Airfoil and gas turbine including same

Abstract

To disclose an air foil with improved cooling efficiency, and a gas turbine comprising the air foil.SOLUTION: An airfoil comprises: an intake surface and a pressure surface in which a cooling hole is formed; at least one main cavity formed in an internal space formed by the intake surface and the pressure surface, and into which cooling fluid flows; an internal cooling flow path formed inside a wall body forming the intake surface and the pressure surface; a plurality of impingement jet holes that is formed on the inside surfaces of the intake surface and the pressure surface, and forces the cooling fluid into the internal cooling flow path for impingement cooling; and a sub-cavity formed so as to surround an internal cooling flow path outlet formed at an end part of the internal cooling flow path, and configured such that the internal cooling flow path outlet and an inlet of the cooling hole are in communication with each other.SELECTED DRAWING: Figure 4

Description

[Technical Field]

[0001] The present invention relates to an airfoil and a gas turbine including the same. [Background technology]

[0002] A turbine is a mechanical device that uses the flow of compressible fluids such as steam or gas to generate rotational force through impulse or reaction force. Examples include steam turbines that use steam and gas turbines that use high-temperature combustion gases.

[0003] Among these, a gas turbine is broadly composed of a compressor, a combustor, and a turbine. The compressor is equipped with an air inlet for introducing air, and multiple compressor vanes and compressor blades are arranged alternately inside the compressor casing.

[0004] The combustor supplies fuel to the compressed air compressed by the compressor and ignites it with a burner, thereby generating high-temperature and high-pressure combustion gas.

[0005] The turbine has a plurality of turbine vanes and turbine blades arranged alternately inside a turbine casing, and a rotor is arranged to penetrate the center of the compressor, combustor, turbine, and exhaust chamber.

[0006] The rotor is rotatably supported by bearings at both ends, and multiple disks are fixed to the rotor, with the blades connected to each disk, and the drive shaft of a generator or the like is connected to the end of the rotor facing the exhaust chamber.

[0007] Such gas turbines do not have a reciprocating mechanism like the pistons in four-stroke engines, so there is no friction between the piston and cylinder, which means that lubricating oil consumption is extremely low. In addition, the amplitude that is characteristic of reciprocating machines is greatly reduced, allowing for high-speed operation.

[0008] To briefly explain the operation of a gas turbine, air compressed by a compressor is mixed with fuel and burned to produce high-temperature combustion gases, which are then injected into the turbine. The injected combustion gases pass through the turbine vanes and turbine blades, generating rotational force that rotates the rotor. [Prior art documents] [Patent documents]

[0009] [Patent Document 1] Republic of Korea Patent Publication No. 10-2010-0064754 (Title: Cooling blade for gas turbine) Summary of the Invention [Problem to be solved by the invention]

[0010] An object of the present invention is to provide an airfoil with improved cooling efficiency and a gas turbine including the same. [Means for solving the problem]

[0011] An airfoil according to an embodiment of the present invention includes a suction surface and a pressure surface having cooling holes formed therein, at least one main cavity formed in an internal space defined by the suction surface and the pressure surface and into which a cooling fluid flows, an internal cooling passage formed within a wall that forms the suction surface and the pressure surface, a plurality of impingement jet holes formed on the inner surfaces of the suction surface and the pressure surface and that cause the cooling fluid to flow into the internal cooling passage for impingement cooling, and a sub-cavity formed to surround an internal cooling passage outlet formed at an end of the internal cooling passage and that connects the internal cooling passage outlet to the inlet of the cooling hole.

[0012] In the airfoil according to the embodiment of the present invention, a plurality of impingement jet holes may be formed along the span direction on the inner surfaces of the suction surface and the pressure surface.

[0013] In the airfoil according to the embodiment of the present invention, the impingement jet holes may be formed on the trailing edge side, and the internal cooling passage outlets may be formed on the leading edge side.

[0014] In the airfoil according to the embodiment of the present invention, a plurality of sub-cavities may be formed along the span direction on the inner surfaces of the suction surface and the pressure surface.

[0015] The airfoil according to the embodiment of the present invention may include a first cooling fin formed to protrude from one surface of the interior of the wall forming the internal cooling passage toward the other surface and spaced apart from the other surface.

[0016] In the airfoil according to the embodiment of the present invention, the first cooling fin may be formed between any one impingement jet hole and its adjacent impingement jet hole.

[0017] The airfoil according to the embodiment of the present invention may include a cooling protrusion formed below the impingement jet hole, protruding in one direction from the other surface of the wall forming the internal cooling passage.

