Parametric Cooling Study of Single-Row Cylindrical Film Holes on Pressure Side of a Rotor Blade

Aerodynamic Loss ◽

Film cooling of 1st stage rotor blades strongly relates to the aerodynamic performance and life expectation of gas turbine. In this piece of work, a parametric study was carried out based on 3D RANS CFD methods to systematically investigate both cooling performance and aerodynamic loss of various film holes’ arrangements on the pressure side (PS) of a realistic rotor blade. In each CFD case, cylindrical film holes were arranged in one row with P/D around 4.0. More than one hundred CFD cases were carried out with various dimensionless streamwise locations (0.1, 0.2, ..., and 0.7), compound angles (−60°, −30°, ..., and 90°), and relative coolant massflows (0.2%, 0.3%, 0.45%, 0.7%, and 1.0% of mainstream massflow). Detailed distributions of adiabatic film cooling effectiveness on blade surface and aerodynamic loss with film cooling were computed for each case. It was found that coolant massflow was the decisive factor of aerodynamic efficiency φ. For near-LE film holes, −60° compound angle gave higher and more uniform η distributions at various coolant massflows. And the separation near LE could be suppressed, improving local cooling at near-hub on PS. For downstream film holes, 60° and −60° compound angles provided equivalent cooling performances, which were also better than other cases. The cooling characteristics of film holes with −30° compound angle were more susceptible to both coolant massflow and the location of ejection, which was mainly because of the blow-off and reattachment of coolant jets. For the cases with 90° compound angle, the distributions of η were generally similar, and the values increased monotonically as coolant massflows get larger.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

Journal of Turbomachinery ◽

10.1115/1.2813012 ◽

2008 ◽

Vol 131 (1) ◽

Cited By ~ 18

Author(s):

Zhihong Gao ◽

Diganta P. Narzary ◽

Je-Chin Han

Keyword(s):

Turbine Blade ◽

Film Cooling ◽

Pressure Ratio ◽

Suction Side ◽

Pressure Side ◽

Blade Surface ◽

Cooling Effectiveness ◽

Pressure Surface ◽

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

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

Volume 3: Turbo Expo 2002, Parts A and B ◽

10.1115/gt2002-30194 ◽

2002 ◽

Cited By ~ 9

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 ◽

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.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

ASME/JSME 2007 Thermal Engineering Heat Transfer Summer Conference, Volume 3 ◽

10.1115/ht2007-32098 ◽

2007 ◽

Cited By ~ 3

Author(s):

Zhihong Gao ◽

Diganta P. Narzary ◽

Je-Chin Han

Keyword(s):

Turbine Blade ◽

Film Cooling ◽

Pressure Ratio ◽

Suction Side ◽

Steady Increase ◽

Pressure Side ◽

Wake Effect ◽

Cooling Effectiveness ◽

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the Pressure Sensitive Paint (PSP). Four rows of fan-shaped, laid-back compound angled cooling holes are distributed on the pressure side while two such rows are provided on the suction side of the blade. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios from 0.4 to 1.5. Presence of wake due to upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals along the pitch-wise direction. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The fan-shaped, laid-back compound angled holes produce good coolant film coverage on the suction side except for those regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and effectiveness level is low on the pressure surface. However, the pressure side acquires relatively uniform film coverage with the design of multiple rows of cooling holes. The presence of stationary upstream wake results in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.4 to 1.5 shows steady increase in effectiveness on the pressure side or the suction side for a given wake rod phases locations. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film cooling effectiveness particularly at higher blowing ratios.

A Combined Experimental and Numerical Study of the Turbine Blade Tip Film Cooling Effectiveness Under Rotation Condition

Volume 5B: Heat Transfer ◽

10.1115/gt2014-25625 ◽

2014 ◽

Cited By ~ 1

Author(s):

M. Rezasoltani ◽

K. Lu ◽

M. T. Schobeiri ◽

J. C. Han

Keyword(s):

