JPH11193701A - Turbine wing - Google Patents

Abstract

PROBLEM TO BE SOLVED: To eliminate a dead zone of a cooling flow passage and avoid excessive heat partially applied to a wing by providing an aero foil portion with a plurality of inner cooling flow passages each having a turbulence accelerator, a plurality of holes for cooling film, and a leading end cap having an opening formed around the rear edge to communicate with the cooling flow passages. SOLUTION: A wing 10 is provided with a dovetail 12 engaged with a turbine rotor and a hollow aero foil portion 16 provided with a platform portion 14 and a leading end cap 18. The aero foil portion 16 includes inner cooling flow passages having turbulence accelerator therein so as to flow a coolant for cooling. A plurality of holes 20 for cooling film are formed in a positive pressure surface 17, which communicate with the inner cooling flow passages and constitute a path through which cooling air admitted through the opening 32 in a base portion is discharged. In cooperation with an opening 46 of the leading end cap 18 positioned at the front portion of the substantially backmost of the cooling passage, the dead zone allowing no air flow can be eliminated, thus avoiding partial excessive heating.

Description

DETAILED DESCRIPTION OF THE INVENTION

FIELD OF THE INVENTION The present invention relates generally to gas turbine engine blades, and more particularly to turbine blades having improved cooling characteristics. BACKGROUND OF THE INVENTION Gas turbine engines include a compressor that pressurizes air through a combustor where it mixes with fuel and ignites to produce hot combustion gases. The combustion gases flow downstream through one or more stages of turbine blades, where energy is extracted from the combustion gases to provide useful work.

Each turbine blade includes a dovetail,
This causes the wings to be mounted around the rotor disk and an integral hollow airfoil extending radially outward from the dovetail. Because the turbine blades are directly exposed to the hot combustion gases, the turbine blades typically have internal cooling circuits that allow coolant, such as compressor bleed air, to flow through the airfoil of the blades. You. The coolant exits the airfoil through a number of film cooling holes distributed on the surface of the airfoil, thereby creating a thin film of cooling air to protect the airfoil from hot combustion gases.

[0003] The aforementioned cooling circuits are typically arranged in a serpentine configuration, in which cooling air enters the airfoil from a base and passes through a first flow path radially outwardly of the wings. It flows to the tip and then turns 180 ° into the next flow path parallel to the first flow path and flows in the opposite direction towards the base of the airfoil. The air stream typically exits the airfoil after flowing through several such channels.

[0004] Frequently, the cooling flow path adjacent to the trailing edge of the turbine blade is a "stop" flow path, that is, a flow path that does not allow air to flow to the next additional flow path. Airflow is maintained primarily through holes in the trailing edge of the wing in communication with the rearmost cooling flow path. In this arrangement, a reduced "dead zone" of airflow may occur at the radially outermost corner in front of the rearmost cooling flow path. This dead zone can cause local overheating of the wing and cause wing failure.

[0005] Accordingly, there is a need for a turbine blade that avoids any such dead zones and provides improved cooling. SUMMARY OF THE INVENTION In accordance with the present invention, a turbine blade having a base portion, a platform portion, and an airfoil portion is disclosed. The airfoil portion includes a plurality of internal passages having angled turbulence enhancers (or "turbulence means"), a plurality of film cooling holes, and a tip cap having a slot near the trailing edge. The combination of these features improves airfoil cooling and reduces airflow inefficiency around the airfoil.

While the subject matter which is considered as the invention is particularly pointed out and distinctly claimed in the concluding portion of the specification, the invention may best be understood by reference to the following description when taken in conjunction with the accompanying drawings wherein: Will be appreciated. DETAILED DESCRIPTION OF THE INVENTION In reference to the drawings, where like reference numerals represent like elements throughout the various views. FIG. 1 shows a turbine blade 10 according to the present invention. The blade 10 has a dovetail 12, a platform portion 14, which engages a mating dovetail feature in the turbine rotor, and a hollow airfoil portion 16 with a tip cap 18, which is preferably an integral part of the cast. I have. The airfoil has a pressure side 17 and a suction side 19. Wing 10
Also includes an internal cooling passage 22 (see FIG. 2) through which a coolant, such as cooling air, is passed through the airfoil section 16 for cooling. The pressure surface 17 has a plurality of radially aligned film cooling holes 20 which flow through the various internal cooling passages 22 of the blade and are introduced through openings 32 in the base portion of the blade 10. It provides an outlet path for the cooled air.

