Heat Transfer and Flow Studies of Different Cooling Configurations for Gas Turbine Rotor Blade

2014 ◽  
Author(s):  
Batchu Suresh ◽  
Ainapur Brijesh ◽  
V. Kesavan ◽  
S. Kishore Kumar
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Metal Temperature ◽  
Coolant Temperature ◽  
Fe Model ◽  
Transfer Coefficients ◽  
Flow Network ◽  
Blade Cooling ◽  
Film Cooling Holes

Military gas turbine engine operates at turbine entry temperatures (TET) of the order of 2000K. Increase in TET improves thermal efficiency and power output. The gas temperature is far above the allowable metal temperature of turbine components. Hence, there is a need to cool the components such as blades and vanes for safe operation. The blades are cooled by combination of internal convective cooling and external film-cooling. Rib tabulators are widely used in blade cooling passages to enhance heat transfer. In the present study, different rib tabulator configurations have been studied. 1D flow network model of blade cooling passages have been modeled using Flowmaster software. Flowmaster software estimates pressure losses, rotational effects and heat transfer of the coolant flow in the blade passages. Cooling passages are modeled as ducts while film cooling holes, impingement holes, tip holes and ejection holes are modeled as orifices. Experimentally measured heat transfer and pressure loss correlations are used in the analysis. The coolant pressure at inlet and sink pressure at exit of film cooling holes are given as input. The heat load coming on to the blade is also given as input for predicting the coolant temperature rise and blade metal temperature. The thermal analysis is carried out with different shaped rib turbulators such as V and W ribs with broken and continuous pattern. It is observed that thermal performance factor for a broken V rib configuration is better than other configurations. The metal temperature for broken V ribbed configuration is 25°C less compared other configurations. The effect of rotation on the blade temperature is also studied. The convective bulk temperatures and convective heat transfer coefficients obtained from 1D flow network are applied on 2D Finite Element (FE) model to obtain nodal temperature distribution.

scholarly journals Comparison of Blade Cooling Performance Using Alternative Fluids

2002 ◽  
Author(s):  
Carlo Carcasci ◽  
Stefano Zecchi ◽  
Gianpaolo Oteri
Keyword(s):  
Heat Transfer ◽  
Carbon Dioxide ◽  
Gas Turbine ◽  
Film Cooling ◽  
Cooling System ◽  
Internal Heat ◽  
Transfer Coefficients ◽  
Cooling Systems ◽  
Heat Transfer Coefficients ◽  
Blade Cooling

CO2 emissions reduction has become an important topic, especially after Kyoto protocol. There are several ways to reduce the overall amount of CO2 discharged into the atmosphere, for example using alternative fluids such as steam or CO2. It is therefore interesting to analyze the consequences of their usage on overall performances of gas turbine and blade cooling systems. The presence of steam can be associated with combined or STIG cycle, whereas pure carbon dioxide or air-carbon dioxide mixtures are present in innovative cycles, where the exhaust gas is recirculated partially or even totally. In this paper we will analyze a commercial gas turbine, comparing different fluids used as working and cooling fluids. The different nature of the fluids involved determines different external heat transfer coefficients (external blade surface), different internal heat transfer coefficients (cooling cavities) and affects film cooling effectiveness, resulting in a change of the blade temperature distribution. Results show that the presence of steam and CO2 could determine a non negligible effect on blade temperature. This means that cooling systems need a deep investigation. A redesign of the cooling system could be required. In particular, results show that steam is well suited for internal cooling, whereas CO2 is better used in film cooling systems.

Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes, as well as elliptical, diffuser, and teardrop shaped dimples, simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of 2 increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2008 ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes as well as elliptical, diffuser, and teardrop shaped dimples simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of two increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

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.

