Apparatus and method for cooling an airfoil for a gas turbine engine

专利摘要:
A hollow airfoil is formed having a body portion, a trench and a plurality of cooling holes in the trench. The body portion extends like a chord between the leading and trailing ends and spans between the radially outer and radially inner surfaces. The trench is arranged on the outer wall along the tip and extends in the span direction and is aligned with stagnation lines extending along the tip.
公开号:KR19990063130A
申请号:KR1019980055494
申请日:1998-12-16
公开日:1999-07-26
发明作者:마틴 지 타비타；제임스 피 다운스；프리드리히 오 쇠칭；토마스 에이 옥시어
申请人:레비스 스테픈 이；유나이티드 테크놀로지스 코포레이션；
IPC主号:

专利说明:

Apparatus and method for cooling an airfoil for a gas turbine engine
FIELD OF THE INVENTION The present invention relates to cooling rotor blades and / or stator vanes for gas turbines, and more particularly, to apparatus and methods for cooling tip and performing film cooling along the surface of the rotor blades or stator vanes.
In the turbine section of a gas turbine engine, the core gas travels through a plurality of stator vanes and rotor blade stages. Each stator vane or rotor blade has one or more internal cavities surrounded by an outer wall. The suction side and the pressurizing side of the outer wall extend between the front end and the rear end of the airfoil. Stator vane airfoils have an inner platform extending like a span between the outer platform and a rotor blade airfoil extends like a span between the platform and the blade tip.
Hot core gas (with air and combination product) impinging on the airfoil tip will diverge around the suction and pressurization sides of the airfoil, or will be impingement trapped at the tip. The point along the leading end at which the velocity of the core gas flow rate is 0 (collision point) is referred to as a stagnation point. There are stagnation points at all spanwise positions along the tip of the airfoil, and the set of points is called the stagnation line. Air impingement on the tip of the airfoil is subsequently diverted around both sides of the airfoil.
The exact location of each stagnation point along the length of the leading edge is a function of the angle of incidence of the core gas with respect to the cordline of the airfoil, for both the rotor air and the stator airfoil. Besides the angle of incidence, the stagnation point of the rotor airfoil is a function of the rotational velocity of the airfoil and the velocity of the core gas. Given the curvature of the tip, the direction and speed of the approaching core gas, and the rotational speed (if any) of the airfoil, the position of the stagnation point along the tip can be readily determined by means well known in the art. In practical cases, rotor speed and core gas velocity vary with engine operating conditions as a function of time and position along the airfoil span. As a result, the stagnation point along the tip of the airfoil (or the stagnation line where they are gathered) will move relative to the tip.
Cooling air, which is usually extracted at the compressor stage at lower temperatures and pressures than the core gas passing through the turbine section, is used to cool the airfoil. Cooler cooler air provides the heat transfer medium and the pressure differential provides the energy required for the cooling air to pass through the stator stage or rotor stage.
In many cases, it is desirable to achieve film cooling along the surface of the stator or rotor airfoil. The film of cooling air moving along the surface of the airfoil transfers thermal energy away from the airfoil, improves uniform cooling and blocks hot core gas passing through the airfoil. However, those skilled in the art will recognize that film cooling is difficult to achieve and maintain in the harsh environment of a gas turbine. In most cases film cooling air is extracted from cooling holes extending through the outer wall of the airfoil. The term extraction reflects a small pressure difference that causes the cooling air to escape the interior cavity of the airfoil.
One problem associated with using a hole to form a cooled air film is the film's sensitivity to pressure differences across the hole. Excessive pressure differentials across the holes cause air to be injected into the passing core gas rather than assisting in the formation of a film of cooling air. If too small a pressure difference appears across the hole, it will slow down the cooling air flow through the hole or slow the introduction of hot core gas. In both cases, the film cooling effect is adversely affected. Another problem associated with achieving film cooling using holes is that the cooling air is distributed at discrete points along the span of the airfoil rather than along a continuous line. The gaps between the holes and the area near the downstream of these gaps are more sensitive to thermal gradients because they are less exposed to cooling air than in the holes and the spaces near the downstream of the holes. Another problem associated with achieving film cooling using holes is the stress density along the holes. The film cooling effect generally increases when the hole is inclined at an acute angle to the outer surface of the airfoil and is tightly enclosed. However, the slanted, tightly enclosed holes form a stress density.
