Pin-impact jet cooling structures for turbine components and gas turbines including them

Abstract

To provide a pin-collision jet cooling structure for a turbine component and a gas turbine including the same.SOLUTION: An air foil for a turbine blade or a turbine vane includes: a cooling flow passage cavity formed inside the air foil; an insert inserted into the cooling flow passage cavity and having a plurality of cooling holes; and a cooling structure formed between an outer surface of the insert and an inner surface of the cooling flow passage cavity. The cooling structure includes: support parts each coming into close contact with the outer surface of the insert and having a plurality of collision jet holes communicating with the plurality of cooling holes; and cooling pins each coupled to a space between a lower surface of the support part and the inner surface of the cooling flow passage cavity.SELECTED DRAWING: Figure 7

Description

Technical Field

[0001] The present invention relates to a pin-impingement jet cooling structure for turbine components and a gas turbine including the same.

Background Art

[0002] A turbine is a mechanical device that obtains rotational force by impulse or reaction force using the flow of a compressible fluid such as steam or gas, and includes a steam turbine using steam and a gas turbine using high-temperature combustion gas.

[0003] Among these, a gas turbine is mainly 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 a compressor housing.

[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 alternately arranged in a turbine housing. Also, a rotor is arranged so as to penetrate the centers of the compressor, the combustor, the turbine, and the exhaust chamber. <x

[0006] Both ends of the rotor are rotatably supported by bearings. A plurality of disks are fixed to the rotor, and each blade is connected thereto. At the same time, a drive shaft such as a generator is connected to the end on the exhaust chamber side.

[0007] Such a gas turbine does not have a reciprocating mechanism like a piston of a four-stroke engine, so there is no mutual friction part like a piston-cylinder, the consumption of lubricating oil is extremely small, the amplitude, which is a characteristic of a reciprocating machine, is significantly reduced, and it has the advantage of being able to perform high-speed movement.

[0008] To briefly explain the operation of a gas turbine, compressed air from a compressor is mixed with fuel and burned to produce high-temperature combustion gases, which are then injected into the turbine. As the injected combustion gases pass through the turbine vanes and blades, they generate rotational force, which causes the rotor to rotate. [Overview of the project] [Problems that the invention aims to solve]

[0009] The present invention aims to provide a pin-impact jet cooling structure for a turbine component and a gas turbine including the same, which can reduce crossflow of impact cooling air and improve cooling efficiency by forming a plurality of support parts and cooling pins in the impact cooling passage of an airfoil. [Means for solving the problem]

[0010] To achieve the above objective, the turbine blade or turbine vane airfoil of the present invention includes a cooling channel cavity formed inside the airfoil, an insert inserted into the cooling channel cavity and including a plurality of cooling holes, and a cooling structure formed between the outer surface of the insert and the inner surface of the cooling channel cavity, wherein the cooling structure includes a support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes communicating with the plurality of cooling holes, and a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling channel cavity.

[0011] The support portion may be formed in the shape of a circular disk of a predetermined thickness.

[0012] The plurality of impact jet holes may include a first jet hole located upstream of the cooling pin and a pair of second jet holes located downstream of the cooling pin.

[0013] When the diameter of the plurality of impact jet holes is d, the height (z) of the cooling pin in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion may be formed to be 2 to 4 times d.

[0014] The support portion may be formed in the shape of a triangular disc of a predetermined thickness with a rounded apex.

[0015] The plurality of impact jet holes may include a first jet hole located upstream of the cooling pin and a pair of second jet holes located downstream of the cooling pin.

[0016] When the diameter of the plurality of impact jet holes is d, the height (z) of the cooling pin in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion may be formed to be 2 to 4 times d.

[0017] The support portion may be formed with an arc-shaped curved surface on its side that surrounds the plurality of impact jet holes and the cooling pin.

[0018] The plurality of impact jet holes may include a first jet hole located upstream of the cooling pin and a pair of second jet holes located downstream of the cooling pin.

[0019] When the diameter of the plurality of impact jet holes is d, the height (z) of the cooling pin in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion may be formed to be 2 to 4 times d.

[0020] The gas turbine of the present invention includes a compressor that inhales and compresses external air, a combustor that mixes fuel with the air compressed by the compressor and burns it, and a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing and the turbine blades are rotated by combustion gases discharged from the combustor, wherein the airfoils of the turbine blades and turbine vanes include a cooling passage cavity formed inside the airfoil, an insert inserted into the cooling passage cavity and including a plurality of cooling holes, and a cooling structure formed between the outer surface of the insert and the inner surface of the cooling passage cavity, wherein the cooling structure includes a support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes communicating with the plurality of cooling holes, and a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling passage cavity.

