KR101875683B1 - Gas turbine blade with internal cooling path in discrete multi-cavity rib and rim impingement cooling for enhancing film cooling effectiveness - Google Patents

Abstract

Disclosed is a gas turbine blade having a segmented multicavity concavo-convex structure. The present invention relates to an airfoil blade body having a blade tip, a spiller tip extending radially outwardly along a blade tip from a leading edge to a trailing edge on the blade tip surface, A cavity concavo-convex structure, and a film cooling hole through which a cooling fluid is discharged on the blade tip surface.
According to the present invention, a multi-cavity concave-convex structure is formed on the surface of a blade and a cooling channel is formed in the inside of the blade, so that the blade can uniformly be sprayed over the entire squarer tip of the blade, It is possible to solve the problem that the cooling flow is not injected evenly on the suction side and the vortex that remains in the inner rim of the squealer tip can be formed by collision cooling using the cooling flow path located inside the concave and convex So that it is possible to prevent the flow reattachment and the swirling flow of the main flow to prevent a high heat load.

Description

Technical Field [0001] The present invention relates to a gas turbine blade, and more particularly, to a gas turbine blade using a gas turbine blade with a discrete multi-cavity rib and a rim impingement cooling enhancing film cooling effect,

An object of the present invention is to prevent breakage of a blade tip and increase durability by preventing a heat load generated in a blade tip by improving a film cooling efficiency, and more particularly, To a gas turbine blade that cools the blade tip and rim in an impinging jet fashion through a segmented multi-cavity concavo-convex structure including a membrane cooling hole.

A gas turbine is a heat engine that uses a high-temperature, high-pressure combustion gas to drive a turbine. Generally, it consists of a compressor, a combustor, and a turbine.

The operating process compresses the air using a compressor, moves the compressed air to the combustion chamber, disperses and burns the fuel, and expands and expands the high temperature and high pressure gas generated at this time to rotate the turbine.

The thermal efficiency of such a gas turbine increases with higher efficiency of the compressor and turbine, higher compression pressure, and higher temperature of the incoming gas entering the turbine.

As mentioned above, the gas turbine blades operate mainly in a high temperature environment because the higher the main flow temperature, the higher the turbine efficiency. As a result of operating in a high temperature environment, a high heat load is generated in the gas turbine, There is a problem that frequent breakage occurs.

In order to compensate for this, conventionally, in addition to the squealer tip structure of the blade, a method of cooling the main stream of high temperature by forming a hole through which the cooling fluid can be generated on the side of the blade body has been utilized, And the formation of a swirling flow, there is still a problem that a high heat load is still generated in the blade despite the cooling.

In addition, there is a method of cooling a spiller tip portion of a blade by providing a plurality of circular film cooling holes by another cooling method. In the case where a plurality of holes are provided, the cooling efficiency is high. However, if cooling is performed by simply increasing the number of holes There is a problem that the efficiency of the entire turbine can be lowered as the number of holes increases, and the cooling fluid coming out of the cooling holes can not evenly cool the entire tip surface.

Among the prior art related to gas turbine blades, Korean Patent No. 10-1573409 describes a method of cooling by providing a plurality of cooling holes. However, unlike the present invention to be described later, a hole array is formed outside the blade tip, and there is no description about the concavo-convex structure, which is different from the present invention.

In addition, US-12486249 is similar to the present invention in that it has a spiller tip and a multi-cavity concavo-convex structure similar to the present invention, So that there is a difference from the present invention.

Korean Patent Laid-Open Publication No. 2015-0134375 discloses a cavity structure capable of cooling to the leading edge of a blade by an impinging jet method, but the positions of a plurality of cavities and cooling flow holes are different from the present invention .

That is, in the conventional technology, it is a reality that a preferable method of increasing the cooling efficiency while maintaining the efficiency of the turbine has not been provided.

01: Korean Registered Patent, Registration No. KR 10-1573409, entitled "Gas Turbine Blade Having Membrane Cooling Hole Array Structure Using Backward Spraying" 02: U.S. Published patent application, Application No.: US-12486249, entitled "Turbine blade squealer tip rail with fence members" 03: Korean Patent Publication No. KR 2015-0134375, entitled " Cooling Apparatus for Turbine Blade of Gas Turbine "

The present invention aims at improving the cooling performance of a gas turbine blade, in particular, by preventing heat load acting on the blade tip, thereby reducing damage to the tip and improving reliability.