What is needed is a device that provides adequate cooling along the tip of the airfoil, which accommodates stagnation lines at various locations, forms a uniform and durable cooling air film downstream of the tip on both sides of the airfoil, and within the airfoil wall To form the minimum stress density.
It is an object of the present invention to provide an airfoil with improved cooling capability along the tip.
Another object of the present invention is to provide an airfoil having a leading end cooling device for accommodating a plurality of stagnation lines.
It is a further object of the present invention to provide an airfoil having a tip cooling device which forms a uniform and durable film cooling downstream flow of the tip on both sides of the airfoil.
It is a further object of the present invention to provide an airfoil having a front end cooling device which forms a minimum stress density in the airfoil wall.
According to the present invention, there is provided a hollow airfoil having a body portion, a trench and a plurality of cooling holes disposed in the trench. The body portion extends like a chord between the leading end and the trailing end and spans between the radially inner and outer surfaces and has an outer wall surrounding the inner cavity. The trench is disposed on the outer wall along the tip and extends in the span direction and extends along the tip to align with the stagnation line.
According to one aspect of the invention, a method of cooling an airfoil is provided wherein a trench is disposed on an outer wall of the airfoil. The trench is aligned with the stagnation line relative to the airfoil.
An advantage of the present invention is that uniform and durable film cooling downstream of the tip is provided on both sides of the airfoil. Cooling air is extracted from the trenches on both sides and forms continuous film cooling downstream of the tip. The trench provides more cooling air for film deployment and maintenance by minimizing the cooling loss characteristics of the cooling holes.
Another advantage of the present invention is that the stress is minimized along the area immediately downstream of the tip and the tip. The trench of cooling air that extends continuously along the tip minimizes thermally induced stresses by eliminating discrete cooling points that are separated by the uncooled area characteristics of conventional cooling devices. A uniform film of cooling air entering and exiting both sides of the trench minimizes thermally induced stresses by eliminating the uncooled zones downstream of the cooling hole characteristics of conventional cooling devices and between.
Another advantage of the present invention is that the tip cooling device accommodates a plurality of stagnation lines. In the most preferred embodiment, the trench is preferably at the center of the stagnation line corresponding to the maximum heat load operating conditions for a given application, the width of the trench being such that the stagnation line will not move outside the sidewalls of the trench under all operating conditions. It is preferred that it is wide enough. As a result, the present invention has improved tip cooling and cooling air film formation as compared to conventional cooling devices.
These and other objects, features and advantages will be apparent from the description of the most preferred embodiments as shown in the accompanying drawings.
1 is a schematic perspective view of a turbine rotor blade for a gas turbine engine,
FIG. 2 is a partial cross sectional view of the airfoil of the rotor blade shown in FIG. 1, with a partial cross sectional view of the airfoil of the stator vane having a core gas inlet line showing the relative position of the trench and the stagnation point of the airfoil
3 is a schematic cross-sectional view of a trench disposed at the tip of the airfoil;
Explanation of symbols for main parts of the drawings
10: rotor blade 12: root portion
14: platform 18: trench
24: outer wall 22: inner cavity
28: suction side 30: pressurized side
26: front end 32: rear end
38: cooling hole 40: stream line
42: stall
Referring to FIG. 1, the gas turbine engine turbine rotor blade 10 includes a root portion 12, a platform 14, an airfoil 16, a trench 18 disposed within the airfoil and a blade tip 20. Equipped. The airfoil 16 has one or more inner cavities 22 (see FIG. 2) surrounded by an outer wall 24, at least one of which is closest to the tip 26 of the airfoil 16. The suction side 28 and the pressurizing side 30 of the outer wall extend like a chord between the leading end 26 and the rear end 32 and as a span between the platform 14 and the blade tip 20. . The tip portion 26 has a smoothly curved outline in which the suction side 28 and the pressing side 30 of the airfoil 16 are coupled.
