Influence of slashface leakage coupled with quasi-labyrinth seal technique on gas turbine endwall aerothermal performance and blade suction side surface phantom cooling

2020 ◽  
pp. 095765092093848
Author(s):  
Kaiyuan Zhang ◽  
Zhigang Li ◽  
Jun Li
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Side Surface ◽  
Suction Side ◽  
Labyrinth Seal ◽  
Surface Cooling ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Cooling Effects ◽  
Back Part

Quasi-labyrinth seal technique was innovatively adopted to control the slashface flow, and the effect of slashface leakage on endwall aerothermal performance and blade suction side surface phantom cooling was numerically investigated. Simulations of five different labyrinth seal lengths ( Ls) and three coolant momentum flux ratios ( I) were conducted by computationally solving the Reynolds-averaged Navier–Stokes equations coupled with the shear stress transport k-ω turbulence model. The results show that the slashface coolant can be accelerated and conveyed downstream inside the slashface carried by the ingested mainstream and passage vortex. By setting up the quasi-labyrinth seal composed of a series of aperture cavities along the slashface, the coolant downstream transportation can be largely weakened, and the labyrinth seal with various lengths all have the potential to move the endwall coolant coverage toward upstream. When I = 0.48, the coolant downstream migration is severest and it has the best coverage on the endwall with seal. The laterally averaged cooling effectiveness is elevated for the fore part endwall but decreased for the back part endwall after sealing, and the highest cooling effectiveness is increased by 5% for 0.6 Ls seal. The 0.6–1.0 Ls seal has the approximate effect and is better than 0.4 Ls seal, and 0.6 Ls seal has the best outcome considering the countering effects of structural strength and endwall cooling performance. For 0.6 Ls sealed case, the high heat transfer level caused by coolant attachment is enhanced by 3.5% when I = 0.48 but decreased by 3.3% and 4.4% when I = 0.96 and 1.42 compared with baseline case. The low heat transfer caused by horseshoe vortex separation is enhanced when I = 1.42 for the sealed case. The quasi-labyrinth seal mainly weakens the back part suction side phantom cooling performance, and the well phantom cooling region moves upstream with the increase of I. The averaged phantom cooling effectiveness is decreased by 0.56%, 0.3%, and 1.44% when I = 0.48, 0.96, and 1.42 for the sealed cases. The results provide the gas turbine designers a better insight into improved slashface leakage control as well as its detailed surface cooling effects.

Numerical Investigation of Gas Turbine Combustor Liner Film Cooling Slots

2018 ◽  
Author(s):  
Firat Kiyici ◽  
Ahmet Topal ◽  
Ender Hepkaya ◽  
Sinan Inanli
Keyword(s):  
Heat Transfer ◽  
Prandtl Number ◽  
Surface Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Turbulent Prandtl Number ◽  
Surface Cooling ◽  
Gas Turbine Combustor ◽  
Cooling Effectiveness ◽  
Combustor Liner

A numerical study, based on experimental work of Inanli et al. [1] is conducted to understand the heat transfer characteristics of film cooled test plates that represent the gas turbine combustor liner cooling system. Film cooling tests are conducted by six different slot geometries and they are scaled-up model of real combustor liner. Three different blowing ratios are applied to six different geometries and surface cooling effectiveness is determined for each test condition by measuring the surface temperature distribution. Effects of geometrical and flow parameters on cooling effectiveness are investigated. In this study, Conjugate Heat Transfer (CHT) simulations are performed with different turbulence models. Effect of the turbulent Prandtl Number is also investigated in terms of heat transfer distribution along the measurement surface. For this purpose, turbulent Prandtl number is calculated with a correlation as a function of local surface temperature gradient and its effect also compared with the constant turbulent Prandtl numbers. Good agreement is obtained with two-layered k–ϵ with modified Turbulent Prandtl number.