In addition, the present invention is intended to compensate for the disadvantage that conventional cooling techniques fail to cool the blade rim and tip evenly, resulting in the presence of a vulnerable area near the suction side of the blade tip and the leading edge of the blade.

The technical object of the present invention is not limited to the above-mentioned technical objects and other technical objects which are not mentioned can be clearly understood by those skilled in the art from the following description will be.

In order to achieve the above-mentioned object, the present invention provides a gas turbine blade comprising: a blade body of an air foil; A spiral tip extending in a radial direction along a surface of the blade side surface composed of the pressure surface and the suction surface of the blade body and formed to have a constant thickness; A cavity concavo-convex structure disposed on the blade tip surface; And a film cooling hole located inside the cavity concavo-convex structure and through which cooling fluid is discharged on the blade tip surface,
The cavity convexo-concave structure is arranged at a certain distance so as not to be attached to any one of both rims of the spiral tip of the airfoil and has a segmented structure so that the cooling fluid can be ejected in the direction of the pressure surface and the suction surface of the spiller tip And at least one flow hole is formed on both segmental surfaces of the cavity concavo-convex structure.

In particular, the cavity concavo-convex structure is a vertical rib, and the height of the lip of the cavity concavo-convex structure is the same as the height of the squeaker tip.

In addition, the plurality of film cooling holes are spaced apart along the mean camber line on the blade tip surface.

The cavity concavo-convex structure is characterized in that at least one or more of the cavity concavo-convex structures are arranged in a direction in which the angle between the plane facing the leading edge of the blade and the average camber line is perpendicular to each other at regular intervals.

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Furthermore, the shape of the flow hole is circular or rectangular.

Further, after the cooling fluid hits the cavity concavo-convex structure, the cooling fluid changes its direction and is divided in the direction of the pressure surface and the suction surface of the spiller tip through the flow hole to cool the rim of the spiller tip by the impulsive jet method .

According to the present invention, a multi-cavity concave-convex structure is formed on the surface of a blade and a cooling flow path is formed in the inside thereof. Thus, evenness can be uniformly distributed over the entire squarer tip of the blade, It is possible to solve the problem that the cooling flow is not injected and the vortex remaining in the inner rim portion of the squeaker tip can be formed through the impingement cooling using the cooling flow path located inside the concavo-convex portion, Adherence and swirling flow can be prevented, and high heat load can be prevented.

1 is a perspective view of a conventional gas turbine blade including a cooling hole at a side of a spiller tip and a blade body.
2 is a perspective view and a cross-sectional view of a gas turbine blade composed of a conventional spiller tip and a membrane cooling hole.
3 is a perspective view of a gas turbine blade including a segmented multi-cavity convexo-concave structure.
Figure 4 is a perspective view of a gas turbine blade showing a flow of cooling flow through a segmented multi-cavity concavo-convex structure.
FIG. 5 is a view showing that a cooling flow emerging from a film cooling hole inside the segmented cavity concavo-convex structure is ejected in the rim direction.
Figure 6 compares the efficiency of cooling techniques with an impingement jet method including a conventional cooling technique and a film cooling hole within a segmented cavity concavo-convex structure.
FIG. 7 is a graph comparing heat transfer coefficients of a blade tip and a rim region according to the presence or absence of a multi-cavity concavo-convex structure.
8 is a view showing the flow of the cooling flow to the blades. Particularly, Fig. 8 (a) shows the flow of cooling flow when only the film cooling hole exists on the surface of the conventional blade tip, and Fig. 8 (b) shows the flow of cooling flow when the film cooling hole is included in the multi- As shown in FIG.
9 is a diagram showing the cooling efficiency of the entire blade. Particularly, Fig. 9 (a) shows the cooling efficiency when only the film cooling holes are present on the surface of the conventional blade tip, and Fig. 9 (b) shows the cooling efficiency when the film cooling holes are included in the multi- FIG.

Hereinafter, the present invention will be described in detail with reference to the drawings. It is to be noted that the same elements among the drawings are denoted by the same reference numerals whenever possible. Further, the detailed description of known functions and configurations that may unnecessarily obscure the subject matter of the present invention will be omitted.

It is to be understood that when an element is referred to as being connected or connected to another element, it may be directly connected or connected to the other element, but it should be understood that there may be other elements in between. Further, when a member is referred to as being "on " another member throughout the specification, this includes not only when a member is in contact with another member but also when another member exists between the two members.

