KR102294770B1 - Metering Plate for Turbine Blade of Turbine Engine - Google Patents

Abstract

Disclosed invention relates to a metering plate which comprises: a leading edge; a trailing edge; and a turbine airfoil having an airfoil cross-sectional shape and having a pressure surface and a suction surface connecting the leading edge and the trailing edge. With respect to a turbine wing having at least one cooling channel inside the turbine airfoil, the metering plate is installed in an inlet of the cooling channel and has a metering hole connected to the cooling channel. The metering hole is disposed in a center portion of every cooling channel, and includes at least one extension slit extended from an edge of the metering hole. A flow cross-sectional area of individual extension slit is smaller than a flow cross-section of the metering hole.

Description

Metering plate for internal cooling of turbine blades for turbine engines {Metering Plate for Turbine Blade of Turbine Engine}

The present invention relates to a metering plate installed in an inlet of a cooling circuit for internal cooling of a turbine blade provided in a turbine engine.

A turbine engine is a mechanical device that uses the flow of a compressive fluid such as steam or gas to obtain rotational force by impact or reaction force, and includes a steam turbine using steam and a gas turbine using high temperature combustion gas.

Among them, the gas turbine is largely composed of a compressor, a combustor, and a turbine. The compressor is provided with an air inlet for introducing air, and a plurality of compressor vanes and compressor blades are alternately arranged in the compressor casing. Air introduced from the outside is gradually compressed as it passes through a plurality of rotating compressor blades and increases to a target pressure.

The combustor generates high-temperature and high-pressure combustion gas by supplying fuel to the compressed air compressed in the compressor and igniting it with a burner.

In the turbine, a plurality of turbine vanes and turbine blades are alternately arranged in a turbine casing. In addition, the rotor is disposed so as to penetrate the compressor, the combustor, the turbine, and the central portion of the exhaust chamber.

Both ends of the rotor are rotatably supported by bearings. A plurality of disks are fixed to the rotor, each blade is connected to each other, and a drive shaft such as a generator is connected to an end of the exhaust chamber side.

Since these gas turbines do not have a reciprocating mechanism such as a piston of a four-stroke engine, there is no mutual friction part such as a piston-cylinder, so the consumption of lubricant is extremely low. There are advantages.

Briefly describing the operation of the gas turbine, the compressed air in the compressor is mixed with fuel and combusted to produce high-temperature combustion gas, and the combustion gas thus produced is injected toward the turbine. As the injected combustion gas passes through the turbine vanes and turbine blades, a rotational force is generated, thereby rotating the rotor.

There are many factors that affect the efficiency of gas turbines. In recent gas turbine development, research is being conducted in various fields, such as improvement of combustion efficiency in a combustor, improvement of thermodynamic efficiency through an increase in turbine inlet temperature, and improvement of aerodynamic efficiency in compressors and turbines.

Classes of industrial gas turbines for power generation can be divided by turbine inlet temperature (TIT). Currently, G-class and H-class gas turbines are taking the lead, and the most recent gas turbine is J-class. There are also examples of reaching The higher the grade of the gas turbine, the higher both the efficiency and the turbine inlet temperature. Since the H-grade gas turbine has a turbine inlet temperature of 1,500℃, the development of heat-resistant materials and the development of cooling technology are required.

Heat resistant design is required throughout gas turbines, which is particularly important in combustors and turbines where hot combustion gases are generated and flowed. Gas turbine cooling is air-cooled using compressed air produced by a compressor, but in the case of a turbine, the design of cooling is more difficult due to the complex structure in which turbine vanes are fixedly arranged between turbine blades rotating over several stages.

In the case of a turbine blade and a turbine vane (hereinafter, collectively referred to as a “turbine blade”), numerous cooling holes and cooling slits are formed to protect the turbine blade from a high temperature thermal stress environment. can be broadly divided into impact cooling and film cooling. Impulse cooling is when high-pressure compressed air directly collides with the surface of a high-temperature member to cause cooling. Film cooling suppresses heat transfer from a high-temperature environment along with cooling by forming a thin layer of air with very low thermal conductivity on the surface of a member exposed to a high-temperature environment. will do Even in the turbine blade, the inner surface of the turbine blade causes collision cooling, and the outer surface of the turbine blade through which the high temperature combustion gas flows performs complex cooling that causes film cooling, thereby protecting the turbine blade in a high temperature environment.

