End-Wall Film Cooling Through Fan-Shaped Holes With Different Area Ratios

2006 ◽  
Vol 129 (2) ◽  
pp. 212-220 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi
Keyword(s):  
Secondary Flow ◽  
Mass Flow ◽  
Film Cooling ◽  
Three Dimensional ◽  
Wide Band ◽  
Lateral Spreading ◽  
Area Ratio ◽  
Adiabatic Effectiveness ◽  
Flow Effects ◽  
End Wall

The present paper reports on the aerothermal performance of a nozzle vane cascade, with film-cooled end walls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two end-wall geometries with different area ratios have been compared. Tests have been carried out at low speed (M=0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through three-dimensional (3D) aerodynamic measurements, by means of a miniaturized five-hole probe. Adiabatic effectiveness distributions have been determined by using the wide-band thermochromic liquid crystals technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performances of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

Endwall Film Cooling Through Fan-Shaped Holes With Different Area Ratios

2006 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi
Keyword(s):  
Secondary Flow ◽  
Mass Flow ◽  
Film Cooling ◽  
Lateral Spreading ◽  
Area Ratio ◽  
Hole Area ◽  
Large Injection ◽  
Adiabatic Effectiveness ◽  
Flow Effects ◽  
Endwall Film Cooling

The present paper reports on the aero-thermal performance of a nozzle vane cascade, with film cooled endwalls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two endwall geometries with different area ratios have been compared. Tests have been carried out at low speed (M = 0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. Adiabatic effectiveness distributions have been determined by using the wide banded thermochromic liquid crystals (TLC) technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performance of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

Film Cooling of a Cylindrical Leading Edge With Injection Through Rows of Compound-Angle Holes

2001 ◽  
Vol 123 (4) ◽  
pp. 645-654 ◽  
Author(s):  
Y.-L. Lin ◽  
T. I.-P. Shih
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Lateral Spreading ◽  
Blind Test ◽  
Stagnation Line ◽  
Sst Model ◽  
Sst Turbulence Model ◽  
Adiabatic Effectiveness ◽  
Compound Angle

Computations, based on the k-ω shear-stress transport (SST) turbulence model in which all conservation equations were integrated to the wall, were performed to investigate the three-dimensional flow and heat transfer about a semi-cylindrical leading edge with a flat afterbody that is cooled by film-cooling jets, injected from a plenum through three staggered rows of compound-angle holes with one row along the stagnation line and two rows along ±25 deg. Results are presented for the surface adiabatic effectiveness, normalized temperature distribution, velocity vector field, and surface pressure. These results show the interactions between the mainstream hot gas and the cooling jets, and how those interactions affect surface adiabatic effectiveness. Results also show how “hot spots” can form about the stagnation zone because of the flow induced by the cooling jets. The computed results were compared with experimental data generated under a blind test. This comparison shows the results generated to be reasonable and physically meaningful. With the SST model, the normal spreading was under predicted from 20 to 50 percent. The lateral spreading was over predicted above the surface, but under predicted on the surface. The laterally averaged surface effectiveness was well predicted.

scholarly journals Leading Edge Film Cooling Effects on Turbine Blade Heat Transfer

1995 ◽  
Author(s):  
Vijay K. Garg ◽  
Raymond E. Gaugler
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Mass Flow ◽  
Film Cooling ◽  
Three Dimensional ◽  
Stagnation Temperature ◽  
Flow Ratio ◽  
Adiabatic Effectiveness ◽  
Mass Flow Ratio

An existing three-dimensional Navier-Stokes code (Arnone et al., 1991), modified to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of spanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. It is found that with the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P/d = 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface due to coolant jet lift-off. For P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dimensional.

Fan-Shaped Hole Effects on the Aero-Thermal Performance of a Film-Cooled Endwall

2005 ◽  
Vol 128 (1) ◽  
pp. 43-52 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Benzoni ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi
Keyword(s):  
Secondary Flow ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Thermal Performance ◽  
Film Cooling ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Endwall Cooling ◽  
Injection Rates

The present paper investigates the effects of a fan-shaped hole endwall cooling geometry on the aero-thermal performance of a nozzle vane cascade. Two endwall cooling geometries with four rows of holes were tested, for different mass flow rate ratios: the first configuration is made of cylindrical holes, whereas the second one features conical expanded exits and a reduced number of holes. The experimental analysis is mainly focused on the variations of secondary flow phenomena related to different injection rates, as they have a strong relationship with the film cooling effectiveness. Secondary flow assessment was performed through downstream 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. The results show that at high injection rates, the passage vortex and the 3D effects tend to become weaker, leading to a strong reduction of the endwall cross flow and to a more uniform flow in spanwise direction. This is of course obtained at the expense of a significant increase of losses. The thermal behavior was then investigated through the analysis of adiabatic effectiveness distributions on the two endwall configurations. The wide-banded thermochromic liquid crystals (TLC) technique was used to determine the adiabatic wall temperature. Using the measured distributions of film-cooling adiabatic effectiveness, the interaction between the secondary flow vortices and the cooling jets can be followed in good detail all over the endwall surface. Fan-shaped holes have been shown to perform better than cylindrical ones: at low injection rates, the cooling performance is increased only in the front part of the vane passage. A larger improvement of cooling coverage all over the endwall is attained with a larger mass flow rate, about 1.5% of core flow, without a substantial increase of the aerodynamic losses.

