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.

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.

Loss Generation in Transonic Turbine Blading

2018 ◽  
Vol 140 (4) ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
Stokes Equations ◽  
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 trends with exit Mach number ranging from 0.8 to 1.4. Assessments using steady Reynolds-averaged Navier--Stokes equations (RANS) computation of the flow in the two turbine blades, complemented with control volume analyses and loss modeling, 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 (TE) 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 the experimental data. The measured loss plateau in blade 1 for exit Mach number of 1–1.4 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 an approximate value of 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.

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.

A Study of Advanced High Loaded Transonic Turbine Airfoils

2004 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Markus Olhofer ◽  
Bernhard Sendhoff ◽  
Friedrich T. Kost ◽  
...  
Keyword(s):  
Numerical Optimization ◽  
High Performance ◽  
Optimization Method ◽  
Turbine Rotor ◽  
Navier Stokes ◽  
Turbine Rotor Blade ◽  
Shock System ◽  
Transonic Turbine ◽  
Exit Mach Number ◽  
Turbine Airfoils

The development of high performance turbine airfoils has been investigated under the condition of a supersonic exit Mach number. In order to obtain a new aerodynamic design concept for a high loaded turbine rotor blade, we employed an evolutionary algorithm for numerical optimization. The target of the optimization method, which is called evolution strategy (ES), was the minimization of the total pressure loss and the deviation angle. The optimization process includes the representation of the airfoil geometry, the generation of the grid for a blade-to-blade CFD analysis, and a 2D Navier-Stokes solver with a low-Re k-ε turbulence model in order to evaluate the performance. Some interesting aspects, for example, a double shock system, an early transition and a re-distribution of aerodynamic loading on blade surface, observed in the optimized airfoil, are discussed. The increased performance of the optimized blade has been confirmed by detailed experimental investigation, using conventional probes, hot-films and L2F system.

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.

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.

A Shock Loss Model for Supersonic Compressor Cascades

1999 ◽  
Vol 121 (1) ◽  
pp. 28-35 ◽  
Author(s):  
G. S. Bloch ◽  
W. W. Copenhaver ◽  
W. F. O’Brien
Keyword(s):  
Mach Number ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Flow Angle ◽  
Loss Model ◽  
Viscous Profile ◽  
Loss Models ◽  
Profile Loss ◽  
Shock Loss ◽  
Good Agreement

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier–Stokes analyses of the LO30-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

PERFORMANCE EVALUATION AND WAKE STUDY OF A MICRO WIND TURBINE

2011 ◽  
Vol 35 (1) ◽  
pp. 101-117 ◽  
Author(s):  
Graeme I. Comyn ◽  
David S. Nobes ◽  
Brian A. Fleck
Keyword(s):  
Wind Tunnel ◽  
Wind Turbine ◽  
Power Output ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Published Data ◽  
Wind Turbine Blades ◽  
Axial Thrust ◽  
Cross Sectional ◽  
Output Performance

In preparation for a study on icing of wind turbine blades, we tested a horizontal axis micro wind turbine in a low speed wind tunnel. The ratio of wind turbine rotor area to wind tunnel cross-sectional area resulted in highly blocked experimental configuration. The turbine was instrumented to measure rotational speed of the rotor, axial thrust and power output. Performance characteristics were calculated and compared with the manufacturer’s published data. In addition, the near wake of the turbine was measured with a Kiel probe. One dimensional axial momentum theory, including a modification that includes channel walls, was applied to determine power extracted from the wind by the rotor. The results were compared to actual power output and show that though the assumptions of the model over-predict power by 50 % the basic trend is followed.

scholarly journals A Shock Loss Model for Supersonic Compressor Cascades

1997 ◽  
Author(s):  
Gregory S. Bloch ◽  
William W. Copenhaver ◽  
Walter F. O’Brien
Keyword(s):  
Mach Number ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Flow Angle ◽  
Loss Model ◽  
Viscous Profile ◽  
Loss Models ◽  
Profile Loss ◽  
Shock Loss ◽  
Good Agreement

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier-Stokes analyses of the L030-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

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