An Examination of the Contributions to Loss on a Transonic Turbine Blade in Cascade

1992 ◽  
Vol 114 (1) ◽  
pp. 155-162 ◽  
Author(s):  
D. J. Mee ◽  
N. C. Baines ◽  
M. L. G. Oldfield ◽  
T. E. Dickens
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Suction Surface ◽  
Near Wake ◽  
Linear Cascade ◽  
Transonic Turbine ◽  
The Individual ◽  
The Way

Experiments to measure losses of a linear cascade of transonic turbine blades are reported. Detailed measurements of the boundary layer at the rear of the suction surface of a blade and examination of wake traverse data enable the individual components of boundary layer, shock and mixing loss to be determined. Results indicate that each component contributes significantly to the overall loss in different Mach number regimes. Traverses in the near wake of the blade indicate the way in which the wake develops and facilitate examination of the development of the mixing loss.

An Examination of the Contributions to Loss on a Transonic Turbine Blade in Cascade

1990 ◽  
Author(s):  
D. J. Mee ◽  
N. C. Baines ◽  
M. L. G. Oldfield ◽  
T. E. Dickens
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Suction Surface ◽  
Near Wake ◽  
Linear Cascade ◽  
Transonic Turbine ◽  
The Individual ◽  
The Way

Experiments to measure losses of a linear cascade of transonic turbine blades are reported. Detailed measurements of the boundary layer at the rear of the suction surface of a blade and examination of wake traverse data enable the individual components of boundary layer, shock and mixing loss to be determined. Results indicate that each component contributes significantly to the overall loss in different Mach number regimes. Traverses in the near wake of the blade indicate the way in which the wake develops and facilitate examination of the development of the mixing loss.

Analysis and Prediction of Transonic Turbine Blade Losses

1991 ◽  
Author(s):  
C. L. S. Farn ◽  
D. K. Whirlow ◽  
S. Chen
Keyword(s):  
Boundary Layer ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Variable Approach ◽  
Transonic Turbine ◽  
Steam Turbine Blades ◽  
Turbine Blade Design ◽  
Empirical Approaches ◽  
Mixing Losses

The use of simple computational means to determine the performance of cascades of turbine blades is attractive because it can quickly and economically yield results that can be used for optimization of classes of blades. Fully viscous flow computations are not at the point where they are economical for use in a routine way, and most computational methods lack the resolution to determine shock losses in the transonic flow regime. There is still a need for approaches that combine computation and empiricism. We describe approaches that combine quasi-three dimensional inviscid codes and boundary layer methods for blade passage flow with empirical approaches for shock losses and base pressure deficits to predict the losses in cascades of blades. Downstream mixing losses are handled by a distributed variable approach that uses the inviscid and boundary layer results to determine the distribution of variables at the trailing edge plane. The method gives accurate predictions for the set of distinctly different steam turbine blades for which it was run, and forms the basis for the development of rule-based turbine blade design.

Effect of Wake-Disturbed Flow on Heat (Mass) Transfer to a Turbine Blade

2010 ◽  
Vol 133 (1) ◽  
Author(s):  
S. Olson ◽  
S. Sanitjai ◽  
K. Ghosh ◽  
R. J. Goldstein
Keyword(s):  
Mass Transfer ◽  
Reynolds Number ◽  
Turbine Blade ◽  
Heat Mass Transfer ◽  
Turbine Blades ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Disturbed Flow ◽  
Linear Cascade ◽  
Naphthalene Sublimation Technique

This study investigates the effect of wakes in the presence of varying levels of background freestream turbulence on the heat (mass) transfer from gas turbine blades. Measurements using the naphthalene sublimation technique provide local values of the mass transfer coefficient on the pressure and suction surfaces of a simulated turbine blade in a linear cascade. Experimental parameters studied include the pitch of the wake-generating blades (vanes), blade-row separation, Reynolds number, and the freestream turbulence level. The disturbed flow strongly affects the mass transfer Stanton number on both sides of the blade, particularly along the suction surface. An earlier transition to a turbulent boundary layer occurs with increased background turbulence, higher Reynolds number, and from wakes shed from vanes placed upstream of the linear cascade. Note that once the effects on mass transfer are known, similar variation on heat transfer can be inferred from the heat/mass transfer analogy.

scholarly journals Detailed Boundary Layer Measurements on a Transonic Turbine Cascade

