Tripod Hole Geometry Performance for a Vane Suction Surface Near Throat Location

2013 ◽  
Author(s):  
Sridharan Ramesh ◽  
Chris LeBlanc ◽  
Srinath V. Ekkad ◽  
Mary Anne Alvin
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Throat Region ◽  
Blowing Ratio ◽  
Hole Geometry ◽  
Film Cooling Holes ◽  
Cooling Air

Film cooling performance depends strongly on the hole exit geometry, blowing ratio, and hole location. The goal of this study is to evaluate film cooling geometries that can provide better protection over the suction surface of the airfoil beyond the throat region. This study compares the performance of standard cylindrical; fan-shaped (10° flare/laidback); tripod hole geometry (15° breakout angle); and tripod holes with shaped exits (5° flare on 15° tripod). Film cooling holes are located just upstream of the throat region on the suction side of an airfoil. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low speed linear cascade wind tunnel. A range of blowing ratios from 0.5 to 2.0 was covered for a cylindrical hole, while ensuring all other hole geometries run under similar mass flow rate conditions. Steady state IR (Infra-Red) technique was employed to measure adiabatic film cooling effectiveness. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations thus providing cooling in the important trailing edge portion of the airfoil.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using PSP Technique

2010 ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Lift Off ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke wheel wake generator) on the modeled rotor blade is studied using the Pressure Sensitive Paint (PSP) mass transfer analogy method. Emphasis of the current study is on the mid-span region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film cooling holes. The blade also has radial shower-head leading edge film cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds Number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side; 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using Pressure Sensitive Paint Technique

2012 ◽  
Vol 134 (8) ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Unsteady Wake ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke-wheel wake generator) on the modeled rotor blade is studied using the pressure sensitive paint (PSP) mass-transfer analogy method. Emphasis of the current study is on the midspan region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film-cooling holes. The blade also has radial shower-head leading edge film-cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side and 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film-cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Film Cooling Performance on Turbine Blade Suction Side With Various Film Cooling Hole Arrangements

2020 ◽  
Author(s):  
Zhiyu Zhou ◽  
Haiwang Li ◽  
Gang Xie ◽  
Ruquan You
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Stokes Equations ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blowing Ratio

Abstract Numerical simulations were carried out to study the film cooling effectiveness distributions of different hole arrangements on the suction side of a high pressure turbine blade under rotating condition. The chord length and the height of the blade are 60mm and 80mm, respectively. Totally 12 models with different hole arrangements and different injection angles were studied. Each blade model has three rows of round holes with diameter of 0.9mm on the suction surface. The first row and the third row are fixed at streamwise location of 12.4% and 34% respectively. Three injection angles, 30°, 45°, and 60°, were investigated. Simulations were conducted under three rotational speeds, 600rpm, 800rpm, 1000rpm, with blowing ratio varying from 0.5 to 2.0. The Mainstream Reynolds numbers corresponding to the rotational speeds are 40560, 54080, and 67600 respectively. The temperature of the mainstream and the coolant is set at 463K and 303K so as to control the density ratio at 1.47. Simulations were performed by using SST turbulence model and were solved by using the three-dimensional Reynolds-averaged Navier–Stokes equations. Results showed that on the rotating turbine blade suction surface, film trajectories are drawn toward the midspan. The film trajectory arrangement may be different from the hole arrangement. Inline film trajectory arrangement can achieve higher film cooling effectiveness with slightly larger injection angle. Staggered film trajectory arrangement is better for uniform film cooling effectiveness distribution in spanwise and can achieve higher film cooling effectiveness with smaller injection angle. A smaller distance between the first row and the second row can achieve better film cooling performance at the downstream. With the increase of rotational speed, the mainstream Reynolds number increases, which improves the film cooling performance with smaller blowing ratio.

