Platform Film Cooling Investigation on an HP Nozzle Vane Cascade With Discrete Shaped Holes and Slot Film Cooling

2020 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
L. Abba ◽  
L. Pestelli
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Heat Transfer Performance ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Intensity Level ◽  
Transfer Performance ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

Abstract The present paper reports on an experimental investigation on the aerodynamic and heat transfer performance of different platform cooling schemes: two based on cylindrical and shaped holes and one featuring a slot located upstream of the leading edge plane simulating the combustor to stator interface gap. Tests were run on a 6-vane cascade operated at an isentropic cascade exit Mach number of 0.4 and a significant inlet turbulence intensity level of about 9%. The cooling schemes were first tested to quantify their impact on secondary flows and related losses for variable injection conditions. Heat transfer performance was then assessed through adiabatic film cooling effectiveness and heat transfer coefficient measurements. The Net Heat Flux Reduction parameter was then computed to critically assess the cooling schemes. When compared with the cylindrical hole scheme, shaped holes outperform for all tested injection rates, while the slot alone is able to thermally protect only the front of the passage. Discrete holes are required to cool the platform region along the whole pressure side and the suction side leading edge region.

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

2015 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Intensity Level ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.

Numerical Prediction of Film Cooling and Heat Transfer on the Leading Edge of a Rotating Blade With Two Rows Holes in a 1-1/2 Turbine Stage at Design and Off Design Conditions

2005 ◽  
Author(s):  
Huitao Yang ◽  
Hamn-Ching Chen ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Rotating Speed ◽  
Rotating Blade

Numerical simulations were performed to predict the film cooling effectiveness and the associated heat transfer coefficient on the leading edge of a rotating blade in a 1-1/2 turbine stage using a Reynolds stress turbulence model together with a non-equilibrium wall function. Simulations were performed for both the design and off-design conditions to investigate the effects of blade rotation on the leading edge film cooling effectiveness and heat transfer coefficient distributions. It was found that the tilt stagnation line on the leading edge of rotor moves from the pressure side to the suction side, and the instantaneous coolant streamlines shift from the suction side to the pressure side with increasing rotating speed. This trend was supported by the experimental results. The result also showed that the heat transfer coefficient increases, but film cooling effectiveness decreases with increasing rotating speed. In addition, the unsteady characteristics of the film cooling and heat transfer at different time phases, as well as different rotating speeds, were also reported.

Heat Transfer and Effectiveness on the Film Cooled Tip and Inner Rim Surfaces of a Turbine Blade

2010 ◽  
Author(s):  
Jun Su Park ◽  
Dong Hyun Lee ◽  
Hyung Hee Cho ◽  
Dong-Ho Rhee ◽  
Shin-Hyung Kang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
High Heat ◽  
Chord Length ◽  
Tip Clearance ◽  
Film Cooling Effectiveness ◽  
High Heat Transfer ◽  
Film Cooling Holes

Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a rotor blade with a squealer rim. The blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, the axial chord length (Cx) of which was 237 mm with a turning angle of 126°. The mainstream Reynolds number based on the axial chord was 1.5×105. The turbulence intensity level at the cascade inlet was approximately 12%. Measurements were made at three different rim heights (H) of about 3%, 6%, and 9% of the axial chord length. The tip clearance (C) ranges were 1–3% of the axial chord length. Also, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and the suction side rim. In particular, a coolant was injected at an incline of 45° from near the suction side film cooling holes. The film cooling experiments were done with a fixed tip clearance and rim height at 1% and 6% of the axial chord length. The blowing rate was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface in some cases, due to the reattachment of tip leakage flow. The peak values moved toward the suction-side edge, and the magnitude and area of high heat transfer increased near the leading edge as the tip clearance increased. The heat transfer decreased on the tip surface with increases in the rim height. In the film-cooling cases, the high heat transfer and film-cooling effectiveness region appeared near the film-cooling holes.

Effect of Flow Parameter Variations on Full Coverage Film-Cooling Effectiveness for a Gas Turbine Blade

2007 ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han ◽  
Ron Rudolph
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Suction Side ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Film Cooling Holes

The effect of film cooling holes placed along the span of a fully-cooled high pressure turbine blade in a stationary, linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Effect of showerhead injection at the leading edge and the presence of compound angled, diffusing holes on the pressure and suction side are also examined. Six rows of compound angled shaped film cooling holes are provided on the pressure side while four such rows are provided on the suction side of the blade. The holes have a laidback and fan-shaped diffusing cross-section. Another three rows of cylindrical holes are drilled at a typical angle on the leading edge to capture the effect of showerhead film coolant injection. The film cooling hole arrangement simulates a typical film cooled blade design used in stage 1 rotor blades for gas turbines used for power generation. A typical blowing ratio is defined for each film hole row and tests are performed for 100%, 150% and 200% of this typical value. Tests are performed for inlet Mach numbers of 0.36 and 0.45 with corresponding exit Mach numbers of 0.51 and 0.68 respectively. The flow remains subsonic in the throat region for both Mach numbers. The corresponding free stream Reynolds number, based on the axial chord length and the exit velocity, are 1.3 million and 1.74 million respectively. Freestream turbulence intensity level at the cascade inlet is 6%. Results show that varying blowing ratios can have a significant impact on film-cooling effectiveness distribution. Large spanwise variations in effectiveness distributions are also observed. Similar distributions were observed for both Mach numbers.

A Numerical Study on the Characteristics of How Heat and Cooling Transfer on the Leading Edge of a Film-Cooling Blade

2011 ◽  
Vol 383-390 ◽  
pp. 3963-3968
Author(s):  
Shao Hua Li ◽  
Li Mei Du ◽  
Wen Hua Dong ◽  
Ling Zhang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

In this paper, a numerical simulation was performed to investigate heat transferring characteristics on the leading edge of a blade with three rows of holes of film-cooling using Realizable k- model. Three rows of holes were located on the suction side leading edge stagnation line and the pressure surface. The difference of the cooling efficiency and the heat transfer of the three rows of holes on the suction side and pressure side were analyzed; the heat transfer and film cooling effectiveness distribution in the region of leading edge are expounded under different momentum rations.The results show that under the same condition, the cooling effectiveness on the pressure side is more obvious than the suction side, but the heat transfer is better on the suction side than the pressure side. The stronger momentum rations are more effective cooling than the heat transfer system.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Suction-Side Gill-Region Film Cooling: Effects of Hole Shape and Orientation on Adiabatic Effectiveness and Heat Transfer Coefficient

2008 ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Thermal Protection ◽  
Suction Side ◽  
Lateral Spreading ◽  
Film Cooling Effectiveness ◽  
Cooling Effects ◽  
Exit Mach Number

The performance of suction-side gill region film cooling is investigated using the University of Utah Transonic Wind Tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on thermal film cooling characteristics are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes only, second row of holes only, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.73 to 1.92 similar to values present in operating gas turbine engines. A mesh grid is used to give a magnitude of longitudinal turbulence intensity of 5.7 percent at the inlet of the test section. Results show that the best overall protection over the widest range of blowing ratios and streamwise locations is provided by either the RC holes, or the RR holes. This result is particularly significant because the RR hole arrangement, which has lower manufacturing costs compared with the RC and SA arrangements, produces better or equivalent levels of performance in terms of magnitudes of adiabatic film cooling effectiveness and heat transfer coefficient. Such improved performance (relative to RA and SA holes) is likely mostly a result of compound angles, which increases lateral spreading. As such, the present results indicate that compound angle appears to be more effective than hole shaping in improving thermal protection relative to that given by RA holes.

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