[0021] The support portion may be formed in the shape of a circular disk of a predetermined thickness.

[0022] The plurality of impact jet holes may include a first jet hole located upstream of the cooling pin and a pair of second jet holes located downstream of the cooling pin.

[0023] When the diameter of the plurality of impact jet holes is d, the height (z) of the cooling pin in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion may be formed to be 2 to 4 times d.

[0024] The support portion may be formed in the shape of a triangular disc of a predetermined thickness with a rounded apex.

[0025] The plurality of impact jet holes may include a first jet hole located upstream of the cooling pin and a pair of second jet holes located downstream of the cooling pin.

[0026] When the diameter of the plurality of impinging jet holes is d, the height (z) of the cooling pins in the jet direction may be formed to be 2 to 4 times that of d, and the thickness (t) of the support portion may be formed to be 2 to 4 times that of d.

[0027] The support portion may be formed in an arc-shaped curved surface shape whose side surrounds the plurality of impinging jet holes and the cooling pins.

[0028] The plurality of impinging jet holes may include one first jet hole disposed upstream of the cooling pins and a pair of second jet holes disposed downstream of the cooling pins.

[0029] [[ID=!1]] When the diameter of the plurality of impinging jet holes is d, the height (z) of the cooling pins in the jet direction may be formed to be 2 to 4 times that of d, and the thickness (t) of the support portion may be formed to be 2 to 4 times that of d.

Advantages of the Invention

[0030] According to the pin-impinging jet cooling structure of the turbine component of the present invention and the gas turbine including the same, by forming a plurality of support portions and cooling pins in the impinging cooling flow path of the airfoil, the cross flow of the impinging cooling air can be reduced and the cooling efficiency can be improved.

Brief Description of the Drawings

[0031] [[ID=2!5]] [Figure 1] It is a partially cut-away perspective view of a gas turbine according to an embodiment of the present invention. [Figure 2] It is a cross-sectional view showing a schematic structure of a gas turbine according to an embodiment of the present invention. [Figure 3] It is a perspective view showing a turbine blade according to an embodiment of the present invention. [Figure 4] It is a perspective view showing a turbine vane according to an embodiment of the present invention. [Figure 5] It is a cross-sectional view showing an airfoil according to an embodiment of the present invention. [Figure 6]This is a cross-sectional view showing the cooling structure arranged between the insert and the outer wall of the cooling channel cavity. [Figure 7] Figure 7(a) is a perspective view showing a cooling structure according to the first embodiment of the present invention, and Figure 7(b) is a top view showing a cooling structure according to the first embodiment. [Figure 8] This is a perspective view showing a cooling structure according to a second embodiment of the present invention. [Figure 9] This is a perspective view showing a cooling structure according to a third embodiment of the present invention. [Figure 10] This photograph shows a comparison of the cooling effect of the conventional technology and the cooling structure of the present invention. [Modes for carrying out the invention]

[0032] While the present invention can have various embodiments through diverse transformations, specific embodiments will be illustrated and described in detail in the detailed description. However, it should be understood that this does not intend to limit the present invention to specific embodiments, but rather includes all transformations, equivalents, or substitutions that fall within the spirit and technical scope of the present invention.

[0033] The terms used in this invention are used solely to describe specific embodiments and are not intended to limit the invention. Singular expressions include plural expressions unless the context clearly indicates otherwise. In this invention, terms such as “includes” or “having” are intended to specify the existence of features, figures, stages, operations, components, parts, or combinations thereof described in the specification, and should be understood not to preemptively exclude the possibility of the existence or addition of one or more other features, figures, stages, operations, components, parts, or combinations thereof.

[0034] Preferred embodiments of the present invention will be described in detail below with reference to the attached drawings. Note that, in the attached drawings, identical components are represented by the same reference numerals whenever possible. Furthermore, detailed descriptions of known functions and configurations that may obscure the gist of the present invention will be omitted. For similar reasons, some components in the attached drawings are exaggerated, omitted, or shown schematically.

[0035] Figure 1 is a partially cut-out perspective view of a gas turbine according to one embodiment of the present invention, and Figure 2 is a cross-sectional view showing the schematic structure of a gas turbine according to one embodiment of the present invention.

