Combustion gas leakage preventing device for turbine blade and gas turbine comprising the same

Abstract

The present invention relates to a combustion gas leakage preventing device for a turbine blade and a gas turbine including the same, the device including: a first damping pin arranged to extend in an axial direction between a platform and a blade root of the turbine blade and inserted into a pin groove; second and third damping pins arranged at lower portions on both sides of the first damping pin, respectively, and inserted into pin grooves; and a retaining plate arranged at lower portions of the second and third damping pins; the second and third damping pins including: a main body portion having a hollow portion; curved extension portions extending outward from peripheral surfaces of both ends of the main body portion; and the curved extension portions including cutout groove portions for inserting contact end portions of the first damping pin on one side and contact end portions of the retaining plate on the other side.

Description

Technical Field

[0001] The present invention relates to a combustion gas leakage prevention device for turbine blades and a gas turbine including the device, and more specifically, to a combustion gas leakage prevention device for turbine blades that provides a damping pin along the periphery of the shank portion of the turbine blade to minimize combustion gas leakage and a gas turbine including the device. Background Technology

[0002] Generally, a turbine is a mechanical device that uses the flow of compressible fluids such as steam and gas to obtain rotational power through impact or reaction forces. It includes steam turbines that use steam and gas turbines that use high-temperature combustion gases.

[0003] The gas turbine mechanism works by mixing and burning air that has been compressed to high pressure in the compressor with fuel, and then injecting the resulting high-temperature, high-pressure combustion gases into the turbine to make it rotate, thereby converting thermal energy into mechanical energy.

[0004] More specifically, a gas turbine includes: a compressor for compressing air; a combustion chamber for burning compressed air supplied from the compressor with fuel to produce combustion gases; and a turbine that rotates the rotor via turbine blades rotating from the high-temperature, high-pressure combustion gases discharged from the combustion chamber.

[0005] In a gas turbine, multiple turbine stator vanes and turbine blades are alternately arranged within the turbine casing. Furthermore, the rotor is positioned centrally, penetrating the compressor, combustion chamber, turbine, and exhaust chamber.

[0006] The rotor is supported by bearings at both ends to enable it to rotate. In addition, multiple discs are fixed to the rotor, each blade is connected to the rotor, and a drive shaft for a generator or the like is connected to the end on the exhaust chamber side.

[0007] Because this type of gas turbine does not have a reciprocating motion mechanism like the piston in a four-stroke engine, there are no frictional parts such as piston-cylinder, resulting in very low lubricant consumption. It also significantly reduces the amplitude of vibration, a characteristic of reciprocating motion, and has the advantage of being able to move at high speeds.

[0008] During engine operation, turbine blades may vibrate. For example, when the flow of high-temperature combustion gases or steam fluctuates, this change in flow can cause the turbine blades to vibrate.

[0009] Therefore, in the design of gas turbines, especially turbines, it is essential to avoid or minimize the dynamic stress caused by resonance, forced response, or aero-elastic instabilities at the natural frequency of turbine blades, in order to control the high cyclic fatigue of turbine blades.

[0010] Therefore, in order to improve the high cyclic fatigue life of turbine blades, multiple damping pins spaced apart from each other are usually installed around the blade root periphery, so as to reduce the vibration response by contacting adjacent turbine blades when vibration occurs.

[0011] However, the combustion gases discharged from the combustion chamber are supplied to the turbine and impact the turbine blades to generate rotational torque. However, since the damping pins used for the above-mentioned vibration damping are arranged at intervals along the periphery of the blade root, there is a problem of combustion gases leaking through the gaps created by the intervals between the damping pins. Summary of the Invention

[0012] The purpose of this invention is to provide a turbine blade combustion gas leakage prevention device and a gas turbine, which minimizes combustion gas leakage and reduces vibration response by forming bent extensions at both ends of the vertical damping pin in the transverse damping pin and the vertical damping pin formed along the circumferential surface of the blade root, and making them contact the two ends of the transverse damping pin for sealing.

