Impact of Wall Temperature on Turbine Blade Tip Aerothermal Performance

2014 ◽  
Vol 136 (5) ◽  
Author(s):  
Q. Zhang ◽  
L. He
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Flow Field ◽  
Temperature Dependence ◽  
Turbine Blade ◽  
Wall Temperature ◽  
High Speed ◽  
Convective Heat Transfer Coefficient ◽  
Secondary Flows ◽  
Blade Tip

Currently the aerodynamics and heat transfer over a turbine blade tip tend to be analyzed separately with the assumption that the wall thermal boundary conditions do not affect the over-tip-leakage (OTL) flow field. There are some existing correlations for correcting the wall temperature effect on heat transfer when scaled to engine realistic conditions. But they were either developed to account for the temperature dependence of fluid properties largely empirically, or based on a boundary-layer model. It would be difficult (if not impossible) to define a boundary layer in many parts of a realistic blade passage with marked three-dimensional (3D) end wall and secondary flows (including those within a blade tip and around it). The questions to be asked here are: is the OTL aerodynamics significantly affected by the wall thermal condition? And if it is, how can we count this effect consistently in turbine blade tip design and analysis using modern CFD methods? In the present study the problem has been examined for typical high-pressure turbine blade tip configurations. An extensively developed RANS code (HYDRA) is employed and validated against the experimental data from a high speed linear cascade testing rig. The numerical analysis reveals that the wall–gas temperature ratio could greatly affect the transonic OTL flow field and there is a strong two-way coupling between aerodynamics and heat transfer. The feedbacks of the thermal boundary condition to aerodynamics behave differently at different flow regimes over the tip, clearly indicating a highly localized dependence of the convective heat transfer coefficient (HTC) upon wall temperatures. This implies that to use HTC for blade metal temperature predictions without resorting a fully conjugate solution, the temperature dependence needs to be corrected locally. A nonlinear correction approach has been adopted in the present work, and the results demonstrate its effectiveness for the transonic turbine tip configurations studied.

Impact of Wall Temperature on Turbine Blade Tip Aero-Thermal Performance

2013 ◽  
Author(s):  
Q. Zhang ◽  
L. He
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Temperature Dependence ◽  
Turbine Blade ◽  
Wall Temperature ◽  
High Speed ◽  
Convective Heat Transfer Coefficient ◽  
Gas Temperature ◽  
Fluid Properties ◽  
Blade Tip

Currently, the aerodynamics and heat transfer over a turbine blade tip tend to be analyzed separately with the assumption that the wall thermal boundary conditions do not affect the Over-Tip-Leakage (OTL) flow field. There are some existing correlations for correcting the wall temperature effect on heat transfer. But they were mainly developed to account for the temperature dependence on fluid properties, and are inherently limited by the empirical nature. The questions arise with regard to: is the OTL aerodynamics significantly affected by the wall thermal condition? And if it is, how can we count this effect consistently in turbine blade tip design and analysis using modern CFD methods? In the present study, the problem has been examined for typical HP turbine blade tip configurations. An extensively developed RANS code (HYDRA) is employed and validated against the experimental data from a high speed linear cascade testing rig. The numerical analysis reveals that the wall-gas temperature ratio could greatly affect the transonic OTL flow field and there is a strong two-way coupling between aerodynamics and heat transfer. The feedbacks of the thermal boundary condition to aerodynamics behave differently at different flow regimes over the tip, clearly indicating a highly localized dependence of the convective heat transfer coefficient (HTC) upon wall temperatures. This implies that to use HTC for blade metal temperature predictions without resorting a fully conjugate solution, the temperature dependence needs to be corrected locally. A nonlinear correction approach has been adopted in the present work, and the results demonstrate its effectiveness for the transonic turbine tip configurations studied.