[0036] As shown in Figure 1, a gas turbine 1000 according to one embodiment of the present invention includes a compressor 1100, a combustor 1200, and a turbine 1300. The compressor 1100 comprises a plurality of radially arranged blades 1110. The compressor 1100 rotates the blades 1110, and air is compressed and moved by the rotation of the blades 1110. The size and installation angle of the blades 1110 can be changed depending on the installation position. In one embodiment, the compressor 1100 can be 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 blades 1110.

[0037] The air compressed in the compressor 1100 moves to the combustor 1200. The combustor 1200 includes a plurality of combustion chambers 1210 arranged in an annular manner and a fuel nozzle module 1220.

[0038] As shown in Figure 2, a gas turbine 1000 according to one embodiment of the present invention is equipped with a housing 1010, and a diffuser 1400 is provided on the rear side of the housing 1010 to discharge the combustion gas that has passed through the turbine. A combustor 1200 is positioned in front of the diffuser 1400 to receive compressed air and burn it.

[0039] To explain using the direction of airflow as a reference, the compressor section 1100 is located upstream of the housing 1010, and the turbine section 1300 is positioned downstream. Between the compressor section 1100 and the turbine section 1300, a torque tube unit 1500 is positioned as a torque transmission member that transmits the rotational torque generated in the turbine section 1300 to the compressor section 1100.

[0040] The compressor section 1100 is equipped with a plurality (for example, 14) of compressor rotor discs 1120, and each of the compressor rotor discs 1120 is fastened by tie rods 1600 so as not to separate in the axial direction.

[0041] Specifically, each of the compressor rotor discs 1120 is aligned axially with respect to each other, with a tie rod 1600, which constitutes the rotation axis, passing through approximately the center of each disc. Here, adjacent compressor rotor discs 1120 are positioned so that their opposing surfaces are pressed together by the tie rod 1600, making relative rotation impossible.

[0042] Multiple blades 1110 are radially connected to the outer circumferential surface of the compressor rotor disk 1120. Each blade 1110 is fastened to the compressor rotor disk 1120 and is equipped with a dovetail portion 1112.

[0043] Between each of the rotor discs 1120, there are vanes (not shown) fixed to the housing. Unlike the rotor discs, the vanes are fixed so as not to rotate and serve to align the flow of compressed air that has passed through the blades of the compressor rotor discs, guiding the air to the blades of the rotor disc located downstream.

[0044] The fastening method for the dovetail portion 1112 can be tangential or axial. This can be selected according to the required structure of the commercial gas turbine and can have the commonly known dovetail or fir-tree shape. In some cases, the blades can be fastened to the rotor disk using other fastening devices other than those described above, such as fasteners such as keys or bolts.

[0045] The tie rod 1600 is positioned to penetrate the centers of the plurality of compressor rotor discs 1120 and turbine rotor discs 1320, and the tie rod 1600 may consist of one or more tie rods. One end of the tie rod 1600 is fastened into the compressor rotor disc located on the upstream side, and the other end of the tie rod 1600 is fastened by a fixing nut 1450.

[0046] The form of the tie rod 1600 can consist of various structures depending on the gas turbine, and is not necessarily limited to the form shown in Figure 2. That is, as shown in the figure, it may have a form in which one tie rod penetrates the center of the rotor disc, or it may have a form in which multiple tie rods are arranged around the circumference, or a combination of these is also possible.

[0047] Although not shown in the diagram, a gas turbine compressor may have vanes that act as guide vanes located after the diffuser to adjust the fluid flow angle of the fluid entering the combustor inlet to the design flow angle after increasing the fluid pressure; these are called deswirlers.

[0048] In the combustor 1200, the incoming compressed air is mixed with fuel and burned to produce high-energy, high-temperature, high-pressure combustion gas, and the temperature of the combustion gas is raised to the heat resistance limit that the combustor and turbine components can withstand during the isobaric combustion process.

[0049] The combustors constituting the combustion system of a gas turbine may be arranged in multiples within a housing formed in a cell shape, and may consist of a burner including fuel injection nozzles, a combustor liner that forms a combustion chamber, and a transition piece that connects the combustor to the turbine.