[0013] The technical challenges of this invention are not limited to the objectives mentioned above. Other objectives and advantages of this invention not mentioned can be understood through the following description and will become more apparent through embodiments of this invention. Furthermore, it will be readily recognized that the objectives and advantages of this invention can be achieved through the solutions and combinations thereof described in the claims.

[0014] To achieve the above objectives, the present invention provides a combustion gas leakage prevention device for turbine blades, comprising: a first damping pin, arranged axially between the platform and the root of the turbine blade and inserted into a pin groove; a second damping pin and a third damping pin, respectively arranged on the lower sides of the first damping pin and inserted into the pin groove; and a retaining plate, arranged below the second damping pin and the third damping pin.

[0015] In this turbine blade combustion gas leakage prevention device, by forming bent extensions at both ends of the vertical damping pin (i.e., the first damping pin) and the second and third vertical damping pins formed along the circumferential surface of the turbine blade root, and making them contact the ends of the lateral damping pin to seal them, the vibration response is reduced by contacting the adjacent turbine blade when vibration occurs, and the leakage of combustion gas is also minimized.

[0016] Therefore, the second damping pin and the third damping pin include: a main body having a hollow portion; a curved extension extending outward from the circumference of both ends of the main body; the curved extension may include a notched groove for inserting the contact end of the first damping pin on one side and the contact end of the retaining plate on the other side.

[0017] According to the turbine blade combustion gas leakage prevention device of the present invention, the main body may include at least one of the following cross-sectional shapes: an annular cross-sectional shape perpendicular to the length direction; and a "C"-shaped cross-sectional shape having a cut portion along the length direction and perpendicular to the length direction.

[0018] According to the present invention, the main body of the turbine blade combustion gas leakage prevention device may include: an upper curved portion that contacts the turbine blade; and a lower curved portion formed on one side of the upper curved portion and inserted into a pin groove.

[0019] According to the turbine blade combustion gas leakage prevention device of the present invention, the curved extension may have an arc-shaped cross section.

[0020] According to the turbine blade combustion gas leakage prevention device of the present invention, the slit groove may include: a first slit portion, which contacts the vertical ends of the first damping pin and the retaining plate respectively, and is cut along a direction corresponding to the length direction of the main body portion; and a second slit portion, which contacts the horizontal ends of the first damping pin and the retaining plate respectively, and is cut along a direction intersecting the direction corresponding to the length direction of the main body portion.

[0021] The turbine blade combustion gas leakage prevention device according to the present invention further includes an auxiliary damping pin inserted into the hollow portion of at least one of the second damping pin and the third damping pin, the auxiliary damping pin being formed to be relatively smaller than the inner diameter of at least one of the second damping pin and the third damping pin.

[0022] According to the turbine blade combustion gas leakage prevention device of the present invention, the second damping pin, the first damping pin and the third damping pin are continuously arranged along the periphery of the blade root. The first damping pin is arranged between the platform and the blade root to extend axially. The second damping pin and the third damping pin can be arranged on both sides of the lower part of the first damping pin in a direction that intersects each other.

[0023] The present invention provides a gas turbine comprising: a compressor for drawing in and compressing external air; a combustion chamber for mixing and burning the air compressed by the compressor with fuel; a turbine having turbine blades and turbine stator blades mounted within a turbine housing, the turbine blades being rotated by combustion gases discharged from the combustion chamber; the gas turbine further comprising: a first damping pin arranged axially between the platform and the root of the turbine blade and inserted into a pin groove; a second damping pin and a third damping pin arranged on both sides of the lower portion of the first damping pin and inserted into the pin groove; and a retaining plate arranged below the second and third damping pins; the second and third damping pins comprising: a body portion having a hollow portion; and a curved extension portion extending outward from the circumference of both ends of the body portion; the curved extension portion including a notched groove portion for inserting a contact end of the first damping pin on one side and a contact end of the retaining plate on the other side.