[0018] The airfoil according to the embodiment of the present invention may include a second cooling fin formed by connecting one surface and the other surface inside the wall forming the internal cooling passage.

[0019] In the airfoil according to the embodiment of the present invention, the second cooling fin may be formed between any one impingement jet hole and its adjacent impingement jet hole.

[0020] The airfoil according to an embodiment of the present invention may further include an impingement cavity that allows the cooling fluid discharged through the internal cooling channel outlet to impinge on a wall surface that forms the suction surface or the pressure surface before being discharged to the outside through the cooling hole.

[0021] A gas turbine according to an embodiment of the present invention includes a turbine including a compressor that compresses incoming air, a combustor that mixes the compressed air from the compressor with fuel and burns the resulting mixture, turbine vanes that generate power using combustion gas from the combustor and guide the combustion gas on a combustion gas path through which the combustion gas passes, and turbine blades that rotate on the combustion gas path by the combustion gas. At least one of the turbine vanes or turbine blades includes an airfoil. The airfoil includes an inlet surface and a pressure surface with cooling holes formed therein, at least one main cavity formed in an internal space defined by the inlet surface and the pressure surface and into which a cooling fluid flows, an internal cooling passage formed within a wall that forms the inlet surface and the pressure surface, a plurality of impingement jet holes formed on inner surfaces of the inlet surface and the pressure surface to direct the cooling fluid into the internal cooling passage for impingement cooling, and a sub-cavity formed to surround an internal cooling passage outlet formed at an end of the internal cooling passage and communicating the internal cooling passage outlet with the inlet of the cooling hole.

[0022] In the gas turbine according to the embodiment of the present invention, a plurality of impingement jet holes may be formed along the span direction on the inner surfaces of the suction surface and the pressure surface.

[0023] In the gas turbine according to the embodiment of the present invention, the impingement jet holes may be formed on the trailing edge side, and the internal cooling flow passage outlets may be formed on the leading edge side.

[0024] In the gas turbine according to the embodiment of the present invention, a plurality of sub-cavities may be formed along the span direction on the inner surfaces of the suction surface and the pressure surface.

[0025] The gas turbine according to the embodiment of the present invention may include a first cooling fin formed to protrude from one surface of the interior of a wall forming an internal cooling flow path toward the other surface and spaced apart from the other surface.

[0026] In the gas turbine according to the embodiment of the present invention, the first cooling fin may be formed between any one of the impingement jet holes and the adjacent impingement jet hole.

[0027] The gas turbine according to the embodiment of the present invention may include a cooling protrusion formed below the impingement jet hole, protruding from the other surface of the wall forming the internal cooling passage toward one surface.

[0028] The gas turbine according to the embodiment of the present invention may include a second cooling fin formed by connecting one surface and the other surface inside the wall forming the internal cooling flow path.

[0029] In the gas turbine according to the embodiment of the present invention, the second cooling fin may be formed between any one impingement jet hole and its adjacent impingement jet hole.

[0030] The gas turbine according to an embodiment of the present invention may further include an impingement cavity that allows the cooling fluid discharged through the internal cooling channel outlet to impinge on a wall surface forming an inlet surface or a pressure surface before being discharged to the outside through the cooling hole.

[0031] Further, specific details of embodiments according to various aspects of the present invention are included in the following detailed description. [Effects of the Invention]

[0032] According to an embodiment of the present invention, the cooling efficiency can be improved by i) cooling by collision of the cooling fluid with the internal cooling channel, ii) increasing the cooling time by internal flow of the internal cooling channel, and iii) increasing the air curtain effect by exhausting through the cooling hole near the leading edge (LE) side. [Brief explanation of the drawings]

[0033] [Figure 1] 1 is a diagram showing the inside of a gas turbine according to an embodiment of the present invention. [Figure 2]1 is a conceptual diagram showing a cross section of a gas turbine according to an embodiment of the present invention. [Figure 3] 1 is a perspective view of a turbine blade including an airfoil according to an embodiment of the present invention; [Figure 4] FIG. 4 is a plan view of the AA' section of FIG. 3 as viewed from above. [Figure 5] 5 is an enlarged plan view of part B of FIG. 4, showing a part of the airfoil according to the first embodiment of the present invention. FIG. [Figure 6] 3A and 3B are diagrams for explaining the cooling effect of the flow of a cooling fluid in the airfoil according to the first embodiment of the present invention. [Figure 7] 5 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a second embodiment of the present invention. FIG. [Figure 8] 5 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a third embodiment of the present invention. FIG. [Figure 9] FIG. 5 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a fourth embodiment of the present invention. [Figure 10] FIG. 5 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a fifth embodiment of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0034] Although the present invention can be implemented in various forms by adding various modifications, specific embodiments will be illustrated and described in detail in the detailed description, but it should be understood that this is not intended to limit the present invention to the specific embodiments, and that the present invention includes all modifications, equivalents, or alternatives that fall within the spirit and technical scope of the present invention.

