scholarly journals Computational study of flow incidence effects on the aeroacoustics of low blade-tip Mach number propellers

2022 ◽  
Vol 120 ◽  
pp. 107275
Author(s):  
Gianluca Romani ◽  
Edoardo Grande ◽  
Francesco Avallone ◽  
Daniele Ragni ◽  
Damiano Casalino
Keyword(s):  
Mach Number ◽  
Computational Study ◽  
Blade Tip

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2009 ◽  
Vol 1 (2) ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Heat Transfer and Flow on the First-Stage Blade Tip of a Power Generation Gas Turbine: Part 2—Simulation Results

1999 ◽  
Vol 122 (2) ◽  
pp. 272-277 ◽  
Author(s):  
A. A. Ameri ◽  
R. S. Bunker
Keyword(s):  
Heat Transfer ◽  
Power Generation ◽  
Computational Study ◽  
Transfer Coefficients ◽  
Large Power ◽  
Heat Transfer Coefficients ◽  
Blade Tip ◽  
Convective Heat Transfer Coefficients ◽  
Edge Separation ◽  
Experimental Rig

A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first-stage blade tip surface for a geometry typical of large power generation turbines (>100 MW). This paper is concerned with the numerical prediction of the tip surface heat transfer. Good comparison with the experimental measured distribution was achieved through accurate modeling of the most important features of the blade passage and heating arrangement as well as the details of experimental rig likely to affect the tip heat transfer. A sharp edge and a radiused edge tip was considered. The results using the radiused edge tip agreed better with the experimental data. This improved agreement was attributed to the absence of edge separation on the tip of the radiused edge blade. [S0889-504X(00)01802-X]

Computational Modelling of Tip Heat Transfer to a Super-Scale Model of an Unshrouded Gas Turbine Blade

2008 ◽  
Author(s):  
Brian M. T. Tang ◽  
Pepe Palafox ◽  
David R. H. Gillespie ◽  
Martin L. G. Oldfield ◽  
Brian C. Y. Cheong
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Nusselt Number ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Computational Study ◽  
Turbulence Models ◽  
Scale Model ◽  
Tip Leakage Flow ◽  
Blade Tip

Control of over-tip leakage flow between turbine blade tips and the stationary shroud is one of the major challenges facing gas turbine designers today. The flow imposes large thermal loads on unshrouded high pressure turbine blades and is significantly detrimental to turbine blade life. This paper presents results from a computational study performed to investigate the detailed blade tip heat transfer on a sharp-edged, flat tip HP turbine blade. The tip gap is engine representative at 1.5% of the blade chord. Nusselt number distributions on the blade tip surface have been obtained from steady flow simulations and are compared to experimental data carried out in a super-scale cascade, which allows detailed flow and heat transfer measurements in stationary and engine representative conditions. Fully structured, multiblock hexahedral meshes were used in the simulations, performed in the commercial solver Fluent. Seven industry-standard turbulence models, and a number of different tip gridding strategies are compared, varying in complexity from the one-equation Spalart-Allmaras model to a seven-equation Reynolds Stress model. Of the turbulence models examined, the standard k-ω model gave the closest agreement to the experimental data. The discrepancy in Nusselt number observed was just 5%. However, the size of the separation on the pressure side rim was underpredicted, causing the position of reattachment to occur too close to the edge. Other turbulence models tested typically underpredicted Nusselt numbers by around 35%, although locating the position of peak heat flux correctly. The effect of the blade to casing motion was also simulated successfully, qualitatively producing the same changes in secondary flow features as were previously observed experimentally, with associated changes in heat transfer to the blade tip.

Computational Modeling of Tip Heat Transfer to a Superscale Model of an Unshrouded Gas Turbine Blade

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Brian M. T. Tang ◽  
Pepe Palafox ◽  
Brian C. Y. Cheong ◽  
Martin L. G. Oldfield ◽  
David R. H. Gillespie
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Nusselt Number ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Computational Study ◽  
Turbine Blades ◽  
Turbulence Models ◽  
Tip Leakage Flow ◽  
Blade Tip

Control of over-tip leakage flow between turbine blade tips and the stationary shroud is one of the major challenges facing gas turbine designers today. The flow imposes large thermal loads on unshrouded high pressure (HP) turbine blades and is significantly detrimental to turbine blade life. This paper presents results from a computational study performed to investigate the detailed blade tip heat transfer on a sharp-edged, flat tip HP turbine blade. The tip gap is engine representative at 1.5% of the blade chord. Nusselt number distributions on the blade tip surface have been obtained from steady flow simulations and are compared with experimental data carried out in a superscale cascade, which allows detailed flow and heat transfer measurements in stationary and engine representative conditions. Fully structured, multiblock hexahedral meshes were used in the simulations performed in the commercial solver FLUENT. Seven industry-standard turbulence models and a number of different tip gridding strategies are compared, varying in complexity from the one-equation Spalart–Allmaras model to a seven-equation Reynolds stress model. Of the turbulence models examined, the standard k-ω model gave the closest agreement to the experimental data. The discrepancy in Nusselt number observed was just 5%. However, the size of the separation on the pressure side rim was underpredicted, causing the position of reattachment to occur too close to the edge. Other turbulence models tested typically underpredicted Nusselt numbers by around 35%, although locating the position of peak heat flux correctly. The effect of the blade to casing motion was also simulated successfully, qualitatively producing the same changes in secondary flow features as were previously observed experimentally, with associated changes in heat transfer with the blade tip.