[0024] According to the present invention, a combustion gas leakage prevention device for turbine blades having the above-described configuration and a gas turbine including the device, by forming bent extensions at both ends of the vertical damping pin in the transverse damping pin and the vertical damping pin formed along the circumferential surface of the blade root, and making them contact the two ends of the transverse damping pin to seal, vibration response is reduced by contacting adjacent turbine blades when vibration occurs, and the effect of minimizing combustion gas leakage is also achieved. Attached Figure Description

[0025] Figure 1 This is a diagram illustrating the schematic structure of a gas turbine according to an embodiment of the present invention.

[0026] Figure 2 This is a perspective view showing the overall structure of the turbine blade involved in the embodiment of the present invention.

[0027] Figure 3 This is a perspective view showing the damping pin involved in the embodiment of the present invention.

[0028] Figure 4 It shows the insertion into Figure 2 A three-dimensional view of the damping pin in the pin groove of a turbine blade.

[0029] Figure 5 and Figure 6 This is a cross-sectional view showing the installation state of damping pins between adjacent turbine blades according to an embodiment of the present invention.

[0030] Figure 7 This is a perspective view showing the damping pin and auxiliary damping pin involved in the embodiments of the present invention.

[0031] Figure 8 This is a cross-sectional view showing the state in which an auxiliary damping pin is inserted into the damping pin according to an embodiment of the present invention.

[0032] Explanation of reference numerals in the attached figures

[0033] 1: Pin groove 10: Damping pin 10a: First damping pin 10b: Second damping pin 10c: Third damping pin 11: Main body 11a: Hollow part 11b: Cut part 11c: Upper curved part 11d: Lower curved part 12: Bending extension part 12a: Cut groove part. Detailed Implementation

[0034] Hereinafter, some embodiments of the present invention will be described in detail with reference to the exemplary accompanying drawings. When adding reference numerals to the constituent elements of the drawings, it should be noted that the same reference numerals should be assigned to the same constituent elements, even if they are shown in different drawings. Furthermore, when describing embodiments of the present invention, detailed descriptions of related well-known structures or functions will be omitted if it is determined that such detailed descriptions would hinder the understanding of the embodiments of the present invention.

[0035] Furthermore, when describing the constituent elements of embodiments of the present invention, terms such as first, second, A, B, (a), and (b) may be used. These terms are only used to distinguish the constituent element from other constituent elements and do not limit the nature, order, or sequence of the constituent elements. When a constituent element is described as being "connected," "combined," or "accessed" to other constituent elements, it should be understood that the constituent element can be directly connected to or accessed to the other constituent elements, but constituent elements can also be "connected," "combined," or "accessed" to other constituent elements.

[0036] Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings.

[0037] Figure 1 This is a diagram illustrating the schematic structure of a gas turbine according to an embodiment of the present invention.

[0038] Reference Figure 1 The gas turbine 100 according to the present invention includes a compressor 110, a combustion chamber 120, and a turbine 130. With reference to the flow direction of the gas (compressed air or combustion gas), the compressor 110 is arranged upstream of the gas turbine 100, and the turbine 130 is arranged downstream. The combustion chamber 120 is arranged between the compressor 110 and the turbine 130.

[0039] [compressor]

[0040] The compressor 110 houses compressor stator vanes 112 and compressor rotor 113 within a compressor housing 111. The turbine 130 houses turbine stator vanes 160 and turbine rotor 170 within a turbine housing 150. These compressor stator vanes 112 and compressor rotor 113 are arranged in a multi-stage configuration along the direction of compressed air flow, and the turbine stator vanes 160 and turbine rotor 170 are also arranged in a multi-stage configuration along the direction of combustion gas flow. The compressor 110 is designed such that its internal space gradually decreases from the front-stage to the rear-stage side in order to compress the intake air; conversely, the turbine 130 is designed such that its internal space gradually increases from the front-stage to the rear-stage side in order to expand the combustion gas supplied from the combustion chamber 120.

[0041] On the other hand, a torque tube 200, serving as a torque transmission component, is arranged between the compressor rotor 113 located on the last stage side of the compressor 110 and the turbine rotor 170 located on the first stage side of the turbine 130, for transmitting the rotational torque generated by the turbine 130 to the compressor 110. The torque tube 200 is as follows... Figure 1 The torque tube 200 can be composed of multiple torque tubes (not shown) consisting of a total of 3 stages, but this is only one of the many embodiments of the present invention. The torque tube 200 can also be composed of multiple torque tubes with 4 or more stages or 2 or fewer stages.