However, in order to cause impingement cooling on the inner surface of the turbine blade and to form film cooling on the surface, it is necessary to create a cooling channel through which compressed air flows in the cavity inside the turbine blade. In addition, when introducing compressed air into the turbine blade, creating a strong jet stream improves cooling efficiency, so a metering plate restricting the flow cross-sectional area is installed at the flow inlet of the turbine blade to increase the flow rate of the compressed air.

At this time, the metering plate arranges the metering hole in the central region of the cooling channel, which causes a flow stagnant region to occur at the leading edge of the turbine blade, resulting in insufficient cooling performance. In addition, fillets for stress dissipation are often formed in the base portion where the airfoil of the turbine blade starts, and as the thickness increases due to the fillet, cooling the fillet becomes difficult, and insufficient cooling of the fillet is caused by thermal stress. It can also cause premature fatigue.

US Patent No. 8,591,189 (Registered on November 26, 2013)

An object of the present invention is to provide a metering plate capable of suppressing the generation of a region where cooling performance is locally insufficient in a turbine blade.

The present invention is provided with a turbine airfoil having an airfoil cross-sectional shape including a leading edge, a trailing edge, a pressure surface connecting the leading edge and the trailing edge, and a suction surface, and at least one cooling channel inside the turbine airfoil. In the metering plate having a metering hole installed at the inlet of the cooling channel and communicating with the cooling channel for a turbine blade having and at least one extension slit extending therefrom, wherein a flow cross-sectional area of each of the extension slits is smaller than a flow cross-sectional area of the metering hole.

Here, the extension slit extends toward the leading edge.

In addition, a plurality of the extension slits may be provided to radially extend toward the leading edge.

In addition, an end of the extension slit may be extended in a diffuser shape.

In addition, the extension slit may extend toward the pressure surface or the suction surface so as to surround the periphery of the metering hole.

In addition, the extension slit may extend toward the leading edge, the pressure surface, and the suction surface while surrounding the metering hole.

In addition, the flow cross-section of the metering hole may have a rectangular shape.

On the other hand, the present invention relates to an internal cooling structure of a turbine blade for a turbine engine, comprising a turbine airfoil having an airfoil cross-sectional shape including a leading edge, a trailing edge, a pressure surface connecting the leading edge and the trailing edge, and a suction surface a turbine blade having at least one cooling channel in the turbine airfoil; and a metering plate installed at the inlet of the cooling channel and having a metering hole communicating with the cooling channel, wherein the metering hole is disposed in a central portion of each cooling channel and extends from an edge of the metering hole. It provides an internal cooling structure of a turbine blade for a turbine engine comprising the above extension slit, wherein the flow cross-sectional area of each of the extension slits is smaller than the flow cross-sectional area of the metering hole.

And, on the other hand, the present invention is a compressor that sucks in and compresses external air; and a combustor for mixing and burning fuel with the air compressed in the compressor; and a turbine in which a turbine blade and a turbine vane are mounted, and the turbine blade rotates by combustion gas discharged from the combustor, wherein at least one turbine blade of the turbine blade and the turbine vane is disposed in the cooling channel a metering plate installed at the inlet of and a flow cross-sectional area of each of the extension slits is smaller than a flow cross-sectional area of the metering hole.

According to the internal cooling structure of the turbine blade for a turbine engine of the present invention having the above configuration, the turbine blade is prone to flow stagnant region by complexly providing the metering hole and one or more extension slits extending from the edge of the metering hole on the metering plate. It is possible to maintain good cooling performance of the leading edge and/or thick fillet.

In particular, since the turbine blade of a turbine engine is a component that is continuously exposed to a high temperature environment, it can cool the turbine blade well, thereby helping to extend the maintenance interval and lifespan of the turbine engine, and through this, It brings positive results.

1 is a cross-sectional view showing a schematic structure of a turbine engine to which an embodiment of the present invention is applied.
Figure 2 is a view schematically showing the internal cooling structure of the turbine blade for a turbine engine according to the present invention.
3 to 7 are views illustrating various embodiments of a metering plate forming an internal cooling structure of a turbine blade for a turbine engine.

Since the present invention can apply various transformations and can have various embodiments, specific embodiments are illustrated and described in detail in the detailed description. However, this is not intended to limit the present invention to specific embodiments, and it should be understood to include all modifications, equivalents and substitutes included in the spirit and scope of the present invention.