[0035] The terms used in the present invention are merely used to describe specific embodiments and are not intended to limit the present invention. A singular expression includes a plural expression unless the context clearly indicates otherwise. It should be understood that, in the present invention, terms such as "comprise" or "have" are intended to specify the presence of features, numbers, steps, operations, components, parts, or combinations thereof described in the specification, and do not preclude the possibility of the presence or addition of one or more other features, numbers, steps, operations, components, parts, or combinations thereof.

[0036] Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. It should be noted that in the accompanying drawings, identical components are denoted by the same reference numerals whenever possible. Detailed descriptions of known functions and configurations that may obscure the gist of the present invention will be omitted. For the same reasons, some components in the accompanying drawings may be exaggerated, omitted, or shown schematically.

[0037] FIG. 1 is a partially cutaway perspective view of a gas turbine according to an embodiment of the present invention, and FIG. 2 is a cross-sectional view showing a schematic structure of the gas turbine according to the embodiment of the present invention.

[0038] As shown in FIG. 1 , a gas turbine 1000 according to a first embodiment of the present invention includes a compressor 1100, a combustor 1200, and a turbine 1300. The compressor 1100 includes a plurality of compressor blades 1110 arranged radially. The compressor 1100 rotates the compressor blades 1110, and the rotation of the compressor blades 1110 compresses and moves air. The size and mounting angle of the compressor blades 1110 may vary depending on the mounting position. In the first embodiment, the compressor 1100 is directly or indirectly connected to the turbine 1300, and can receive a portion of the power generated by the turbine 1300 and use it to rotate the compressor blades 1110.

[0039] The air compressed by the compressor 1100 travels to a combustor 1200. The combustor 1200 includes multiple combustion chambers 1210 and fuel nozzle modules 1220 arranged in an annular configuration.

[0040] 2, a gas turbine 1000 according to the first embodiment of the present invention includes a housing 1010, and a diffuser 1400, through which combustion gas that has passed through the turbine is discharged, is provided on the rear side of the housing 1010. A combustor 1200, which receives and combusts compressed air, is disposed in front of the diffuser 1400.

[0041] In terms of the air flow direction, compressor 1100 is located upstream of housing 1010, and turbine 1300 is located downstream. Between compressor 1100 and turbine 1300, torque tube 1500 is located as a torque transmission member that transmits the rotational torque generated by turbine 1300 to compressor 1100.

[0042] The compressor 1100 is provided with a plurality of (for example, 14) compressor rotor disks 1120, and the compressor rotor disks 1120 are fastened together by tie rods 1600 so as not to be spaced apart in the axial direction.

[0043] Specifically, each compressor rotor disk 1120 is aligned axially with a tie rod 1600, which constitutes a rotation axis, penetrating substantially the center of the disk. The opposing surfaces of adjacent compressor rotor disks 1120 are pressed together by the tie rod 1600, and the disks are arranged so that relative rotation is not possible.

[0044] A plurality of compressor blades 1110 are radially coupled to the outer circumferential surface of the compressor rotor disk 1120. Each compressor blade 1110 is fastened to the compressor rotor disk 1120.

[0045] Compressor vanes (not shown) are fixed to the housing and positioned between each rotor disk 1120. The compressor vanes are fixed so as not to rotate unlike the rotor disks, and serve to align the flow of compressed air that has passed through the compressor blades 1110 of the compressor rotor disks 1120 and guide the air to the compressor blades 1110 of the rotor disk 1120 located downstream.

[0046] Tie rod 1600 is arranged to penetrate the center of multiple compressor rotor disks 1120 and turbine rotor disk 1320, and may be made up of one or more tie rods. One end of tie rod 1600 is fastened into the compressor rotor disk located most upstream, and the other end of tie rod 1600 is fastened by fixing nut 1450.

[0047] The shape of the tie rod 1600 may have various structures depending on the gas turbine, and is not necessarily limited to the shape shown in Fig. 2. That is, as shown in the figure, it may have a shape in which one tie rod penetrates the center of the rotor disk, or it may have a shape in which multiple tie rods are arranged circumferentially, or a combination of these may be used.