[0042] The compressor rotor 113 includes a compressor disc 114 and compressor blades 115. Multiple (e.g., 14) compressor discs 114 are disposed inside the compressor housing 111, and each compressor disc 114 is fastened axially without separating by tie rods 210. More specifically, the compressor discs 114 are arranged axially relative to each other with the tie rods 210 passing through their centers. Furthermore, the opposing surfaces of adjacent compressor discs 114 are pressed together by the tie rods 210, thereby preventing them from rotating relative to each other.

[0043] Multiple compressor blades 115 are radially arranged on the outer circumferential surface of the compressor disc 114. Furthermore, between the compressor blades 115, multiple compressor stator vanes 112 are arranged in a ring on the inner circumferential surface of the compressor housing 111, based on the same stage. Unlike the compressor disc 114, the compressor stator vanes 112 remain fixed and do not rotate. Their function is to rectify the flow of compressed air after passing through the compressor blades 115 and guide the compressed air to the downstream compressor blades 115. In this case, the compressor housing 111 and the compressor stator vanes 112 can be collectively referred to as the compressor stator to distinguish them from the compressor rotor 113.

[0044] The tie rod 210 is arranged to pass through the center of multiple compressor discs 114 and the turbine disc 180 described later. One end of the tie rod is fastened to the compressor disc 114 located at the front end of the compressor 110, and the other end is fastened by a fastening nut 116.

[0045] The form of the tie rod 210 can vary depending on the gas turbine 100, and is therefore not necessarily limited to a certain type. Figure 1 The configuration shown is as follows. That is, it can have a configuration where a single pull rod 210 passes through the center of the compressor disc 114 and the turbine disc 180, as shown in the figure; it can also have a configuration where multiple pull rods 210 are arranged circumferentially; or it can be a combination of both.

[0046] Although not shown, in the compressor 110 of the gas turbine 100, a deswirler that acts as a guide vane can be installed to adjust the flow angle of the fluid entering the inlet of the combustion chamber 120 after the fluid pressure is increased to the design flow angle.

[0047] In the combustion chamber 120, the incoming compressed air is mixed with fuel and burned to produce high-energy, high-temperature, high-pressure combustion gas. The temperature of the combustion gas is raised to the heat resistance limit that the combustion chamber 120 and turbine 130 can withstand through the isobaric combustion process.

[0048] [Combustion Chamber]

[0049] The combustion chamber 120 constituting the combustion system of the gas turbine 100 can be arranged in a shell in the form of a canister, including: a burner containing fuel injection nozzles, etc.; a combustion chamber liner forming the combustion chamber; and a transition piece serving as the connection between the combustion chamber 120 and the turbine 130.

[0050] Specifically, the bushing can provide a combustion space so that fuel injected by the fuel injection nozzle mixes with and burns the compressed air of the compressor 110. Such a bushing may include: a flame tube that provides a combustion space for the combustion of the fuel mixed with air; and a bushing annular flow path that surrounds the flame tube and forms an annular space.

[0051] In addition, a fuel nozzle is attached to the front end of the bushing, and an igniter can be attached to the side wall.

[0052] On the other hand, a transition section is connected to the rear end of the bushing to deliver the combustion gases to the turbine 130 side. Similar to the bushing, the outer wall of this transition section is cooled by compressed air supplied from the compressor 110 to prevent damage due to the high temperature of the combustion gases.

[0053] For this purpose, the transition section is provided with holes for cooling so that air can be injected into the interior. After the compressed air cools the interior body through these holes, it flows to the bushing side.

[0054] On the other hand, the high-temperature, high-pressure combustion gas discharged from the combustion chamber 120 is supplied to the turbine 130. The high-temperature, high-pressure combustion gas supplied to the turbine 130 expands as it passes through the turbine 130 and impacts the turbine blades 190, generating a reaction force and thus producing rotational torque. This rotational torque is transmitted to the compressor 110 via the torque tube 200. Excess power beyond what is required to drive the compressor 110 can be used to drive a generator, etc.