The terms used in the present invention are only used to describe specific embodiments, and are not intended to limit the present invention. The singular expression includes the plural expression unless the context clearly dictates otherwise. In the present invention, terms such as 'comprising' or 'having' are intended to designate that the features, numbers, steps, operations, components, parts, or combinations thereof described in the specification exist, but one or more other features It should be understood that this does not preclude the existence or addition of numbers, steps, operations, components, parts, or combinations thereof.

Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. In this case, it should be noted that in the accompanying drawings, the same components are denoted by the same reference numerals as much as possible. In addition, detailed descriptions of well-known functions and configurations that may obscure the gist of the present invention will be omitted. For the same reason, some components are exaggerated, omitted, or schematically illustrated in the accompanying drawings.

1, an example of a turbine engine 100 to which an embodiment of the present invention is applied is shown. The turbine engine 100 takes a gas turbine as an example, and includes a housing 102 , and a diffuser 106 through which combustion gas passing through the turbine is discharged is provided on the rear side of the housing 102 . In addition, a combustor 104 for receiving and burning compressed air in front of the diffuser 106 is disposed.

Referring to the flow direction of the air, the compressor section 110 is positioned on the upstream side of the housing 102 , and the turbine section 120 is disposed on the downstream side. A torque tube 130 is disposed between the compressor section 110 and the turbine section 120 as a torque transmitting member for transmitting the rotational torque generated in the turbine section to the compressor section.

The compressor section 110 is provided with a plurality of (for example, 14 pieces) compressor rotor disks 140 , and each of the compressor rotor disks 140 is fastened so as not to be spaced apart in the axial direction by the tie rods 150 . .

Specifically, each of the compressor rotor disks 140 are aligned along the axial direction with each other with the tie rod 150 passing through the center. Here, each of the adjacent compressor loader disks 140 are arranged so that the opposite surfaces are compressed by the tie rods 150 so that relative rotation is impossible.

A plurality of blades 144 are radially coupled to the outer circumferential surface of the compressor rotor disk 140 . Each blade 144 has a root portion 146 and is fastened to the compressor rotor disk 140 .

A vane (not shown) fixed to the housing is positioned between each rotor disk 140 . Unlike the rotor disk, the vane is fixed and does not rotate, and serves to align the flow of compressed air passing through the blades of the compressor rotor disk to guide the air to the blades of the rotor disk located on the downstream side.

The fastening method of the root part 146 includes a tangential type and an axial type. This may be selected according to the required structure of a commercially available gas turbine, and may have a commonly known dovetail or fir-tree shape. In some cases, the blade may be fastened to the rotor disk by using a fastening device other than the above type, for example, a fastener such as a key or a bolt.

The tie rod 150 is disposed to pass through the central portion of the plurality of compressor rotor disks 140 , and one end is fastened in the compressor rotor disk located at the upstream side, and the other end is fixed in the torque tube 130 .

Since the shape of the tie rod 150 may have various structures depending on the gas turbine, it is not necessarily limited to the shape shown in FIG. 1 . That is, as shown, one tie rod may have a shape passing through the central portion of the rotor disk, or a plurality of tie rods may have a shape arranged in a circumferential shape, and a mixture thereof may be used.

Although not shown, in the compressor of the gas turbine, a vane serving as a guide vane may be installed at a position next to the diffuser to adjust the flow angle of the fluid entering the combustor inlet to the design flow angle after increasing the fluid pressure. and this is called a deswirler.

The combustor 104 mixes and combusts the introduced compressed air with fuel to produce high-energy, high-temperature, high-pressure combustion gas, and the combustion gas temperature is raised to the limit of heat resistance that the combustor and turbine parts can withstand through an isostatic combustion process.

A plurality of combustors constituting the combustion system of the gas turbine may be arranged in a casing formed in a cell shape, a burner including a fuel injection nozzle, etc., a combustor liner forming a combustion chamber, and a combustor And it is configured to include a transition piece (Transition Piece) that becomes a connection part of the turbine.

Specifically, the liner provides a combustion space in which the fuel injected by the fuel nozzle is mixed with the compressed air of the compressor and combusted. Such a liner may include a flame passage that provides a combustion space in which fuel mixed with air is combusted, and a flow sleeve that surrounds the flame barrel and forms an annular space. In addition, the fuel nozzle is coupled to the front end of the liner, and the igniter is coupled to the side wall.