Influence of Coolant on Cooling Performance Sensitivity of Internally Convective Turbine Vane

2021 ◽  
Author(s):  
Thanapat Chotroongruang ◽  
Prasert Prapamonthon ◽  
Rungsimun Thongdee ◽  
Thanapat Thongmuenwaiyathon ◽  
Zhenxu Sun ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Internal Cooling ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Turbine Engine ◽  
Performance Sensitivity ◽  
Turbine Vane ◽  
Cooling Air

Abstract Based on the Brayton cycle for gas-turbine engines, the high thermal efficiency and power output of a gas-turbine engine can be obtainable when the gas-turbine engine operates at high turbine inlet temperatures. However, turbine components e.g., inlet guide vane, rotor blade, and stator vane request high cooling performance. Typically, internal cooling and film cooling are two effective techniques that are widely used to protect high thermal loads for the turbine components in a state-of-the-art gas turbine. Consequently, the high thermal efficiency and power output can be obtained, and the turbine lifespan can be prolonged, also. On top of that, a comprehensive understanding of flow and heat transfer phenomena in the turbine components is very important. As a result, both experiments and simulations have been used to improve the cooling performance of the turbine components. In fact, the cooling air used in the internal cooling and film cooling is partially extracted from the compressor. Therefore, variations in the cooling air affect the cooling performance of the turbine components directly. This paper presents a numerical study on the influence of the cooling air on cooling-performance sensitivity of an internally convective turbine vane, MARK II using the computational fluid dynamics (CFD)/conjugate heat transfer (CHT) with the SST k-ω turbulence model. Result comparisons are conducted in terms of pressure, temperature, and cooling effectiveness under the effects of the inlet temperature, mass flow rate, turbulence intensity, and flow direction of the cooling air. The cooling-performance sensitivity to the coolant parameters is shown through variations of local cooling effectiveness, and area and volume-weighted average cooling effectiveness.

Experimental Investigations of SYCEE Film Cooling Performance on a Plate and a Tested Vane of an F-Class Gas Turbine

2014 ◽  
Author(s):  
Chang Han ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Round Hole ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Cooling Technology ◽  
Compound Angle

Film cooling is widely used in modern gas turbines for the protection of the hot components against hot gases from the combustion process. Film cooling directly influences the thermal efficiency of the gas turbine, as the cooling gas is extracted from the compressor and mixed with the mainstream in the hot component. Huge efforts by industry as well as research organizations have been undertaken to improve the film cooling effectiveness. It can been concluded that there are two key points for the improvement of film cooling effectiveness, constraining the blow-off of cooling ejection and extending the lateral coverage of cooling gas. The paper presents a new cooling technology, which reaches high film-cooling effectiveness as a result of a well-designed cooling hole, named SYCEE film cooling technology (SFCT). Plate film cooling experiments of SYCEE tested by pressure sensitive paint (PSP) are carried out in this work, and traditional shape-hole are included as well for baselines. It is resulted that SFCT has a better film cooling performance than shape-hole in the same conditions, and the gap of the averaged film cooling effectiveness between them continuously enlarges as the blowing ratio increases. Furthermore, an application of SFCT on the first stage vane of an F-class gas turbine is studied as well. A two-dimension cascade has been employed to measure the cooling performance of SFCT using pressure sensitive paint (PSP) as well, and the tested vanes separately with round-hole and shape-hole are considered again for baselines. The different kinds of film holes separately locate on the pressure and suction side, while the showerhead in different cases are kept the same, arranged with round-holes. The cooling air is ejected at inclination angle 45° with compound-angle 90° in the showerhead and inclination angle 35°∼45° without compound-angle on the pressure side and suction side. The detailed local cooling effectiveness distributions as well as the span-averaged effectiveness over the vane surface are presented. As expected, the film cooling performance of round-hole is the worst due to the lift-off of the cooling ejection. SFCT has better film cooling performance than shape-hole on the pressure side, but the advantage decreases along the mainstream direction. However, the span-averaged film cooling effectiveness of SYCEE is similar with that of the shape-hole on the suction side. This may be due to enhanced impact of mainstream flow derived from the pressure gradient in the turbine passage, and consequently weakening the effect of film hole on the suction side.

