine ◽  
Highly Loaded

A shock loss reduction method for highly loaded transonic turbine blades with convergent passages is presented. The method is illustrated with an improved blade profile that employs a negative curvature curve on its uncovered suction side. The improved profile and a conventional baseline profile are applied to two cascades with the same solidity, chord and aspect ratio respectively. The numerical simulation results for the two cascades show that a reduction of 4.58% in the total pressure loss coefficient is obtained for the improved profile at the design condition. The effects of back pressures on the performance of both cascades are also presented, and the improved blade profile shows a much better part-load performance. The paper compares the flow fields of the baseline and the improved blade profiles to understand loss reduction mechanism especially by analyzing the shock interactions downstream of the trailing edge. It is found that, for the improved profile, the reflected shock of pressure side leg of the trailing-edge shock rotates forward and the suction side leg of the trailing-edge shock rotates backward. Therefore, the two shocks delay their intersection points where they merge into a relatively strong shock, and consequently produce less shock losses than those of the baseline profile.

scholarly journals The Influence of Trailing Edge Ejection on the Base Pressure in Transonic Turbine Cascades

1982 ◽  
Author(s):  
C. H. Sieverding
Keyword(s):  
Turbine Blades ◽  
Density Ratio ◽  
Main Flow ◽  
Coolant Flow ◽  
Base Pressure ◽  
Trailing Edge ◽  
Number Range ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Transonic Cascade

This paper summarizes the results of base pressure studies on transonic turbine blades in presence of an ejection of coolant flow from a slot in the trailing edge. The first part of the paper reports on tests carried out on a enlarged model of the overhang section of a typical transonic cascade. This model provides valuable information about the detailed trailing edge pressure distribution and points to an asymmetric evolution of the base pressure on both sides of the slot in presence of a bleed. The second part of the paper presents experimental results from cascade tests covering an outlet Mach number range from M2, is = 0.5 to 1.35. These experiments underline the importance of the coolant flow impact on the base pressure and confirm the asymmetry of the base pressure with respect to the cooling slot. Tests with different coolant flow gases point to the significance of a proper simulation of the density ratio between coolant flow and main flow.

The Influence of Trailing Edge Ejection on the Base Pressure in Transonic Turbine Cascades

1983 ◽  
Vol 105 (2) ◽  
pp. 215-222 ◽  
Author(s):  
C. H. Sieverding
Keyword(s):  
Turbine Blades ◽  
Density Ratio ◽  
Main Flow ◽  
Coolant Flow ◽  
Base Pressure ◽  
Trailing Edge ◽  
Number Range ◽  
Transonic Turbine ◽  
Turbine Cascades ◽  
Transonic Cascade

This paper summarizes the results of base pressure studies on transonic turbine blades in the presence of an ejection of coolant flow from a slot in the trailing edge. The first part of the paper reports on tests carried out on a enlarged model of the overhang section of a typical transonic cascade. This model provides valuable information about the detailed trailing edge pressure distribution and points to an asymmetric evolution of the base pressure on both sides of the slot in the presence of a bleed. The second part of the paper presents experimental results from cascade tests covering an outlet Mach number range M2, is = 0.5 to 1.35. These experiments underline the importance of the coolant flow impact on the base pressure and confirm the asymmetry of the base pressure with respect to the cooling slot. Tests with different coolant flow gases point to the significance of a proper simulation of the density ratio between coolant flow and main flow.

The Aerodynamics of Trailing-Edge-Cooled Transonic Turbine Blades: Part 2 — Theoretical and Computational Approach

1997 ◽  
Author(s):  
Mathias Deckers ◽  
John D. Denton
Keyword(s):  
Transonic Flow ◽  
Theoretical Study ◽  
Computational Study ◽  
Control Volume ◽  
Turbine Blades ◽  
Computational Approach ◽  
Trailing Edge ◽  
Time Marching ◽  
Transonic Turbine ◽  
Volume Method

A theoretical and computational study into the aerodynamics of trailing-edge-cooled transonic turbine blades is described in this part of the paper. The theoretical study shows that, for unstaggered blades with coolant ejection, the base pressure and overall loss can be determined exactly by a simple control volume analysis. This theory suggests that a thick, cooled trailing edge with a wide slot can be more efficient than a thin, solid trailing edge. An existing time-marching finite volume method is adapted to calculate the transonic flow with trailing edge coolant ejection on a structured, quasi-orthogonal mesh. Good overall agreement between the present method, inviscid and viscous, and experimental evidence is obtained.