[0055] [turbine]

[0056] The structure of the turbine 130 is basically similar to that of the compressor 110. That is, the turbine 130 also has multiple turbine rotors 170 similar to the compressor rotor 113 of the compressor 110. Therefore, the turbine rotor 170 also includes a turbine disk 180 and multiple turbine blades 190 arranged radially from the turbine disk.

[0057] Between the turbine blades 190, based on the same stage, multiple turbine stator vanes 160 are also arranged in a ring on the turbine housing 150. The turbine stator vanes 160 guide the flow direction of the combustion gas after passing through the turbine blades 190. At this time, the turbine housing 150 and the turbine stator vanes 160 can also be defined as the turbine stator 140 in order to distinguish them from the turbine rotor 170.

[0058] Figure 2 This is a perspective view showing the overall structure of the turbine blade involved in the embodiment of the present invention. Figure 3 This is a perspective view showing the damping pin according to an embodiment of the present invention. Figure 4 It shows the insertion into Figure 2 A three-dimensional view of the damping pin in the pin groove of a turbine blade. Figure 5 and Figure 6 This is a cross-sectional view showing the installation state of damping pins between adjacent turbine blades according to an embodiment of the present invention.

[0059] The turbine disk 180 is disc-shaped, with multiple turbine disk grooves (not shown) formed on its outer circumference.

[0060] Reference Figure 2The turbine blade 190 can be mounted radially outside the turbine disk 180. The turbine blade 190 may include: a root member (not shown) coupled to the turbine disk groove; a platform 192 coupled radially outside the root member; and an airfoil 193 coupled radially outside the platform 192 and rotated by the combustion gas.

[0061] In addition, the side of platform 192 is in contact with the side of the platform 192 of other adjacent turbine blades 190, thereby maintaining the spacing between turbine blades 190.

[0062] A root member is provided on the bottom surface of platform 192 to fasten to the turbine disk groove. The root member is formed to correspond to the shape of the curved surface formed on the turbine disk groove. This can be selected according to the required structure of gas turbine 100, and can have a known dovetail shape or fir-tree shape.

[0063] The root components can be fastened in two ways: tangential type, where the component is inserted into the turbine disk groove along the tangential direction of the outer circumferential surface of the turbine disk 180; and axial type, where the component is inserted into the turbine disk groove along the axial direction of the turbine disk 180. Depending on the circumstances, other fastening devices besides the above-mentioned types may also be used, such as using fasteners like keys or bolts to fasten the turbine blade 190 to the turbine disk 180.

[0064] An airfoil 193 is provided on the upper surface of the platform 192. The airfoil 193 is formed as an airfoil optimized according to the specifications of the gas turbine 100, and has a leading edge arranged on the upstream side and a trailing edge arranged on the downstream side based on the direction of combustion gas flow.

[0065] Here, unlike the compressor blades 115, the turbine blades 190 are in direct contact with the high-temperature, high-pressure combustion gases. Since the temperature of the combustion gases reaches as high as 1700°C, a cooling mechanism is required. For this purpose, an extraction passage (not shown) is provided to draw compressed air from a portion of the compressor 110 and supply it to the turbine blades 190.

[0066] The air extraction passage can extend from the outside of the housing or through the interior of the compressor disc 114, and both external and internal flow paths can be used simultaneously. Multiple film cooling holes (not shown) can be formed on the surface of the blade 193, which are connected to the cooling flow path (not shown) formed inside the blade 193, serving to supply compressed air to the surface of the blade 193.

[0067] Reference Figure 2 Turbine blades 190 are mounted adjacent to each other along the circumference of turbine disk 180, and a plurality of pin slots 1 are formed on the corresponding surfaces of a platform 192 of one turbine blade 190 and another adjacent platform 192.