Dependence of External Heat Transfer Coefficient and Aerodynamics on Wall Temperature for 3-D Turbine Blade Passage

2014 ◽  
Author(s):  
R. Maffulli ◽  
L. He
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Wall Temperature ◽  
Convective Heat Transfer Coefficient ◽  
Secondary Flows ◽  
Strong Impact ◽  
Blade Passage

The dependency of convective heat transfer coefficient (HTC) on wall temperature has been recognized in some previous works but existing corrections are confined to either empirically based correlations or based on a boundary layer approach. A recent study by the present authors on a 2D configuration highlights upstream flow history has a strong impact on HTC for a non-adiabatic blade surface, and such an effect cannot be adequately corrected by the use of existing empirical correlations. A boundary layer based approach may be used in a 2D case for the correction as attempted previously. However, it is strongly argued that a boundary layer based method would become very difficult, if not impossible, to apply for complex 3D flows as those in endwall and secondary flow regions of a turbine blade passage. The present work is aimed to examine how the HTC and main 3D passage aerodynamic features of interest may be affected by the wall temperature condition. A systematic computational study has been firstly carried out for a 3D NGV passage. The impacts of wall temperature on the secondary flows, trailing edge shock waves and the passage flow capacity are discussed, underlining the connection and interactions between the wall temperature and the external aerodynamics of the 3D passage. The local errors in HTC in these 3D flow regions can be as high as 30–40% if the wall temperature dependence is not corrected. The effort is then directed to a new 3-point non-linear correction method. The benefit of the 3-point method in reducing errors in HTC is clearly demonstrated. A further study illustrates that the new method also offers much enhanced robustness in the HTC procedure, particularly relevant when the wall thermal condition is shown to influence the laminar-turbulent transition as exhibited by two well-established transition models adopted in the present work.

scholarly journals A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

1994 ◽  
Author(s):  
Y. W. Kim ◽  
W. Abdel-Messeh ◽  
J. P. Downs ◽  
F. O. Soechting ◽  
G. D. Steuber ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Secondary Flows ◽  
Test Surface ◽  
Leakage Flow ◽  
Test Technique ◽  
Coolant Injection ◽  
Blade Tip

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger who has made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work that has been performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow which is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

scholarly journals CFD Simulation of Convective Heat Transfer on Vernacular Sustainable Architecture: Validation and Application of Methodology

2019 ◽  
Vol 11 (15) ◽  
pp. 4231
Author(s):  
Wenzhou Zhong ◽  
Tong Zhang ◽  
Tetsuro Tamura
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Flow Field ◽  
Convective Heat Transfer ◽  
Convective Heat ◽  
Convective Heat Transfer Coefficient ◽  
Vernacular Architecture ◽  
Building Design ◽  
Building Model

The global background of energy shortages and climate deterioration demands bioclimatic sustainable buildings. Vernacular architecture can provide a useful resource of passive strategies and techniques for creating inner comfort conditions with minimum heating, ventilation, and air conditioning (HVAC) assistance. The identification and verification of such knowledge are essential for climate responsive or energy passive building design. Among the methods, computational fluid dynamics (CFD) is a useful tool for simulating convective heat transfer of vernacular architecture and predicting the convective heat transfer coefficient (CHTC) and flow field. Geometric complexity and diversity of building samples are crucial in the development of an effective simulation methodology in terms of computational cost and accuracy. Therefore, this paper presents high-resolution 3D steady Reynolds-averaged Navier–Stokes (RANS) CFD simulations of convective heat transfer on Japanese vernacular architecture, namely, “machiya.” A CFD validation study on the CHTC is performed based on wind-tunnel experiments on a cube heated by constant heat flux and placed in a turbulent channel flow with a Reynolds number of 3.3 × 104. Three steady RANS models and two boundary layer modeling approaches are compared and discussed. Results show that the SST k-ω model applied with low Reynolds number modeling approach is suitable for CHTC simulations on a simplified building model. The RNG k-ε model applied with wall functions is an appropriate choice for simulating flow field of a complicated building model. Overall, this study develops a methodology involving RANS model selection, boundary layer modeling, and target model fitting to predict the convective heat transfer on vernacular architecture.

scholarly journals Ramp Heating in High-Speed Transient Thermal Measurement With Reduced Uncertainty

2015 ◽  
Author(s):  
H. Ma ◽  
Z. Wang ◽  
L. Wang ◽  
Q. Zhang ◽  
Z. Yang ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
Convective Heat Transfer Coefficient ◽  
Thermal Measurement ◽  
Design Guideline ◽  
Heating Method ◽  
Speed Flow ◽  
Step Heating ◽  
Blade Tip ◽  
Transient Thermal

The uncertainty associated with the convective heat transfer coefficient (HTC) obtained in transient thermal measurement is often high, especially in high speed flow. The present study demonstrates that the experimental accuracy could be much improved by an actively controlled ramp heating instead of the conventional step heating approach. A general design guideline for the proposed ramp heating method is derived theoretically and further demonstrated by simulation cases. This paper also presents a detailed experimental study for transonic turbine blade tip heat transfer. Repeatable, high-resolution tip HTC contour was obtained through transient IR measurement with the proposed ramp heating method. Detailed uncertainty analysis shows that the resulting HTC uncertainty level is much lower than the experimental data currently available in the open literature. The ramp heating approach is specially recommended to the high-speed heat transfer experimental research community to improve the accuracy of the transient thermal measurement technique.