New Supersonic Loss Model for the Preliminary Design of Transonic Turbine Blades and the Influence of Pitch

2020 ◽  
Vol 142 (4) ◽  
Author(s):  
Luis Teia
Keyword(s):  
Mach Number ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Published Data ◽  
Efficient Design ◽  
Operational Parameters ◽  
Loss Model ◽  
Shock System ◽  
Transonic Turbine ◽  
Shock Loss

Abstract In order to produce a more efficient design of a compact turbine driving a cryogenic engine turbo-pump for a satellite delivering rocket, a new supersonic loss model is proposed. The new model was constructed based on high-quality published data, composed of Schlieren photographs and experimental measurements, that combined provided a unique insight into the mechanisms driving supersonic losses. Using this as a cornerstone, model equations were formulated that predict the critical Mach number and shock loss and shock-induced mixing loss as functions of geometrical (i.e., blade outlet and uncovered turning angle and trailing edge thickness) and operational parameters (i.e., exit Mach number). A series of highly resolved CFD numerical simulations were conducted on an in-house designed state-of-the-art transonic turbine rotor row (around unity aspect ratio (AR)) to better understand changes in the shock system for varying parameters. The main outcome showed that pitch to chord ratio has a powerful impact on the shock system, and thus on the manner by which shock loss and shock-induced mixing loss is distributed to compose the overall supersonic losses. The numerical loss estimates for two pitch to chord ratios—t⁄c = 0.70 and t⁄c = 0.98—were compared with absolute loss data of a previously published similar blade with satisfactory agreement. Calibrated equations are provided to allow hands-on integration into existing overall turbine loss models, where supersonic losses play a key role, for further enhancement of preliminary turbine design.

Numerical Study of Heat Transfer in Novel Wavy Trailing Edge Design for Gas Turbine Airfoils

2019 ◽  
Author(s):  
Sarwesh Parbat ◽  
Li Yang ◽  
Minking Chyu ◽  
Sin Chien Siw ◽  
Ching-Pang Lee
Keyword(s):  
Heat Transfer ◽  
Steady State ◽  
Gas Turbine ◽  
Numerical Study ◽  
Turbine Blades ◽  
Suction Side ◽  
Inlet Temperature ◽  
Protective Film ◽  
Pressure Side ◽  
Trailing Edge

Abstract The strive to achieve increasingly higher efficiencies in gas turbine power generation has led to a continued rise in the turbine inlet temperature. As a result, novel cooling approaches for gas turbine blades are necessary to maintain them within the material’s thermal mechanical performance envelope. Various internal and external cooling technologies are used in different parts of the blade airfoil to provide the desired levels of cooling. Among the different regions of the blade profile, the trailing edge (TE) presents additional cooling challenges due to the thin cross section and high thermal loads. In this study, a new wavy geometry for the TE has been proposed and analyzed using steady state numerical simulations. The wavy TE structure resembled a sinusoidal wave running along the span of the blade. The troughs on both pressure side and suction side contained the coolant exit slots. As a result, consecutive coolant exit slots provided an alternating discharge between the suction side and the pressure side of the blade. Steady state conjugate heat transfer simulations were carried out using CFX solver for four coolant to mainstream mass flow ratios of 0.45%, 1%, 1.5% and 3%. The temperature distribution and film cooling effectiveness in the TE region were compared to two conventional geometries, pressure side cutback and centerline ejection which are widely used in vanes and blades for both land-based and aviation gas turbine engines. Unstructured mesh was generated for both fluid and solid domains and interfaces were defined between the two domains. For turbulence closer, SST-kω model was used. The wall y+ was maintained < 1 by using inflation layers at all the solid fluid interfaces. The numerical results depicted that the alternating discharge from the wavy TE was able to form protective film coverage on both the pressure and suction side of the blade. As a result, significant reduction in the TE metal was observed which was up to 14% lower in volume averaged temperature in the TE region when compared to the two baseline conventional configurations.

The Unsteady Effect of Passing Shocks on Pressure Surface Versus Suction Surface Heat Transfer in Film-Cooled Transonic Turbine Blades

2003 ◽  
Author(s):  
A. C. Smith ◽  
A. C. Nix ◽  
T. E. Diller ◽  
W. F. Ng
Keyword(s):  
Heat Transfer ◽  
Shock Wave ◽  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Surface Heat ◽  
Suction Surface ◽  
Time Resolved ◽  
Pressure Surface ◽  
Transonic Turbine

This paper documents the measurement of the unsteady effects of passing shock waves on film cooling heat transfer on both the pressure and suction surfaces of first stage transonic turbine blades with leading edge showerhead film cooling. Experiments were performed for several cooling blowing ratios with an emphasis on time-resolved pressure and heat flux measurements on the pressure surface. Results without film cooling on the pressure surface demonstrated that increases in heat flux were a result of shock heating (the increase in temperature across the shock wave) rather than shock interaction with the boundary layer or film layer. Time-resolved measurements with film cooling demonstrated that the relatively strong shock wave along the suction surface appears to retard coolant ejection there and causes excess coolant to be ejected from pressure surface holes. This actually causes a decrease in heat transfer on the pressure surface during a large portion of the shock passing event. The magnitude of the decrease is almost as large as the increase in heat transfer without film cooling. The decrease in coolant ejection from the suction surface holes did not appear to have any effects on suction surface heat transfer.

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