[0048] Although not shown, the compressor of the gas turbine may be provided with a vane that acts as a guide vane located next to the diffuser in order to adjust the flow angle of the fluid that enters the combustor inlet after increasing the pressure of the fluid to the design flow angle; this is called a deswirler.

[0049] The combustor 1200 mixes the incoming compressed air with fuel and burns it to produce high-energy, high-temperature, high-pressure combustion gases, and raises the temperature of the combustion gases through a constant-pressure combustion process to the heat limit that the combustor and turbine components can withstand.

[0050] The combustors constituting the combustion system of a gas turbine may be arranged in a plurality of units in a housing formed in a cellular form, and are composed of a burner including a fuel injection nozzle, a combustor liner forming a combustion chamber, and a transition piece connecting the combustor and the turbine.

[0051] Specifically, the liner provides a combustion space where fuel injected by a fuel nozzle is mixed with compressed air from the compressor and burned. Such a liner may include a flame tube that provides the combustion space where the fuel mixed with air is burned, and a flow sleeve that surrounds the flame tube to form an annular space. In addition, a fuel nozzle is coupled to the front end of the liner, and a spark plug is coupled to the side wall.

[0052] A transition piece is connected to the rear end of the liner so that combustion gases burned by the spark plug can be sent to the turbine side. The outer wall of this transition piece is cooled by compressed air supplied from the compressor to prevent damage due to the high temperature of the combustion gases.

[0053] For this reason, the transition piece is provided with cooling holes that allow air to be injected into the interior, and the compressed air passes through the holes to cool the main body inside before flowing to the liner side.

[0054] Cooling air that has cooled the transition piece flows through the annular space of the liner, and compressed air is provided from outside the flow sleeve through cooling holes provided in the flow sleeve and collides with the outer wall of the liner as cooling air.

[0055] Meanwhile, high-temperature, high-pressure combustion gases discharged from the combustor are supplied to the turbine 1300. As the supplied high-temperature, high-pressure combustion gases expand, they collide with the turbine rotors, generating a reaction force and rotational torque. The rotational torque thus obtained is transmitted to the compressor via the torque tube 1500, and any power exceeding that required to drive the compressor is used to drive a generator, etc.

[0056] The turbine 1300 is basically similar in structure to a compressor. That is, the turbine 1300 also includes a plurality of turbine rotor disks 1320 similar to the compressor rotor disks of a compressor. Therefore, the turbine rotor disks 1320 also include a plurality of turbine blades 1310 arranged radially. The turbine blades 1310 can be connected to the turbine rotor disk 1320 by a method such as a dovetail. In addition, turbine vanes 1330 fixed to the turbine casing 1350 are provided between the turbine blades 1310 to guide the flow direction of the combustion gas that has passed through the turbine blades 1310.

[0057] An airfoil according to an embodiment of the present invention may be an airfoil applied to at least one of a compressor blade 1110, a compressor vane, a turbine blade 1310, and a turbine vane 1330. In the following description, an airfoil applied to a turbine blade 1310 of a gas turbine will be described as an example. Furthermore, the technical concept described in this specification is not limited to gas turbines, and can be applied to devices equipped with airfoils, such as steam turbines.

[0058] FIG. 3 is a perspective view of a turbine blade including an airfoil according to an embodiment of the present invention, and FIG. 4 is a plan view taken along the line AA' of FIG. 3 and viewed from above.

[0059] A turbine blade 1310 according to an embodiment of the present invention includes a root portion 1312 and an airfoil 2000 .

[0060] Referring to FIG. 3, the turbine blade 1310 is mounted on a turbine rotor disk 1320 to rotate the turbine using high-pressure combustion gas. A root portion 1312 is formed on the lower side to be connected to the turbine rotor disk 1320, and an airfoil 2000 that rotates using air pressure is integrally connected to the upper side of the root portion 1312, so that the turbine rotates using the pressure difference between the front and rear surfaces of the airfoil 2000.

[0061] The outer surface of the root portion 1312 is formed with a shank and a platform that protrude outward to provide a strong fixation. The root portion 1312 is formed with an inlet 1312a through which a cooling fluid flows into the airfoil 2000. The cooling fluid is a portion of the compressed air compressed by the compressor 1100, and is supplied from the compressor 1100 to the root portion 1312 of the turbine blade 1310, and flows into the airfoil 2000 through the inlet 1312a to cool the turbine blade 1310. Alternatively, the cooling fluid is supplied to the root portion 1312 through an internal flow path (not shown) connecting the compressor 1100 to the turbine 1300, and flows into the airfoil 2000 through the inlet 1312a to cool the turbine blade 1310.

