US20110097191A1

US20110097191A1 - Method and structure for cooling airfoil surfaces using asymmetric chevron film holes

Info

Publication number: US20110097191A1
Application number: US12/607,586
Authority: US
Inventors: Ronald Scott Bunker
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2009-10-28
Filing date: 2009-10-28
Publication date: 2011-04-28
Also published as: DE102010037050A1; CH702107A2; CH702107B1; CN102052092B; JP5738555B2; CN102052092A; JP2011094609A

Abstract

A film-cooled turbine structure is configured with one or more asymmetric chevron film cooling holes for improving film cooling for a variety of airfoil surfaces or airfoil regions, particularly in regions and applications where the surface fluid streamline curvature is significant.

Description

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH & DEVELOPMENT

This invention was made with U.S. Government Department of Energy support under contract number DE-FC26-05NT42643. The Government has certain rights in the invention.

BACKGROUND

The invention relates generally to film-cooled parts and more particularly to a method of film cooling common locations on airfoil surfaces using asymmetric chevron film holes.
Gas turbines and other high-temperature equipment use film cooling extensively for effective protection of the hot gas path components, such as turbine blades. Film cooling refers to a technique for cooling a part in which cool air is discharged through a plurality of small holes in the external walls of the part to provide a thin, cool barrier along the external surface of the part and prevent or reduce direct contact with hot gasses.
Common locations employed to film cool vane and blade airfoils include, among others, the leading edge, pressure side, and suction side, as well as endwall film cooling, including the inner and outer vane endwalls and the blade platforms. Film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces. This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience. There is significant migration of film cooling across the flow passage, making injection and efficient cooling very difficult.
Injection film holes are generally either round or diffuser shaped. These holes are oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling in others.
In view of the foregoing, it would be advantageous to provide a structure and method of injecting film cooling on a surface in the presence of a strong lateral pressure gradient seeking to move the film coolant away from the intended region to be protected. The structure and method should keep the film coolant in the desired region without creating undue mixing losses caused by simply injecting the flow crossways to the flow of main hot gas.

BRIEF DESCRIPTION

Briefly, in accordance with one embodiment, a film-cooled airfoil or airfoil region is configured with one or more asymmetric chevron film cooling holes.
According to another embodiment, a film-cooled turbine structure comprises at least one asymmetric chevron film cooling hole such that one side of the chevron is dominant over the other side of the chevron with respect to directing a portion of injected coolant onto a surface of the film-cooled turbine structure.

DRAWINGS

These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:

FIG. 1 is a top view illustrating a chevron film-cooling hole known in the art;

FIG. 2 is a side view of the film-cooling hole depicted in FIG. 1;

FIG. 3 is a frontal view of the film-cooling hole depicted in FIG. 1;

FIG. 4 is a perspective view illustrating film-cooling holes disposed in the endwall region of a turbine vane;

FIG. 5 is a perspective view illustrating film-cooling holes disposed in the endwall region of a turbine blade;

FIG. 6 illustrates injection film holes oriented with injection along the approximate direction of the local surface streamline for the endwall region of a turbine vane;

FIG. 7 illustrates an asymmetric chevron film hole suitable for use on airfoils or endwalls; and

FIG. 8 illustrates two asymmetric chevron regions in which each chevron region comprises a pair of wing troughs with dissimilar geometries respect to one another.

While the above-identified drawing figures set forth alternative embodiments, other embodiments of the present invention are also contemplated, as noted in the discussion. In all cases, this disclosure presents illustrated embodiments of the present invention by way of representation and not limitation. Numerous other modifications and embodiments can be devised by those skilled in the art which fall within the scope and spirit of the principles of this invention.