Numerical Study on Flow and Heat Transfer of a Cooled First Stage Vane of a Gas Turbine

2016 ◽  
Author(s):  
Lv Ye ◽  
Zhao Liu ◽  
Xiangyu Wang ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Numerical Study ◽  
Coolant Flow ◽  
The Other ◽  
Edge Region ◽  
Flow Rates ◽  
Flow And Heat Transfer ◽  
Film Cooling Holes

This paper presents a numerical simulation of composite cooling on a first stage vane of a gas turbine, in which gas by fixed composition mixture is adopted. To investigate the flow and heat transfer characteristics, two internal chambers which contain multiple arrays of impingement holes are arranged in the vane, several arrays of pin-fins are arranged in the trailing edge region, and a few arrays of film cooling holes are arranged on the vane surfaces to form the cooling film. The coolant enters through the shroud inlet, and then divided into two parts. One part is transferred into the chamber in the leading edge region, and then after impinging on the target surfaces, it proceeds further to go through the film cooling holes distributed on the vane surface, while the other part enters into the second chamber immediately and then exits to the mainstream in two ways to effectively cool the other sections of the vane. In this study, five different coolant flow rates and six different inlet pressure ratios were investigated. All the cases were performed with the same domain grids and same boundary conditions. It can be concluded that for the internal surfaces, the heat transfer coefficient changes gradually with the coolant flow rate and the inlet total pressure ratio, while for the external surfaces, the average cooling effectiveness increases with the increase of coolant mass flow rates while decreases with the increase of the inlet stagnation pressure ratios within the study range.

Evaluation of Numerical Methods to Predict Temperature Distributions of an Experimentally Investigated Convection-Cooled Gas-Turbine Blade

2017 ◽  
Author(s):  
E. Findeisen ◽  
B. Woerz ◽  
M. Wieler ◽  
P. Jeschke ◽  
M. Rabs
Keyword(s):  
Heat Transfer ◽  
Numerical Methods ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Cooling System ◽  
Fe Model ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Temperature Distributions

This paper presents two different numerical methods to predict the thermal load of a convection-cooled gas-turbine blade under realistic operating temperature conditions. The subject of the investigation is a gas-turbine rotor blade equipped with an academic convection-cooling system and investigated at a cascade test-rig. It consists of three cooling channels, which are connected outside the blade, so allowing cooling air temperature measurements. Both methods use FE models to obtain the temperature distribution of the solid blade. The difference between these methods lies in the generation of the heat transfer coefficients along the cooling channel walls which serve as a boundary condition for the FE model. One method, referred to as the FEM1D method, uses empirical one-dimensional correlations known from the available literature. The other method, the FEM2D method, uses three-dimensional CFD simulations to obtain two-dimensional heat transfer coefficient distributions. The numerical results are compared to each other as well as to experimental data, so that the benefits and limitations of each method can be shown and validated. Overall, this paper provides an evaluation of the different methods which are used to predict temperature distributions in convection-cooled gas-turbines with regard to accuracy, numerical cost and the limitations of each method. The temperature profiles obtained in all methods generally show good agreement with the experiments. However, the more detailed methods produce more accurate results by causing higher numerical costs.

Heat Transfer Coefficient Measurements of Film-Cooling Holes With Expanded Exits

1998 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Film Cooling ◽  
Cylindrical Hole ◽  
Superposition Method ◽  
Transfer Coefficients ◽  
Cooling Performance ◽  
Heat Transfer Coefficients ◽  
Film Cooling Holes ◽  
Injection Location

Detailed measurements of heat transfer coefficients in the nearfield of three different film-cooling holes are presented. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fan-shaped and a laidback fanshaped hole). They were tested over a range of blowing ratios M = 0.25…1.75 at an external crossflow Mach number of 0.6 and a coolant-to-mainflow density ratio of 1.85. Additionally, the effect of the internal coolant supply Mach number is addressed. Temperatures of the diabatic surface downstream of the injection location are measured by means of an infrared camera system. They are used as boundary conditions for a finite element analysis to determine surface heat fluxes and heat transfer coefficients. The superposition method is applied to evaluate the overall film-cooling performance of the hole geometries investigated. As compared to the cylindrical hole, both expanded holes show significantly lower heat transfer coefficients downstream of the injection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the injected coolant than the fanshaped hole which leads to lower laterally averaged heat transfer coefficients. Coolant passage crossflow Mach number affects the flowfield of the jet being ejected from the hole and, therefore, has an important impact on film-cooling performance.

