Second Law Analysis of Aerodynamic Gains Associated with Simple Angle and Compound Angle Full Coverage Film Cooling

2021 ◽  
Vol 42 (11) ◽  
Author(s):  
Sneha Reddy Vanga ◽  
Phillip M. Ligrani
Keyword(s):  
Film Cooling ◽  
Second Law ◽  
Full Coverage ◽  
Second Law Analysis ◽  
Compound Angle

Second Law Analysis of Aerodynamic Losses: Results for a Cambered Vane With and Without Film Cooling

2012 ◽  
Author(s):  
Phil Ligrani ◽  
Jae Sik Jin
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Chord Length ◽  
Second Law ◽  
Exergy Destruction ◽  
Freestream Turbulence ◽  
Second Law Analysis ◽  
Aerodynamic Losses

Results of second law analysis of experimentally-measured aerodynamic losses are presented for a cambered vane with and without film cooling, including comparisons with similar results from a symmetric airfoil. Included are distributions of local entropy creation, as well as mass-averaged magnitudes of global exergy destruction. The axial chord length of the cambered vane is 4.85 cm, the true chord length is 7.27 cm, and the effective pitch is 6.35 cm. Data are presented for three airfoil Mex distributions (including one wherein the flow is transonic), magnitudes of inlet turbulence intensity from 1.1 percent to 8.2 percent, and ks/cx surface roughness values of 0, 0.00108, and 0.00258. The associated second law aerodynamics losses are presented for two different measurement locations downstream of the vane trailing edge (one axial chord length and 0.25 axial chord length). The surface roughness, when present, simulates characteristics of the actual roughness which develops on operating turbine airfoils from a utility power engine, over long operating times, due to particulate deposition and to spallation of thermal barrier coatings (TBCs). Quantitative surface roughness characteristics which are matched include equivalent sandgrain roughness size, as well as the irregularity, non-uniformity, and the three-dimensional irregular arrangement of the roughness. Relative to a smooth, symmetric airfoil with no film cooling at low Mach number and low freestream turbulence intensity, overall, the largest increases in exergy destruction occur with increasing Mach number, and increasing surface roughness. Important variations are also observed as airfoil camber changes. Progressively smaller mass-averaged exergy destruction increases are then observed with changes of freestream turbulence intensity, and different film cooling conditions. In addition, the dependences of overall exergy destruction magnitudes on mainstream turbulence intensity and freestream Mach number are vastly different as level of vane surface roughness changes. When film cooling is present, overall mass-averaged exergy destruction magnitudes are significantly less than values associated with increased airfoil surface roughness for both the cambered vane and the symmetric airfoil. Exergy destruction values (associated with wake aerodynamic losses) for the symmetric airfoil with film cooling are then significantly higher than data from the cambered vane with film cooling, when compared at a particular blowing ratio.

Investigation of Film Cooling Effectiveness of Full-Coverage Inclined Multihole Walls With Different Hole Arrangements

2003 ◽  
Author(s):  
Yuzhen Lin ◽  
Bo Song ◽  
Bin Li ◽  
Gaoen Liu ◽  
Zhiyong Wu
Keyword(s):  
Experimental Data ◽  
Numerical Simulation ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Inclination Angles ◽  
Combustor Liner ◽  
Compound Angle

An experimental and numerical investigation of adiabatic film cooling effectiveness was conducted on four full-coverage inclined multihole walls with different hole arrangements. The hole geometrical patterns and the test conditions were chosen to be representative of film cooling designs for modern aeroengine combustor liners. The four hole arrangements were grouped into two types based on lateral hole pitch ( P ) and streamwise row spacing ( S ). One type included two test plates which had the same S and P (S/P = 2) and compound angle (β = 0 deg) but different hole inclination angles ( α ) (30 and 150 deg ). The other type included two test plates which had the same S and P (but S/P = 1) and inclination angle (α = 30 deg) but different compound angles (0 deg and 50 deg). Heat-mass transfer analogy method was employed to investigate the adiabatic film cooling effectiveness of these multihole walls with typical blowing ratios for aeroengine combustors. The numerical simulation was performed to characterize the flowfield and temperature distribution, aiming to further understand the film cooling mechanisms. The experimental results indicated that blowing ratio within the range from 1 to 4 had negligible influence on adiabatic film cooling effectiveness (η) in the case of concurrent coolant injection while hole arrangement had large effect on η. But the blowing ratio within the range from 1 to 4 had large effect on the film cooling effectiveness for the counterflow film cooling scheme. The numerical results were compared with experimental data and fairly good agreement was obtained. The numerical simulation revealed the flow structure, particularly exhibiting significant influence of the interaction between mainstream flow and coolant jets on η. With validation by experimental data, film cooling numerical simulation seems quite helpful in selecting optimum multihole arrangement for modern combustor liner design.