[0062] The airfoil 2000 is disposed above the root portion 1312. On the other hand, when the airfoil 2000 is formed on the turbine vane 1330, the airfoil 2000 is formed between an outer shroud and an inner shroud, and the cooling fluid flows in through a cooling fluid flow passage formed on the outer shroud side or a cooling fluid flow passage formed on the inner shroud side.

[0063] The airfoil 2000 has a suction side (2001) that protrudes outward and curves at the front where the combustion gas flows in, and a pressure side (2002) that is recessed toward the suction side (2001) at the rear, thereby maximizing the pressure difference between the front and rear of the airfoil 2000 and ensuring smooth airflow.

[0064] The airfoil 2000 includes a leading edge (LE) and a trailing edge (TE) that are ends where the pressure surface 2002 and the suction surface 2001 meet, and the leading edge LE refers to the end of the front part of the airfoil 2000 that meets the fluid flowing through the airfoil 2000, and the trailing edge TE refers to the end of the rear part of the airfoil 2000. The direction from the root part toward the tip of the airfoil is called the span direction.

[0065] The airfoil 2000 includes a plurality of cooling holes 2003 formed through the suction surface 2001 or the pressure surface 2002. The cooling fluid is injected through the cooling holes 2003 and acts like an air curtain on the outer surface of the airfoil, thereby cooling the outer surface of the airfoil 2000 by a so-called film cooling method.

[0066] Referring to FIG. 4, an airfoil 2000 according to an embodiment of the present invention includes a main cavity 2100, a cooling channel portion 2200, and a sub-cavity 2300.

[0067] At least one main cavity 2100 may be formed in an internal space formed by the suction surface 2001 and the pressure surface 2002 of the airfoil 2000. The main cavity 2100 may be formed along the span direction. When a plurality of main cavities 2100 are formed, the main cavity 2100 may be divided into a plurality of cavities along the longitudinal direction extending from the leading edge LE to the trailing edge TE. Although the drawing shows an example in which two main cavities are formed, the present invention is not limited to this.

[0068] The cooling fluid flowing through the main cavity 2100 flows into and through the cooling channel portion 2200, thereby cooling the suction surface 2001 and the pressure surface 2002. The cooling channel portion 2200 is formed inside the wall forming the suction surface 2001 and the pressure surface 2002. The cooling channel portion 2200 can be realized in various forms, which will be described later with reference to FIGS. 5 to 10.

[0069] The sub-cavities 2300 are formed to protrude by a predetermined size from the inner surfaces of the suction surface 2001 and the pressure surface 2002. Similar to the main cavity 2100, a plurality of sub-cavities 2300 are formed along the span direction on the inner surfaces of the suction surface 2001 and the pressure surface 2002. The sub-cavities 2300 are formed to surround the cooling channel outlet (2230, see FIG. 5) and the cooling hole inlet (2003a, see FIG. 5), so that the cooling channel outlet 2230 and the cooling hole inlet 2003a are fluidly isolated from the main cavity 2100.

[0070] FIG. 5 is an enlarged plan view of part B of FIG. 4, showing a part of the airfoil according to the first embodiment of the present invention, and FIG. 6 is a diagram for explaining the cooling effect due to the flow of cooling fluid in the airfoil according to the first embodiment of the present invention.

[0071] Referring to FIG. 5, the cooling channel portion 2200 includes an impingement jet hole 2210 , an internal cooling channel 2220 , and a cooling channel outlet 2230 .

[0072] A plurality of impingement jet holes 2210 may be formed along the span direction on the inner surfaces of the suction surface 2001 and the pressure surface 2002. The cooling fluid that flows in through the inlet 1312a of the root portion 1312 can flow into the impingement jet holes 2210 while flowing along the main cavity 2100 in the span direction.

[0073] Internal cooling passage 2220 is formed inside the wall forming suction surface 2001 and pressure surface 2002. Internal cooling passage 2220 may be formed inside the wall extending in the longitudinal direction of the wall (the direction connecting the leading edge and the trailing edge).

[0074] A cooling channel outlet 2230 is formed at the end of the internal cooling channel 2220. The cooling channel outlet 2230 may be formed on the same surface as the impingement jet holes 2210. That is, the cooling channel outlet 2230 may be formed on the inner surfaces of the suction surface 2001 and the pressure surface 2002.