[0050] Specifically, the liner provides a combustion space in which fuel injected by a fuel nozzle is mixed with compressed air from a compressor and burned. Such a liner may include a flame tube that provides a combustion space in which the fuel mixed with air is burned, and a flow sleeve that surrounds the flame tube and forms an annular space. A fuel nozzle is coupled to the front end of the liner, and a spark plug is coupled to the side wall.

[0051] Meanwhile, a transition piece is connected to the rear end of the liner to allow combustion gases, which are burned by the spark plug, to be sent to the turbine. The outer wall of such a transition piece is cooled by compressed air supplied from the compressor to prevent damage from the high temperature of the combustion gases.

[0052] For this purpose, the transition piece is provided with cooling holes that allow air to be injected into it. The compressed air cools the main body inside through the holes before flowing towards the liner.

[0053] Cooling air, which has cooled the transition piece described above, flows through the annular space of the liner, and compressed air provided outside the flow sleeve as cooling air can collide with the outer wall of the liner through cooling holes provided in the flow sleeve.

[0054] Meanwhile, the high-temperature, high-pressure combustion gases discharged from the combustor are supplied to the turbine 1300 described above. The supplied high-temperature, high-pressure combustion gases expand and collide with the turbine blades, generating a reaction force and producing rotational torque. This rotational torque is transmitted to the compressor via the torque tube described above, and any power exceeding the power required to drive the compressor is used to drive a generator or the like.

[0055] The turbine 1300 is basically similar in structure to that of a compressor. That is, the turbine 1300 is also equipped with a plurality of turbine rotor discs 1320 similar to the compressor rotor disc of a compressor. Therefore, the turbine rotor discs 1320 also include a plurality of turbine blades 1340 arranged radially. The turbine blades 1340 can also be coupled to the turbine rotor discs 1320 by a method such as a dovetail. In addition, turbine vanes 1330 fixed to the housing are provided between the blades 1340 of the turbine rotor discs 1320 to guide the flow direction of the combustion gases that have passed through the blades.

[0056] Figure 3 is a perspective view showing a turbine blade according to one embodiment of the present invention. Unlike Figure 2, Figure 3 shows the drawing number "100" on the turbine blade.

[0057] The turbine blade 100 includes an airfoil 110 at its upper end that rotates due to the pressure of the combustion gases, a platform portion 120 integrally formed at the lower part of the airfoil, and a root portion 130 integrally formed at the lower part of the platform portion and coupled to the turbine rotor disk 1320. An inlet may be formed inside the platform portion 120 for supplying cooling fluid to an internal flow path formed inside the airfoil 110.

[0058] The airfoil 110 includes an intake surface 112 that bulges outward on one side into which combustion gases flow, and a pressure surface 111 that is recessed on the opposite side of the intake surface. The front corner where the pressure surface 111 and the intake surface 112 meet constitutes a leading edge 113, and the rear corner forms a trailing edge 114. An internal flow path (not shown) may be formed inside the airfoil 110 through which cooling air flowing in through the inlet flows.

[0059] The platform portion 120 plays a role in maintaining the spacing between adjacent turbine blades by having the platform portion 120 and its side surface come into contact with each other.

[0060] The root portion 130 may have an axial-type configuration, being inserted along the axial direction of the turbine rotor disc into a coupling slot formed on the outer circumferential surface of the turbine rotor disc 1320. The root portion 130 may have a substantially dovetail or fir tree-shaped bend, which may be formed to correspond to the shape of the bend formed in the coupling slot.

[0061] Figure 4 is a perspective view showing a turbine vane according to one embodiment of the present invention. Unlike Figure 2, Figure 4 shows the drawing number "200" on the turbine vane.

[0062] The turbine vane 200 may include an airfoil 210 fixed between the turbine blades 100 to guide the flow direction of combustion gases passing through the turbine blades, an inner end wall 220 formed radially inward of the airfoil, and an outer end wall 230 formed radially outward of the airfoil.

[0063] The airfoil 210 of the turbine vane 200, like the airfoil 110 of the turbine blade 100, includes a recessed pressure surface 211, an outwardly bulging intake surface 212, a leading edge 213, and a trailing edge 214.

[0064] Inside the airfoil 210, a cooling channel cavity 240 may be formed, separated by one or more partition walls. The cooling channel cavity 240 may be formed radially elongated inside the airfoil 210.

[0065] Figure 5 is a cross-sectional view showing an airfoil according to one embodiment of the present invention, and Figure 6 is a cross-sectional view showing a cooling structure arranged between the insert and the outer wall of the cooling channel cavity.

