Effect of Upstream Purge Slot on a Transonic Turbine Blade Passage: Part 1 — Aerodynamic Performance

2013 ◽  
Author(s):  
Dorian M. Blot ◽  
Arnab Roy ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Andrew S. Lohaus ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Incidence Angle ◽  
Flow Characteristics ◽  
Experimental Results ◽  
Backward Facing Step ◽  
Film Cooling Effectiveness

In this paper, detailed experimental results of total pressure loss and secondary flow field are presented for a high turning (127°) airfoil passage in presence of an upstream purge slot (with and without coolant injection). The experiments were performed at Virginia Tech’s quasi 2D linear turbine cascade operating at transonic conditions. Measurements were made at design exit Mach number 0.88 and design incidence angle. The selected coolant to mainstream mass flow ratio (MFR) was set at 1.0%. In order to match engine representative inlet/exit blade loading, a diverging endwall was utilized where the span increased from the inlet to the exit at a 13 degree angle. A 5-hole probe traverse was used to measure exit total pressure. Pressure loss coefficients were calculated both along pitchwise and spanwise directions at 0.1 axial chord downstream of the blade trailing edge. CFD studies were conducted to compliment the experimental results. The backward facing step present with the upstream slot affects the approaching boundary layer and influences the passage horse-shoe vortex strength. The addition of coolant from the purge slot further increased the aerodynamic losses. However, the backward facing step of the upstream slot seems to be the predominant factor in affecting pressure losses when compared to with or without blowing cases. These results provide further understanding of the passage secondary flow characteristics and aid towards improved design of endwall passages. The heat transfer experiments, designed to find the heat transfer coefficient and the film cooling effectiveness are described in detail in part II of this paper [1].

Numerical investigation of blowing ratio, density ratio and axial position of film holes on the vane endwall film cooling effectiveness with upstream step

2021 ◽  
pp. 095441002110165
Author(s):  
Qingzong Xu ◽  
Qiang Du ◽  
Pei Wang ◽  
Xiangtao Xiao ◽  
Jun Liu
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Density Ratio ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Endwall Film Cooling

The aerothermal performance of interrupted slot and film holes was numerically investigated. Previous study indicates that the interrupted slot performs better compared to the conventional slot. In the meanwhile, the step formed along with the interrupted slot affects the film cooling characteristics. In this article, a row of film holes is arranged downstream of the step, and the mass flow rate for the interrupted slot is constant at 1%. Blowing ratio (BR) from 0.5 to 1.5 and density ratio from 1 to 2 were studied for the film holes. Endwall film cooling effectiveness distribution indicates that film cooling is easily affected by the secondary flow inside passage and the upstream step. Coolant traces are split into two parts due to the effects of step vortex and transverse flow. For different density ratios, increasing BR shows a different trend of film cooling effectiveness due to the variation of coolant momentum. The coolant jet is easily affected by the secondary flow when its momentum is low, but tends to liftoff when its momentum is too high. As a result, it is better to position the film holes far away from the upstream step. The total pressure loss coefficient distribution at the passage exit indicates that the coolant injection increases the total pressure loss. But density ratio has smaller effect on the loss variation. Besides, two axial positions of cooling holes were studied to improve the endwall cooling performance. Without the effect of step vortex, the film effectiveness of cooling holes is improved.

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.

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.

Sweeping Jet Film Cooling on a Turbine Vane

2018 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Aerothermodynamics of a High-Pressure Turbine Blade With Very High Loading and Vortex Generators

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
Reinaldo A. Gomes ◽  
Reinhard Niehuis
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Flow Separation ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Vortex Generators ◽  
High Pressure Turbine ◽  
Very High

AITEB-2 is a project where aerothermal challenges of modern high pressure turbine designs are analyzed. One of the scopes of the project is to allow for new gas turbine designs with less parts and lighter jet engines by increasing the blade pitch and therefore the aerodynamic blade loading. For transonic profiles, this leads to very high velocities on the suction side and shock induced separation is likely to occur. The total pressure loss increase due to flow separation and strong shocks, as well as the underturning of the flow, limits the increase of the blade pitch. In this paper, experiments using a linear turbine blade cascade with high aerodynamic loading are presented. The blade pitch is increased such that at design conditions, a strong separation occurs on the suction side. The experiments were run at high subsonic exit Mach numbers and at Reynolds numbers of 390,000 and 800,000. In order to reduce the flow separation and the aerodynamic losses, air jet vortex generators are used, which create streamwise vortices prior to the separation start. Since in high pressure turbine blades film cooling is widely used, also the influence of film cooling both with and without using vortex generators is analyzed. Film cooling is provided on the suction side by two rows of cylindrical holes. This paper provides an analysis of the influence of different main flow conditions, film cooling, and vortex generators on total pressure loss, heat transfer and film cooling effectiveness. The experiments show that the vortex generators, as well as the film cooling reduce flow separation and total pressure losses. The effects are also seen in the local heat transfer, especially with enhanced heat transport in the region with flow separation. The cases presented in this paper deal with complex flow phenomena, which are challenging to be predicted with modern numerical tools correctly. Therefore, the experimental data serve as a comprehensive database for validation of simulation tools in the AITEB-2 project.

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