[0075] Referring to FIG. 6, the cooling fluid that flows into the impingement jet hole 2210 impinges (impinges jet) on one side of the internal cooling channel 2220 to perform primary cooling on the suction surface 2001 or the pressure surface 2002, and then flows along the internal cooling channel 2220 to perform secondary cooling, and is then discharged through the cooling channel outlet 2230.

[0076] In this case, it is preferable that the impingement jet holes 2210 are formed on the trailing edge TE side and the cooling channel outlets 2230 are formed on the leading edge LE side so that the flow direction of the cooling fluid flowing through the internal cooling channel 2220 is opposite to the flow direction of the high-temperature combustion gas HG. In this configuration, the cooling fluid is discharged to the outside of the airfoil 2000 through the cooling holes 2003 closer to the leading edge LE side, thereby providing a greater air curtain effect.

[0077] According to the airfoil according to the first embodiment of the present invention, the cooling efficiency can be improved by i) cooling due to collision of the cooling fluid with the internal cooling passage 2220, ii) increasing the cooling time due to the internal flow of the internal cooling passage 2220, and iii) increasing the air curtain effect due to exhaust through the cooling hole 2003 near the leading edge LE.

[0078] FIG. 7 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a second embodiment of the present invention.

[0079] 7, the cooling channel part 2200 of this embodiment includes an impingement jet hole 2210, an internal cooling channel 2220, a cooling channel outlet 2230, and a first cooling fin 2240. The impingement jet hole 2210, the internal cooling channel 2220, and the cooling channel outlet 2230 are substantially the same as those in the first embodiment, and therefore, a repeated description will be omitted.

[0080] The first cooling fins 2240 are formed to protrude by a predetermined size from one surface of the interior of the wall forming the internal cooling channel 2220 toward the other surface, and are spaced apart from the other surface. Figure 7 shows an example in which the first cooling fins 2240 are formed to protrude from the upper surface (inner surface) of the interior of the wall forming the internal cooling channel 2220 to the lower surface (outer surface). In addition, the first cooling fins 2240 are formed between any one of the impingement jet holes 2210 and the adjacent impingement jet hole 2210.

[0081] A portion of the cooling fluid flowing through the internal cooling channel 2220 collides with the first cooling fins 2240 to form a vortex, while the remainder of the cooling fluid flows straight toward the cooling channel outlet 2230. That is, the first cooling fins 2240 impart vortices and straightness to the cooling fluid, thereby improving the cooling efficiency.

[0082] According to the airfoil according to the second embodiment of the present invention, the cooling efficiency can be improved by i) cooling due to collision of the cooling fluid with the internal cooling passage 2220, ii) improving the cooling efficiency by the first cooling fin 2240, and iii) increasing the air curtain effect by discharging through the cooling hole 2003 near the leading edge LE.

[0083] FIG. 8 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a third embodiment of the present invention.

[0084] 8, the cooling channel part 2200 of this embodiment includes impingement jet holes 2210, internal cooling channels 2220, cooling channel outlets 2230, and cooling protrusions 2250. The impingement jet holes 2210, internal cooling channels 2220, and cooling channel outlets 2230 are substantially the same as those in the first embodiment, and therefore, repeated description will be omitted.

[0085] The cooling protrusion 2250 is formed below the impingement jet hole 2210. The cooling protrusion 2250 is formed to protrude from the other surface (lower surface) of the wall forming the internal cooling channel 2220 toward one surface (upper surface) by a predetermined size.

[0086] The cooling fluid that flows in vertically through the collision jet holes 2210 collides with the cooling protrusions 2250 (collision jet) and spreads radially, performing primary cooling on the suction surface 2001 or the pressure surface 2002, and then flows along the internal cooling flow path 2220 to perform secondary cooling, before being discharged through the cooling flow path outlet 2230.

[0087] At this time, the distance between the impingement jet hole 2210 and the internal cooling channel 2220 is shortened by the size of the cooling protrusion 2250, thereby shortening the collision distance and improving the collision cooling efficiency. Also, the contact area of ​​the cooling fluid is increased by the size of the cooling protrusion 2250, thereby improving the collision cooling efficiency.

[0088] According to the airfoil according to the third embodiment of the present invention, the cooling efficiency can be improved by: i) shortening the collision distance by the cooling protrusions 2250, thereby improving the cooling efficiency; ii) expanding the collision area by the cooling protrusions 2250, thereby improving the cooling efficiency; and iii) increasing the air curtain effect by discharging through the cooling holes 2003 near the leading edge LE.