[0066] In the present invention, the airfoil 110 or 210 of the turbine blade 100 or turbine vane 200 includes a cooling channel cavity 140 or 240 formed inside the airfoil, inserts 150, 250 inserted inside the cooling channel cavity and containing a plurality of cooling holes 155, and a plurality of cooling structures 300 formed between the outer surface of the insert and the inner surface of the cooling channel cavity.

[0067] The cooling channel cavities 140 or 240 may be formed in two or more parts within the airfoil 110 or 210, partitioned by partition walls.

[0068] The inserts 150 and 250 may be formed in a manner corresponding to the inner circumferential surface of the cooling channel cavity 140 or 240 and inserted and fitted into the cooling channel cavity 140 or 240. Multiple cooling holes 155 are formed through the side walls of the inserts 150 and 250 so that cooling air inside the insert can flow to the outside of the insert through the multiple cooling holes 155. Alternatively, the airfoil 110 or 210 may be manufactured such that a structure corresponding to the inserts 150 and 250 is integrally formed inside the cooling channel cavity 140 or 240.

[0069] The cooling structures 300 can be integrally formed so as to be arranged in multiples between the outer surfaces of the inserts 150, 250 and the inner surfaces of the cooling channel cavities 140 or 240. Multiple cooling structures 300 may be formed to be integrally connected between the outer surfaces of the inserts 150, 250 and the inner surfaces of the cooling channel cavities 140 or 240, and may be arranged to guide the cooling air flowing through the multiple cooling holes 155 and enhance the impact cooling effect.

[0070] As shown in Figure 6, the cooling structure 300 may include a support portion 320 that is in close contact with the outer surface of the insert 150 and includes a plurality of impact jet holes 330, 340 that communicate with a plurality of cooling holes 155, and a cooling pin 310 that is connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity 140.

[0071] The support portion 320 may be formed to be in close contact with the outer surface of the insert 150 and to surround the multiple impact jet holes 330, 340 formed at positions corresponding to the multiple cooling holes 155. The support portion 320 may be formed to have a thickness (t) of approximately half the height (h) of the space between the outer surface of the insert 150 and the inner surface of the cooling flow channel cavity 140.

[0072] Multiple impact jet holes 330, 340 can be formed by penetrating the support portion 320 in the thickness direction, both in front of and behind the cooling pin 310, with respect to the direction of cooling air flow.

[0073] The cooling pin 310 can be connected to the lower surface of the support portion 320 approximately in the center. Furthermore, the lower end of the cooling pin 310 can be integrally connected to the inner circumferential surface of the outer wall 145 of the cooling flow path cavity 140.

[0074] The sidewall of the insert 150, which has multiple cooling holes 155 formed therein, can form a jet plate in the impact cooling channel, and the inner circumferential surface of the outer wall 145 of the cooling channel cavity 140 can form a target surface into which the cooling air impacts.

[0075] Figure 7(a) is a perspective view showing a cooling structure according to the first embodiment of the present invention, and Figure 7(b) is a top view showing a cooling structure according to the first embodiment.

[0076] In the cooling structure 300 according to the first embodiment of the present invention, the support portion 320 may be formed in the shape of a circular disk of a predetermined thickness. Multiple impact jet holes 330, 340 can be formed through the support portion 320 in the longitudinal direction of the cooling pin 310, both in front of and behind the cooling pin 310.

[0077] The multiple impact jet holes 330, 340 may include one first jet hole 330 located upstream of the cooling pin 310 and a pair of second jet holes 340 located downstream of the cooling pin.

[0078] The first jet hole 330 and the second jet hole 340 may be formed in a circular hole shape having an inner diameter smaller than the diameter of the cooling pin 310. The first jet hole 330 may be positioned at a predetermined distance upstream from the center of the cooling pin 310 with respect to the direction of cooling air flow. A pair of second jet holes 340 may be positioned at a predetermined distance downstream from the center of the cooling pin 310. The pair of second jet holes 340 may be formed at predetermined distances apart on either side of a straight line passing through the center of the first jet hole 330 and the center of the cooling pin 310. In other words, the center of the first jet hole 330 and the center of the pair of second jet holes 340 may be positioned at the vertices of an isosceles triangle.