Film Cooling ◽

Numerical Study ◽

Experimental Investigations ◽

Pressure Side ◽

Cooling Effectiveness ◽

Side Edge ◽

Blade Tip ◽

The Impact ◽

Detailed numerical and experimental investigations of film cooling effectiveness were conducted on the blade tips of the first rotor row pertaining to a three-stage research turbine. Four different blade tip ejection configurations were utilized to determine the impact of the hole arrangements on the film cooling effectiveness. plane tip with tip hole cooling, squealer tip with tip hole cooling, plane tip with pressure-side-edge compound angle hole cooling and squealer tip with pressure-side-edge compound angle hole cooling. To avoid rotor imbalance, every pair is installed radially. Film cooling effectiveness measurements were performed for three blowing ratios (M) of 0.75, 1.25 and 1.75. Film cooling data was also obtained for three rotational speeds; 3000 rpm (reference condition), 2550 rpm and 2000 rpm. Film cooling measurements were performed using pressure sensitive paint (PSP) technique. In a parallel effort, extensive numerical investigations of the above configurations were performed to give a better view of flow behavior using a commercially available code. The experimental investigations were performed in the three-stage multi-purpose turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University.

3D Flow Prediction and Improvement of Holes Arrangement of a Film-Cooled Turbine Blade Using a Feature-Based Jet Model

Volume 3: Heat Transfer, Parts A and B ◽

10.1115/gt2006-91073 ◽

2006 ◽

Cited By ~ 1

Author(s):

Andre´ Burdet ◽

Reza S. Abhari

Keyword(s):

Turbine Blade ◽

Film Cooling ◽

Good Prediction ◽

Pressure Side ◽

3D Flow ◽

Cooling Effectiveness ◽

Numerical Predictions ◽

Flow Prediction ◽

Feature Based

A feature-based jet model has been proposed for use in 3D CFD prediction of turbine blade film cooling. The goal of the model is to be able to perform computationally efficient flow prediction and optimization of film-cooled turbine blades. The model reproduces in the near hole region the macro flow features of a coolant jet within a Reynolds-Averaged Navier Stokes (RANS) framework. Numerical predictions of the 3D flow through a linear transonic film-cooled turbine cascade are carried out with the model, with a low computational overhead. Different cooling holes arrangement are computed and the prediction accuracy is evaluated versus experimental data. It shown that the present model provides a reasonably good prediction of the adiabatic film-cooling effectiveness and Nusselt number around the blade. A numerical analysis of the interaction of coolant jets issuing from different rows of holes on the blade pressure side is carried out. It is shown that the upward radial migration of the flow due to the passage secondary flow structure has an impact on the spreading of the coolant and the film cooling effectiveness on the blade surface. Based on this result, a new arrangement of the cooling holes for the present case is proposed that leads to a better spanwise covering of the coolant on the blade pressure side surface.

Assessment of a Double Hole Film Cooling Geometry Using S-PIV and PSP

Volume 3B: Heat Transfer ◽

10.1115/gt2013-94614 ◽

2013 ◽

Cited By ~ 7

Author(s):

Lesley M. Wright ◽

Stephen T. McClain ◽

Charles P. Brown ◽

Weston V. Harmon

Keyword(s):

Film Cooling ◽

Density Ratio ◽

Three Dimensional ◽

Field Measurements ◽

Secondary Flows ◽

Cylindrical Hole ◽

Cooling Effectiveness ◽

Hole Geometry ◽

A novel, double hole film cooling configuration is investigated as an alternative to traditional cylindrical and fanshaped, laidback holes. This experimental investigation utilizes a Stereo-Particle Image Velocimetry (S-PIV) to quantitatively assess the ability of the proposed, double hole geometry to weaken or mitigate the counter-rotating vortices formed within the jet structure. The three-dimensional flow field measurements are combined with surface film cooling effectiveness measurements obtained using Pressure Sensitive Paint (PSP). The double hole geometry consists of two compound angle holes. The inclination of each hole is θ = 35°, and the compound angle of the holes is β = ± 45° (with the holes angled toward one another). The simple angle cylindrical and shaped holes both have an inclination angle of θ = 35°. The blowing ratio is varied from M = 0.5 to 1.5 for all three film cooling geometries while the density ratio is maintained at DR = 1.0. Time averaged velocity distributions are obtained for both the mainstream and coolant flows at five streamwise planes across the fluid domain (x/d = −4, 0, 1, 5, and 10). These transverse velocity distributions are combined with the detailed film cooling effectiveness distributions on the surface to evaluate the proposed double hole configuration (compared to the traditional hole designs). The fanshaped, laidback geometry effectively reduces the strength of the kidney-shaped vortices within the structure of the jet (over the entire range of blowing ratios considered). The three-dimensional velocity field measurements indicate the secondary flows formed from the double hole geometry strengthen in the plane perpendicular to the mainstream flow. At the exit of the double hole geometry, the streamwise momentum of the jets is reduced (compared to the single, cylindrical hole), and the geometry offers improved film cooling coverage. However, moving downstream in the steamwise direction, the two jets form a single jet, and the counter-rotating vortices are comparable to those formed within the jet from a single, cylindrical hole. These strong secondary flows lift the coolant off the surface, and the film cooling coverage offered by the double hole geometry is reduced.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator–Rotor Purge Flow and Compound-Angle Holes