Darryl E. Metzger Memorial Session Paper: A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

1995 ◽  
Vol 117 (1) ◽  
pp. 1-11 ◽  
Author(s):  
Y. W. Kim ◽  
J. P. Downs ◽  
F. O. Soechting ◽  
W. Abdel-Messeh ◽  
G. D. Steuber ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Secondary Flows ◽  
Test Surface ◽  
Leakage Flow ◽  
Test Technique ◽  
Coolant Injection ◽  
Blade Tip

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger, who made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work, which was performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow, which is related strongly to the airfoil pressure loading distribution, and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

Experimental Characterization of the Vane Heat Flux Under Pulsating Trailing-Edge Blowing

2016 ◽  
Author(s):  
J. Saavedra ◽  
G. Paniagua ◽  
B. H. Saracoglu
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Shock Wave ◽  
Heat Flux ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Convective Heat Transfer Coefficient ◽  
Trailing Edge ◽  
Power Extraction ◽  
The Impact

The steady improvement of aircraft engine performance has led towards more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 and 6 million). The steady and unsteady heat flux was retrieved using thin-film 2-layered gauges. The complexity of the tests required the development of an original heat transfer post-processing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant.

scholarly journals Experimental and Numerical Investigation of Optimized Blade Tip Shapes: Part I — Turbine Rainbow Rotor Testing and CFD Methods

2018 ◽  
Author(s):  
Bogdan Cernat ◽  
Marek Pátý ◽  
Cis De Maesschalck ◽  
Sergio Lavagnoli
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
High Performance ◽  
Numerical Data ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Design Strategies ◽  
Flow Physics ◽  
Leakage Flows ◽  
Blade Tip

Blade tip design and tip leakage flows are crucial aspects for the development of modern aero-engines. The inevitable clearance between stationary and rotating parts in turbine stages generates high-enthalpy unsteady leakage flows that strongly reduce the engine efficiency and can cause thermally induced blade failures. An improved understanding of the tip flow physics is essential to refine the current design strategies and achieve increased turbine aerothermal performance. However, while past studies have mainly focused on conventional tip shapes (flat tip or squealer geometries), the open literature suffers from a shortage of experimental and numerical data on advanced blade tip configurations of unshrouded rotors. This work presents a complete numerical and experimental investigation on the unsteady flow field of a high-pressure turbine, adopting three different blade tip profiles. The aerothermal characteristics of two novel high-performance tip geometries, one with a fully contoured shape and the other presenting a multi-cavity squealer-like tip with partially open external rims, are compared against the baseline performance of a regular squealer geometry. The turbine stage is tested at engine-representative conditions in the high-speed turbine facility of the von Karman Institute. A rainbow rotor is mounted for simultaneous aerothermal testing of multiple blade tip geometries. On the rotor disk, the blades are arranged in sectors operating at two different clearance levels. A numerical campaign of full-stage simulations was also conducted on all the investigated tip designs to model the secondary flows development and identify the tip loss and heat transfer mechanisms. In the first part of this work, we describe the experimental setup, instrumentation and data processing techniques used to measure the unsteady aerothermal field of multiple blade tip geometries using the rainbow rotor approach. We report the time-average and time-resolved static pressure and heat transfer measured on the shroud of the turbine rotor. The experimental data are compared against CFD predictions. These numerical results are then used in the second part of the paper to analyze the tip flow physics, model the tip loss mechanisms and quantify the aero-thermal performance of each tip geometry.

scholarly journals Complementary Velocity and Heat Transfer Measurements in a Rotating Cooling Passage With Smooth Walls