[0079] As shown in Figure 7(b), the angle (θ) between the center of the first jet hole 330 and the center of the second jet hole 340 with respect to the center of the cooling pin 310 may be 140 to 160 degrees, particularly about 150 degrees. In this case, the pair of second jet holes 340 may be positioned far apart such that the distance between their centers is greater than the outer diameter of the cooling pin 310.

[0080] As shown in Figure 6, when the diameter of the multiple impact jet holes 330, 340 is d, the height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion 320 may be formed to be 2 to 4 times d.

[0081] The inner diameter (d) of the first jet hole 330 and the second jet hole 340 can be formed to be the same as that of the cooling hole 155 of the insert 150 that constitutes the jet plate. The inner diameter (d) of the first jet hole 330 and the second jet hole 340 may be formed to be 0.8 to 1.2 mm.

[0082] The height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340. The thickness (t) of the support portion 320 may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340.

[0083] As a result, the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140 that constitutes the target surface may be formed to be 4 to 8 times the diameter (d) of the impact jet holes 330 and 340. In other words, the height (z) of the cooling pin 310 and the thickness (t) of the support portion 320 may be formed to be approximately half the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140.

[0084] Figure 8 is a perspective view showing a cooling structure according to a second embodiment of the present invention.

[0085] In the cooling structure 300 according to the second embodiment of the present invention, the support portion 320 may be formed in the shape of a triangular disk of a predetermined thickness with a rounded apex. Multiple impact jet holes 330, 340 can be formed through the support portion 320 in the longitudinal direction of the cooling pin 310, both in front of and behind the cooling pin 310.

[0086] The multiple impact jet holes 330, 340 may include one first jet hole 330 located upstream of the cooling pin 310 and a pair of second jet holes 340 located downstream of the cooling pin. The first jet hole 330 and the pair of second jet holes 340 may be formed so that they are located inside the three rounded vertices of the support portion 320. In other words, the support portion 320 may be formed in the shape of an isosceles triangular disk with rounded vertices that encloses the three impact jet holes 330, 340.

[0087] The first jet hole 330 and the second jet hole 340 may be formed in a circular hole shape having an inner diameter smaller than the diameter of the cooling pin 310. The first jet hole 330 may be positioned at a predetermined distance upstream from the center of the cooling pin 310 with respect to the direction of cooling air flow. A pair of second jet holes 340 may be positioned at a predetermined distance downstream from the center of the cooling pin 310. The pair of second jet holes 340 may be formed at predetermined distances apart on either side of a straight line passing through the center of the first jet hole 330 and the center of the cooling pin 310. In other words, the center of the first jet hole 330 and the center of the pair of second jet holes 340 may be positioned at the vertices of an isosceles triangle.

[0088] Referring to Figure 7(b), the angle (θ) between the center of the first jet hole 330 and the center of the second jet hole 340 with respect to the center of the cooling pin 310 may be 140 to 160 degrees, particularly about 150 degrees. In this case, the pair of second jet holes 340 may be positioned far apart such that the distance between their centers is greater than the outer diameter of the cooling pin 310.

[0089] As shown in Figure 6, when the diameter of the multiple impact jet holes 330, 340 is d, the height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion 320 may be formed to be 2 to 4 times d.

[0090] The inner diameter (d) of the first jet hole 330 and the second jet hole 340 can be formed to be the same as that of the cooling hole 155 of the insert 150 that constitutes the jet plate. The inner diameter (d) of the first jet hole 330 and the second jet hole 340 may be formed to be 0.8 to 1.2 mm.

[0091] The height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340. The thickness (t) of the support portion 320 may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340.

[0092] As a result, the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140 that constitutes the target surface may be formed to be 4 to 8 times the diameter (d) of the impact jet holes 330 and 340. In other words, the height (z) of the cooling pin 310 and the thickness (t) of the support portion 320 may be formed to be approximately half the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140.

[0093] Figure 9 is a perspective view showing a cooling structure according to a third embodiment of the present invention.

[0094] In the cooling structure 300 according to the third embodiment of the present invention, the support portion 320 may be formed in an arc-shaped curved surface on its side surface that surrounds a plurality of impact jet holes 330, 340 and a cooling pin 310. The plurality of impact jet holes 330, 340 can be formed through the support portion 320 in the longitudinal direction of the cooling pin 310, both in front of and behind the cooling pin 310.