1990 ◽  
Author(s):  
D. J. Mee ◽  
N. C. Baines ◽  
M. L. G. Oldfield
Keyword(s):  
Thin Films ◽  
Boundary Layer ◽  
Boundary Layers ◽  
Turbine Blade ◽  
Large Scale ◽  
Turbine Cascade ◽  
Engine Operation ◽  
Turbulent Layer ◽  
Linear Cascade ◽  
Transonic Turbine

The boundary layers of a transonic turbine blade have been measured in detail. The full velocity profiles have been measured at a number of stations on both the suction and pressure surfaces, at conditions representative of engine operation, using a pitot traverse technique and a large scale (300 mm chord) linear cascade. This information has made it possible to follow the development of the boundary layers, initially laminar, through a region of natural transition to a fully developed turbulent layer. Comparisons with other, less detailed, measurements on the same profile using pitot traverse and surface mounted thin films confirm the essential features of the boundary layers.

Detailed Boundary Layer Measurements on a Transonic Turbine Cascade

1992 ◽  
Vol 114 (1) ◽  
pp. 163-172 ◽  
Author(s):  
D. J. Mee ◽  
N. C. Baines ◽  
M. L. G. Oldfield
Keyword(s):  
Thin Films ◽  
Boundary Layer ◽  
Boundary Layers ◽  
Turbine Blade ◽  
Large Scale ◽  
Turbine Cascade ◽  
Engine Operation ◽  
Turbulent Layer ◽  
Linear Cascade ◽  
Transonic Turbine

The boundary layers of a transonic turbine blade have been measured in detail. The full velocity profiles have been measured at a number of stations on both the suction and pressure surfaces, at conditions representative of engine operation, using a Pilot traverse technique and a large-scale (300 mm chord) linear cascade. This information has made it possible to follow the development of the boundary layers, initially laminar, through a region of natural transition to a fully developed turbulent layer. Comparisons with other, less detailed, measurements on the same profile using Pilot traverse and surface-mounted thin films confirm the essential features of the boundary layers.

Flow Instability in Low-Pressure Turbine Blade Passages

1984 ◽  
Author(s):  
D. K. Whirlow ◽  
T. H. McCloskey ◽  
J. Davids ◽  
S. Chen ◽  
J. R. Kadambi ◽  
...  
Keyword(s):  
Mach Number ◽  
Flow Instability ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
Spatial Integration ◽  
Suction Surface ◽  
Number Range ◽  
Pressure Transducers ◽  
Actual Field ◽  
The Individual

The unsteady interaction among shocks, spontaneous condensation and boundary layers were measured in a cascade of tip section turbine blades with steam conditions similar to those that would be experienced by the blade under actual field conditions. Test conditions covered the Mach number range of 1.0 to 1.4, based upon isentropic exit conditions. Inlet temperature conditions were such as to provide fully dry runs, runs with shock and condensation being approximately coincident, and runs with the condensation occurring upstream of the shock. Unsteady pressure data were provided by six miniature pressure transducers embedded in the suction surface of one of the blades in the region of the impingement of the trailing edge shock from the adjacent blade. Spatial integration of the individual transducer signals provided fluctuating forces over the region measured. Unsteady forces were generally low but the nature and level of the fluctuating forces were different near a Mach number of 1.0, where several distinct spikes were present, and responses for some frequencies reached nearly one percent of the steady forces.

Effect of Wake-Disturbed Flow on Heat (Mass) Transfer to a Turbine Blade

2009 ◽  
Author(s):  
S. J. Olson ◽  
S. Sanitjai ◽  
K. Ghosh ◽  
R. J. Goldstein
Keyword(s):  
Mass Transfer ◽  
Reynolds Number ◽  
Turbine Blade ◽  
Heat Mass Transfer ◽  
Turbine Blades ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Disturbed Flow ◽  
Linear Cascade ◽  
Naphthalene Sublimation Technique

This study investigates the effect of wakes in the presence of varying levels of background freestream turbulence on the heat (mass) transfer from gas turbine blades. Measurements using the naphthalene sublimation technique provide local values of the mass transfer coefficient on the pressure and suction surfaces of a simulated turbine blade in a linear cascade. Experimental parameters studied include the pitch of the wake-generating blades (vanes), blade-row separation, Reynolds number and the freestream turbulence level. The disturbed flow strongly affects the mass transfer Stanton number on both sides of the blade, particularly along the suction surface. An earlier transition to a turbulent boundary layer occurs with increased background turbulence, higher Reynolds number and from wakes shed from vanes placed upstream of the linear cascade. Note that once the effects on mass transfer are known, similar variation on heat transfer can be inferred from the heat/mass transfer analogy.

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

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.

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.

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