Effect of a Cutback Squealer and Cavity Depth on Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Shantanu Mhetras ◽  
Diganta Narzary ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cavity Depth ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling effectiveness from shaped holes on the near tip pressure side and cylindrical holes on the squealer cavity floor is investigated. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. Effects of varying blowing ratios and squealer cavity depth are also examined on film-cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure side and suction side rim walls using pressure sensitive paint technique. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film-cooling holes. Two rows of cylindrical film-cooling holes are arranged offset to the suction side profile and along the camber line on the tip. Another row of shaped film-cooling holes is arranged along the pressure side just below the tip. The average blowing ratio of the cooling gas is controlled to be 0.5, 1.0, 1.5, and 2.0. A five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5% is used to perform the experiments. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,480,000 and the inlet and exit Mach numbers were 0.23 and 0.65, respectively. A blowing ratio of 1.0 is found to give best results on the pressure side, whereas the tip surfaces forming the squealer cavity give best results for M=2. Results show high film-cooling effectiveness magnitudes near the trailing edge of the blade tip due to coolant accumulation from upstream holes in the tip cavity. A squealer depth with a recess of 2.1mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2mm.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Experimental Investigation of Turbine Phantom Cooling on Suction Side With Combustor-Turbine Leakage Gap Flow and Endwall Film Cooling

2012 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Suction Side ◽  
Leakage Flow ◽  
Airfoil Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Passage Vortex

The film cooling injection on Hp turbine component surface is strongly affected by the complex flow structure in the nozzle guide vane or rotor blade passages. The action of passage vortex near endwall surface could dominate the film cooling effectiveness distribution on the component surfaces. The film cooling injections from endwall and airfoil surface are mixed with the passage vortex. Considering a small part of the coolant injection from endwall will move towards the airfoil suction side and then cover some area, the interaction between the coolants injected from endwall and airfoil surface is worth investigating. Though the temperature of coolant injection from endwall increases after the mixing process in the main flow, the injections moving from endwall to airfoil suction side still have the potential of second order cooling. This part of the coolant is called “Phantom cooling flow” in the paper. A typical scale-up model of GE-E3 Hp turbine NGV is used in the experiment to investigate the cooling performance of injection from endwall. Instead of the endwall itself, the film cooling effectiveness is measured on the airfoil suction side. This paper is focused on the combustor-turbine interface gap leakage flow and the coolant from fan-shaped holes moving from endwall to airfoil suction side. The coolant flow is injected at a 30deg angle to the endwall surface both from a slot and four rows of fan-shaped holes. The film cooling holes on the endwall and the leakage flow are used simultaneously. The blowing ratio and incidence angle are selected to be the parameters in the paper. The experiment is completed with the blowing ratio changing from M = 0.7 to M = 1.3 and the incidence angle varying from −10deg to +10deg, with inlet Reynolds numbers of Re = 3.5×105 and an inlet Mach number of Ma = 0.1.

Influence of Coolant Density on Turbine Blade Film Cooling at Transonic Cascade Flow Conditions Using the Pressure Sensitive Paint Technique

2021 ◽  
Author(s):  
Izhar Ullah ◽  
Sulaiman M. Alsaleem ◽  
Lesley M. Wright ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Abstract This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

Effects of Blowing Ratio on Multiple Shallow Angle Film Cooling Holes

2012 ◽  
Vol 225 ◽  
pp. 49-54 ◽  
Author(s):  
Kamil Abdullah ◽  
Ken Ichi Funazaki
Keyword(s):  
Temperature Field ◽  
Surface Temperature ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Ir Camera ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Hole Configuration ◽  
The Common ◽  
Film Cooling Holes

This paper presents the investigation on the effects of the blowing ratio of multiple shallow angle film cooling holes. Multiple film cooling holes having a shallow hole angle (θ = 20°), arranged to perform in-line hole configuration has been considered in the present study. The investigation focuses on the effects of high blowing ratio of the film cooling effectiveness which have been carried out at ReD= 3100 and BR = 2.0, 3.0 and 4.0. The experiments make use of the IR camera in capturing the surface temperature to determine the film cooling effectiveness. The contours of the film cooling effectiveness distribution together with plots on laterally average film cooling effectiveness along the x/D are presented. The discussions have been made with a support of the temperature field captured at x/D = 3, 13, 23, and 33. The results clearly show the benefit of the employment of shallow hole angle (θ = 20°) at high blowing ratio which is much more superior in comparison to the common hole configuration (θ = 35°).

