Gas turbine airfoil film cooling hole

Abstract

A gas turbine airfoil (1) with a pressure sidewall and a suction sidewall (6) comprises several internal cooling passages, through which cooling air flows, and several film cooling holes (7) that extend from the internal cooling passages to the outer surface (11) of the suction sidewall (6). According to the invention the film cooling holes (7) are oriented in the radial outward direction as well as toward the trailing edge of the airfoil (1). They each comprise a diffused section with sidewalls that are diffused in the radial outward direction, in the radial inward direction as well as toward downstream direction. Cooling air flowing through the diffused film cooling hole (7) onto the outer surface (6) of the airfoil (1) forms an air curtain preventing plugging of the film cooling hole by particles entrained in the hot gas flow.

Description

FIELD OF INVENTION

This invention pertains to gas turbine airfoils and in particular to a cooling construction with film cooling holes and the prevention of contamination and plugging of the film cooling holes.

BACKGROUND ART

The airfoils of gas turbines, turbine rotor blades and stator vanes, require extensive cooling in order to keep the metal temperature below a certain allowable level and prevent damage due to overheating from the hot gas flow. Typically, such airfoils are designed with hollow spaces and a plurality of passages and cavities within the airfoil for cooling fluid to flow through. The cooling fluid is typically air bled from the compressor having a higher pressure and lower temperature compared to the gas traveling through the turbine. The higher pressure forces the air through the cavities and passages as it transports the heat away from the airfoil walls. The cooling construction further comprises film cooling holes leading from the hollow spaces within the airfoil to the external surfaces of the leading and trailing edge as well as to the suction and pressure sidewalls. The cooling fluid flows through the film cooling holes to the airfoil outer surface and flows along the outer surfaces forming a film of cooling air of a given penetration depth.

One problem encountered in the design of the gas turbine airfoils is caused by particles entrained in the hot gas flow. During engine operation small particles remain entrained in the gas flow while larger and heavier particles impinge on the airfoil surface and can cause local impact damage. This problem is encountered in particular on the suction side of the airfoil and downstream from the leading edge. The particles can furthermore plug the film cooling holes especially on the suction side. When the particles strike the surface of the airfoil they solidify on the cooled airfoil wall and accumulate between the holes and eventually plug up the holes.

The particles hence not only cause impact damage but can also prevent the air from reaching the airfoil outer surface and cause secondary damage due to loss of film cooling and resultant overheating of the airfoil. Plugging of the film cooling holes is particularly likely if the film cooling holes are oriented in the axial direction as the particles travel in the axial as well as radial direction due to centrifugal forces.

U.S. Pat. No. 5,688,104 discloses an airfoil with a cooling construction comprising internal cavities for the cooling air to flow and film cooling holes with a metering section and a diffusing section. FIG. 4 of the patent disclosure shows an axial diffusion hole with a diffusing angle with respect to the hole axis and in the plane of the shown cross-section that is approximately parallel to the root of the airfoil.

SUMMARY OF INVENTION

It is an object of the invention to provide a cooling construction for the suction side of a gas turbine airfoil that prevents contamination and plugging of film cooling holes by hot particles entrained in the gas flow.

According to the invention a gas turbine airfoil with a pressure sidewall and a suction sidewall extending from a root to a tip and from a leading edge to trailing edge comprises several internal cooling passages within through which cooling air can flow and cool the airfoil. Several film cooling holes lead from the internal cooling passages to the outer surfaces on the suction sidewall providing cooling air to flow onto the airfoil outer surface. The film cooling holes on the suction side of the airfoil each comprise a metering section of cylindrical shape and a diffused section. The sidewalls of the diffused section are diffused with respect to the film cooling hole axis in the radial direction as well as in the downstream direction toward the trailing edge of the airfoil.

The film cooling hole according to the invention causes the cooling air flow exiting from the film cooling holes onto the surface of the suction sidewall to diffuse in the radial as well as in the downstream direction. The downstream direction is defined here as the direction along the tangent of the airfoil at the point of the exit port of the film cooling hole and in the plane perpendicular to the radial direction. The diffusion in the radial as well as in the downstream direction brings about an improved film coverage, which seals the airfoil surface from hot particles in the manner of an air curtain.

Particles in the gas stream that approach a film cooling hole will flow around or over this air curtain. Thus the film cooling hole is protected from particles accumulating in its vicinity and contamination and plugging of the film cooling hole by such particles is prevented.

In a particular embodiment of the invention the sidewalls of the film cooling holes are diffused in the radially outward direction toward the tip of the airfoil as well as in the radially inward direction toward the root of the airfoil. This further improves film coverage between neighboring film cooling hole exit ports in the radial direction.

In a further particular embodiment of the invention the diffusion angle of the sidewall with respect to the film cooling hole axis and in the radially outward direction toward the tip of the airfoil is in the range of 3 to 70, preferably about 5°. This sidewall is diffused with respect to the film cooling hole axis and in the radially inward direction toward the root of the airfoil by an angle in the range of 7 to 12°, preferably about 10°.