DETAILED DESCRIPTION

Chevron film holes have proven beneficial for improving film effectiveness on airfoil surfaces. Present chevron film holes are always based on symmetric designs about the film hole centerline.
Film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces, as stated herein before. This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience. The airfoil surfaces do however experience such secondary flow effects, but generally to a much lesser degree, except in regions where the airfoils meet the endwall regions. There is significant migration of film cooling across the flow passage, making injection and efficient cooling very difficult.
Injection film holes are generally either round or diffuser shaped. These holes are usually oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling in others.
Asymmetric chevron film hole embodiments achieving similar fluid flow benefits in regions and uses where the surface fluid streamline curvature is significant are described herein. These embodiments, while still using a single circular through-hole, alter the two halves of the chevron footprint to have differing sizes or orientations of troughs. This asymmetry advantageously makes one side of the chevron dominant over the other side with respect to directing a portion of the injected coolant onto the surface to be cooled. The dominant or major side of the chevron should be directed/oriented to counteract the streamline curvature imposed by the hot gases.
A discussion of symmetric chevron film-cooling holes is first presented herein with reference to FIGS. 1-3 to provide a better understanding of the asymmetric film-cooling hole principles and embodiments described thereafter. FIG. 1 is a top view illustrating a symmetric chevron film-cooling hole 10 known in the art. The ridge 12 is outwardly convex laterally in depth between the two wing toughs 14. The convex ridge 12 is arcuate and generally triangular in profile, and diverges in the downstream direction between the inlet bore 16 and the junction of its downstream end with the outer surface 18. The trailing edge of the ridge 12 blends flush with the outer surface 18 along a laterally arcuate downstream end of the chevron outlet, with the convex trailing edge being bowed upstream toward the inlet hole 16. The curved form of the compound chevron film cooling hole 10 enjoys the advantages of compound inclination angles A, B as illustrated in further detail in FIG. 2 that depicts a side view of chevron film-cooling hole 10 relative to the flow of hot gas 20 in which the chevron outlet diverges aft from the inlet hole 16 differently inclined at an inclination angle A. More specifically, inclination angles A and B are the two limiting angles, one along the centerline and the other in each trough.
FIG. 3 is a frontal view of the symmetric film-cooling hole 10 depicted in FIG. 1. This frontal view is in the direction of the hot gas 20 shown in FIG. 2. Chevron film hole 10 is based on a symmetric design about the film hole 10 centerline shown in FIG. 1, and illustrates further details of ridge 12 and wing troughs 14.
FIG. 4 is a perspective view illustrating film-cooling holes disposed in the endwall regions 22 of a turbine vane 24; while FIG. 5 is a perspective view illustrating film-cooling holes 26 disposed in the endwall region of a turbine blade 28. As stated above, film cooling for the endwall regions of turbine airfoils differs from that of the airfoils themselves in that the endwalls experience the complete range of static pressure distribution seen by both the airfoil pressure and suction side surfaces, as stated herein before. The airfoil surfaces do however experience such secondary flow effects, but generally to a much lesser degree, except in regions where the airfoils meet the endwall regions, also stated above. This complete pressure field drives significant secondary flow patterns affecting the injected film cooling that the airfoil surfaces do not experience, causing significant migration of film cooling across the flow passage, making injection and efficient cooling very difficult.
FIG. 6 illustrates round injection film holes oriented with injection along the approximate direction of the local surface streamline for the endwall region of a turbine vane 29. Although round injection film holes are shown in FIG. 6, the injection film holes are generally either round or diffuser shaped. These holes are oriented with injection along the approximate direction of the local surface streamline to minimize mixing losses. This often results in the accumulation of film cooling in certain regions, and the associated lack of film cooling other regions, as stated above.
FIG. 7 is a top view illustrating an asymmetric chevron film hole 30 suitable for use in the endwall regions depicted in FIGS. 4 and 5, according to one embodiment of the invention. A flat ridge 32 is increasing laterally in width between the two wing troughs 34, 36. The ridge 32 is flat and generally triangular in profile, and diverges in the downstream direction between an inlet bore 38 and the junction of its downstream end with the outer surface 40. The trailing edge of the ridge 32 blends flush with the outer surface 40 along a laterally flat downstream end of the chevron outlet. The size of wing trough 34 is different from the size of wing trough 36 such that the wing troughs 34, 36 each blend into the surrounding portions of the inlet bore 38 differently with respect to one another.
A curved form of the asymmetric chevron film cooling hole 30, not shown, having compound inclination angles in which the chevron outlet diverges aft from the inlet bore 38 will enjoy similar advantages to those described above for compound inclination angels B, C shown in FIG. 2 for a symmetric chevron film-cooling hole 10.
Asymmetric chevron film cooling hole 30 is substantially different from symmetric film cooling hole 10 in that particular asymmetric film cooling hole 30 embodiments my include differing tough depths, differing trough widths, differing trough diffusion angles, differing trough shaping, and so on such as depicted in FIG. 8. FIG. 8 illustrates two asymmetric chevron regions in which each chevron region comprises a pair of wing troughs with dissimilar geometries respect to one another.
According to one embodiment, for example, wing trough 34 has a depth that is different from the depth of wing trough 36. According to another embodiment, wing trough 34 has a width that is different from the width of wing trough 36. According to yet another embodiment, wing trough 34 has a diffusion angle B that is different from diffusion angle C. According to still another embodiment, wing trough 34 has a shape that is different from the shape of wing trough 36.
The foregoing asymmetry between the wing troughs 34, 36 then becomes a similar transitional area as with the symmetric chevron, with particular embodiments being flat or multi-planar or asymmetrically arcuate in shape. According to one embodiment, asymmetric chevron film cooling hole 30 employs a single round through-hole feeding the coolant to the chevron region.
In summary explanation, an asymmetric chevron film cooling hole is described herein for improving film cooling for a variety of airfoil surfaces, particularly in regions and applications where the surface fluid streamline curvature is significant. This use of asymmetric film cooling holes allows for more efficient injection of film cooling on a surface in the presence of a strong lateral pressure gradient seeking to move the film coolant away from the intended region to be protected, and keeps the film coolant in the desired region(s) without creating undue mixing losses. Higher mixing losses result from conventional film holes that simply inject the flow crossways to the main hot gas to attempt to counteract the pressure gradients. More efficient coolant use leads to higher efficiency engines such as industrial engines with longer lives.
Asymmetric chevron film cooling holes offer advantages beyond those achievable using known trial and error placement of film holes until a compromise of cooling adequacy and losses is found, or beyond those achievable simply by adding diffuser shaping to round film holes, and possibly a compound angle on the diffuser to help direct coolant in the desired direction. Asymmetric chevron film cooling holes further offer advantages beyond those achievable by simply altering the shape of the endwall itself to help mitigate secondary flows and pressure gradients, rather than modifying the film holes.
While only certain features of the invention have been illustrated and described herein, many modifications and changes will occur to those skilled in the art. It is, therefore, to be understood that the appended claims are intended to cover all such modifications and changes as fall within the true spirit of the invention.