[0089] FIG. 9 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a fourth embodiment of the present invention.

[0090] 9, the cooling channel part 2200 of this embodiment includes an impingement jet hole 2210, an internal cooling channel 2220, a cooling channel outlet 2230, and a second cooling fin 2260. The impingement jet hole 2210, the internal cooling channel 2220, and the cooling channel outlet 2230 are substantially the same as those in the first embodiment, and therefore, a repeated description will be omitted.

[0091] The second cooling fins 2260 can generate turbulence in the flow of cooling fluid flowing inside the internal cooling channel 2220. The second cooling fins 2260 can improve the structural rigidity of the airfoil wall surfaces (pressure surface, suction surface) whose rigidity has been weakened by the internal cooling channel 2220. The second cooling fins 2260 can increase the heat transfer area and improve the cooling efficiency of the cooling fluid.

[0092] Specifically, the second cooling fin 2260 may be a fin structure formed across one surface of the internal cooling channel 2220 and the opposing surface thereof. The fin structure may be formed in various shapes such as a polygonal shape, a circular shape, an X shape, etc. The second cooling fin 2260 may be formed between any one of the impingement jet holes 2210 and the adjacent impingement jet hole 2210.

[0093] According to the airfoil according to the fourth embodiment of the present invention, the cooling efficiency can be improved by i) cooling due to collision of the cooling fluid with the internal cooling passage 2220, ii) improving the cooling efficiency due to turbulence caused by the second cooling fin 2260, and iii) increasing the air curtain effect due to exhaust through the cooling hole 2003 near the leading edge LE.

[0094] FIG. 10 is an enlarged plan view of part B of FIG. 4, showing a part of an airfoil according to a fifth embodiment of the present invention.

[0095] 10, a cooling channel portion 2200 of the fifth embodiment includes an impingement jet hole 2210, an internal cooling channel 2220, a cooling channel outlet 2230, and an impingement cavity 2270. The impingement jet hole 2210, the internal cooling channel 2220, and the cooling channel outlet 2230 are substantially the same as those of the first embodiment, and therefore a repeated description will be omitted. The fifth embodiment can be applied in combination with at least one of the first to fourth embodiments.

[0096] The collision cavity 2270 includes a collision induction member 2271 of a predetermined shape formed in the internal space of the sub-cavity 2300, and a collision induction inlet 2272 formed by opening a part of the collision induction member 2271.

[0097] Collision induction member 2271 may be a polyhedral or hemispherical plate member formed to surround cooling hole entrance 2003a. Collision induction inlet 2272 may be formed by opening the upper portion of cooling hole entrance 2003a in the plate member.

[0098] The impingement cavity 2270 allows the cooling fluid discharged through the cooling channel outlet 2230 to collide again with the wall surface forming the suction surface 2001 or the pressure surface 2002 before being discharged to the outside through the cooling hole 2003, thereby performing secondary impingement cooling.

[0099] At this time, in order to facilitate the discharge of the cooling fluid that has been secondarily impinged, an inclined surface 2273 having a wide upper portion and a narrow lower portion may be formed around the inlet 2003a of the cooling hole.

[0100] According to the airfoil according to the fifth embodiment of the present invention, in addition to the effects of the first to fourth embodiments described above, the cooling fluid undergoes secondary collision cooling with the collision cavity 2270 before being discharged through the cooling hole 2003, thereby further improving the cooling efficiency.

[0101] The internal structure of the airfoil described above (first to fifth embodiments) was previously impossible to manufacture due to its complex shape, but by applying recently developed metal 3D printing technology, it is now possible to manufacture an airfoil as a single, continuous structure, rather than by manufacturing and assembling separate parts.

[0102] Although the embodiments of the present invention have been described above, a person having ordinary skill in the art may modify and change the present invention in various ways by adding, changing, deleting or adding components within the scope of the concept of the present invention as set forth in the claims, and this also falls within the scope of the present invention. [Explanation of symbols]

[0103] 1000: Gas turbine 1100: Compressor 1200: Combustor 1300: Turbine 2000: Airfoil 2001: Intake surface, 2002: Pressure surface 2003: Cooling Hall 2100: Main cavity 2200: Cooling channel section 2210: Impingement jet hole, 2220: Internal cooling channel 2230: cooling channel outlet, 2240: first cooling fin 2250: Cooling protrusion, 2260: Second cooling fin 2270: Collision Cavity