In the present application, the term " comprises " or " having " or the like is intended to specify the presence of stated features, integers, steps, operations, elements, parts, or combinations thereof, But do not preclude the presence or addition of other features, numbers, steps, operations, components, parts, or combinations thereof.

1 is a perspective view of a conventional gas turbine blade including a cooling hole at a side of a spiller tip and a blade body.

As a conventional method of cooling a gas turbine blade, there is a method of cooling a hot main stream by arranging a cooling flow hole connected from the inside of the blade to a blade body side 12 including a squealer tip 20 .

However, in the above-described cooling technique, the high-temperature flow is reattached to the tip portion due to the spiller tip included in the blade, and a swirling flow is generated, so that a high heat load is still generated in the blade tip portion there is a problem.

2 is a perspective view and a cross-sectional view of a gas turbine blade composed of a conventional spiller tip and a membrane cooling hole.

In the following description of the position of the gas turbine blades, it is assumed that the root of the blade is referred to as a root, and a side surface composed of the pressure surface 21 and the suction surface 22 is defined along the blade body side surface 12, The area parallel to the root is referred to as the blade tip and tip surface 11.

In addition to the aforementioned cooling methods, a plurality of film cooling holes 40 may be disposed in the blade tip surface 11 so that the cooling flow discharged through the film cooling holes 40 cools the tip and rim 23 of the blade There is also a method.

The cooling method using the plurality of film cooling holes 40 can solve the problem of re-adhering the high-temperature flow of the previous cooling method or generating the swirl flow. However, this method can also solve the problem of uniformly spraying the cooling flow over the blade tip There is a problem that the heat load weak portion 60 exists even in the tip.

The heat load vulnerable region refers to a region having a low cooling efficiency near the suction surface 22 and the blade leading edge of the blade tip of FIG. 9, which will be described later, according to the conventional cooling method. The heat load weak region is a region where the cooling fluid flowing out of the film cooling hole 40 is not uniformly injected, and a risk of corrosion of the blade tip is high due to a heat load acting thereon.

In order to solve the problems of the conventional cooling method, the present invention provides a gas turbine blade comprising a blade body 10 of an airfoil, a pressure surface 21 of the blade body 10 and a suction surface 22 A squeaker tip (20) extending and formed with a constant radial thickness along the blade side surface (12), a cavity concave and convex structure (30) disposed on the blade tip surface (11); And a film cooling hole (40) through which cooling fluid is discharged on the blade tip surface (11).

Means that the spiller tip 20 is formed to extend and form a certain thickness in the radial direction along the surface of the blade side surface 12 composed of the pressure surface 21 and the suction surface 22 of the blade body 10, 2 and FIG. Means that it is formed so as to be connected to the trailing edge at a certain thickness while being divided into two forks along the outermost edge of the blade side 12 from the leading edge of the blade.

Also, the cavity concavo-convex structure 30 may be manufactured in various shapes, but it preferably has a concave-convex structure of a vertical rib shape.

The expression that the cavity concavo-convex structure 30 is a vertical lip refers to a form in which a rectangular column having a large slenderness ratio is laid out on a plane as shown in Figs. 3 to 5, which will be described later. However, the rectangular shape is generally rectangular in cross section, but it may be triangular or semi-circular in cross section according to the embodiment.

In addition, the height of the vertical lip is preferably equal to the height of the squeaker tip 20 of the blade tip surface 11. However, the present invention is not limited thereto, and the height may vary according to the embodiment.

As the cavity convexo-concave structure 30 takes the vertical lip structure, the cooling flow collides with the upper end 31 of the cavity concavo-convex structure, and the flow direction is changed toward the rim 23 of the blade to enable collision jet cooling. However, details related to this will be described later.

In the present invention, a plurality of film cooling holes 40 may be disposed on the blade tip surface 11 at regular intervals along a mean camber line to allow the cooling fluid to flow out of the blades.

Particularly, the film cooling holes 40 may be arranged so that three to five of the film cooling holes 40 are spaced apart from each other on the blade tip surface 11 considering that the cavity concavo-convex structure 30 is arranged. However, the number of the film cooling holes 40 is not limited thereto, but may be arranged with a plurality of rows along an average camber line, as shown in the figure.

The spacing in which the film cooling holes 40 are disposed may vary depending on the arrangement of the cavity concavo-convex structure 30. However, it is preferable that the position of the film cooling hole 40 is located on the average camber line which is the center of the cavity concavity and convexity so that the impinging jet cooling described later can act on the pressure surface 21 and the suction surface 22 side of the blade will be.