Fan-Shaped Hole Effects on the Aero-Thermal Performance of a Film Cooled Endwall

2005 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Benzoni ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi
Keyword(s):  
Secondary Flow ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Thermal Performance ◽  
Film Cooling ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Endwall Cooling ◽  
Injection Rates

The present paper investigates the effects of a fan-shaped hole endwall cooling geometry on the aero-thermal performance of a nozzle vane cascade. Two endwall cooling geometries with four rows of holes were tested, for different mass flow rate ratios: the first configuration is made of cylindrical holes, whereas the second one features conical expanded exits and a reduced number of holes. The experimental analysis is mainly focused on the variations of secondary flow phenomena related to different injection rates, as they have a strong relationship with the film cooling effectiveness. Secondary flow assessment was performed through downstream 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. The results show that at high injection rates, the passage vortex and the 3D effects tend to become weaker, leading to a strong reduction of the endwall cross flow and to a more uniform flow in spanwise direction. This is of course obtained at the expense of a significant increase of losses. The thermal behavior was then investigated through the analysis of adiabatic effectiveness distributions on the two endwall configurations. The wide banded TLC’s technique was used to determine the adiabatic wall temperature. Using the measured distributions of film cooling adiabatic effectiveness, the interaction between the secondary flow vortices and the cooling jets can be followed in good detail all over the endwall surface. Fan-shaped holes have been shown to perform better than cylindrical ones: at low injection rates, the cooling performance is increased only in the front part of the vane passage. A larger improvement of cooling coverage all over the endwall is attained with a larger mass flow rate, about 1.5% of core flow, without a substantial increase of the aerodynamic losses.

scholarly journals The Design of an Improved Endwall Film-Cooling Configuration

1998 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
High Static Pressure ◽  
Endwall Film Cooling ◽  
Aerodynamic Losses

The endwall film-cooling cooling configuration investigated by Friedrichs et al. (1996, 1997) had in principle sufficient cooling flow for the endwall, but in practice, the redistribution of this coolant by secondary flows left large endwall areas uncooled. This paper describes the attempt to improve upon this datum cooling configuration by redistributing the available coolant to provide a better coolant coverage on the endwall surface, whilst keeping the associated aerodynamic losses small. The design of the new, improved cooling configuration was based on the understanding of endwall film-cooling described by Friedrichs et al. (1996, 1997). Computational fluid dynamics were used to predict the basic flow and pressure field without coolant ejection. Using this as a basis, the above described understanding was used to place cooling holes so that they would provide the necessary cooling coverage at minimal aerodynamic penalty. The simple analytical modelling developed in Friedrichs et al. (1997) was then used to check that the coolant consumption and the increase in aerodynamic loss lay within the limits of the design goal. The improved cooling configuration was tested experimentally in a large scale, low speed linear cascade. An analysis of the results shows that the redesign of the cooling configuration has been successful in achieving an improved coolant coverage with lower aerodynamic losses, whilst using the same amount of coolant as in the datum cooling configuration. The improved cooling configuration has reconfirmed conclusions from Friedrichs et al. (1996, 1997); firstly, coolant ejection downstream of the three-dimensional separation lines on the endwall does not change the secondary flow structures; secondly, placement of holes in regions of high static pressure helps reduce the aerodynamic penalties of platform coolant ejection; finally, taking account of secondary flow can improve the design of endwall film-cooling configurations.

Film Cooling From Novel Sister Shaped Single-Holes

2014 ◽  
Author(s):  
Siavash Khajehhasani ◽  
Bassam Jubran
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Lateral Spreading ◽  
Navier Stokes ◽  
The Novel ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
Lift Off

A numerical investigation of the film cooling performance from novel sister shaped single-holes (SSSH) is presented in this paper and the obtained results are compared with a single cylindrical hole, a forward diffused shaped hole, as well as discrete sister holes. Three types of the novel sister shaped single-hole schemes namely downstream, upstream and up/downstream SSSH, are designed based on merging the discrete sister holes to the primary hole in order to reduce the jet lift-off effect and increase the lateral spreading of the coolant on the blade surface as well as a reduction in the amount of coolant in comparison with discrete sister holes. The simulations are performed using three-dimensional Reynolds-Averaged Navier Stokes analysis with the realizable k–ε model combined with the standard wall function. The upstream SSSH demonstrates similar film cooling performance to that of the forward diffused shaped hole for the low blowing ratio of 0.5. While it performs more efficiently at M = 1, where the centerline and laterally averaged effectiveness results improved by 70% and 17%, respectively. On the other hand, the downstream and up/downstream SSSH schemes show a considerable improvement in film cooling performance in terms of obtaining higher film cooling effectiveness and less jet lift-off effect as compared with the single cylindrical and forward diffused shaped holes for both blowing ratios of M = 0.5 and 1. For example, the laterally averaged effectiveness for the downstream SSSH configuration shows an improvement of approximately 57% and 110% on average as compared to the forward diffused shaped hole for blowing ratios of 0.5 and 1, respectively.