Numerical Investigation of Aerothermal Characteristics of the Blade Tip and Near-Tip Regions of a Transonic Turbine Blade

2015 ◽  
Vol 137 (9) ◽  
Author(s):  
A. Arisi ◽  
S. Xue ◽  
W. F. Ng ◽  
H. K. Moon ◽  
L. Zhang
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Rotor Blade ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Tip Leakage Flow ◽  
Blade Tip ◽  
Exit Mach Number

In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the exit Mach number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at exit Mach numbers of Mexit = 0.85 (Reexit = 9.75 × 105) and Mexit = 1.0 (Reexit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The exit Reynolds number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with exit Mach number due to high turbulence generation in the tip gap and flow reattachment. While increase in exit Mach number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in exit Mach number was found to also induce strong flow relaminarization on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in exit Mach number but significant increase in local heat transfer was noted in this region.

scholarly journals An improved model for tip clearance loss in transonic axial compressors

2017 ◽  
Vol 232 (4) ◽  
pp. 295-314
Author(s):  
Shubo Ye ◽  
Qingjun Zhao ◽  
Weiwei Cui ◽  
Guang Xi ◽  
Jianzhong Xu
Keyword(s):  
Flow Control ◽  
Computational Study ◽  
Control Volume ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Incoming Flow ◽  
Flow Angle ◽  
Tip Clearance Flow ◽  
Clearance Flow ◽  
Blade Tip

An improved compressible model for estimating tip clearance loss in transonic compressors is presented with the emphasis on the effects of blade tip loading distribution and double leakage flow. Tip clearance flow is treated as three parts along the chord and the progressive relations from upstream to downstream part is revealed to be responsible for the formation of tip clearance flow. Control volume method is applied to simplify the mixing process and calculate the mixed-out loss for the three parts, separately. Computational study shows that mass flow of the incoming flow entering the control volume is consistent with that passing through an equivalent area of about half of tip leakage vortex region. The new model reveals that the second part of tip clearance flow has a larger mixed-out loss capacity than the two other parts. This difference is attributed to two factors: larger injection flow angle and more enrolled incoming flow, and both factors tend to increase the mixed-out loss. The success of the model implies that blade design or flow control strategies turning the tip clearance/main flow interface’s arrival onto blade tip pressure side downstream and the shock’s impingement point onto blade tip suction side upstream may be beneficial in desensitizing compressor performance to tip clearance size, without trading off pressure rise.

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine: Part 1 — Experimental Results

1999 ◽  
Author(s):  
Ronald S. Bunker ◽  
Jeremy C. Bailey ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Power Generation ◽  
Mach Number ◽  
Transfer Coefficient ◽  
Pressure Distribution ◽  
Turbulence Intensity ◽  
Experimental Results ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines (>100MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment which develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57•106, and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heal transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

scholarly journals Computational study of flow around a NACA 0012 by using Roe FVM scheme and davis-yee TVD scheme

2019 ◽  
Vol 16 (2) ◽  
pp. 1018
Author(s):  
Fatimah Yusop ◽  
Zamri Omar ◽  
Bambang Basuno ◽  
Nik Normunira Mat Hassan
Keyword(s):  
Mach Number ◽  
Numerical Scheme ◽  
Computational Study ◽  
Pressure Coefficient ◽  
Flow Analysis ◽  
Computer Code ◽  
Tvd Scheme ◽  
Finite Volume Scheme ◽  
Airfoil Surface ◽  
Naca 0012

<p>Currently CFD had been considered as an important tool for solving engineering problems. The application of CFD had been used intensively in aircraft industries in design a new aircraft or in the effort of improvement on the exiting aircraft. In term of CFD computer code, the CFD code differs with any others may due to the difference in the numerical scheme have been used. Therefore, the present work presents the comparison result between two developed computer codes with ANSYS-FLUENT software to the case of transonic steady flow past through airfoil NACA 0012. The first computer code used a finite difference method with numerical scheme according to Davis-Yee TVD scheme. Meanwhile, the second computer code used a Roe’s cell centre finite volume scheme. The flow analysis is carried out at two Mach number, M (0.65 &amp; 0.8). Each Mach number applied to two different angles of attacks (0° &amp; 5°).  The flow domain discretized by use of C-topology with 193x63 grid points. The comparison in term of the pressure coefficient, along the airfoil surface are presented. From the result, indicated that developed computer code is able to capture the presence of shock wave in the flow field.</p>

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