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.

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.

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 Three-Dimensional Pressure Patterns in a Recess of a Turbulent and Compressible Hybrid Journal Bearing

2003 ◽  
Vol 125 (2) ◽  
pp. 301-308 ◽  
Author(s):  
Mathieu Helene ◽  
Mihai Arghir ◽  
Jean Frene
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Journal Bearing ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Wire Mesh ◽  
Primary Concern ◽  
Reference Parameters ◽  
Exit Mach Number

The present work investigates the flow in the feeding recess of a hybrid journal bearing. Numerical integration of the complete Navier-Stokes equations was performed with an appropriate turbulence model. Of primary concern is the pressure field on the rotating journal surface that is commonly known as the recess pressure pattern. The goal of the work is to determine the influences of fluid compressibility, operating conditions and recess geometry. Reference parameters selected for this study comprise feeding Reynolds number Rea of 2.105, sliding Reynolds number Rec of 5.103 and recess depth over film thickness ratio e/H of 2.2. Compressibility was considered first. Three values of the axial exit Mach number were selected for computation, namely 0.2, 0.45, and 0.7. As no significant variation was found, the Mach number was fixed at 0.45 in subsequent studies concerning other parameters:     Feeding Reynolds number, Rea       2.104,2.105,4.105     Recess depth, e/H           0, 2.2, 8     Feedhole axis inclination        90°, 135°, 165°     Feedhole location (Figs. 1(a) and 13)   centered, downstream offset. As each parameter is varied, wire mesh plot of pressure and its sectional profiles are examined and effects of varying various parameters are discussed in reference to flow processes as they may affect the support characteristics of the hybrid journal bearing.

Receptivity of the boundary layer to vibrations of the wing surface

2013 ◽  
Vol 723 ◽  
pp. 480-528 ◽  
Author(s):  
A. I. Ruban ◽  
T. Bernots ◽  
D. Pryce
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Stokes Equations ◽  
Wing Surface ◽  
Main Concern ◽  
Elastic Vibrations ◽  
Stokes Layer ◽  
Characteristic Wavelength ◽  
Tollmien Schlichting ◽  
Pressure Perturbations

AbstractIn this paper we study the generation of Tollmien–Schlichting waves in the boundary layer due to elastic vibrations of the wing surface. The subsonic flow regime is considered with the Mach number outside the boundary layer $M= O(1)$. The flow is investigated based on the asymptotic analysis of the Navier–Stokes equations at large values of the Reynolds number, $\mathit{Re}= {\rho }_{\infty } {V}_{\infty } L/ {\mu }_{\infty } $. Here $L$ denotes the wing section chord; and ${V}_{\infty } $, ${\rho }_{\infty } $ and ${\mu }_{\infty } $ are the free stream velocity, air density and dynamic viscosity, respectively. We assume that in the spectrum of the wing vibrations there is a harmonic that comes in to resonance with the Tollmien–Schlichting wave on the lower branch of the stability curve; this happens when the frequency of the harmonic is a quantity of the order of $({V}_{\infty } / L){\mathit{Re}}^{1/ 4} $. The wavelength, $\ell $, of the elastic vibrations of the wing is assumed to be $\ell \sim L{\mathit{Re}}^{- 1/ 8} $, which has been found to represent a ‘distinguished limit’ in the theory. Still, the results of the analysis are applicable for $\ell \gg L{\mathit{Re}}^{- 1/ 8} $ and $\ell \ll L{\mathit{Re}}^{- 1/ 8} $; the former includes an important case when $\ell = O(L)$. We found that the vibrations of the wing surface produce pressure perturbations in the flow outside the boundary layer, which can be calculated with the help of the ‘piston theory’, which remains valid provided that the Mach number, $M$, is large as compared to ${\mathit{Re}}^{- 1/ 4} $. As the pressure perturbations penetrate into the boundary layer, a Stokes layer forms on the wing surface; its thickness is estimated as a quantity of the order of ${\mathit{Re}}^{- 5/ 8} $. When $\ell = O({\mathit{Re}}^{- 1/ 8} )$ or $\ell \gg {\mathit{Re}}^{- 1/ 8} $, the solution in the Stokes layer appears to be influenced significantly by the compressibility of the flow. The Stokes layer on its own is incapable of producing the Tollmien–Schlichting waves. The reason is that the characteristic wavelength of the perturbation field in the Stokes layer is much larger than that of the Tollmien–Schlichting wave. However, the situation changes when the Stokes layer encounters a wall roughness, which are plentiful in real aerodynamic flows. If the longitudinal extent of the roughness is a quantity of the order of ${\mathit{Re}}^{- 3/ 8} $, then efficient generation of the Tollmien–Schlichting waves becomes possible. In this paper we restrict our attention to the case when the Stokes layer interacts with an isolated roughness. The flow near the roughness is described by the triple-deck theory. The solution of the triple-deck problem can be found in an analytic form. Our main concern is with the flow behaviour downstream of the roughness and, in particular, with the amplitude of the generated Tollmien–Schlichting waves.