Experimental Investigation of Louver Cooling Scheme on Gas Turbine Vane Pressure Side

2010 ◽  
Author(s):  
T. Elnady ◽  
W. Saleh ◽  
I. Hassan ◽  
L. Kadem ◽  
T. Lucas
Keyword(s):  
Heat Transfer ◽  
Experimental Investigation ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Side ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
The Heat Transfer Coefficient

An experimental investigation has been performed to measure the cooling performance of the louver scheme using a two-dimensional cascade simulating the scaled vane of a high-pressure gas turbine. Two rows of an axially oriented louver scheme are distributed in a stagger arrangement over the pressure side. The effect of hole location on the cooling performance is investigated for each row individually, then the row interaction is investigated for both rows. The temperature distribution on the vane is mapped using a transient Thermochromic Liquid Crystal (TLC) technique to obtain the local distributions of the heat transfer coefficient and film cooling effectiveness. The performance of the louver scheme for each case is compared with that of two similar rows with a standard cylindrical exit at 0.9 density ratio. The exit Reynolds number based on the true chord is 1.5E5 and exit Mach number is 0.23. The local distributions of the effectiveness and the heat transfer coefficient are presented at four different blowing ratios ranging from 1 to 2. The louver scheme shows a superior cooling effectiveness than that of the cylindrical holes at all blowing ratios in terms of protection and lateral coverage. The row location highly affects the cooling performance for both the louver and cylindrical scheme due to the local pressure change and the variation of the surface curvature.

Experimental Investigation of Louver Cooling Scheme on Gas Turbine Vane Suction Side

2011 ◽  
Author(s):  
T. Elnady ◽  
O. Hassan ◽  
I. Hassan ◽  
L. Kadem ◽  
T. Lucas
Keyword(s):  
Experimental Investigation ◽  
Gas Turbine ◽  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Thermochromic Liquid Crystal ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Cylindrical Holes

An experimental investigation has been performed to measure the film cooling performance of louver scheme over a scaled vane of high-pressure gas turbine using a two-dimensional cascade. Two rows of axially oriented louver scheme are used to cool the suction side and their performance is compared with two similar rows of standard cylindrical holes. The effect of hole location on the cooling performance is investigated for each row individually, then the row interaction is investigated for both rows at four different blowing ratios ranging from 1 to 2 with a 0.9 density ratio. The exit Reynolds number based on the true chord is 1.5E5 and exit Mach number is 0.23. The temperature distribution on the vane is mapped using a transient Thermochromic Liquid Crystal (TLC) technique to obtain the local distributions of the heat transfer coefficient and film cooling effectiveness. The louver scheme shows a superior cooling effectiveness than that of the cylindrical holes at all blowing ratios in terms of protection and lateral coverage. The row location highly affects the cooling performance for both the louver and cylindrical scheme.

The Effects of Obstructions on Film Cooling Effectiveness on the Suction Side of a Gas Turbine Vane

2006 ◽  
Author(s):  
Patricia Demling ◽  
David G. Bogard
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Temperature Profiles ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness

The effects of obstructions on film cooling performance on a scaled-up 1st stage turbine vane will be discussed. Experimental results show that obstructions located upstream or inside of a film cooling hole will degrade adiabatic effectiveness up to 80% of the levels found with no obstructions. Downstream obstructions had little effect on performance. The location where the upstream obstructions ceased to degrade adiabatic effectiveness was determined and temperature profiles were constructed to determine how the upstream obstructions were affecting the mainstream and coolant flow.

Predictions of the Effect of Roughness on Heat Transfer From Turbine Airfoils

1998 ◽  
Author(s):  
Anil K. Tolpadi ◽  
Michael E. Crawford
Keyword(s):  
Heat Transfer ◽  
Side Surface ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Surface Finishing ◽  
Transfer Coefficients ◽  
Average Roughness ◽  
Wide Range ◽  
Turbine Airfoils ◽  
Good Agreement