Journal of Thermal Science and Engineering Applications ◽

10.1115/1.4026964 ◽

2014 ◽

Vol 6 (4) ◽

Cited By ~ 8

Author(s):

Kevin Liu ◽

Shang-Feng Yang ◽

Je-Chin Han

Keyword(s):

Film Cooling ◽

Density Ratio ◽

Cooling Effectiveness ◽

Fundamental Parameters ◽

Blowing Ratio ◽

Purge Flow ◽

Cylindrical Holes ◽

Exit Mach Number ◽

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator–rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study—a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and was based on the exit velocity and chord length of the blade. The measurement technique adopted was the conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with an increase of blowing ratio. Results also indicate that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

Journal of Turbomachinery ◽

10.1115/1.1622712 ◽

2003 ◽

Vol 125 (4) ◽

pp. 648-657 ◽

Cited By ~ 31

Author(s):

Jae Su Kwak ◽

Je-Chin Han

Keyword(s):

Heat Transfer ◽

Gas Turbine ◽

Turbine Blade ◽

Film Cooling ◽

Rotor Blade ◽

Transfer Coefficients ◽

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.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

Volume 7B: Heat Transfer ◽

10.1115/gt2020-14443 ◽

2020 ◽

Author(s):

Joao Vieira ◽

John Coull ◽

Peter Ireland ◽

Eduardo Romero

Keyword(s):

High Pressure ◽

Film Cooling ◽

Turbine Blades ◽

Reynolds Numbers ◽

High Pressure Turbine ◽

Cooling Effectiveness ◽

Aerodynamic Loss ◽

Blade Tip ◽

Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

Investigations of improved cooling effectiveness for ramp film cooling with compound angle film cooling jets

Aircraft Engineering and Aerospace Technology ◽

10.1108/aeat-05-2020-0082 ◽

2021 ◽

Vol ahead-of-print (ahead-of-print) ◽

Author(s):

Krishna Anand Vasu Devan Nair Girija Kumari ◽

Parammasivam Kanjikoil Mahali

Keyword(s):

Film Cooling ◽

Flat Surface ◽

Flow Characteristics ◽

Test Surface ◽

Cooling Effectiveness ◽

Content Type ◽

Ramp Test ◽

Cooling Hole ◽

Purpose This paper aims to investigate the film cooling effectiveness (FCE) and mixing flow characteristics of the flat surface ramp model integrated with a compound angled film cooling jet. Design/methodology/approach Three-dimensional numerical simulation is performed on a flat surface ramp model with Reynolds Averaged Navier-Stokes approach using a finite volume solver. The tested model has a fixed ramp angle of 24° and a ramp width of two times the diameter of the film cooling hole. The coolant air is injected at 30° along the freestream direction. Three different film hole compound angles oriented to freestream direction at 0°, 90° and 180° were investigated for their performance on-ramp film cooling. The tested blowing ratios (BRs) are in the range of 0.9–2.0. Findings The film hole oriented at a compound angle of 180° has improved the area-averaged FCE on the ramp test surface by 86.74% at a mid-BR of 1.4% and 318.75% at higher BRs of 2.0. The 180° film hole compound angle has also produced higher local and spanwise averaged FCE on the ramp test surface. Originality/value According to the authors’ knowledge, this study is the first of its kind to investigate the ramp film cooling with a compound angle film cooling hole. The improved ramp model with a 180° film hole compound angle can be effectively applied for the end-wall surfaces of gas turbine film cooling.