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.

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.

Account of Film Turbulence for Predicting Film Cooling Effectiveness in Gas Turbine Combustors

1979 ◽  
Author(s):  
G. J. Sturgess
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Wide Range ◽  
Engine Combustion ◽  
Turbulence Effects ◽  
The Impact ◽  
Combustor Liner

The paper deals with a small but important part of the overall gas turbine engine combustion system and continues earlier published work on turbulence effects in film cooling to cover the case of film turbulence. Film cooling of the gas turbine combustor liner imposes certain geometric limitations on the coolant injection device. The impact of practical film injection geometry on the cooling is one of increased rates of film decay when compared to the performance from idealized injection geometries at similar injection conditions. It is important to combustor durability and life estimation to be able to predict accurately the performance obtainable from a given practical slot. The coolant film is modeled as three distinct regions, and the effects of injection slot geometry on the development of each region are described in terms of film turbulence intensity and initial circumferential non-uniformity of the injected coolant. The concept of the well-designed slot is introduced and film effectiveness is shown to be dependent on it. Only slots which can be described as well-designed are of interest in practical equipment design. A prediction procedure is provided for well-designed slots which describes growth of the film downstream of the first of the three film regions. Comparisons of predictions with measured data are made for several very different well-designed slots over a relatively wide range of injection conditions, and good agreement is shown.

Account of Film Turbulence for Predicting Film Cooling Effectiveness in Gas Turbine Combustors

1980 ◽  
Vol 102 (3) ◽  
pp. 524-534 ◽  
Author(s):  
G. J. Sturgess
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Wide Range ◽  
Engine Combustion ◽  
Turbulence Effects ◽  
The Impact ◽  
Combustor Liner

The paper deals with a small but important part of the overall gas turbine engine combustion system and continues earlier published work on turbulence effects in film cooling to cover the case of film turbulence. Film cooling of the gas turbine combustor liner imposes certain geometric limitations on the coolant injection device. The impact of practical film injection geometry on the cooling is one of increased rates of film decay when compared to the performance from idealized injection geometries at similar injection conditions. It is important to combustor durability and life estimation to be able to predict accurately the performance obtainable from a given practical slot. The coolant film is modeled as three distinct regions, and the effects of injection slot geometry on the development of each region are described in terms of film turbulence intensity and initial circumferential non-uniformity of the injected coolant. The concept of the well-designed slot is introduced and film effectiveness is shown to be dependent on it. Only slots which can be described as well-designed are of interest in practical equipment design. A prediction procedure is provided for well-designed slots which describes growth of the film downstream of the first of the three film regions. Comparisons of predictions with measured data are made for several very different well-designed slots over a relatively wide range of injection conditions, and good agreement is shown.

Heat Transfer Coefficient and Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Rotor Blade ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The detailed distributions of heat transfer coefficient and film cooling effectiveness on a gas turbine blade tip were measured using a hue detection based transient liquid crystal technique. Tests were performed on a five-bladed linear cascade with blow down facility. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. 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 97.7°. The overall pressure ratio was 1.32 and the inlet and exit Mach number were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. The blade model was equipped with a single row of film cooling holes at both the tip portion along the camber line and near the tip region of the pressure-side. All measurements were made at the three different tip gap clearances of 1%, 1.5%, and 2.5% of blade span and the three blowing ratios of 0.5, 1.0, and 2.0. Results showed that, in general, heat transfer coefficient and film effectiveness increased with increasing tip gap clearance. As blowing ratio increased, heat transfer coefficient decreased, while film effectiveness increased. Results also showed that adding pressure-side coolant injection would further decrease blade tip heat transfer coefficient but increase film effectiveness.

Gas Turbine Combustor Cooling by Augmented Backside Convection

1979 ◽  
Vol 101 (1) ◽  
pp. 109-115 ◽  
Author(s):  
D. M. Evans ◽  
M. L. Noble
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Exhaust Gas ◽  
Confined Space ◽  
Gas Turbine Combustor ◽  
Exhaust Gas Emission ◽  
Cooling Air ◽  
Cooling Methods

Traditionally, gas turbine combustor walls have been cooled by one or more of the various film cooling methods. The current motivation to control exhaust gas emission composition has led to the serious consideration of backside convection wall cooling, where the cooling air is introduced to the main gas stream not prior to the dilution zone. Due to the confined space and the severe nature of the wall cooling problem, it is essential to maximize the heat transfer/pumping power characteristic, which suggests an augmented convection technique. A particular heat transfer design of a combustor cooled by means of transverse rib turbulence promoters applied to the exterior wall of the annular spaces surrounding the primary and secondary zones is described. Analytical methods for designing such a cooling system are reviewed and a comparison between analytical and experimental results is presented.

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 and Numerical Study of Heat Transfer in a Gas Turbine Combustor Liner

2003 ◽  
Vol 125 (4) ◽  
pp. 994-1002 ◽  
Author(s):  
J. C. Bailey ◽  
J. Intile ◽  
T. F. Fric ◽  
A. K. Tolpadi ◽  
N. V. Nirmalan ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Numerical Study ◽  
Cross Flow ◽  
Turbulence Models ◽  
Flow Distribution ◽  
Transfer Coefficients ◽  
Gas Turbine Combustor ◽  
Heat Transfer Coefficients ◽  
Combustor Liner

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat-plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady-state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.

scholarly journals Conjugate Heat Transfer Investigation on Swirl-Film Cooling at the Leading Edge of a Gas Turbine Vane

Entropy ◽  
2019 ◽  
Vol 21 (10) ◽  
pp. 1007 ◽  
Author(s):  
Du ◽  
Mei ◽  
Zou ◽  
Jiang ◽  
Xie
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Leading Edge ◽  
Pressure Side ◽  
Cooling Efficiency ◽  
Cooling Effectiveness ◽  
Turbine Vane ◽  
Cooling Chamber

Numerical calculation of conjugate heat transfer was carried out to study the effect of combined film and swirl cooling at the leading edge of a gas turbine vane with a cooling chamber inside. Two cooling chambers (C1 and C2 cases) were specially designed to generate swirl in the chamber, which could enhance overall cooling effectiveness at the leading edge. A simple cooling chamber (C0 case) was designed as a baseline. The effects of different cooling chambers were studied. Compared with the C0 case, the cooling chamber in the C1 case consists of a front cavity and a back cavity and two cavities are connected by a passage on the pressure side to improve the overall cooling effectiveness of the vane. The area-averaged overall cooling effectiveness of the leading edge () was improved by approximately 57%. Based on the C1 case, the passage along the vane was divided into nine segments in the C2 case to enhance the cooling effectiveness at the leading edge, and was enhanced by 75% compared with that in the C0 case. Additionally, the cooling efficiency on the pressure side was improved significantly by using swirl-cooling chambers. Pressure loss in the C2 and C1 cases was larger than that in the C0 case.

Sign in / Sign up

Export Citation Format

Share Document