Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes

2021 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Advanced Technique ◽  
Cooling Effectiveness ◽  
Uniform Pattern ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Heat Transfer Enhancement for Gas Turbine Blade Leading Edge Cooling Using Curved Double-Wall/Vortex Cooling With Various Disturbing Objects

2019 ◽  
Author(s):  
Xiaojun Fan ◽  
Liang Li ◽  
Jiefeng Wang ◽  
Fan Wu
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Wall Impingement ◽  
Pin Fins ◽  
Cooling Passage ◽  
Double Wall

Abstract A new double-wall cooling configuration combined with the vortex cooling is established to study the cooling behavior for the gas turbine blade leading edge. This configuration consists of multiple nozzles, a curved inner cooling passage, a row of bridge holes and a curved outer cooling passage with 4 kinds of disturbing objects (namely smooth wall, pin-fins, dimples and protrusions). Numerical simulations are performed based on the 3D viscous steady Reynolds Averaged Navier-Stokes (RANS) equations and the k-ω turbulence model. The cooling behavior of the Double-wall/vortex cooling configuration is compared with the Double-wall/impingement cooling configuration at the same conditions. Generally, the Double-wall/vortex cooling configuration has a better cooling performance. It is found the Nusselt number of the inner surface for the Double-wall/vortex cooling configuration is 46.7% higher. However, the Double-wall/impingement cooling configuration has a smaller friction coefficient and a total pressure loss. Different disturbing objects have significant influences on the heat transfer performance of the outer surface. The Nusselt number of disturbing objects (pin-fins, dimples and protrusions) is much higher than the smooth wall, and the value is 1.27–2.22 times larger. Configuration with protrusions has the highest globally-averaged Nusselt number. For the heat transfer performance of the inner surface and the total pressure loss coefficient, disturbing objects have no obvious influence. As bridge holes row increases, the overall cooling performance is improved. The globally-averaged Nusselt number of the outer target is enhanced while the total pressure loss is reduced.

Effect of Non-Axisymmetric Endwall Profiling on Heat Transfer and Film Cooling Effectiveness of a Transonic Rotor Blade

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Jinjin Li ◽  
Xin Yan ◽  
Kun He
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Rotor Blade ◽  
Stanton Number ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Transonic Rotor

Abstract Effects of non-axisymmetric endwall profiling on total pressure loss, heat transfer, and film cooling effectiveness of a transonic rotor blade were numerically investigated. The numerical methods, including the turbulence model and grid sensitivity, were validated with the existing experimental data. To reduce the thermal load on endwall, non-axisymmetric endwall profiling near leading edge and at pressure-side corner area was performed with a range of contour amplitudes. Heat transfer and flow fields near the profiled endwalls were analyzed and also compared with the plain endwall configuration. On the profiled endwall, three kinds of cooling holes, i.e., cylindrical holes, rounded-rectangular holes, and elliptical holes, were arranged, and film cooling effect was investigated at three blowing ratios. Results indicate that, with endwall profiling, the area-averaged Stanton number on endwall is reduced by 7.71% and total pressure loss in cascade is reduced by 11.07%. Among three kinds of cooling holes, the arrangement of the elliptical hole performs the best film cooling effect on the profiled endwall. Compared with the plain endwall, non-axisymmetric endwall with elliptical cooling holes improves film cooling coverage by 10.87%, reduces the Stanton number by 8.88%, and increases the net heat flux reduction performance by 4% at M = 0.7.

Effect of Non-Axisymmetric Endwall Profiling on Heat Transfer and Film Cooling Effectiveness of a Transonic Rotor Blade

2019 ◽  
Author(s):  
Jinjin Li ◽  
Xin Yan ◽  
Kun He
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Rotor Blade ◽  
Stanton Number ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Transonic Rotor

Abstract Effects of non-axisymmetric endwall profiling on total pressure loss, heat transfer and film cooling effectiveness of a transonic rotor blade were numerically investigated. The numerical methods, including the turbulence model and grid sensitivity, were validated with the existing experimental data. To reduce thermal load on endwall, non-axisymmetric endwall profiling near leading edge and at pressure-side corner area were performed with a range of contour amplitudes. Heat transfer and flow fields near the profiled endwalls were analyzed and also compared to the plain endwall configuration. On the profiled endwall, three kinds of cooling-holes, i.e. cylindrical holes, rounded-rectangular holes and elliptical holes, were arranged, and film cooling effect was investigated at three blowing ratios. Results indicate that, with endwall profiling, the area-averaged Stanton number on endwall is reduced by 7.71% and total pressure loss in cascade is reduced by 11.07%. Among three kinds of cooling holes, arrangement of elliptical hole performs the best film cooling effect on profiled endwall. Compared with plain endwall, non-axisymmetric endwall with elliptical cooling holes improves film cooling coverage by 10.87%, reduces the Stanton number by 8.88% and increases the net heat flux reduction performance by 4% at M = 0.7.