[0095] The multiple impact jet holes 330, 340 may include one first jet hole 330 located upstream of the cooling pin 310 and a pair of second jet holes 340 located downstream of the cooling pin. The first jet hole 330 and the pair of second jet holes 340 may be formed so that they are located inside the three rounded vertices of the support portion 320. In other words, the figure formed by connecting the centers of the three impact jet holes 330, 340 can form an isosceles triangle.

[0096] The contour of the support portion 320 may be formed in a manner in which three arc-shaped curved surfaces having a constant radius of curvature surround the centers of the three impact jet holes 330, 340 and a pair of arc-shaped curved surfaces having a constant radius of curvature surround the center of the cooling pin 310 and are connected to each other.

[0097] The first jet hole 330 and the second jet hole 340 may be formed in a circular hole shape having an inner diameter smaller than the diameter of the cooling pin 310. The first jet hole 330 may be positioned at a predetermined distance upstream from the center of the cooling pin 310 with respect to the direction of cooling air flow. A pair of second jet holes 340 may be positioned at a predetermined distance downstream from the center of the cooling pin 310. The pair of second jet holes 340 may be formed at predetermined distances apart on either side of a straight line passing through the center of the first jet hole 330 and the center of the cooling pin 310. In other words, the center of the first jet hole 330 and the center of the pair of second jet holes 340 may be positioned at the vertices of an isosceles triangle.

[0098] Referring to Figure 7(b), the angle (θ) between the center of the first jet hole 330 and the center of the second jet hole 340 with respect to the center of the cooling pin 310 may be 140 to 160 degrees, particularly about 150 degrees. In this case, the pair of second jet holes 340 may be positioned far apart such that the distance between their centers is greater than the outer diameter of the cooling pin 310.

[0099] As shown in Figure 6, when the diameter of the multiple impact jet holes 330, 340 is d, the height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times d, and the thickness (t) of the support portion 320 may be formed to be 2 to 4 times d.

[0100] The inner diameter (d) of the first jet hole 330 and the second jet hole 340 can be formed to be the same as that of the cooling hole 155 of the insert 150 that constitutes the jet plate. The inner diameter (d) of the first jet hole 330 and the second jet hole 340 may be formed to be 0.8 to 1.2 mm.

[0101] The height (z) of the cooling pin 310 in the jet direction may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340. The thickness (t) of the support portion 320 may be formed to be 2 to 4 times the diameter (d) of the impact jet holes 330 and 340.

[0102] As a result, the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140 that constitutes the target surface may be formed to be 4 to 8 times the diameter (d) of the impact jet holes 330 and 340. In other words, the height (z) of the cooling pin 310 and the thickness (t) of the support portion 320 may be formed to be approximately half the height (h) between the outer surface of the insert 150 and the inner surface of the outer wall 145 of the cooling channel cavity 140.

[0103] Figure 10 is a photograph comparing the cooling effect of the conventional technology and the cooling structure of the present invention.

[0104] In conventional technology, the jet plate only has multiple cooling holes, and the impact cooling channel structure does not have cooling structures such as cooling pins.

[0105] In the first embodiment, the support portion of the cooling structure is formed in the shape of a circular disk.

[0106] In the second embodiment, the support portion of the cooling structure is formed in a rounded disk shape, with the vertices of an isosceles triangle surrounding the impact jet holes.

[0107] In the third embodiment, the support portion of the cooling structure is formed such that its contour line is a connected arc surrounding the impact jet hole and the cooling pin.

[0108] In conventional technology, a very large temperature difference is observed between the upstream and downstream sides of the impingement cooling channel. This is likely due to a large amount of cross-flow of the impingement cooling air, which reduces cooling efficiency.

[0109] In contrast, according to the cooling structure according to the embodiment of the present invention, the cross-flow of the impact cooling air is reduced by approximately 40% or more compared to the conventional technology, and the overall cooling efficiency is improved by approximately 30% or more compared to the conventional technology.