1998 ◽  
Author(s):  
Jeffrey P. Bons ◽  
Jack L. Kerrebrock
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Nusselt Number ◽  
Turbine Blade ◽  
Secondary Flows ◽  
Integral Model ◽  
Velocity Data ◽  
Transfer Data ◽  
Cooling Passage ◽  
Heated Case

An experimental investigation was conducted on the internal flowfield of a simulated smooth-wall turbine blade cooling passage. The square cross-sectioned passage was manufactured from quartz for optical accessibility. Velocity measurements were taken using Particle Image Velocimetry for both heated and non-heated cases. Thin film resistive heaters on all four exterior walls of the passage allowed heat to be added to the coolant flow without obstructing laser access. Under the same conditions, an infrared detector with associated optics collected wall temperature data for use in calculating local Nusselt number. The test section was operated with radial outward flow and at values of Reynolds number and Rotation number typical of a small turbine blade. The density ratio was 0.27. Velocity data for the non-heated case document the evolution of the coriolis-induced double vortex. The vortex has the effect of disproportionately increasing the leading side boundary layer thickness. Also, the streamwise component of the coriolis acceleration creates a considerably thinned side wall boundary layer. Additionally, these data reveal a highly unsteady, turbulent flowfield in the cooling passage. Velocity data for the heated case show a strongly distorted streamwise profile indicative of a buoyancy effect on the leading side. The coriolis vortex is the mechanism for the accumulation of stagnant flow on the leading side of the passage. Heat transfer data show a maximum factor of two difference in the Nusselt number from trailing side to leading side. A first-order estimate of this heat transfer disparity based on the measured boundary layer edge velocity yields approximately the same factor of two. A momentum integral model was developed for data interpretation which accounts for coriolis and buoyancy effects. Calculated streamwise profiles and secondary flows match the experimental data well. The model, the velocity data, and the heat transfer data combine to strongly suggest the presence of separated flow on the leading wall starting at about five hydraulic diameters from the channel inlet for the conditions studied.

Experimental and Numerical Investigation of Optimized Blade Tip Shapes—Part I: Turbine Rainbow Rotor Testing and Numerical Methods

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
Bogdan C. Cernat ◽  
Marek Pátý ◽  
Cis De Maesschalck ◽  
Sergio Lavagnoli
Keyword(s):  
Heat Transfer ◽  
High Speed ◽  
High Performance ◽  
Numerical Data ◽  
Secondary Flows ◽  
Design Strategies ◽  
Flow Physics ◽  
Leakage Flows ◽  
Numerical Predictions ◽  
Blade Tip

Blade tip design and tip leakage flows are crucial aspects for the development of modern aero-engines. The inevitable clearance between stationary and rotating parts in turbine stages generates high-enthalpy unsteady leakage flows that strongly reduce the engine efficiency and can cause thermally induced blade failures. An improved understanding of the tip flow physics is essential to refine the current design strategies and achieve increased turbine aerothermal performance. However, while past studies have mainly focused on conventional tip shapes (flat tip or squealer geometries), the open literature suffers from a shortage of experimental and numerical data on advanced blade tip configurations of unshrouded rotors. This work presents a complete numerical and experimental investigation on the unsteady flow field of a high-pressure turbine, adopting three different blade tip profiles. The aerothermal characteristics of two novel high-performance tip geometries, one with a fully contoured shape and the other presenting a multicavity squealer-like tip with partially open external rims, are compared against the baseline performance of a regular squealer geometry. The turbine stage is tested at engine-representative conditions in the high-speed turbine facility of the von Karman Institute. A rainbow rotor is mounted for simultaneous aerothermal testing of multiple blade tip geometries. On the rotor disk, the blades are arranged in sectors operating at two different clearance levels. A numerical campaign of full-stage simulations was also conducted on all the investigated tip designs to model the secondary flows development and identify the tip loss and heat transfer mechanisms. In the first part of this work, we describe the experimental setup, instrumentation, and data processing techniques used to measure the unsteady aerothermal field of multiple blade tip geometries using the rainbow rotor approach. We report the time-average and time-resolved static pressure and heat transfer measured on the shroud of the turbine rotor. The experimental data are compared against numerical predictions. These numerical results are then used in the second part of the paper to analyze the tip flow physics, model the tip loss mechanisms, and quantify the aero-thermal performance of each tip geometry.

Sign in / Sign up

Export Citation Format

Share Document