Referring to FIG. 2, the trench 18 has a base bottom 34 at the outer wall 24 and a pair of side walls 36 along the tip 26, preferably near the entire span of the airfoil (FIG. 1). Reference). The plurality of cooling holes 38 form a passage between the trench 18 and the foremost inner cavity 22 containing the cooling air. The shape of the cooling hole 38 and the position in the trench 18 will vary depending on the application. In FIG. 2, there is a stream line 40 indicating the core gas along the core gas path to show the direction of the core gas relative to the airfoil 16.
As noted above, the stagnation point 42 at a predetermined position along the span will move in accordance with direct engine operating conditions. This trench 18 is at the center of the stagnation point 42 corresponding to the maximum heat load operating condition for a given application, and the stagnation line 42 extends out of the sidewall 36 of the trench 18 in all operating conditions. It is preferable that the width of the trench 18 is wide enough so as not to. However, if it is impossible to make the width of the trench 18 wide enough to accommodate the positions of all possible stagnation lines 42, the width 44 and the position of the trench 18 correspond to the maximum heat load operating conditions. It is selected to accommodate the largest number of stagnation lines 42. The most appropriate trench width 44 and depth 46 for a given application can be determined by empirical studies. Referring to FIG. 3, for example, empirical studies show that the trench 18 for the rotor airfoil 16 has a depth 46 and three cooling holes 38 approximately equal to the diameter D of one cooling hole 38. Having a width 44 substantially equal to the diameter 3D of the cooling hole 38, the cooling hole 38 is disposed in the trench 18 and preferably cools the tip portion 26 and forms a downward cooling air film.
In operation according to the invention, the cooling air normally extracted from the compressor stage (not shown) is oriented into the airfoil 16 of the rotor blade 10 (or stator vanes) by means known in the art. do. Cooling air disposed in the inner cavity 22 near the tip 26 of the airfoil 16 is at a lower temperature and pressure than the core gas flowing through the outer wall 24 of the airfoil 16. The pressure difference across the airfoil outer wall 24 causes the internal cooling air to enter the cooling aperture 38 and then pass through the trench 18 disposed in the outer wall 24 along the tip 26. Cooling air entering and exiting the cooling holes 38 is mixed with air already present in the trenches 18 and distributed in the trenches 18. Cooling air continues to enter and exit the entire sidewall 36 of the trench 18 almost uniformly. The entry and exit flow forms a film of cooling air on both sides of the trench 18 extending downward.
One advantage of distributing cooling air in the trench 18 is that the characteristic problems with the pressure differentials of conventional cooling holes (not shown) are minimized. For example, the pressure difference across the cooling aperture 38 is a function of the pressure in the local internal cavity 22 and the local core gas pressure adjacent the aperture 38. These positive pressures vary as a function of time. If the core gas pressure is high and the internal cavity pressure is small near a particular cooling hole of a conventional device (not shown), undesirable hot core gas inflow can occur. The present invention minimizes undesirable inflow opportunities because the cooling air from all holes 38 is uniformly distributed and increased into trenches 18, thereby reducing the chance of low pressure bands occurring. Likewise, the distribution of cooling air in trench 18 may avoid spikes due to the pressure of cooling air that injects cooling air into the core gas rather than adding cooling air to the cooling air downstream membrane in a conventional manner.
While the invention has been shown and described with respect to particular embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail may be made therein without departing from the spirit and scope of the invention. For example, FIG. 2 shows a partial cross-sectional view of airfoil 16. Airfoil 16 may be an airfoil of stator vanes or rotor blades.
According to the invention, the body portion extends like a chord between the tip and the rear end and like a span between the radially outer surface and the radially inner surface, and the trench is disposed on the outer wall along the tip portion and extends in the span direction and extends along the tip portion. Aligned with stagnation lines it is possible to achieve film cooling along the surface of the stator or rotor airfoil.