Comparison of Film Effectiveness and Cooling Uniformity of Conical and Cylindrical-Shaped Film Hole With Coolant-Exit Temperature Correction

2010 ◽  
Author(s):  
Cuong Q. Nguyen ◽  
Perry L. Johnson ◽  
Bryan C. Bernier ◽  
Son H. Ho ◽  
Jayanta S. Kapat
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Temperature Correction ◽  
Coolant Temperature ◽  
Transfer Model ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Film Cooling Holes

Data from conical-shaped film cooling holes is extremely sparse in open literature, especially the cooling uniformity characteristic, an important criterion for evaluating any film cooling design. The authors will compare the performance of conical-shaped holes to cylindrical-shaped holes. Cylindrical-shaped holes are often considered a baseline in terms of film cooling effectiveness and cooling uniformity coefficient. The authors will study two coupons with conical-shaped holes, which have 3° and 6° diffusion angles, named CON3 and CON6 respectively. A conjugate heat transfer computational fluid dynamics model and an experimental wind tunnel will be used to study these coupons. The three configurations: cylindrical baseline, CON3, and CON6, have a single row of holes with an inlet metering diameter of 3mm, length-to-nominal diameter of 4.3, and an injection angle of 30°. In this study, the authors will also take into account the heat transfer into the coolant flow from the coolant channel. In other words, coolant temperature at the exit of the coolant hole will be different than that measured at the inlet, and the conjugate heat transfer model will be used to correct for this difference. For the numerical model, the realizable k-ε turbulent model will be applied with a second order of discretization and enhanced wall treatment to provide the highest accuracy available. Grid independent studies for both cylindrical-shaped film cooling holes and conical-shaped holes will be performed and the results will be compared to data in open literature as well as in-house experimental data. Results show that conical-shaped holes considerably outperform cylindrical-shaped holes in film cooling effectiveness at all blowing ratios. In terms of cooling uniformity, conical-shaped holes perform better than cylindrical-shaped holes for low and mid-range blowing ratios, but not at higher levels.

scholarly journals Rotating Heat Transfer Experiments With Turbine Airfoil Internal Flow Passages

1986 ◽  
Author(s):  
Joel H. Wagner ◽  
Jay C. Kim ◽  
Bruce V. Johnson
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Large Scale ◽  
Flow Direction ◽  
Internal Flow ◽  
Operating Conditions ◽  
Coolant Temperature ◽  
Rossby Number ◽  
Transfer Coefficients ◽  
Bulk Fluid

Internal convective cooling is used to maintain acceptable gas turbine rotor blade temperatures. The heat transfer from the blade coolant passage walls is governed by forced convection, Coriolis forces and buoyance due to wall and coolant temperature differences. Currently little data is available to designers regarding the combined effects of these three parameters. To obtain required data, a rotating heat transfer facility was developed for experiments with large scale models and run at rotation and flow parameters typical of current gas turbine operating conditions. Analysis of the equations of motion showed that the perinent nondimensional parameters were Reynolds number, Rossby number, the difference in wall fluid and bulk fluid density and geometric ratios. The models were instrumented to measure average heat transfer rates on the coolant passage wall elements, and with pressure taps for friction data. An initial set of experiments have been conducted with rough wall geometries, typical of those used in blades. Results from the rotating experiments showed large heat transfer coefficient increases and decreases on the coolant passage leading and trailing surfaces compared to nonrotating heat transfer coefficients. The heat transfer was shown to be a function of inward or outward flow direction and Rossby number for the experiments conducted.

scholarly journals Full Coverage Discrete Hole Wall Cooling: Cooling Effectiveness

1984 ◽  
Author(s):  
G. E. Andrews ◽  
M. L. Gupta ◽  
M. C. Mkpadi
Keyword(s):  
Heat Transfer ◽  
Low Temperature ◽  
Film Cooling ◽  
Test Facility ◽  
Coolant Temperature ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Full Coverage ◽  
Hole Wall

The development of a test facility for investigating full coverage discrete hole wall cooling for gas turbine combustion chamber wall cooling is described. A low temperature test condition of 750K mainstream temperature and 300K coolant temperature was used to investigate the influence of coolant flow rate at a constant cross flow Mach number. Practical combustion conditions of 2100K combustor temperature and 700K coolant temperature are investigated to establish the validity of applying the low temperature results to practical conditions. For both situations a heat balance programme, taking into account the heat transfer within the wall was used to compute the film heat transfer coefficients. The mixing of the coolant air with the mainstream gases was studied through boundary layer temperature and CO2 profiles. It was shown that entrainment of hot flame gases between the injection holes resulted in a very low ‘adiabatic’ film cooling effectiveness.

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