The wing 10 is made of a suitable high-temperature metal, such as a nickel-base superalloy, which exhibits suitable strength when operated at high temperatures, and is a one-piece casting of a dovetail 12, a platform 14, an airfoil 16 and a tip cap 18. It is preferred to mold as. The platform portion 14 of the wing 10 generally has four corners, a corner 24 in front of the pressure plane,
It has a rectangular surface with a corner 26 behind the pressure side, a corner 28 in front of the suction side, and a corner 30 behind the suction side. Fig. 3-5
As can best be seen, the platform portion 14 of the wing 10 is contoured to improve the flow of hot gas over the platform between adjacent wings. In particular, the corner 2 in front of the pressure surface of the platform portion 14
4 and the angle 30 behind the suction surface are slightly deflected inward, while the angle 26 behind the pressure surface and the angle 28 ahead of the suction surface.
Is slightly deflected outward. As a result, in the assembled turbine rotor, the hot gas entering the turbine stage is reduced by the angle 28 in front of the suction surface of one blade platform.
"Step-down" when passing from the pressure side of the next blade to the front corner 24 of the next blade (see FIG. 4), and again from the corner 26 behind the pressure plane of one blade platform to the rear of the suction surface of the next blade. Again "step down" when passing through the corner 30 (see FIG. 5). With this "platform step" configuration, the hot gas does not undergo an abrupt transition from one platform portion 14 to the next. This feature prevents hot gas from impinging on the edge of the wing platform portion 14 and causing local overheating.

FIG. 2 shows the internal shape of the wing 10.
The wing casting has a number of internal channels 22 to allow cooling air to flow.
Are arranged in a meandering pattern. Cooling air is introduced into the wing through openings 32 in dovetail 12 of wing 10. The air then flows around the meandering channels 22 and exits through the film cooling holes 20 in the pressure surface 17 of the blade. The rearmost cooling channel 34 is supplied with air through its own opening 32 in the base of the wing 10. Unlike the other channels 22, the last channel 34 is "stop", i.e., the radially outer end of the channel 34 is closed in the sense that flow therethrough does not enter the adjacent cooling channel. I have. As can be seen in FIG. 6, the rearmost channel 34 can be radially divided into a front portion 36 and a rear portion 38. Last channel 34
Has a layered array of cylindrical pins 40 extending from the inner wall of the pressure surface 17 to the inner wall of the suction surface 19, which promotes turbulence and provides additional surface area for heat transfer inside the airfoil 10. Provided. The air flow flowing through the rearmost flow path 34 is applied to the film cooling hole 2 on the pressure surface 17 of the blade 10.
0, exhaust through the radially aligned holes 42 in the trailing edge 44 of the wing 10 and through openings 46 in the tip cap 18 out of the wing 10.

Most of the film cooling holes 20 are formed on the pressure surface 1 of the blade 10.
7, one or more cooling holes 48 in the pressure surface 17 of the blade 10 at the radially outwardly forward corner of the rearmost cooling channel 34 (see FIG. 1). Is placed. This additional hole 48 (one or more) cooperates with an opening 46 in the tip cap 18, which is located substantially in the forward portion 36 of the rearmost cooling channel 34, so that airflow is Through the cooling channel 34 to the radially outer end, eliminating a "dead zone" with little or no airflow at the radially outwardly forward corner of the last cooling channel. Make sure that.

The tip cap opening 46 is positioned as forward as possible, substantially in the forward portion 36 of the rearmost channel 34. Due to the limited width at the narrow aft portion of wing 10, opening 46 is preferably non-circular in shape so that its area can be larger than the area of the circular hole. Openings 46 are positioned to communicate with a relatively low pressure air flow outside of wing tip 18. This creates a pressure differential across the opening 46, whereby the cooling air flow is directed to the rearmost cooling channel 3
At 4 it is flowed through a zone that would otherwise be a dead zone.

FIG. 2 also shows a turbulence enhancer or turbulence means 5.
The turbulence enhancer is an elongated rib molded on the inner wall of the cooling channel 22 as part of the wing casting. The turbulence means 50 serves to promote turbulence and increase the cooling efficiency of the blade 10. In the wing design, it is beneficial to keep the pressure drop as low as possible and keep the heat transfer rate high. If the turbulence means is angled, an improvement or reduction of the pressure drop can be expected. Since the pressure drop is proportional to the coefficient of fluid friction, reducing the angle of the turbulent flow means from 90 ° in the direction of flow reduces the flow resistance or friction and reduces the pressure drop. However, when the angle of the turbulence means is moved away from 90 ° in the direction of flow, the heat transfer rate also decreases. Setting this rib at 0 ° (parallel to the air flow) minimizes both heat transfer and friction. In practice, the turbulence means 50 is placed at a non-perpendicular angle to the flow to compromise between optimal heat transfer and minimal flow loss. Non-right angle is about 40 °
A range from about 90 ° to about 90 ° is preferred.

Having described what is considered a preferred exemplary embodiment of the present invention, other modifications of the present invention will be apparent to those skilled in the art in light of the teachings herein. It is hoped that all modifications that fall within the true spirit and scope of the invention are set forth in the following claims.