Claims

1. A film-cooled airfoil or airfoil region configured with one or more asymmetric chevron film cooling holes.

2. The film-cooled airfoil or airfoil region according to claim 1, wherein the airfoil region comprises an endwall region of a turbine vane or turbine blade.

3. The film-cooled airfoil or airfoil region according to claim 1, wherein the airfoil comprises a turbine airfoil.

4. The film-cooled airfoil or airfoil region according to claim 1, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar depths with respect to one another.

5. The film-cooled airfoil or airfoil region according to claim 1, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar widths with respect to one another.

6. The film-cooled airfoil or airfoil region according to claim 1, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar diffusion angles with respect to one another.

7. The film-cooled airfoil or airfoil region according to claim 1, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar shapes with respect to one another.

8. The film-cooled airfoil or airfoil region according to claim 1, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar geometries with respect to one another.

9. The film-cooled airfoil or airfoil region according to claim 8, wherein the trough region geometry is selected from one or more of trough depth, tough width, trough diffusion angle, and trough shaping.

10. The film-cooled airfoil or airfoil region according to claim 1, wherein each asymmetric chevron film cooling hole comprises a single round through-hole feeding coolant to a corresponding chevron region.

11. A film-cooled turbine structure comprising at least one asymmetric chevron film cooling hole such that one side of each chevron is dominant over the other side of the chevron with respect to directing a portion of injected coolant onto a surface of the film-cooled turbine structure.

12. The film-cooled turbine structure according to claim 11, wherein the size of one side of each chevron is a different size with respect to the opposite side of the chevron.

13. The film-cooled turbine structure according to claim 11, wherein the orientation of one side of each chevron is oriented differently than the opposite side of the chevron.

14. The film-cooled turbine structure according to claim 11, wherein the dominant side of each chevron is oriented to counteract a streamline curvature imposed by hot gases flowing over the surface of the film-cooled turbine structure.

15. The film-cooled turbine structure according to claim 11, wherein the structure comprises an endwall region of the turbine vane or turbine blade.

16. The film-cooled turbine structure according to claim 11, wherein the structure comprises a turbine airfoil.

17. The film-cooled turbine structure according to claim 11, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar depths with respect to one another.

18. The film-cooled turbine structure according to claim 11, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar widths with respect to one another.

19. The film-cooled turbine structure according to claim 11, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar diffusion angles with respect to one another.

20. The film-cooled turbine structure according to claim 11, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar shapes with respect to one another.

21. The film-cooled turbine structure according to claim 11, wherein at least one asymmetric chevron film cooling hole comprises a first trough region and a second trough region, the first trough region and the second trough region comprising dissimilar geometries with respect to one another.