[0068] [Pin slot]

[0069] Based on platform 192, multiple pin grooves 1 are formed in the lower part of platform 192, and a blade root 197 is formed to provide space for cooling air flow.

[0070] Multiple pin slots 1 are formed along the axial direction of the gas turbine 100 on the lower part of the platform 192. Corresponding pin slots 1 are also formed on the lower part of the platform 192 of other adjacent turbine blades 190 along the circumferential direction on the turbine disk 180. Damping pins 10 can be inserted into the pin slots 1.

[0071] [Damping pin]

[0072] When the turbine blades 190 vibrate during the operation of the gas turbine 100, the damping pin 10 reduces the vibration energy by contacting other adjacent turbine blades 190.

[0073] like Figure 2 As shown, the damping pin 10 includes a first damping pin 10a, which is arranged to extend axially between the platform 192 and the blade root 197. The first damping pin 10a is a transverse damping pin that extends laterally along the platform 192.

[0074] In addition, the damping pin 10 also includes a second damping pin 10b and a third damping pin 10c, which are arranged on both sides of the lower part of the first damping pin 10a in a direction that intersects with each other.

[0075] Here, the direction in which the lower portions of the two sides of the first damping pin 10a intersect each other can be defined as being arranged vertically relative to the extension direction of the first damping pin 10a. Therefore, the second damping pin 10b and the third damping pin 10c can be arranged along a direction having a vertical component relative to the first damping pin 10a. However, this is only a relative direction based on the extension direction of the first damping pin 10a, and the vertical direction does not necessarily refer to the vertical direction. That is, the second damping pin 10b and the third damping pin 10c can extend in a direction inclined relative to the first damping pin 10a, arranged to extend downward from the outer portion of the first damping pin 10a, and protrude outward from these two points respectively.

[0076] With the first damping pin 10a to the third damping pin 10c arranged along the periphery of the blade root 197 in a state of being inserted into the pin groove 1, a sealing effect can be achieved between one platform 192 and the adjacent platform 192.

[0077] Reference Figure 4 Damping pins 10 are also installed at corresponding positions on the lower part of the platform 192 of the other turbine blades 190 that are circumferentially adjacent to the turbine disk 180. The damping pins 10 thus set on the turbine blades 190 protrude toward the opposite surface of the other adjacent turbine blades 190 to form a surface contact state, thereby reducing the vibration response of the turbine blades 190.

[0078] Observing the structure of the pin groove 1 involved in the embodiment of the present invention, it is formed as a recessed semi-circular groove structure for inserting and placing a damping pin 10 with a circular cross-section, and has a length corresponding to the damping pin 10. When the damping pin 10 is inserted into the pin groove 1, the damping pin 10 is not easily dislodged from the pin groove 1.

[0079] like Figure 3 As shown, the damping pin 10 involved in the embodiments of the present invention not only includes structures installed on both sides of the transverse damping pin, i.e., the first damping pin 10a, to reduce vibration response, but also includes a second damping pin 10b and a third damping pin 10c for preventing combustion gas leakage.

[0080] Therefore, the damping pin 10 involved in the embodiments of the present invention will be mainly described around the second damping pin 10b and the third damping pin 10c.

[0081] like Figure 5 As shown, the second damping pin 10b and the third damping pin 10c involved in the embodiments of the present invention are configured to include a main body 11 having a hollow portion 11a and an annular cross-sectional shape perpendicular to the length direction.

[0082] In addition, such as Figure 6 As shown, the second damping pin 10b and the third damping pin 10c according to another embodiment of the present invention include a main body 11, which has a cut portion 11b cut along the length direction on one side and has a "C" shaped cross-section perpendicular to the length direction.

[0083] Referring to the accompanying drawings, the damping pin 10, including the second damping pin 10b and the third damping pin 10c, reduces vibration responsiveness by contacting the adjacent turbine blade 190 when vibration occurs, while also acting as a flow resistance to the combustion gas to prevent combustion gas leakage.

[0084] The main body 11 is the part that plays a damping role to reduce vibration response, and it is formed to be long along the length direction.

