High Coverage Blade Tip Film Cooling

2004 ◽  
Author(s):  
M. Kuwabara ◽  
Keizo Tsukagoshi ◽  
T. Arts
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Separation Bubble ◽  
Density Ratio ◽  
Tip Clearance ◽  
Experimental Program ◽  
Film Cooling Effectiveness ◽  
Blade Tip

More sophisticated cooling schemes are required for the turbine blade due to the demand of increased turbine temperature for improved performance. Although the tip portion of a turbine blade is one of the most critical portions in a gas turbine, there are few studies on cooling this portion compared to those for airfoil, especially film cooling strategies. Industrial gas turbines have a more uniform gas temperature profile than aero engines. For these applications, it is more important to understand the characteristics of tip film cooling to improve the blade durability and gas turbine performance by reducing cooling air. A numerical and experimental program was initiated to study film cooling effectiveness on a flat blade tip as a function of tip gap and mass flux ratios. Flow visualization tests were conducted with and without film cooling to verify the numerical CFD findings. The predictions and visualization results showed that a separation bubble forms at the pressure side edge that increases with tip gap. Film effectiveness measurements were carried out on a 1.3X scale blade model in a low speed test while simulating the normalized pressure distribution typical of an engine design. The engine density ratio of the coolant to mainstream was replicated in the film cooling tests to provide the best simulation of the engine. Two rows of holes were placed near the tip of the blade to provide high film coverage prior to the flowing over the tip. The data shows that film effectiveness increases with decreasing tip clearance. Blowing ratio provides an improvement due to the added mass flow, which was shown by a non-dimensionalized correlation.

Film-Cooling Effectiveness on Squealer Cavity and Rim Walls of Gas-Turbine Blade Tip

2006 ◽  
Vol 22 (4) ◽  
pp. 898-899 ◽  
Author(s):  
Shantanu Mhetras ◽  
Huitao Yang ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbine Blade ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

Film Cooling Performance in a Simulated Turbine Blade Tip Geometry

2005 ◽  
Author(s):  
Jae H. Yoon ◽  
Ricardo F. Martinez-Botas
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Thermal Loading ◽  
Separation Bubble ◽  
Transfer Coefficients ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blade Tip

One of the most problematic areas in gas turbine engines is the blade tip region, especially near the trailing edge, where it is very difficult to provide sufficient cooling. In all configurations with unshrouded tips, a clearance gap exists between the turbine blade and the outer shroud. The pressure difference between the suction and pressure sides of the blade drives a sink-like flow through this gap. The combination of leakage flow from the freestream and coolant flow induces high convective heat transfer coefficients on the blade tip surface. The resultant thermal loading can be significant and detrimental to the turbine blade tip durability, leading to early failure. Film cooling can be provided by means of a series of holes located on the tip itself providing protection not dissimilar to film cooling of the main blade. However, the interaction of coolant and the separation bubble resulted in a significantly different film cooling performance to that of non-tip cases. An experimental investigation of the simulated turbine blade tip is presented in here. The first section discusses PIV flow field measurements, the second covers the measurement of film cooling effectiveness and the third heat transfer measurements. All three parts investigate the effect of using different film cooling injection points and blowing ratios for injection on the blade tip itself, close to the pressure surface corner. Additionally, the effect of varying the corner radii between the pressure surface and the tip is reported. The experimental method uses the steady state liquid crystal technique. A Reynolds number of 30,000 based on the clearance gap hydraulic diameter for the main flow was used.

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

2003 ◽  
Vol 125 (3) ◽  
pp. 494-502 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
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 crystals technique. Tests were performed on a five-bladed linear cascade with blow-down facility. 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 1.2 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%. 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 1.0%, 1.5%, and 2.5% of blade span and the three blowing ratios of 0.5, 1, and 2. 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 the blade tip heat transfer coefficient but increase film-cooling effectiveness.

scholarly journals The Model V84.3 Shop Tests: Tip Clearance Measurements and Evaluation

1994 ◽  
Author(s):  
M. Janssen ◽  
J. Seume ◽  
H. Zimmermann
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Cold Start ◽  
Tip Clearance ◽  
Test Facility ◽  
Experimental Program ◽  
Tip Leakage Flow ◽  
Hot Start ◽  
Blade Tip ◽  
Design Calculations

