Aerothermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
D. O. O’Dowd ◽  
Q. Zhang ◽  
L. He ◽  
B. C. Y. Cheong ◽  
I. Tibbott
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
High Speed ◽  
Tip Clearance ◽  
Test Facility ◽  
Pressure Probe ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Loss Coefficients

This paper presents an experimental investigation of the aerothermal performance of a cooled winglet tip under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27 × 106, based on axial chord). Spatially resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

Aero-Thermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

2011 ◽  
Author(s):  
D. O. O’Dowd ◽  
Q. Zhang ◽  
L. He ◽  
B. C. Y. Cheong ◽  
I. Tibbott
Keyword(s):  
Nusselt Number ◽  
Thermal Performance ◽  
Film Cooling ◽  
High Speed ◽  
Tip Clearance ◽  
Test Facility ◽  
Pressure Probe ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Loss Coefficients

This paper presents an experimental investigation of the aero-thermal performance of a cooled winglet tip, under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27×106, based on axial chord). Spatially-resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially-made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

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 ◽  
Film 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.

Effects of Trenched Holes on Film Cooling of a Contoured Endwall Nozzle Vane

2010 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Energy Loss ◽  
Film Cooling ◽  
Loss Coefficient ◽  
Flow Rate Ratio ◽  
Pressure Probe ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Coolant Injection ◽  
Cylindrical Holes

The present paper investigates the effects of the application of trenched holes in the front part of a contoured film cooled endwall. Two trench configurations were tested, changing the trench depth. Tests have been carried out at low speed (M2is = 0.2) and low inlet turbulence intensity level, with coolant mass flow rate ratio varied within the 0.5–2.5% range. Pressure probe traverses were performed downstream of the vane trailing edge to show the secondary flow field modifications and to evaluate trench additional losses. Endwall distributions of film cooling effectiveness have been obtained by TLC technique. For each injection condition energy loss coefficient and film cooling effectiveness distributions were analyzed and compared to the ones obtained from rows of cylindrical holes. Laterally and area averaged effectiveness as well as pitch and mass averaged kinetic energy loss coefficient were computed to enlighten any change induced by the introduction of trenched holes. A uniform and high thermal coverage was obtained in the region just downstream of the trench, but it quickly decayed, because of enforced mixing of coolant with main-flow. Compared to the cylindrical hole configuration, trenches are able to provide a higher global cooling effectiveness, but a larger amount of coolant injection is required. The introduction of both trenches is responsible for a secondary thermodynamic loss increase of about 0.7%, at low coolant injection rates. Increasing blowing rates, the additional loss is going to vanish.

Film-Cooled Turbine Endwall in a Transonic Flow Field: Part I—Aerodynamic Measurements

2001 ◽  
Vol 123 (4) ◽  
pp. 709-719 ◽  
Author(s):  
Friedrich Kost ◽  
Martin Nicklas
Keyword(s):  
Flow Field ◽  
Transonic Flow ◽  
Film Cooling ◽  
Coolant Flow ◽  
Pressure Probe ◽  
Thermal Methods ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Oil Flow ◽  
Oil Flow Visualization

Thermodynamic and aerodynamic measurements were carried out in a linear turbine cascade with transonic flow field. Heat transfer and adiabatic film-cooling effectiveness resulting from the interaction of the flow field and the ejected coolant at the endwall were measured and will be discussed in two parts. The investigations were performed in the Windtunnel for Straight Cascades (EGG) at DLR, Go¨ttingen. The film-cooled NGV endwall was operated at representative dimensionless engine conditions of Mach and Reynolds number Ma2is=1.0 and Re2=850,000 respectively. Part I of the investigation discusses the aerodynamic measurements. Detailed aerodynamic measurements were carried out in the vicinity of a turbine stator endwall using conventional pressure measurements and a Laser-2-Focus (L2F) device. The L2F served as a velocimeter measuring 2D-velocity vectors and turbulence quantities and as a tool to determine the concentration of coolant ejected through a slot and through holes at the endwall. Pressure distribution measurements provided information on the endwall pressure field and its variation with coolant flow rate. Pressure probe measurements delivered cascade performance data. Oil flow visualization and laser velocimetry gave a picture of the near endwall flow field and its interference with the coolant. A strikingly strong interaction of coolant air and secondary flow field could be identified. The measurement of coolant concentration downstream of the ejection locations provided a detailed picture of the coolant flow convection and its mixing with the main flow. The relative coolant concentration in the flow field is directly comparable to the adiabatic film-cooling effectiveness measured by thermal methods at the wall.