In a preferred embodiment of the invention the axis of the film cooling hole is at an angle with respect to the streamwise direction that is in the range of 45 to 55° and preferably about 50°. The streamwise direction is defined as the direction perpendicularly away from the outer surface of the suction side, perpendicular to the radial direction.

The film cooling axis is furthermore oriented at an angle with respect to the downstream direction that is in the range of 35 to 45° and preferably about 40°0. The downstream direction is here defined as the direction along the tangent to the suction side at the point of the exit port of the film cooling hole and pointing away from the leading edge and toward the trailing edge of the airfoil.

In a further particular embodiment of the invention the sidewall of the film cooling hole is diffused with respect to the film cooling axis and in the downstream direction toward the trailing edge of the airfoil by an angle that is in the range of 7 to 12°, preferably about 10°.

In a further embodiment of the invention the film cooling holes are arranged in rows extending radially from the root to the tip of the airfoil. Some of the rows are arranged in the impact zone or so-called strike-zone of the airfoil, which can be approximated by knowing the wheel speed and hot gas axial velocity.

In a further preferred embodiment of the invention the film coverage in the rows of film cooling holes in the strike zone is at least 75%. That is, the ratio of the film hole break-out length to the filk hole spacing is at least 0.75.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 shows view of a gas turbine airfoil and its suction side with several rows of film cooling holes and an enlarged view of the film cooling holes on the suction side of the airfoil,

FIG. 1a shows an enlarged view of the exit ports of the film cooling holes on the suction side of the airfoil,

FIG. 2 shows a cross-section of the airfoil along line II-II and the film cooling holes with their diffused sidewalls in the radial direction,

FIG. 3 a cross-section of the airfoil along line III-III and an film cooling hole with the diffused sidewall in the streamline direction,

FIG. 3a a detailed view of the film cooling hole in the cross-section along line III-III,

FIG. 4 shows a perspective of an individual film cooling hole with diffused sidewalls in the radial and streamline directions.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 shows a typical airfoil 1 of a gas turbine extending from a root 2 to a tip 3 and frown a leading edge 4 to a trailing edge 5. The figure shows the airfoil from its suction side 6. Several film cooling holes 7 are arranged In radially extending rows 8a, 8b, and 8c on the suction side 6. The film cooling holes 7 are realized with diffused sidewalls resulting in an irregularly shaped exit port 9 as shown in an enlarged view in FIG. 1a.

The film cooling holes extend from internal cooling passages within the airfoil 1 through the suction sidewall 6 to its outer surface 11. They provide convective cooling of the sidewall from within as well as film cooling of the sidewall outer surface 11.

In one embodiment of the invention the film cooling holes with diffused sidewalls are arranged in one or more radially extending rows on the suction side of the airfoil.

The number of rows of film cooling holes and their position is determined based on the metal temperature required for the particular airfoil. Some of the rows of film cooling are also placed in the so-called strike zone of the suction side of the airfoil where the particles and in particular the larger particles of approximately 40 mils (approximately 1 mm) diameter, strike the airfoil's suction side. The smaller particles, with a size of 2 mils (approximately 0.05 mm) for example, tend to travel along with the hot gas flow. The larger particles however, impact the airfoil suction side surface downstream of the leading edge and in particular in the radially outer half of the airfoil due to centrifugal force (beyond the 50% span).

The arrangement of the film cooling holes in rows brings about an improved film coverage of this strike zone and prevents heavy particles from damaging the airfoil. The particular design of the film cooling hole with its orientation of the film cooling hole axis and the diffusion of the hole sidewalls result in an air curtain that prevents the plugging of the holes just in this zone.

The rows 8a and 8b are positioned in the so-called strike-zone of the airfoil where the larger and heavier particles entrained in the hot gas flow typically strike the airfoil. Row 8c is placed far from this strike-zone near the trailing edge of the airfoil.

In the rows 8a and 8b the ratio of the film hole break-out length to the film hole spacing is preferably at least 0.75.

FIG. 2 shows a cross-section of the suction sidewall 6 of the airfoil with three of the several film cooling holes 7. They extend from an internal cooling space 10 through the sidewall 6 to the outer surface 11 of the sidewall. The film cooling holes 7 each comprise a first metering section 12 with cooling hole sidewalls 13 that run parallel to the film cooling hole axis 14. They further comprise a second, diffused section 15 designed according to the invention to create an air curtain on the surface 11 suction sidewall and to prevent plugging of the film cooling hole by hot particles. The diffused section 15 has a sidewall 16a that is diffused with respect to the film cooling hole longitudinal axis 14 in the outward radial direction A toward the tip of the airfoil by an angle α. This angle α is in the range of 3 to 7°. A sidewall 16b of the film cooling hole diffused section 15 that is closer to the root of the airfoil is oriented at angle β toward the root of the airfoil in the radially inward direction B with respect to the film cooling hole axis 13. This angle β is in the range of 7 to 12°. Furthermore, the axis 14 of the film cooling hole is oriented at an angle γ with respect to the direction C, which is the steamline direction. The angle γ is in the range of 45 to 55°. Since a heavy particle is generally observed to occur in the outboard 50% airfoil span, the particle trajectory angles can be estimated to be in the range of 45 to 55° with respect to the direction C. The preferable angle for α is 5° outward from the film hole axis 14, the preferable angle for β is 10° inward from the film hole axis 14, and the preferable angle for γ is 50°. With film cooling holes of such dimensions an air curtain is formed in the range of 40 to 55° with respect to the direction C, which is equal or greater than the hot particle trajectory angles.