Compressibility and Variable Density Effects in Turbulent Boundary layers

2006 ◽  
Vol 129 (4) ◽  
pp. 441-448 ◽  
Author(s):  
Kunlun Liu ◽  
Richard H. Pletcher
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Boundary Layers ◽  
Wall Temperature ◽  
Stokes Equations ◽  
Variable Density ◽  
Turbulent Boundary Layers ◽  
Reynolds Analogy ◽  
Density Effects

Two compressible turbulent boundary layers have been calculated by using direct numerical simulation. One case is a subsonic turbulent boundary layer with constant wall temperature for which the wall temperature is 1.58 times the freestream temperature and the other is a supersonic adiabatic turbulent boundary layer subjected to a supersonic freestream with a Mach number 1.8. The purpose of this study is to test the strong Reynolds analogy (SRA), the Van Driest transformation, and the applicability of Morkovin’s hypothesis. For the first case, the influence of the variable density effects will be addressed. For the second case, the role of the density fluctuations, the turbulent Mach number, and dilatation on the compressibility will be investigated. The results show that the Van Driest transformation and the SRA are satisfied for both of the flows. Use of local properties enable the statistical curves to collapse toward the corresponding incompressible curves. These facts reveal that both the compressibility and variable density effects satisfy the similarity laws. A study about the differences between the compressibility effects and the variable density effects associated with heat transfer is performed. In addition, the difference between the Favre average and Reynolds average is measured, and the SGS terms of the Favre-filtered Navier-Stokes equations are calculated and analyzed.

Loss reduction on ultra high lift low-pressure turbine blades using selective roughness and wake unsteadiness

2007 ◽  
Vol 111 (1118) ◽  
pp. 257-266 ◽  
Author(s):  
R. J. Howell ◽  
K. M. Roman
Keyword(s):  
Boundary Layer ◽  
Steady Flow ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Low Pressure ◽  
Surface Boundary ◽  
Suction Surface ◽  
High Lift ◽  
Profile Losses ◽  
Lp Turbine

This paper describes how it is possible to reduce the profile losses on ultra high lift low pressure (LP) turbine blade profiles with the application of selected surface roughness and wake unsteadiness. Over the past several years, an understanding of wake interactions with the suction surface boundary layer on LP turbines has allowed the design of blades with ever increasing levels of lift. Under steady flow conditions, ultra high lift profiles would have large (and possibly open) separation bubbles present on the suction side which result from the very high diffusion levels. The separation bubble losses produced by it are reduced when unsteady wake flows are present. However, LP turbine blades have now reached a level of loading and diffusion where profile losses can no longer be controlled by wake unsteadiness alone. The ultra high lift profiles investigated here were created by attaching a flap to the trailing edge of another blade in a linear cascade — the so called flap-test technique. The experimental set-up used in this investigation allows for the simulation of upstream wakes by using a moving bar system. Hotwire and hotfilm measurements were used to obtain information about the boundary-layer state on the suction surface of the blade as it evolved in time. Measurements were taken at a Reynolds numbers ranging between 100,000 and 210,000. Two types of ultra high lift profile were investigated; ultra high lift and extended ultra high lift, where the latter has 25% greater back surface diffusion as well as a 12% increase in lift compared to the former. Results revealed that distributed roughness reduced the size of the separation bubble with steady flow. When wakes were present, the distributed roughness amplified disturbances in the boundary layer allowing for more rapid wake induced transition to take place, which tended to eliminate the separation bubble under the wake. The extended ultra high lift profile generated only slightly higher losses than the original ultra high lift profile, but more importantly it generated 12% greater lift.