Claims

1. a suction surface and a pressure surface having cooling holes formed therein; at least one main cavity formed in an internal space defined by the suction surface and the pressure surface, into which a cooling fluid flows; an internal cooling passage formed within the wall forming the suction surface and the pressure surface; a plurality of impingement jet holes formed on inner surfaces of the suction surface and the pressure surface, which allow the cooling fluid to flow into the internal cooling passage and perform impingement cooling; a sub-cavity formed to surround an internal cooling passage outlet formed at an end of the internal cooling passage, the sub-cavity connecting the internal cooling passage outlet and the cooling hole inlet; Including, airfoil.

2. The impingement jet holes are formed in a plurality along the span direction on the inner surfaces of the suction surface and the pressure surface, the impingement jet hole is formed on a trailing edge side, and the internal cooling flow passage outlet is formed on a leading edge side, The airfoil according to claim 1 , wherein a plurality of the sub-cavities are formed along a span direction on the inner surfaces of the suction surface and the pressure surface.

3. 2. The airfoil according to claim 1, further comprising a first cooling fin formed inside the wall defining the internal cooling passage and projecting from one surface toward the other surface and spaced apart from the other surface.

4. 4. The airfoil of claim 3, wherein the first cooling fin is formed between any one impingement jet hole and an adjacent impingement jet hole.

5. 2. The airfoil according to claim 1, further comprising a cooling protrusion formed below the impingement jet hole, the cooling protrusion protruding in one direction from the other surface of the interior of the wall forming the internal cooling passage.

6. 2. The airfoil according to claim 1, further comprising a second cooling fin formed by connecting one surface and the other surface of the interior of the wall forming the internal cooling passage.

7. 7. The airfoil of claim 6, wherein the second cooling fin is formed between any one impingement jet hole and an adjacent impingement jet hole.

8. 2. The airfoil according to claim 1, further comprising an impingement cavity that allows the cooling fluid discharged through the internal cooling channel outlet to impinge on a wall surface that forms the suction surface or the pressure surface before being discharged to the outside through the cooling hole.

9. a compressor for compressing incoming air; a combustor that mixes the compressed air from the compressor with fuel and burns the mixture; a turbine that generates power using combustion gas from the combustor, the turbine including turbine vanes that guide the combustion gas on a combustion gas path through which the combustion gas passes, and turbine blades that rotate by the combustion gas on the combustion gas path, At least one of the turbine vane or the turbine blade includes an airfoil, the airfoil comprising: a suction surface and a pressure surface having cooling holes formed therein; at least one main cavity formed in an internal space defined by the suction surface and the pressure surface, into which a cooling fluid flows; an internal cooling passage formed within the wall forming the suction surface and the pressure surface; a plurality of impingement jet holes formed on inner surfaces of the suction surface and the pressure surface, which allow the cooling fluid to flow into the internal cooling passage and perform impingement cooling; a sub-cavity formed to surround an internal cooling passage outlet formed at an end of the internal cooling passage, the sub-cavity connecting the internal cooling passage outlet and the cooling hole inlet; , including a gas turbine.

10. The impingement jet holes are formed in a plurality along the span direction on the inner surfaces of the suction surface and the pressure surface, the impingement jet hole is formed on a trailing edge side, and the internal cooling flow passage outlet is formed on a leading edge side, The gas turbine according to claim 9 , wherein a plurality of the sub-cavities are formed along the span direction on the inner surfaces of the suction surface and the pressure surface.

11. a first cooling fin formed to protrude from one surface of the wall forming the internal cooling passage toward the other surface and spaced apart from the other surface; The gas turbine according to claim 9 , wherein the first cooling fin is formed between any one impingement jet hole and an adjacent impingement jet hole.

12. The gas turbine according to claim 9 , further comprising a cooling protrusion formed below the impingement jet hole, the cooling protrusion protruding in one direction from the other surface inside the wall forming the internal cooling flow path.

13. The gas turbine according to claim 9 , further comprising a second cooling fin formed by connecting one surface and the other surface inside the wall forming the internal cooling flow passage.

14. The gas turbine according to claim 13 , wherein the second cooling fin is formed between any one impingement jet hole and an adjacent impingement jet hole.

15. 10. The gas turbine according to claim 9, further comprising an impingement cavity that allows the cooling fluid discharged through the internal cooling channel outlet to impinge on a wall surface that forms the suction surface or the pressure surface before being discharged to the outside through the cooling hole.

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