However, the film cooling hole 40 may be positioned on the side of the suction surface 22 having a relatively low cooling efficiency, rather than the center of the cavity concavo-convex structure 20, in order to improve the cooling efficiency of the heat load weak region.

In the present invention, the cavity concavo-convex structure 30 may also be arranged on the blade tip surface 11 in a plurality of the same manner as the film cooling holes 40.

It is preferable that the position where the cavity concavo-convex structure 30 is disposed is arranged in a direction in which the angle formed between the average camber line and the face facing the leading edge of the cavity concavo-convex structure 30 is vertical. In another embodiment, the surface facing the leading edge of the concavoconvex structure 30 is not perpendicular to the average camber line, but may be arranged so that the angle between the long edge portion of the rectangle and the average camber line is vertical.

In the multi-cavity concavo-convex structure (30), the number of the cavity concavo-convex structures (30) arranged on the blade surface (11) is the number of the film cooling holes (40) arranged at regular intervals along the average camber line May be 3 to 5 in correspondence with the number of pixels. However, the number of the concavo-convex structures is not limited to the above-described number, and the number of concavo-convex structures can be adjusted according to the embodiment of the present invention.

However, the number of the cavity concave-convex structures 30 and the number of the film cooling holes 40 are not necessarily corresponded to each other. That is, in an embodiment of the invention, only the film cooling hole 40 is disposed on the surface 11 of the blade without the cavity concavo-convex structure 30, so that the number of the film cooling holes 40 is smaller than the number of the cavity concavo- ). ≪ / RTI >

3 is a perspective view of a gas turbine blade including a segmented multi-cavity convexo-concave structure.

Figure 4 is a perspective view of a gas turbine blade showing a flow of cooling flow through a segmented multi-cavity concavo-convex structure.

As one of the key techniques of the present invention, the plurality of cavity concave-convex structures 30 may provide a gas turbine blade arranged to include the film cooling hole 40 therein.

As a result of the inclusion of the film cooling holes 40 in the cavity irregularities 30, a cooling flow is ejected from the blades into the film cooling holes 40 as shown in FIG. 4, 4, the cooling flow moves along the uneven structure.

FIG. 5 is a view showing that a cooling flow emerging from a film cooling hole inside the segmented cavity concavo-convex structure is ejected in the rim direction.

Preferably, the cavity concavo-convex structure 30 is disposed at a certain distance from the rim 23 of the spiral tip 20 of the airfoil and has a segmented structure.

Since the distance between the rim 23 of the spill tip and the concave and convex structure 30 is only a segmented structure, the distance between the concave and convex structures 30 does not matter. However, in one embodiment, When tested, it would be desirable to be spaced apart by 0.005 to 0.02 times the chord line length, which is more efficient than the other spacing.

The distance by the above width is to secure a space in which the cooling fluid is ejected to the rim 23 and is not necessarily limited to this, and if impinging jet cooling is possible in the direction of the pressure surface 21 and the suction surface 22 The distance between the rim 23 and the concave and convex structure 30 is irrelevant.

In order to eject the cooling fluid in the direction of the pressure surface 21 and the suction surface 22 of the squeaker tip 20 in the cavity convexo-concave structure 30, And a flow hole (50).

The shape of the flow hole 50 may be various and preferably circular or rectangular. However, the present invention is not limited thereto, and any shape can be used as long as the cooling flow can be ejected toward the rim 23 of the blade.

Further, although it is common that one of the flow holes 50 is disposed on the segment surface 32, it is also possible to dispose one or more flow holes 50 according to the embodiment. That is, a porous channel shape is also possible.

When the multi-cavity concavo-convex structure 30, which is a key technical feature of the present invention, is segmented by a certain distance from the blade rim 23 and the flow hole 50 is disposed on the segmental surface 32, The cooling fluid coming from the film cooling hole 40 contained in the concave-convex structure 30 collides with the upper end 31 of the concave-convex structure, and then flows toward the segmented surface 32 to change direction. And the cooling fluid may be ejected through the flow holes 50 contained in the segmented surface 32 to cool the rim 23 on both the pressure surface 21 and the suction surface 22 of the blade in an impinging jet fashion .

After the cooling fluid impinges on the rim 23 of the blade tip through the cooling of the impinging jet system above, it is evenly sprayed on the blade surface 11 between the multi-cavity concavo-convex structures 30, And also forms a vortex that remains in the inner rim 23 of the squeaker tip as the cooling flow is circulated in the tip 20, It is possible to prevent the reattachment and swirling flow from occurring.