FIGS. 3 and 3a show a further cross-section of the gas turbine airfoil 1 perpendicular to the cross-section of FIG. 2 and illustrates the diffusion of a sidewall 16c of the film cooling hole 7 on the suction side 6 in the plane shown. Again the longitudinal axis 14 of the film cooling hole is shown. The sidewall 16c closer to the trailing edge 5 of the airfoil 1 is diffused with respect to the film cooling hole axis 14 at an angle δ that is in the range of 7 to 12° and preferably about 10°. The axis 14 of the film cooling hole is oriented at an angle ε with respect to the streamline direction D and toward the trailing edge 5. This angle ε is in the range of 35 to 45°, preferably 40°. The streamline direction D follows the tangent to the airfoil 1 at the point of the exit port of the film cooling hole and in the plane of FIG. 3.

FIG. 4 shows for a better understanding of the “multiple-diffusion” film cooling hole, a perspective view of the hole. It shows the straight-walled metering section and the diffused section with the exit port 9. The diffused sidewalls 16a, 16b are shown with the diffusion angles α, β, with respect to the film cooling hole axis 14.

TERMS USED IN THE FIGURES

• 1 gas turbine airfoil
• 2 root of the airfoil
• 3 tip of the airfoil
• 4 leading edge
• 5 trailing edge
• 6 suction side
• 7 film cooling hole
• 8a,8b,8c rows of film cooling holes
• 9 exit port of film cooling hole
• 10 internal cooling space
• 11 outer surface of suction side
• 12 metering section
• 13 film cooling hole sidewall in metering section
• 14 film cooling hole longitudinal axis
• 15 diffused section
• 16a sidewall diffused in outward radial direction
• 16b sidewall diffused in inward radial direction
• 16c sidewall diffused in streamline direction
• α angle of diffusion in outward radial direction
• β angle of diffusion in inward radial direction
• γ angle of orientation of longitudinal axis
• δ angle of diffusion in streamline direction
• ε angle of orientation of longitudinal axis in streamline direction
• A radial outward direction
• B radial inward direction
• C streamline direction
• D downstream direction

Claims

1. Gas turbine airfoil with a pressure sidewall and a suction sidewall extending from a root to a tip and from a leading edge to a trailing edge of the airfoil comprises several internal cooling passages, through which cooling air flows, and several film cooling holes that extend from the internal cooling passages to the outer surface of the suction sidewall

wherein

the film cooling holes each comprise a metering section of cylindrical shape and a diffused section with sidewalls that are diffused in the radial direction and in the streamline direction both with respect to the longitudinal axis of the film cooling hole.

2. Gas turbine airfoil according to claim 1

wherein

the sidewalls of the film cooling holes are diffused in the radial outward direction toward the tip of the airfoil and in the radial inward direction toward the root of the airfoil.

3. Gas turbine airfoil according to claim 2

wherein

the sidewall closest to the tip of the airfoil is diffused with respect to the longitudinal axis of the film cooling hole in the radially outward direction by an angle that is in the range of 3 to 7°, preferably about 5°, and that the sidewall closest to the root of the airfoil is diffused with respect to the longitudinal axis of the film cooling hole in the radially inward direction by an angle that is in the range of 7 to 12°, preferably about 10°.

4. Gas turbine airfoil according to claim 3

wherein

the axis of the film cooling hole is oriented at an angle with respect to the streamwise direction that is in the range of 45 to 55°, preferably about 50°, where the streamwise direction is perpendicular to the radial direction and perpendicularly away from the outer surface of the suction side of the airfoil

5. Gas turbine airfoil according to claim 4

wherein

the longitudinal axis of the film cooling hole is oriented at an angle with respect to the downstream direction and toward the trailing edge of the airfoil, where the angle ε is in the range of 35 to 45° and preferably about 40°.

6. Gas turbine airfoil according to claim 5

wherein

the sidewall of the film cooling hole that is closer to the trailing edge of the airfoil is diffused at an angle with respect to the longitudinal axis of the film cooling hole that is in the range of 7 to 12° and preferably about 10°.

7. Gas turbine airfoil according to claim 1

wherein

the film cooling holes with diffused sidewalls are positioned in one or more radially extending rows on the suction side of the airfoil.

8. Gas turbine airfoil according to claim 7

wherein some rows of film cooling holes are placed in the strike-zone of the airfoil where heavy particles primarily strike the suction side of the airfoil.

9. Gas turbine airfoil according to claim 8

wherein

the film coverage for the rows of film cooling holes in strike-zone is at least 75%.