Meanwhile, at the rear end of the liner, a transition piece is connected to send the combustion gas to the turbine side. The transition piece is cooled by the compressed air supplied from the compressor to prevent damage caused by the high temperature of the combustion gas.

To this end, holes for cooling are provided in the transition piece so that air can be injected into the transition piece, and compressed air flows toward the liner after cooling the body inside through the holes.

Cooling air cooled by the above-described transition piece flows in the annular space of the liner, and compressed air from the outside of the flow sleeve is provided as cooling air through cooling holes provided in the flow sleeve to the outer wall of the liner to collide.

On the other hand, the high-temperature, high-pressure combustion gas from the combustor is supplied to the turbine section 120 described above. As the supplied high-temperature and high-pressure combustion gas expands, it collides with the rotor blades of the turbine and gives a reaction force to cause rotational torque. The power is used to drive a generator, etc.

The turbine section is basically similar in structure to the compressor section. That is, the turbine section 120 is also provided with a plurality of turbine rotor disks 180 similar to the compressor rotor disks of the compressor section. Thus, the turbine rotor disk 180 also includes a plurality of radially arranged turbine blades 184 . The turbine blade 184 may also be coupled to the turbine rotor disk 180 in a dovetail or the like manner. In addition, a vane (not shown) fixed to the housing is provided between the blades 184 of the turbine rotor disk 180 to guide the flow direction of the combustion gas passing through the blades.

2 is a view schematically showing the internal cooling structure of the turbine blade 300 for a turbine engine according to the present invention. Hereinafter, the internal cooling structure of the turbine blade 300 for a turbine engine of the present invention will be described in detail with reference to the accompanying drawings.

Here, the object to which the metering plate 400 of the present invention is applied is a turbine blade 300 for a turbine engine, and the turbine blade 300 includes both a turbine blade, a rotor blade, and a turbine vane, a stator blade, provided in the turbine section 120 . it will be called That is, it should be understood that the present invention can be applied to aerodynamic parts in the form of blades provided in the turbine section 120 .

The turbine blade 300 is an airfoil including a leading edge 312 and a trailing edge 314 , and a pressure surface 316 and a suction surface 318 connecting the leading edge 312 and the trailing edge 314 . A turbine airfoil 310 having a cross-sectional shape is provided. The turbine airfoil 310 serves to induce the flow of high-temperature and high-pressure fluid or to extract power therefrom. In addition, at least one cooling channel 320 is provided inside the turbine airfoil 310 . The cooling channel 320 is a flow path through which a cooling fluid (eg, compressed air) flows. In FIG. 2 , two cooling channels 320 are schematically illustrated. Of course, the number and shape of the cooling channels 320, the flow path structure, etc. are implemented in various embodiments.

A metering plate 400 is provided at the inlet of the cooling channel 320 . The metering plate 400 is installed for the purpose of distributing the cooling fluid for each cooling channel 320 , and increasing the speed of the cooling fluid introduced into the cooling channel 320 to obtain a sufficient collision cooling effect. A metering hole 410 is formed in the metering plate 400 to guide and limit the flow of the cooling fluid as designed.

The metering hole 410 serves to evenly distribute the cooling fluid for each cooling channel 320 by being disposed in the central portion of each cooling channel 320 . However, as the metering hole 410 is disposed in the central region of the cooling channel 320 , a flow stagnant region is generated on the leading edge 312 side of the turbine blade 300 , which tends to cause insufficient cooling performance. . In addition, a fillet for stress dissipation is often formed in the base portion where the turbine airfoil 310 of the turbine blade 300 starts. Insufficient cooling may cause premature fatigue due to thermal stress. The present invention is to solve the problems of such a conventional metering plate.

The metering plate 400 of the present invention includes at least one extension slit 412 extending from an edge of the metering hole 410 . 3 shows an embodiment of a metering plate 400 in which an extension slit 412 is formed. The extension slit 412 basically serves as an auxiliary passage formed for distributing a cooling fluid to a region where cooling performance tends to be insufficient in the turbine blade 300 .

In particular, the flow cross-sectional area of each extension slit 412 is designed to be smaller than the flow cross-sectional area of the metering hole 410 , which ensures that the metering hole 410 responsible for the overall cooling of the turbine blade 300 is wide, while local cooling The extension slit 412 responsible for the flow is to have an appropriate flow rate to have a sufficient flow rate.