[0110] Although one embodiment of the present invention has been described above, a person with ordinary skill in the art can modify and change the present invention in various ways by adding, changing, deleting, or adding components, without departing from the spirit of the invention as described in the claims, and this is also included within the scope of the rights of the present invention. [Explanation of symbols]

[0111] 1000: Gas turbine, 1010: Housing 1100: Compressor, 1110: Blade 1112: Dovetail section, 1120: Compressor rotor disc 1200: Combustor, 1210: Combustion chamber 1220: Fuel nozzle module 1300: Turbine, 1320: Turbine rotor disc 1330: Turbine vanes, 1340: Turbine blades 1400: Diffuser, 1450: Fixing nut 1500: Torque tube unit, 1600: Tie rod 100: Turbine blades, 110: Airfoils 111: Pressure surface, 112: Intake surface 113: Leading edge, 114: Trailing edge 120: Platform section, 130: Route section 140: Cooling channel cavity, 145: Outer wall 150: Insert, 155: Cooling hole 200: Turbine vane, 210: Airfoil 211: Pressure surface, 212: Intake surface 213: Leading edge, 214: Trailing edge 220: Inner end wall, 230: Outer end wall 240: Cooling channel cavity, 245: Outer wall 250: Insert 300: Cooling structure 310: Cooling pin, 320: Support part 330: First jet hole, 340: Second jet hole

Claims

1. In the airfoil of a turbine blade or turbine vane, A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, The aforementioned multiple impact jet holes are A first jet hole positioned upstream of the cooling pin, An airfoil including a pair of second jet holes positioned downstream of the cooling pins.

2. The support portion is formed in the shape of a circular disc of a predetermined thickness, as described in claim 1.

3. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The airfoil according to claim 2, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

4. The airfoil according to claim 1, wherein the support portion is formed in the shape of a triangular disc of a predetermined thickness with a rounded apex.

5. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The airfoil according to claim 4, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

6. The airfoil according to claim 1, wherein the support portion has an arc-shaped curved surface on its side that surrounds the plurality of impact jet holes and the cooling pins.

7. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The airfoil according to claim 6, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

8. In an airfoil for a turbine blade or turbine vane, A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The airfoil is formed with a support portion thickness (t) that is 2 to 4 times d.

9. In an airfoil for a turbine blade or turbine vane, A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, The support portion is an airfoil, which has a triangular disc shape with a predetermined thickness and a rounded apex, or an arc-shaped curved surface on its side that surrounds the plurality of impact jet holes and the cooling pins.

10. A compressor that takes in outside air and compresses it, A combustor that mixes fuel with the air compressed by the aforementioned compressor and burns it, In a gas turbine, which includes a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing, and the turbine blades are rotated by combustion gases discharged from the combustor, The airfoils of the turbine blades and turbine vanes are A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, The aforementioned multiple impact jet holes are A first jet hole positioned upstream of the cooling pin, A gas turbine including a pair of second jet holes located downstream of the cooling pins.

11. The gas turbine according to claim 10, wherein the support portion is formed in the shape of a circular disk of a predetermined thickness.

12. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The gas turbine according to claim 11, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

13. The gas turbine according to claim 10, wherein the support portion is formed in the shape of a triangular disk of a predetermined thickness with a rounded apex.

14. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The gas turbine according to claim 13, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

15. The gas turbine according to claim 10, wherein the support portion has an arc-shaped curved surface on its side that surrounds the plurality of impact jet holes and the cooling pins.

16. When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The gas turbine according to claim 15, wherein the thickness (t) of the support portion is formed to be 2 to 4 times d.

17. A compressor that draws in and compresses external air, A combustor that mixes fuel with the air compressed by the aforementioned compressor and burns it, In a gas turbine, which includes a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing, and the turbine blades are rotated by combustion gases discharged from the combustor, The airfoils of the turbine blades and turbine vanes are A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, When the diameter of the plurality of impact jet holes is d, The height (z) of the cooling pin in the jet direction is formed to be 2 to 4 times d. The thickness (t) of the support portion is formed to be 2 to 4 times d in the gas turbine.

18. A compressor that draws in and compresses external air, A combustor that mixes fuel with the air compressed by the aforementioned compressor and burns it, In a gas turbine, which includes a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing, and the turbine blades are rotated by combustion gases discharged from the combustor, The airfoils of the turbine blades and turbine vanes are A cooling channel cavity formed inside the airfoil, An insert is inserted into the cooling channel cavity and includes a plurality of cooling holes, The cooling structure includes the outer surface of the insert and the inner surface of the cooling channel cavity, The aforementioned cooling structure is A support portion that is in close contact with the outer surface of the insert and includes a plurality of impact jet holes that communicate with a plurality of cooling holes, It includes a cooling pin connected between the lower surface of the support portion and the inner surface of the cooling flow path cavity, The support portion is a triangular disc shape with a predetermined thickness and a rounded apex, or a gas turbine in which the side surface is formed into an arc-shaped curved surface that surrounds the plurality of impact jet holes and the cooling pins.

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