Heat Transfer in Film-Cooled Turbine Blades

1993 ◽  
Author(s):  
Vijay K. Garg ◽  
Raymond E. Gaugler
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Film Cooling ◽  
Stokes Equations ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Explicit Finite Difference ◽  
Control Volumes

In order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine blades, a three-dimensional Navier-Stokes code has been developed. An existing code (Chima and Yokota, 1990) has been modified for the purpose. The code is an explicit finite difference code with an algebraic turbulence model. The thin-layer Navier-Stokes equations are solved using a general body-fitted coordinate system. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the blade surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison with experimental data is fair. Further validation of the code is required, however, and in this respect, there is an urgent need for detailed experimental data on actual turbine blades.

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.

Calculation of Heat Transfer to Convection-Cooled Gas Turbine Blades

1985 ◽  
Vol 107 (3) ◽  
pp. 620-627 ◽  
Author(s):  
W. Rodi ◽  
G. Scheuerer
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Complex Nature ◽  
Surface Boundary ◽  
Transfer Coefficients ◽  
Free Stream Turbulence ◽  
Gas Turbine Blades ◽  
Heat Transfer Coefficients

A mathematical model is presented for calculating the external heat transfer coefficients around gas turbine blades. The model is based on a finite-difference procedure for solving the boundary-layer equations which describe the flow and temperature field around the blades. The effects of turbulence are simulated by a low-Reynolds number version of the k-ε turbulence model. This allows calculation of laminar and transitional zones and also the onset of transition. Applications of the calculation method are presented to turbine-blade situations which have recently been investigated experimentally. Predicted and measured heat transfer coefficients are compared and good agreement with the data is observed. This is true especially for the pressure-surface boundary layer which is of a rather complex nature because it remains in a transitional state over the full blade length. The influence of various flow phenomena like laminar-turbulent transition and of the boundary conditions (pressure gradient, free-stream turbulence) on the predicted heat transfer rates is discussed.

About the Distribution of the Mach Number and Losses Under Consideration of the Diffusion Factor: Profile Distribution and Losses for Low Pressure Turbine Profiles With High Diffusion Factors Under Steady Flow Conditions

2016 ◽  
Author(s):  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbulence Intensity ◽  
Flow Separation ◽  
Total Pressure ◽  
Suction Side ◽  
Diffusion Factor ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Exit Mach Number

The results of this investigation consist of two linear cascades at high diffusion factors. The present measurements for each low pressure turbine profile were conducted at midspan under a range of Reynolds- and exit Mach numbers. The exit Mach number was varied in a range covering low subsonic up to values where a transonic flow regime on the suction side of the blade could be expected. The variation of the exit Mach number was also used to create different locations of the maximum Mach number and to evaluate the resulting total pressure losses. This work focuses on two profiles with a diffusion factor in a range of 0.18 ≤ DF ≤ 0.22, which is considered as a comparable level for the two cascades. The profile A is a front-loaded design and has shown no obvious flow separation on the suction side of the blade. Compared to the profile A the design B is a more aft-loaded profile which indicates flow separation on the suction side for all investigated Reynolds numbers. The integral total pressure losses were evaluated by wake traverses downstream of the profile. To determine the isentropic Mach numbers and the character of the boundary layer along the suction side of the profile, static pressure tappings and measurements with a flattened Pitot probe were carried out. Numerical studies were also conducted to investigate further the influence of a reduced turbulence intensity on the boundary layer of the suction side of design B. The results show that the optimum of the integral total pressure losses are significantly dependent on the Reynolds number. Therefore a correlation between the maximum Mach number on the suction side and the integral total pressure losses has been successfully established. A significant change of the turbulence intensity at the inlet of the cascade leads to shift of the location of the maximum Mach number. It also results in an equivalent change of the total pressure losses, which has been predicted by the trend line. However, the trend lines, which are based on the data of the integral total pressure losses of an attached boundary layer, are not able to predict the integral total pressure loss or the location of the maximum Mach number on the suction side of the blade since an open separation bubble occurs.

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