In addition, in an embodiment of the present invention, the flow cross-section of the metering hole 410 has a rectangular shape. This is to ensure a sufficient flow area by making the shape of the metering hole 410 as simple and similar to the cross-sectional shape of the individual cooling channels 320 as possible.

And, referring to FIG. 3 , the extension slit 412 extends toward the leading edge 312 in order to solve the problem of lack of cooling performance caused by the formation of a flow stagnant region on the side of the leading edge 312 . Also, FIG. 4 shows an embodiment in which a plurality of extension slits 412 are formed. In this case, the plurality of extension slits 412 radially extend toward the leading edge 312 .

The extension slit 412 extending from the metering hole 410 toward the leading edge 312 supplements the flow rate of the cooling fluid that may be insufficient due to flow stagnation, thereby reducing the cooling performance in the area of the leading edge 312 . is improved In particular, since the flow cross-sectional area of each of the extension slits 412 is small compared to the metering hole 410, the flow rate itself is not biased to the locally biased leading edge 410 region, and the flow rate is fast enough, so the cooling performance over the entire area becomes more uniform. In addition, when the plurality of extension slits 412 are radially extended toward the leading edge 312 , the cooling fluid is supplied in the shape of a brush to effectively cool the entire area of the leading edge 312 .

Meanwhile, FIG. 5 shows an embodiment in which the extension slit 412 extends toward the pressure surface 316 and/or the suction surface 318 to surround the metering hole 410 . This is to additionally cool the fillet area of the pressure surface 316 and the suction surface 318 occupying the largest area in the turbine airfoil 310 . Similar to the case of the extended slit 412 provided toward the leading edge 312 , the extended slit 412 extending close to the wall surfaces of the pressure side 316 and the suction side 318 of the turbine airfoil 310 is The thicker the fillet area, the stronger the cooling effect.

Here, the configuration of the extension slit 412 facing the leading edge 312 and the extension slit 412 facing the pressure surface 316 and/or the suction surface 318 can be combined with each other, FIG. 6 shows the extension It shows a complex configuration in which the slit 412 extends toward the leading edge 312 and the pressure surface 316 and the suction surface 318 while surrounding the metering hole 410 . Detailed design matters such as the number and location of the extension slits 412 and the flow area may be appropriately and variously implemented according to the specifications of the actual turbine blade 300 .

However, in the present invention, the extension slit 412 is not disposed on the trailing edge 314 portion of the turbine airfoil 310, which is the size of the fillet as the trailing edge 314 itself has a very thin thickness. This is because the practical benefit of disposing the separate extension slit 412 is small because it is also small.

And, the end of the extension slit 412, as shown in FIG. 7, may form a kind of funnel shape extending in the form of a diffuser. The end of the extension slit 412 is the portion closest to the wall surfaces of the leading edge 312, the pressure surface 316, and the suction surface 318 of the turbine airfoil 310, and the end of the extension slit 412 is a diffuser. By expanding into the shape, the area where the cooling fluid collides with the wall surface is widened, thereby reducing the occurrence of variations in the cooling effect.

On the other hand, the present invention provides a turbine blade 300 and a turbine engine 100 including the metering plate 400 of the above configuration. The turbine engine 100 is equipped with a compressor 110 that sucks in and compresses external air, a combustor 104 that mixes fuel with the compressed air in the compressor 110 and burns it, and a turbine blade and a turbine vane therein. and a turbine 120 in which the turbine blades are rotated by the combustion gas discharged from the combustor 104 . As described above, when the turbine blade and the turbine vane are collectively referred to as the turbine blade 300 , the above-described metering plate 400 is provided for at least one of the turbine blade and the turbine vane, and is provided on the metering plate 400 . Thickness among the leading edge 312 of the turbine blade 300 and/or the turbine airfoil 310, where a flow stagnation region is likely to occur by forming an extension slit 412 having a smaller flow area for the metering hole 410 to be can maintain good cooling performance of the thick fillet area.

In the above, although an embodiment of the present invention has been described, those of ordinary skill in the art can add, change, delete or add components within the scope that does not depart from the spirit of the present invention described in the claims. It will be possible to variously modify and change the present invention by, etc., such modifications and changes will also be included within the scope of the present invention.