权利要求:
Claims (20)
[1" claim-type="Currently amended] In the hollow airfoil,
A body part having an outer wall surrounding the inner cavity and a tip extending like a span;
A trench disposed on the outer wall along the tip and extending in the span direction and having a stagnation line extending along the tip;
And a plurality of cooling holes disposed in the trench and extending through the outer wall to provide a cooling air passage between the inner cavity and the trench.
[2" claim-type="Currently amended] The method of claim 1,
The trench includes a first sidewall, a second sidewall, a base extending between the first sidewall and the second sidewall, and the stagnation line is disposed between the first sidewall and the second sidewall.
[3" claim-type="Currently amended] The method of claim 2,
The hollow airfoil is a hollow airfoil that is part of a stator vane.
[4" claim-type="Currently amended] The method of claim 2,
The hollow airfoil is a hollow airfoil is a part of the rotor blade.
[5" claim-type="Currently amended] The method of claim 4, wherein
Each of said cooling holes having a diameter, said trenches having a depth substantially equal to said diameter and a width approximately equal to three times said diameter.
[6" claim-type="Currently amended] The method of claim 1,
The trench includes a first sidewall, a second sidewall, a base extending between the first sidewall and the second sidewall, and the stagnation line is disposed between the first sidewall and the second sidewall.
[7" claim-type="Currently amended] The method of claim 6,
The hollow airfoil is a hollow airfoil that is part of a stator vane.
[8" claim-type="Currently amended] The method of claim 6,
The hollow airfoil is a hollow airfoil is a part of the rotor blade.
[9" claim-type="Currently amended] A method of cooling an airfoil exposed to a core gas in a gas turbine engine, the method of cooling an airfoil having an outer wall surrounding the inner cavity and a body having a tip portion extending like a span.
Providing a trench disposed in the outer wall along the tip,
Aligning the trench with a congestion line;
Providing a plurality of cooling devices disposed within the trenches and penetrating into internal cavities;
Providing cooling air at a lower temperature and higher pressure than the core gas in the inner cavity.
[10" claim-type="Currently amended] The method of claim 9,
And said trench comprises a first sidewall, a second sidewall, and a base extending between said first sidewall and said second sidewall.
[11" claim-type="Currently amended] The method of claim 10,
Determining the stagnation line corresponding to the maximum heat load capacity for a given application;
And centering the trench on the stagnation line corresponding to the maximum heat load capacity for a given application.
[12" claim-type="Currently amended] The method of claim 11,
And the stagnation line lies between the sidewalls under all airfoil operating conditions.
[13" claim-type="Currently amended] The method of claim 10,
The stagnation line moves transversely under different airfoil operating conditions between the first transverse limit and the second transverse limit,
And the trench sidewalls are disposed laterally outside a transverse limit of the stagnation line.
[14" claim-type="Currently amended] The method of claim 10,
And wherein the stagnation lines have a plurality of transverse positions under different airfoil operating conditions, and wherein almost all of the transverse positions of the stagnation lines lie between the first sidewall and the second sidewall of the trench.
[15" claim-type="Currently amended] A method for manufacturing an airfoil of a coolable gas turbine engine having an outer wall, the airfoil surrounding an inner cavity, and a body having a tip extending like a span,
Providing a trench disposed on the outer wall along the tip, the trench having a width and a depth extending laterally;
Determining a congestion line for each of the plurality of selected airfoil operating conditions;
Aligning the trench with the stagnation line;
Providing a plurality of cooling devices disposed within the trenches and penetrating into the interior cavities, wherein the cooling holes provide passages for cooling the air moving between the interior cavities and the trenches.
[16" claim-type="Currently amended] The method of claim 15,
Determining the stagnation line corresponding to the maximum heat load capacity for a given airfoil application;
And centering the trench on the stagnation line corresponding to the maximum heat load capacity.
[17" claim-type="Currently amended] The method of claim 16,
Determining a first transverse limit and a second transverse limit with respect to the stagnation line for the plurality of selected airfoil operating conditions such that the stagnation line is between a first transverse limit and a second transverse limit;
Allowing the trench to have a pair of sidewalls and allowing the width to extend between the sidewalls; and
And the trench sidewalls disposed laterally outside the first and second lateral limits of the stagnation line such that all of the stagnation lines are between the trench sidewalls.
[18" claim-type="Currently amended] The method of claim 16,
Determining a first transverse limit and a second transverse limit with respect to the stagnation line for the plurality of selected airfoil operating conditions such that the stagnation line is between a first transverse limit and a second transverse limit;
Allowing the trench to have a pair of sidewalls and allowing the width to extend between the sidewalls; and
And the trench sidewalls are disposed in the outer wall near the first and second lateral limits of the stagnation line such that all of the stagnation lines are between the trench sidewalls.
[19" claim-type="Currently amended] The method of claim 15,
Determining a first transverse limit and a second transverse limit with respect to the stagnation line for the plurality of selected airfoil operating conditions such that the stagnation line is between a first transverse limit and a second transverse limit;
Allowing the trench to have a pair of sidewalls and allowing the width to extend between the sidewalls; and
And the trench sidewalls disposed transversely in the outer wall outside the first and second lateral limits of the stagnation line such that all of the stagnation lines are between the trench sidewalls.
[20" claim-type="Currently amended] The method of claim 16,
Determining a first transverse limit and a second transverse limit with respect to the stagnation line for the plurality of selected airfoil operating conditions such that the stagnation line is between a first transverse limit and a second transverse limit;
Allowing the trench to have a pair of sidewalls and allowing the width to extend between the sidewalls; and
And the trench sidewalls are disposed in the outer wall near the first and second lateral limits of the stagnation line such that all of the stagnation lines are between the trench sidewalls.

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引用文献:

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公开号 | 申请日 | 公开日 | 申请人 | 专利标题

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法律状态:
1997-12-17|Priority to US08/992,322

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1997-12-17|Priority to US08/992,322

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1997-12-17|Priority to US8/992,322

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1998-12-16|Application filed by 레비스 스테픈 이, 유나이티드 테크놀로지스 코포레이션

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1999-07-26|Publication of KR19990063130A

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2006-08-30|Application granted

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2006-08-30|Publication of KR100581301B1

优先权:

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申请号 | 申请日 | 专利标题

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US08/992,322|US6050777A|1997-12-17|1997-12-17|Apparatus and method for cooling an airfoil for a gas turbine engine|

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US08/992,322|1997-12-17|

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US8/992,322|1997-12-17|