[Brief description of the drawings]

FIG. 1 is a schematic diagram depicting a turbine blade of the present invention.

FIG. 2 is an elevation view of a side cross section of a wing.

FIG. 3 is a schematic diagram of two adjacent wings.

FIG. 4 is a cross-sectional view illustrating a pattern of airflow around the forward portions of two adjacent wing platforms.

FIG. 5 is a cross-sectional view illustrating a pattern of air flow around the rear portions of two adjacent wing platforms.

FIG. 6 is an enlarged side cross-sectional view of a radially outer rear corner of the wing.

[Explanation of symbols]

 10: turbine blade 12: dovetail 14: platform 16: airfoil 17: positive pressure surface 18: tip cap 19: negative pressure surface 20: film cooling hole 22: internal cooling flow channel 24: positive pressure surface front corner 26: positive pressure surface Back corner 28: Front corner of the suction surface 30: Back corner of the suction surface 32: Opening 34: Cooling channel at the rear 36: Front part 38: Back part 40: Pin 42: Hole 44: Trailing edge 46: Opening 48: Cooling hole 50: Turbulent flow means

Continued on the front page (72) Inventor Robert Cooper Simons Glen Falls Court, Cincinnati, Ohio, USA, 11847

Claims (12)

[Claims] 1. A dovetail for mounting wings to a rotor disk, a platform joined to the dovetail, an airfoil extending outwardly from the platform and including a tip cap and laterally opposed pressure and suction surfaces, and an airfoil. A plurality of internal cooling passages arranged in the foil include an internal cooling passage provided with a front portion and a rear portion in the rearmost cooling passage, and an opening is formed in the tip cap, The turbine blade, wherein the opening is substantially located in the front portion in communication with the rearmost cooling flow path. 2. The turbine blade according to claim 1, wherein said openings are non-circular slots. 3. The turbine blade according to claim 1, wherein said tip cap is an integral part of said airfoil. 4. A plurality of film cooling holes are formed in the airfoil, and at least one of the film cooling holes is formed.
One is positioned on the pressure side in communication with the rearmost cooling flow path, the at least one film cooling hole is positioned substantially in the front portion, and substantially the rearmost cooling flow The turbine blade of claim 1, wherein the turbine blade is located at a radially outermost corner of the path. 5. The turbine blade according to claim 1, wherein the rearmost cooling flow path has a plurality of pins formed therein. 6. The system according to claim 6, wherein said platform is a corner in front of a pressure surface,
A back surface corner, a suction surface front corner, and a suction surface rear corner, wherein the suction surface front corner is disposed radially outward of the pressure surface front corner and the suction surface rear corner. The turbine blade of claim 1, wherein the platform is contoured such that a corner of the platform is disposed radially outward of a corner behind the suction surface. 7. A plurality of turbulence enhancers are formed in the internal cooling passage, and the turbulence enhancers are formed at a non-perpendicular angle to the direction of flow through each cooling passage. The turbine blade according to claim 1, wherein the turbine blade is arranged. 8. The turbine blade according to claim 7, wherein said non-perpendicular angle is between about 40 ° and about 90 °. 9. A dovetail for mounting wings to a rotor disk, a platform joined to the dovetail, and an airfoil extending outwardly from the platform, the airfoil having a tip cap and a laterally opposed pressure surface. And a suction surface, wherein the tip cap is an integral part of the airfoil and the platform defines a corner in front of the pressure surface, a corner in rear of the pressure surface, a corner in front of the suction surface and a corner in back of the suction surface. The suction surface front corner is disposed radially outward of the pressure surface front corner and the pressure surface rear corner is disposed radially outward of the suction surface rear corner. The platform is contoured, and the airfoil has a plurality of internal cooling passages disposed therein, the cooling passage being the last one of the cooling passages. There is a part and a rear part,
An opening is formed in the tip cap, and the opening is substantially located in the front portion in communication with the rearmost cooling channel, and a plurality of film cooling holes are formed in the airfoil. At least one of the film cooling holes is located on the pressure surface in communication with the rearmost cooling flow path, and the at least one film cooling hole is substantially at the front portion. And substantially positioned at a radially outermost corner of the rearmost cooling flow path, wherein a plurality of turbulence enhancers are formed within the internal cooling flow path; and Turbine blades wherein the enhancers are positioned at a non-perpendicular angle to the direction of flow through the respective cooling passages. 10. The turbine blade according to claim 9, wherein said openings are non-circular slots. 11. The turbine blade of claim 9, wherein said non-perpendicular angle is between about 40 ° and about 90 °. 12. The turbine blade according to claim 9, wherein the rearmost cooling passage has a plurality of pins formed therein.