The heat transfer and aerodynamic performance of turbine airfoils are greatly influenced by the gas side surface finish. In order to operate at higher efficiencies and to have reduced cooling requirements, airfoil designs require better surface finishing processes to create smoother surfaces. In this paper, three different cast airfoils were analyzed: the first airfoil was grit blasted and codep coated, the second airfoil was tumbled and aluminide coated, and the third airfoil was polished further. Each of these airfoils had different levels of roughness. The TEXSTAN boundary layer code was used to make predictions of the heat transfer along both the pressure and suction sides of all three airfoils. These predictions have been compared to corresponding heat transfer data reported earlier by Abuaf et al. (1997). The data were obtained over a wide range of Reynolds numbers simulating typical aircraft engine conditions. A three-parameter full-cone based roughness model was implemented in TEXSTAN and used for the predictions. The three parameters were the centerline average roughness, the cone height and the cone-to-cone pitch. The heat transfer coefficient predictions indicated good agreement with the data over most Reynolds numbers and for all airfoils-both pressure and suction sides. The transition location on the pressure side was well predicted for all airfoils; on the suction side, transition was well predicted at the higher Reynolds numbers but was computed to be somewhat early at the lower Reynolds numbers. Also, at lower Reynolds numbers, the heat transfer coefficients were not in very good agreement with the data on the suction side.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Conjugate Heat Transfer Analysis of a High Pressure Air-Cooled Gas Turbine Vane

2010 ◽  
Author(s):  
Zhenfeng Wang ◽  
Peigang Yan ◽  
Hongyan Huang ◽  
Wanjin Han
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Boundary Conditions ◽  
Gas Turbine ◽  
Turbulence Model ◽  
Conjugate Heat Transfer ◽  
Suction Side ◽  
Turbulence Characteristic ◽  
Transition Flow ◽  
Characteristic Boundary

The ANSYS-CFX software is used to simulate NASA-Mark II high pressure air-cooled gas turbine. The work condition is Run 5411 which have transition flow characteristics. The different turbulence models are adopted to solve conjugate heat transfer problem of this three-dimensional turbine blade. Comparing to the experimental results, k-ω-SST-γ-θ turbulence model results are more accurate and can simulate accurately the flow and heat transfer characteristics of turbine with transition flow characteristics. But k-ω-SST-γ-θ turbulence model overestimates the turbulence kinetic energy of blade local region and makes the heat transfer coefficient higher. It causes that local region temperature of suction side is higher. In this paper, the compiled code adopts the B-L algebra model and simulates the same computation model. The results show that the results of B-L model are accurate besides it has 4% temperature error in the suction side transition region. In addition, different turbulence characteristic boundary conditions of turbine inner-cooling passages are given and K-ω-SST-γ-θ turbulence model is adopted in order to obtain the effect of turbulence characteristic boundary conditions for the conjugate heat transfer computation results. The results show that the turbulence characteristic boundary conditions of turbine inner-cooling passages have a great effect on the conjugate heat transfer results of high pressure gas turbine. ANSYS is applied to analysis the thermal stress of Mark II blade which has ten radial cooled passages and the results of Von Mises stress show that the temperature gradient results have a great effect on the results of blade thermal stress.

Conjugate Heat Transfer Simulations to Evaluate the Effect of Anisotropic Thermal Conductivity on Overall Cooling Effectiveness

2021 ◽  
pp. 1-26
Author(s):  
Carol E. Bryant ◽  
James L. Rutledge
Keyword(s):  
Heat Transfer ◽  
Thermal Conductivity ◽  
Gas Turbine ◽  
Conjugate Heat Transfer ◽  
Ceramic Matrix Composites ◽  
Conductivity Anisotropy ◽  
Cooling Effectiveness ◽  
External Cooling ◽  
Anisotropic Thermal Conductivity ◽  
Overall Effectiveness

Abstract Ceramic matrix composites (CMCs) show promise as higher temperature capable alternatives to traditional metallic components in gas turbine engine hot gas paths. However, CMC components will still require both internal and external cooling, such as film cooling. The overall cooling effectiveness is determined not only by the design of coolant flow, but also by the conduction through the materiel itself. CMCs have anisotropic thermal conductivity, giving rise to heat flow that differs somewhat relative to what we have come to expect from experience with traditional metallic components. Conjugate heat transfer computational fluid dynamics (CFD) simulations were performed in order to isolate the effect anisotropic thermal conductivity has on a cooling architecture consisting of both internal and external cooling. Results show the specific locations and unique effects of anisotropic thermal conduction on overall effectiveness. Thermal conductivity anisotropy is shown to have a significant effect on the resulting overall effectiveness. As CMCs begin to make their way into gas turbine engines, care must be taken to ensure that anisotropy is characterized properly and considered in the thermal analysis.

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