Film-Cooling Effectiveness on a Gas Turbine Blade Tip Using Pressure-Sensitive Paint

2005 ◽  
Vol 127 (5) ◽  
pp. 521-530 ◽  
Author(s):  
Jaeyong Ahn ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Oxygen Concentration ◽  
Film Cooling ◽  
Pressure Sensitive Paint ◽  
Nitrogen Gas ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Effects of the presence of squealer, the locations of the film-cooling holes, and the tip-gap clearance on the film-cooling effectiveness were studied and compared to those for a plane (flat) tip. The film-cooling effectiveness distributions were measured on the blade tip using the pressure-sensitive paint technique. Air and nitrogen gas were used as the film-cooling gases, and the oxygen concentration distribution for each case was measured. The film-cooling effectiveness information was obtained from the difference of the oxygen concentration between air and nitrogen gas cases by applying the mass transfer analogy. Plane tip and squealer tip blades were used while the film-cooling holes were located (a) along the camber line on the tip or (b) along the tip of the pressure side. The average blowing ratio of the cooling gas was 0.5, 1.0, and 2.0. Tests were conducted with a stationary, five-bladed linear cascade in a blow-down facility. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,138,000, and the inlet and the exit Mach numbers were 0.25 and 0.6, respectively. Turbulence intensity level at the cascade inlet was 9.7%. All measurements were made at three different tip-gap clearances of 1%, 1.5%, and 2.5% of blade span. Results show that the locations of the film-cooling holes and the presence of squealer have significant effects on surface static pressure and film-cooling effectiveness, with film-cooling effectiveness increasing with increasing blowing ratio.

Film Cooling Effectiveness Distribution on First-Stage Vane Endwall With and Without Leading-Edge Fillets: Part I—Effect of Leading Edge Geometry

2011 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Flow Direction ◽  
Leading Edge ◽  
Suction Side ◽  
Axial Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Geometry ◽  
Blade Cascade

The paper is focused on the effect of leading edge airfoil geometry on endwall film cooling. Fillets placed at the junctions of the leading edge and the endwall are used in investigation. Three types of fillet profiles are tested, and the results are compared with baseline geometry without fillet. The design of the fillet is based on the suggestion by previous literature data indicating that sharp is effective in controlling the secondary flow. Three types of sharp slope fillet with the length to height ratio of 2.8, 1.2 and 0.5 are made using stereo lithography (SLA) and assessed in the experiment. Distributed with the approximately inviscid flow direction, four rows of compound angle laidback fan-shaped holes are arranged on the endwall to form full covered coolant film. The four rows of fanshaped holes are inclined 30 deg to the endwall surface and held an angle of 0, 30, 45 and 60 deg to axial direction respectively. The fanshaped holes have a lateral diffusion angle of 10 deg from the hole-centerline and a forward expansion angle of 10 deg to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5*105, and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet) while the blowing ratio of the coolant through the discrete holes varies from 0.4 to 1.2. The film-cooling effectiveness distributions are obtained using the PSP (pressure sensitive paint) technique, by which the effect of different fillet geometry on passage induced flow and coolant is shown. The present paper compares the film cooling effectiveness distributions in a baseline blade cascade with three similar blades with different leading edge by adding fillets. The results show that with blowing ratio increasing, the film cooling effectiveness increases on the endwall. For specific blowing ratio, the effects of leading edge geometries could be illustrated as follows. The baseline geometry provides the best film cooling performance near leading edge pressure side. As for the leading edge suction side, the best leading edge geometry depends on the blowing ratio. The longfillet is the more effective in controlling horseshoe vortex at low blowing ratio, but for the high blowing ratio shortfillet and mediumfillet are better.

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