Investigations of Heat Transfer and Film Cooling Effect on a Worn Squealer Tip

2020 ◽  
Author(s):  
Mingliang Ye ◽  
Xin Yan
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Wear Depth ◽  
Original Design ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cavity Floor

Abstract Wear damage commonly occurs in modern gas turbine rotor blade tip due to relative movements and expansions between rotating and stationary parts. Tip wear has a significant impact on the aerodynamic, heat transfer and cooling performance of rotor blades, thus threatening the economy and safety of whole gas turbine system. Based on a simple linear wear model, this paper numerically investigates the aerodynamic, heat transfer and film cooling performance of a worn squealer tip with three starting-locations of wear (sl = 25%Cax, 50%Cax and 75%Cax) and five wear-depths (wd = 0.82%, 1.64%, 2.46%, 3.28% and 4.10%). Firstly, based on the existing experimental data, numerical methods and grid independence are examined carefully. Then, three dimensional flow fields, total pressure loss distributions, heat transfer coefficients and film cooling effectiveness in worn squealer tip region are computed, which are compared with the original design case. The results show that, with the increase of wear depth and the movement of wear starting-location to the leading edge, the scale and intensity of cavity vortex are increased, which results in the extended high heat transfer area on cavity floor near the leading edge. Wear makes more coolant flow out of the cavity, and reduces the area-averaged film cooling effectiveness at the bottom of cavity, but increases the film cooling effectiveness on pressure-side rim. The increase of wear depth makes more flow leak through the tip gap, thus increasing the scale and intensity of leakage vortex and further increasing the total pressure loss in the tip gap. Compared with the original design case, as the wear depth is increased from 0.82% to 4.10%, the mass-averaged total pressure loss in cascade is increased by 0.3–6.7%, the area-averaged heat transfer coefficient on cavity floor is increased by 1.7–29.1% while on squealer rim it is decreased by 3.1–26.3%, and the area-averaged film cooling effectiveness on cavity floor is decreased by 0.035 at most while on squealer rim it is increased by 0.064 at most.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

Conjugate Heat Transfer Characteristics of Double Wall Cooling on a Film Plate With Gradient Thickness

2020 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Weilun Zhou ◽  
Wei He ◽  
Tieyu Gao ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Conjugate Heat Transfer ◽  
Flow Direction ◽  
Plate Thickness ◽  
Heat Transfer Characteristics ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Impingement Cooling ◽  
Blowing Ratio ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is an advanced cooling method of gas turbines. In this paper, the flow and conjugate heat transfer characteristics of double wall cooling which has a film plate with gradient thickness are analyzed numerically. The detailed overall cooling effectiveness distributions are obtained by solving steady three dimensional Reynolds-averaged Navier-Stokes equations. In the double wall cooling scheme, seven vertical film holes and six impingement holes are staggered with same diameter (D), and the hole pitch of them are both set to 6D in flow direction and lateral direction. The gradient thickness along the flow direction is realized by setting the angle (α) between the lower surface of the film plate and the horizontal plane at −1.5 deg and 1.5 deg respectively. By comparing the results of four broadly used turbulence models with experimental data, SST k-ω is selected as the optimal turbulence model for double wall cooling analysis in this paper. In addition, the number of grids are finally determined to be 5.2 million by grid sensitivity calculation. The influence of the thickness gradient on the overall cooling effectiveness is revealed by comparing with the constant thickness film plate (Baseline 1 and 2), and all the cases are performed under four various coolant mass flow rates, which correspond to blowing ratios ranging from 0.25 to 1.5. The calculated results show that the thickening of the film plate downstream is beneficial to improve overall cooling effectiveness at low blowing ratio, which is benefit from two aspects. One is the thicken film plate weakens the flow separation in film hole and velocity of film hole outlet, another is the thicken film plate makes the impingement channels convergence, and impingement cooling is strengthened to some extent. However, with the increase of blowing ratio, the increasing trend gradually weakens due to the jet-off and limited impinge ability. For thickening film plate, the variations of the double wall cooling configurations are considered at initial film plate thickness tf of 2D and 3D, it is found that the ability to improve the overall cooling effectiveness by thickening the film plate downstream decrease as the initial film plate thickness increases, which is due to the increase of heat transfer resistance, and another finding is the cooling effectiveness of downstream thickening film plate with initial thickness of 2D is higher than that of 3D, which will provide a theoretical foundation both for improving cooling performance and reducing turbine blade weight at the same time. The influence of initial impingement gap H is also observed, and the study come to the fact that the best cooling performance occurred in H = 2D.