300: turbine blade 310: turbine airfoil
312: leading edge 314: trailing edge
316: pressure side 318: suction side
320: cooling channel 400: metering plate
410: metering hole 412: extension slit

Claims (20)

A turbine airfoil having an airfoil cross-sectional shape including a leading edge, a trailing edge, a pressure surface connecting the leading edge and the trailing edge, and a suction surface, the turbine airfoil having at least one cooling channel inside the turbine airfoil With respect to the turbine blade, in the metering plate is installed in the inlet of the cooling channel is formed a metering hole communicating with the cooling channel,
The metering hole includes at least one extension slit disposed at a central portion of each cooling channel and extending from an edge of the metering hole,
The flow cross-sectional area of each of the extension slits is smaller than the flow cross-sectional area of the metering hole,
The extension slit is a metering plate, characterized in that extending toward at least one of the leading edge, the pressure surface, the suction surface. According to claim 1,
and the extension slit extends toward the leading edge. 3. The method of claim 2,
A metering plate, characterized in that a plurality of extension slits extend radially toward the leading edge. 4. The method according to any one of claims 1 to 3,
The metering plate, characterized in that the end of the extension slit is expanded in the form of a diffuser. According to claim 1,
The extension slit extends toward the pressure surface or the suction surface to surround the metering hole. 6. The method of claim 5,
The metering plate, characterized in that the end of the extension slit is expanded in the form of a diffuser. According to claim 1,
The extension slit surrounds the metering hole and extends toward the leading edge, the pressure surface, and the suction surface. According to claim 1,
A metering plate, characterized in that the flow cross-section of the metering hole forms a rectangular shape. A turbine airfoil having an airfoil cross-sectional shape including a leading edge, a trailing edge, a pressure surface connecting the leading edge and the trailing edge, and a suction surface, wherein at least one cooling channel is provided in the turbine airfoil turbine blades; and
a metering plate installed at the inlet of the cooling channel and having a metering hole communicating with the cooling channel;
The metering hole includes at least one extension slit disposed at a central portion of each cooling channel and extending from an edge of the metering hole,
The flow cross-sectional area of each of the extension slits is smaller than the flow cross-sectional area of the metering hole,
The extension slit is an internal cooling structure of a turbine blade for a turbine engine, characterized in that it extends toward at least one of the leading edge, the pressure surface, and the suction surface. 10. The method of claim 9,
The extension slit is an internal cooling structure of a turbine blade for a turbine engine, characterized in that extending toward the leading edge. 11. The method of claim 10,
The internal cooling structure of the turbine blade for a turbine engine, characterized in that the plurality of extension slits extend radially toward the leading edge. 12. The method according to any one of claims 9 to 11,
The internal cooling structure of the turbine blade for a turbine engine, characterized in that the end of the extension slit is expanded in the form of a diffuser. 10. The method of claim 9,
The extension slit is an internal cooling structure of a turbine blade for a turbine engine, characterized in that it extends toward the pressure surface or the suction surface so as to surround the periphery of the metering hole. 14. The method of claim 13,
The internal cooling structure of the turbine blade for a turbine engine, characterized in that the end of the extension slit is expanded in the form of a diffuser. 10. The method of claim 9,
The extension slit surrounds the metering hole and extends toward the leading edge, the pressure surface, and the suction surface. 10. The method of claim 9,
The internal cooling structure of a turbine blade for a turbine engine, characterized in that the flow cross-section of the metering hole has a rectangular shape. a compressor that sucks in and compresses outside air;
a combustor for mixing fuel with the air compressed in the compressor and burning; and
A turbine blade and a turbine vane are mounted therein, and the turbine blade is rotated by the combustion gas discharged from the combustor; includes,
Here, at least one turbine blade of the turbine blade and the turbine vane,
A turbine airfoil having an airfoil cross-sectional shape including a leading edge, a trailing edge, a pressure surface connecting the leading edge and the trailing edge, and a suction surface, wherein at least one cooling channel is provided in the turbine airfoil and
and a metering plate installed at the inlet of the cooling channel and having a metering hole communicating with the cooling channel,
The metering hole includes at least one extension slit disposed at a central portion of each cooling channel and extending from an edge of the metering hole,
The flow cross-sectional area of each of the extension slits is smaller than the flow cross-sectional area of the metering hole,
The extension slit extends toward at least one of the leading edge, the pressure surface, and the suction surface. 18. The method of claim 17,
The extension slit extends toward the leading edge. 18. The method of claim 17,
The extension slit extends toward the pressure surface or the suction surface so as to surround the periphery of the metering hole. 20. The method of claim 18 or 19,
The end of the extension slit is a turbine engine, characterized in that expanded in the form of a diffuser.

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