The design of high-performance gas turbines requires the reliable prediction of blade tip clearances. Excess clearances allow a portion of the hot gas to flow over the blade tips without performing useful work. The tip leakage flow disturbs the flow field which results in additional losses. Moreover, insufficient blade tip clearance may cause interference which can reduce turbine life. In conventional turbomachines, the blade tip clearances vary markedly with the operating condition of the turbine, essentially as a result of variations in gas temperatures and rotor speed. Siemens tests prototype gas turbines in its own test facility. An extensive experimental program is devised to verify design calculations regarding strength, aerodynamics and thermodynamics. Among other measurements, the minimum operating tip clearance is measured by abrasion pins. Electro-mechanical sensors measure transient tip clearance during a selected duty cycle consisting of turning-gear operation, cold start, idle operation, as well as part-load, full-load, and most importantly, hot-start. In the present paper, the compressor and turbine tip clearances measured during such a load cycle are compared with calculated predictions. The experimental instrumentation for the prototype gas turbine, as well as design calculations, are presented. The results show that the new Model V84.3 gas turbine does not exhibit critically small clearances during cold start nor during hot-start due to the careful matching of magnitude and the time constants of the thermal expansion of the blades, discs, blade-ring carriers and casing.

Conjugate Heat Transfer and Film Cooling of a Multi-Stage Cooling Scheme

2008 ◽  
Author(s):  
M. Ghorab ◽  
S. I. Kim ◽  
I. Hassan
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
Conjugate Heat Transfer ◽  
Density Ratio ◽  
Design Parameters ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Multi Stage ◽  
Internal Gap

Cooling techniques play a key role in improving efficiency and power output of modern gas turbines. The conjugate technique of film and impingement cooling schemes is considered in this study. The Multi-Stage Cooling Scheme (MSCS) involves coolant passing from inside to outside turbine blade through two stages. The first stage; the coolant passes through first hole to internal gap where the impinging jet cools the external layer of the blade. Finally, the coolant passes through the internal gap to the second hole which has specific designed geometry for external film cooling. The effect of design parameters, such as, offset distance between two-stage holes, gap height, and inclination angle of the first hole, on upstream conjugate heat transfer rate and downstream film cooling effectiveness performance are investigated computationally. An Inconel 617 alloy with variable properties is selected for the solid material. The conjugate heat transfer and film cooling characteristics of MSCS are analyzed across blowing ratios of Br = 1 and 2 for density ratio, 2. This study presents upstream wall temperature distributions due to conjugate heat transfer for different gap design parameters. The maximum film cooling effectiveness with upstream conjugate heat transfer is less than adiabatic film cooling effectiveness by 24–34%. However, the full coverage of cooling effectiveness in spanwise direction can be obtained using internal cooling with conjugate heat transfer, whereas adiabatic film cooling effectiveness has narrow distribution.

Recent Advances in the Study of Film Cooling on the Gas Turbine Blade

2014 ◽  
Vol 971-973 ◽  
pp. 143-147 ◽  
Author(s):  
Ping Dai ◽  
Shuang Xiu Li
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Performance ◽  
Leading Edge ◽  
Development Trend ◽  
Research Progress ◽  
Gas Turbine Blade ◽  
New Generation

The development of a new generation of high performance gas turbine engines requires gas turbines to be operated at very high inlet temperatures, which are much higher than the allowable metal temperatures. Consequently, this necessitates the need for advanced cooling techniques. Among the numerous cooling technologies, the film cooling technology has superior advantages and relatively favorable application prospect. The recent research progress of film cooling techniques for gas turbine blade is reviewed and basic principle of film cooling is also illustrated. Progress on rotor blade and stationary blade of film cooling are introduced. Film cooling development of leading-edge was also generalized. Effect of various factor on cooling effectiveness and effect of the shape of the injection holes on plate film cooling are discussed. In addition, with respect to progress of discharge coefficient is presented. In the last, the future development trend and future investigation direction of film cooling are prospected.

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.

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