The Effect of Area Ratio Change via Increased Hole Length for Shaped Film Cooling Holes With Constant Expansion Angles

2017 ◽  
Author(s):  
Shane Haydt ◽  
Stephen Lynch ◽  
Scott Lewis
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Area Ratio ◽  
Lateral Expansion ◽  
Blowing Ratio ◽  
High Area ◽  
Expansion Angle ◽  
Ratio Change ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

Shaped film cooling holes are used as a cooling technology in gas turbines to reduce metal temperatures and improve durability, and they generally consist of a small metering section connected to a diffuser that expands in one or more directions. The area ratio of these holes is defined as the area at the exit of the diffuser, divided by the area at the metering section. A larger area ratio increases the diffusion of the coolant momentum, leading to lower average momentum of the coolant jet at the exit of the hole and generally better cooling performance. Cooling holes with larger area ratios are also more tolerant of high blowing ratio conditions, and the increased coolant diffusion typically better prevents jet liftoff from occurring. Higher area ratios have traditionally been accomplished by increasing the expansion angle of the diffuser while keeping the overall length of the hole constant. The present study maintains the diffuser expansion angles and instead increases the length of the diffuser, which results in longer holes. Various area ratios have been examined for two shaped holes: one with forward and lateral expansion angles of 7° (7-7-7 hole) and one with forward and lateral expansion angles of 12° (12-12-12 hole). Each hole shape was tested at numerous blowing ratios to capture trends across various flow rates. Adiabatic effectiveness measurements indicate that for the baseline 7-7-7 hole, a larger area ratio provides higher effectiveness, especially at higher blowing ratios. Measurements also indicate that for the 12-12-12 hole, a larger area ratio performs better at high blowing ratios but the hole experiences ingestion at low blowing ratios. Steady RANS simulations did not accurately predict the levels of adiabatic effectiveness, but did predict the trend of improving effectiveness with increasing area ratio for both hole shapes. Flowfield measurements with PIV were also performed at one downstream plane for a low and high area ratio case, and the results indicate an expected decrease in jet velocity due to a larger diffuser.

Effects on Film Cooling Performance in the Showerhead From Geometric Parameterization of Shaped Hole Designs

2021 ◽  
Author(s):  
Jacob D. Moore ◽  
Christopher C. Easterby ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Thermal Protection ◽  
Leading Edge ◽  
Field Measurements ◽  
High Heat ◽  
Area Ratio ◽  
Cooling Performance ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Marginal Improvement

Abstract The high heat loads at the leading-edge regions of turbine vanes and blades necessitate the most robust thermal protection, typically accomplished via a dense array of film cooling holes, nicknamed the “showerhead.” Although research has shown that film cooling using shaped holes provides more reliable thermal protection than that using cylindrical holes, the effects on cooling performance from varying the geometric details of the shaped hole design are not well characterized. In this study, adiabatic effectiveness and off-the-wall thermal field measurements were conducted for two shaped hole geometries designed as successors to a baseline hole geometry presented in a previous study. One geometry with a 40% increase in area ratio exhibited only a marginal improvement in adiabatic effectiveness (∼10%). A second design with a 12° forward and lateral expansion angle with a breakout area 40% larger performed marginally worse than its matched area ratio counterpart (∼15% lower), suggesting a negative sensitivity to breakout area. Such changes in performance for different shaped hole designs were small compared to the boost in performance gained by switching from a cylindrical hole to a shaped hole, which suggests cooling performance is insensitive to specific shaped hole details provided the exterior coolant flow is well-attached.

An Experimental Study of Heat Transfer and Film Cooling on Low Aspect Ratio Turbine Nozzles

1990 ◽  
Vol 112 (3) ◽  
pp. 488-496 ◽  
Author(s):  
K. Takeishi ◽  
M. Matsuura ◽  
S. Aoki ◽  
T. Sato
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Film Cooling ◽  
Three Dimensional ◽  
Horseshoe Vortex ◽  
Three Dimensional Flow ◽  
Flow Passage ◽  
Pressure Surface ◽  
Low Aspect Ratio ◽  
Complex Phenomena

The effects of the three-dimensional flow field on the heat transfer and the film cooling on the endwall, suction, and pressure surface of an airfoil were studied using a low speed, fully annular, low aspect h/c = 0.5 vane cascade. The predominant effects on the horseshoe vortex, secondary flow, and nozzle wake of increases in the heat transfer and decreases in the film cooling on the suction vane surface and the endwall were clearly demonstrated. In addition, it was demonstrated that secondary flow has little effect on the pressure surface. Pertinent flow visualization of the flow passage was also carried out for better understanding of these complex phenomena. Heat transfer and film cooling on the fully annular vane passage surface are discussed.

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