Full coverage film cooling using compound angle

2009 ◽  
Vol 337 (6-7) ◽  
pp. 562-572 ◽  
Author(s):  
Brice Michel ◽  
Pierre Gajan ◽  
Alain Strzelecki ◽  
Nicolas Savary ◽  
Azeddine Kourta ◽  
...  
Keyword(s):  
Film Cooling ◽  
Full Coverage ◽  
Compound Angle

Experimental Evaluation of Large Spacing Compound Angle Full-Coverage Film Cooling Arrays: Adiabatic Film Cooling Effectiveness

2016 ◽  
Vol 138 (7) ◽  
Author(s):  
Greg Natsui ◽  
Roberto Claretti ◽  
Mark A. Ricklick ◽  
Jayanta S. Kapat ◽  
Michael E. Crawford ◽  
...  
Keyword(s):  
Film Cooling ◽  
Temperature Sensitive ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Full Coverage ◽  
Inclination Angles ◽  
Temperature Sensitive Paint ◽  
Compound Angle ◽  
Surface Angle

Adiabatic film cooling effectiveness contours are obtained experimentally with the use of temperature sensitive paint (TSP) on low thermal conductivity film cooled surfaces. The effects of blowing ratio, surface angle, and hole spacing are observed by testing four full-coverage arrays composed of cylindrical staggered holes all compounded at 45 deg, which parametrically vary the inclination angles, 30 deg and 45 deg, and the spacing of the holes 14.5 and 19.8 times the diameter. Local film cooling effectiveness is obtained throughout these largely spaced arrays to 23 rows for the 19.8 diameter spacing array and 30 rows for the 14.5 diameter spacing array. The coolant takes several rows to merge and begin to interact with lateral holes at these large spacings; however, at downstream rows the film merges laterally and provides high effectiveness in the gaps between injections. At low blowing, each individual jet remains discrete throughout the array. At higher blowing rates, the profile is far more uniform due to jets spreading as they reattach with the wall. Laterally averaged values of effectiveness approach 0.3 in most cases with some high blowing low spacing, even reaching 0.5.

Second Law Analysis of Aerodynamic Losses: Results for a Cambered Vane With and Without Film Cooling

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Phil Ligrani ◽  
Jae Sik Jin
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Chord Length ◽  
Second Law ◽  
Exergy Destruction ◽  
Freestream Turbulence ◽  
Second Law Analysis ◽  
Aerodynamic Losses

Results of second law analysis of experimentally-measured aerodynamic losses are presented for a cambered vane with and without film cooling, including comparisons with similar results from a symmetric airfoil. Included are distributions of local entropy creation, as well as mass-averaged magnitudes of global exergy destruction. The axial chord length of the cambered vane is 4.85 cm, the true chord length is 7.27 cm, and the effective pitch is 6.35 cm. Data are presented for three airfoil Mex distributions (including one wherein the flow is transonic), magnitudes of inlet turbulence intensity from 1.1% to 8.2%, and ks/cx surface roughness values of 0, 0.00108, and 0.00258. The associated second law aerodynamics losses are presented for two different measurement locations downstream of the vane trailing edge (one axial chord length and 0.25 axial chord length). The surface roughness, when present, simulates characteristics of the actual roughness which develops on operating turbine airfoils from a utility power engine, over long operating times, due to particulate deposition and to spallation of thermal barrier coatings. Quantitative surface roughness characteristics which are matched include equivalent sandgrain roughness size, as well as the irregularity, nonuniformity, and the three-dimensional irregular arrangement of the roughness. Relative to a smooth, symmetric airfoil with no film cooling at low Mach number and low freestream turbulence intensity, overall, the largest increases in exergy destruction occur with increasing Mach number, and increasing surface roughness. Important variations are also observed as airfoil camber changes. Progressively smaller mass-averaged exergy destruction increases are then observed with changes of freestream turbulence intensity, and different film cooling conditions. In addition, the dependences of overall exergy destruction magnitudes on mainstream turbulence intensity and freestream Mach number are vastly different as level of vane surface roughness changes. When film cooling is present, overall mass-averaged exergy destruction magnitudes are significantly less than values associated with increased airfoil surface roughness for both the cambered vane and the symmetric airfoil. Dimensional exergy destruction values (associated with wake aerodynamic losses) for the symmetric airfoil with film cooling are then significantly higher than data from the cambered vane with film cooling, when compared at a particular blowing ratio.

scholarly journals Large Eddy Simulation of Film Cooling Involving Compound Angle Holes: Comparative Study of LES and RANS

Processes ◽  
2021 ◽  
Vol 9 (2) ◽  
pp. 198
Author(s):  
Seung Il Baek ◽  
Joon Ahn
Keyword(s):  
Large Eddy Simulation ◽  
Film Cooling ◽  
Orientation Angle ◽  
Turbulence Models ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blowing Ratio ◽  
Eddy Simulation ◽  
Large Eddy ◽  
Compound Angle

A large eddy simulation (LES) was performed for film cooling in the gas turbine blade involving spanwise injection angles (orientation angles). For a streamwise coolant injection angle (inclination angle) of 35°, the effects of the orientation angle were compared considering a simple angle of 0° and 30°. Two ratios of the coolant to main flow mass flux (blowing ratio) of 0.5 and 1.0 were considered and the experimental conditions of Jung and Lee (2000) were adopted for the geometry and flow conditions. Moreover, a Reynolds averaged Navier–Stokes simulation (RANS) was performed to understand the characteristics of the turbulence models compared to those in the LES and experiments. In the RANS, three turbulence models were compared, namely, the realizable k-ε, k-ω shear stress transport, and Reynolds stress models. The temperature field and flow fields predicted through the RANS were similar to those obtained through the experiment and LES. Nevertheless, at a simple angle, the point at which the counter-rotating vortex pair (CRVP) collided on the wall and rose was different from that in the experiment and LES. Under the compound angle, the point at which the CRVP changed to a single vortex was different from that in the LES. The adiabatic film cooling effectiveness could not be accurately determined through the RANS but was well reflected by the LES, even under the compound angle. The reattachment of the injectant at a blowing ratio of 1.0 was better predicted by the RANS at the compound angle than at the simple angle. The temperature fluctuation was predicted to decrease slightly when the injectant was supplied at a compound angle.

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