Figure 6 compares the efficiency of cooling techniques with an impingement jet method including a conventional cooling technique and a film cooling hole within a segmented cavity concavo-convex structure.

As can be seen from the graph of FIG. 6, it can be confirmed that the average cooling efficiency was increased by 260% as compared with the conventional tip and the blade tip having only the circular cooling hole. That is, the technique of including the film cooling hole 40 in the cavity concavo-convex structure 30 of the present invention is superior in cooling effect to the conventional cooling technique.

FIG. 7 is a diagram comparing the heat transfer coefficients of the blade tip and the rim region according to the presence or absence of the multi-cavity convexo-concave structure.

When the heat transfer coefficient is measured while adjusting the number and arrangement angle of the cavity irregularities through the experiment, the heat transfer coefficient of the blade tip and the rim is generally higher during the cooling through the present invention compared to the conventional cooling technique, Able to know.

8 is a view showing the flow of the cooling flow to the blades. Particularly, Fig. 8 (a) shows the flow of cooling flow when only the film cooling hole exists on the surface of the conventional blade, and Fig. 8 (b) shows the flow of cooling flow when the film cooling hole is included in the multi- Fig.

Finally, FIG. 9 is a graph showing the cooling efficiency of the entire blade. Particularly, Fig. 9 (a) shows the cooling efficiency when only the film cooling hole is present on the surface of the conventional blade, and Fig. 9 (b) shows the cooling efficiency when the film cooling hole is included in the multi- to be.

8 to 9, the cooling flow according to the conventional cooling method is such that the cooling flow is not evenly injected in the vicinity of the leading edge of the squeaker tip 20, so that the vicinity of the leading edge and the suction surface 22, There is a problem that a high heat load is generated in the above-mentioned vulnerable portion of the heat load. However, according to the cooling method of the present invention, the cooling flow is evenly distributed to the blade tip and the rim.

Accordingly, the present invention provides a gas turbine blade capable of solving the problems of the conventional blade cooling system, preventing damage to the blade, and securing reliability based on the technical features as described above.

While the present invention has been particularly shown and described with reference to exemplary embodiments thereof, it is to be understood that the invention is not limited to the disclosed exemplary embodiments, It will be understood by those skilled in the art that various changes in form and details may be made therein without departing from the spirit and scope of the present invention.

Furthermore, the terms used in the present invention are used only to describe specific embodiments and are not intended to limit the present invention. The singular expressions include plural expressions unless the context clearly dictates otherwise.

The scope of protection of the present invention should be construed according to the following claims, and all technical ideas within the scope of equivalents should be construed as falling within the scope of the present invention.

10: blade body
11: blade tip surface
12: Blade body side
20: Spiller tips
21: Pressure side
22: suction face
23: rim
30: cavity uneven structure
31: top of concave structure
32:
40: Film cooling hole
50: flow hole

Claims (9)

In a gas turbine blade,
A blade body of an airfoil;
A spiral tip extending in a radial direction along a surface of the blade side surface composed of the pressure surface and the suction surface of the blade body and formed to have a constant thickness;
A cavity concavo-convex structure disposed on the blade tip surface; And
At least one or more film cooling holes which are located inside the cavity concavo-convex structure and through which cooling fluid is discharged on the blade tip surface,
Wherein the cavity concavo-convex structure is arranged at a predetermined distance so as not to be attached to any one of both rims of the spiral tip of the airfoil,
Wherein at least one flow hole is formed on both segmental surfaces of the cavity concavo-convex structure so that the cooling fluid can be ejected in the direction of the pressure surface and the suction surface of the spiller tip.
The method according to claim 1,
Wherein the cavity concavo-convex structure is a vertical rib, and the height of the lip of the cavity concavo-convex structure is equal to the height of the spiller tip.
The method according to claim 1,
Wherein said plurality of film cooling holes are spaced apart along a mean camber line on a blade tip surface.
The method according to claim 1,
Wherein at least one of the cavity concavo-convex structure is disposed at a predetermined interval in a direction in which an angle between a plane that faces the leading edge of the blade and an average camber line is vertical.
delete delete delete The method according to claim 1,
Wherein the shape of the flow hole is circular or square.
The method according to claim 1,
Characterized in that the cooling fluid is deflected in the direction of the pressure surface and the suction surface of the spiller tip through the flow hole by changing the direction after it hits the upper end of the cavity concavo-convex structure to cool the rim of the spiller tip by the impulsive jet method Turbine blades.

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