The Influence of Aero-Derivative Combustor Turbulence and Reynolds Number on Vane Aerodynamics Losses, Secondary Flows, and Wake Growth

2006 ◽  
Author(s):  
F. E. Ames ◽  
J. D. Johnson ◽  
N. J. Fiala
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Reynolds Numbers ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Boundary ◽  
High Turbulence ◽  
Turbulence Condition

Exit surveys detailing total pressure loss, turning angle, and secondary velocities have been acquired for a fully loaded vane profile in a large scale low speed cascade facility. Exit surveys have been taken over a four-to-one range in Reynolds numbers based on exit conditions and for both a low turbulence condition and a high turbulence condition. The high turbulence condition was generated using a mock aero-derivative combustor. Exit loss, angle, and secondary velocity measurements were acquired in the facility using a five-hole cone probe at two stations representing axial chord spacings of 0.25 and 0.50. Substantial differences in the level of losses, distribution of losses, and secondary flow vectors are seen with the different turbulence conditions and at the different Reynolds numbers. The higher turbulence condition produces a significantly broader wake than the low turbulence case and shows a measurable total pressure loss in the region outside the wakes. Generally, total pressure losses are about 0.02 greater for the high turbulence case compared with the low turbulence case primarily due to the state of the suction surface boundary layers. Losses decrease moderately with increasing Reynolds number. Cascade inlet velocity distributions have been previously documented in an endwall heat transfer study of this same geometry. These exit survey measurements support our understanding of the endwall heat transfer distributions, the secondary flows in the passage, and the origin of losses.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

2009 ◽  
Vol 131 (4) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Film Cooling Holes

This paper is focused on the effect of film-hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film-cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film-cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film-cooling holes are divided into four rows and loosely distributed on the platform. Four film-cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete-hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film-cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

2008 ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Film Cooling Holes

This paper is focused on the effect of film hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film cooling holes are divided into four rows and loosely distributed on the platform. Four film cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film cooling effectiveness distributions on the platform was obtained using pressure sensitive paint (PSP) technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

Elliptical Pin Fins as an Alternative to Circular Pin Fins for Gas Turbine Blade Cooling Applications: Part 1 — Endwall Heat Transfer and Total Pressure Loss Characteristics

2001 ◽  
Author(s):  
Oğuz Uzol ◽  
Cengiz Camci
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Major Axis ◽  
Total Pressure Loss ◽  
Turbine Blade Cooling ◽  
Pin Fin ◽  
Blade Cooling ◽  
Pin Fins ◽  
Pin Fin Arrays

Detailed experimental investigation of the wall heat transfer enhancement and total pressure loss characteristics for two alternative elliptical pin fin arrays is conducted and the results are compared to the conventional circular pin fin arrays. Two different elliptical pin fin geometries with different major axis lengths are tested, both having a minor axis length equal to the circular fin diameter and positioned at zero degrees angle of attack to the free stream flow. The major axis lengths for the two elliptical fins are 1.67 and 2.5 times the circular fin diameter, respectively. The pin fin arrays with H/D = 1.5 are positioned in a staggered 2 row configuration with 3 fins in the first row and 2 fins in the second row with S/D = X/D = 2. Endwall heat transfer and total pressure loss measurements are performed two diameter downstream of the pin fin arrays (X/D = 2) in a rectangular cross-section tunnel with an aspect ratio of 4.8 and for varying Reynolds numbers between 10000 and 47000 based on the inlet velocity and the fin diameter. Liquid Crystal Thermography is used for the measurement of convective heat transfer coefficient distributions on the endwall inside the wake. The results show that the wall heat transfer enhancement capability of the circular pin fin array is about 25–30% higher than the elliptical pin fin arrays in average. However in terms of total pressure loss, the circular pin fin arrays generate 100–200% more pressure loss than the elliptical pin fin arrays. This makes the elliptical fin arrays very promising cooling devices as an alternative to conventional circular pin fin arrays used in gas turbine blade cooling applications.

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