Effect of Hole Diameter on Nozzle Endwall Film Cooling and Associated Phantom Cooling

2015 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Kevin Liu ◽  
Moon Hee-Koo
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Leading Edge ◽  
Suction Side ◽  
Hole Diameter ◽  
Two Dimensional ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blowing Ratio

Flow fields near the turbine nozzle endwall are highly complex due to the passage vortices and endwall cross flows. Consequently, it is challenging to provide proper cooling to the endwall surfaces. An effective way to cool the endwall is to have film cooling holes forward of the leading edge, often called “inlet-film cooling”. This paper presents the results of an experimental investigation on how the film hole diameter affects the film effectiveness on nozzle endwall and associated phantom cooling effectiveness on airfoil suction side. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Double staggered rows of film injections, which were located upstream from the nozzle leading edge, provided cooling to the contoured endwall surfaces. Film cooling effectiveness on the endwall surface and corresponding phantom cooling effectiveness on the airfoil suction side were measured separately with a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Four different film hole diameters with the same injection angle and the same pitch to diameter ratio were studied for up to six different MFR’s (mass flow ratios). Two dimensional film effectiveness distributions on the endwall surface and two dimensional phantom cooling distributions on the airfoil suction side are presented. Film/phantom cooling effectiveness distributions are pitchwise/spanwise averaged along the axial direction and also presented. The results indicate that both the endwall film effectiveness and the suction side phantom cooling effectiveness increases with the hole diameter (as decreases in blowing ratio for a given MFR) up to a specific diameter, then starts decreasing. An optimal value of the film hole diameter (blowing ratio) for the given injection angle is also suggested based on current study.

scholarly journals Film Cooling With Compound Angle Holes: Adiabatic Effectiveness

1994 ◽  
Author(s):  
Donald L. Schmidt ◽  
Basav Sen ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Density Ratio ◽  
Flux Ratio ◽  
Test Facility ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Baseline Case ◽  
Adiabatic Effectiveness ◽  
Compound Angle

Film cooling effectiveness was studied experimentally in a flat plate test facility with zero pressure gradient using a single row of inclined holes which injected high density, cryogenically cooled air. Round holes and holes with a diffusing expanded exit were directed laterally away from the freestream direction with a compound angle of 60°. Comparisons were made with a baseline case of round holes aligned with the freestream. The effects of doubling the hole spacing to six hole diameters for each geometry were also examined. Experiments were performed at a density ratio of 1.6 with a range of blowing ratios from 0.5 to 2.5 and momentum flux ratios from 0.16 to 3.9. Lateral distributions of adiabatic effectiveness results were determined at streamwise distances from 3 D to 15 D downstream of the injection holes. All hole geometries had similar maximum spatially averaged effectiveness at a low momentum flux ratio of I = 0.25, but the round and expanded exit holes with compound angle had significantly greater effectiveness at larger momentum flux ratios. The compound angle holes with expanded exits had a much improved lateral distribution of coolant near the hole for all momentum flux ratios.

Study of Film Cooling Effectiveness on a Double-Walled Effusion-Cooled Turbine Blade in a High-Speed Flow Using Pressure Sensitive Paint

2019 ◽  
Author(s):  
Gladys C. Ngetich ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
High Speed ◽  
High Porosity ◽  
Pressure Sensitive Paint ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Double Wall

Abstract A detailed analysis of film cooling performance on a double-walled effusion-cooled blade is essential for both the coolant consumption optimization and assessment of the film to offer the desired levels of the turbine blade protection. Yet there are hardly any film effectiveness studies on double-wall full-coverage film cooled turbine blades. This paper presents a detailed film cooling effectiveness study over the full surface of a double-walled effusion-cooled high-pressure turbine rotor blade using Pressure Sensitive Paint (PSP). PSP permitted a non-intrusive and conduction-errors-free means of obtaining clean and distinct local distribution of film effectiveness on the blade surface making it possible to extract valuable film cooling effectiveness performance data on the whole blade surface. Three large-scale circular pedestal double-wall blade designs with varying pedestal height, pedestal diameter and cooling hole diameter were tested in a high-speed stationary single-blade linear cascade running at engine-representative Mach and Reynolds numbers. All the blades were tested within a range of representative modern engine coolant mass flow, ṁc to mainstream, ṁg ratios; 1.6% < ṁc/ṁ∞ < 5.5%. High porosity blade exhibited a better flow distribution and was found to consistently perform the best.

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.

Investigation on the Effect of Different Lands on Trailing Edge Slot Film Cooling

2018 ◽  
Vol 0 (0) ◽  
Author(s):  
Yang Xu ◽  
Hui-ren Zhu ◽  
Wei-jiang Xu ◽  
Jian-sheng Wei
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Turbine Blades ◽  
Experimental Studies ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Wall Functions ◽  
Film Cooling Effectiveness ◽  
Slot Film Cooling

Abstract Trailing edge slot film cooling is a widely used method for protecting the trailing edge of turbine blades from hot gas impingement. The structures that separate the slots, known as “lands,” come in a variety of configurations. This paper presents the effects of the trailing edge with different lands on the film cooling performance. Experimental studies are conducted on the film cooling effectiveness and Nusselt number with different lands. Four trailing edge configurations, including the straight lands, the beveling lands, the fillet lands and the tapered lands are considered under four blowing ratios (0.5, 0.7, 1.0 and 1.5). The Reynolds numbers of mainstream is fixed as 375,000. Film cooling effectiveness and Nusselt number performances are measured by transient liquid crystal measurement technique. Reynolds-averaged Navier-Stokes (RANS) simulation with realizable k-ε turbulence model and enhanced wall functions are performed using a commercial code Fluent. In each case, the slot height is kept constant. It is shown that the beveling lands, the fillet lands and the tapered lands have higher cooling effectiveness and lower Nusselt number compared with the straight lands. Under higher blowing ratios, the trailing edges of all four lands have higher cooling effectiveness and higher Nusselt number.

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