[0085] The main body 11 may adopt an annular cross-section structure and a "C" shaped cross-section, taking into account damping characteristics, but is not limited to these, and can be deformed in various ways.

[0086] However, it can be defined as a damping structure that reduces vibration responsiveness by contacting adjacent turbine blades 190, and in embodiments of the invention it may include annular cross-sectional shapes and "C" shaped cross-sectional shapes.

[0087] In addition, the main body 11 includes a curved extension 12, which is used to block the flow path of combustion gases so as to act as a flow resistance to combustion gases.

[0088] The curved extension 12 may include an arc-shaped cross section extending outward from the circumference of both ends of the main body 11 with reference to both ends of the main body 11. That is, the curved extension 12 may be an extension portion extending outward from the upper curved portion 11c of the main body 11, which will be described later.

[0089] In addition, the main body 11 includes a lower curved portion 11d with a "U"-shaped cross-section formed on one side of the upper curved portion 11c, which is inserted into and fixed in a groove-shaped pin groove 1.

[0090] The main body 11 does not include the curved extension 12, and may be a cylindrical part other than the curved extension 12.

[0091] The upper curved portion 11c can be a contact surface that contacts the adjacent turbine blade 190. This upper curved portion 11c, by contacting the adjacent turbine blade 190, can reduce vibration response.

[0092] As described above, the curved extension 12 may be a U-shaped portion that extends integrally outward from both ends of the main body 11. The curved extension 12 may be a portion whose one end contacts the end of the first damping pin 10a and whose other end contacts the retainer plate 194 mounted on the turbine blade 190.

[0093] A notch 12a may be formed in the curved extension 12, which is used to insert the contact end that contacts the first damping pin 10a and the contact end that contacts the retaining plate 194.

[0094] Here, the ends of the first damping pin 10a and the retaining plate 194 are formed by a rectangular cross-section structure, which has a shorter vertical end and a horizontal end that is relatively longer than the vertical end.

[0095] A portion of the notch groove 12a contacts the horizontal ends of the first damping pin 10a and the retaining plate 194, while another portion of the notch groove 12a contacts the vertical ends of the first damping pin 10a and the retaining plate 194, thereby blocking the leakage of combustion gases without interfering with each other.

[0096] The notch groove 12a includes a first notch 12a-1, which contacts the vertical ends of the first damping pin 10a and the retaining plate 194 respectively, and is cut along a direction corresponding to the length direction of the main body 11.

[0097] Furthermore, the notch groove 12a includes a second notch 12a-2, which contacts the horizontal ends of the first damping pin 10a and the retaining plate 194 respectively, and is cut in a direction that intersects with the direction corresponding to the length direction of the main body 11.

[0098] That is, the first cut portion 12a-1 and the second cut portion 12a-2 in the cut groove portion 12a where the horizontal and vertical ends of the first damping pin 10a and the retaining plate 194 meet can be formed to intersect each other.

[0099] The second damping pin 10b and the third damping pin 10c, which are thus provided with the curved extension 12, can be arranged in directions that intersect each other, with the first damping pin 10a as a reference.

[0100] The cutouts 12a of the curved extensions 12 formed on the second damping pin 10b arranged in the cross direction on one side of the first damping pin 10a and the third damping pin 10c arranged in the cross direction on the other side of the first damping pin 10a can contact the two ends of the first damping pin 10a and the retaining plate 194.

[0101] Thus, when the notched groove 12a of the curved extension 12 formed on the second damping pin 10b and the third damping pin 10c contacts the first damping pin 10a and the two ends of the retaining plate 194, a sealing effect can be ensured to prevent leakage of combustion gas flowing between the platform 192 and the blade root 197 of the turbine blade 190.

[0102] This is because when the bent extension 12 of the damping pin 10 contacts the end of the first damping pin 10a and the end of the retaining plate 194, it blocks the exhaust path of the combustion gas, thus playing the role of resistance to the flow of combustion gas.

[0103] In other words, the first damping pin 10a and the retaining plate 194 are in contact with the cut grooves 12a formed on the curved extensions 12 of the second and third damping pins 10b and 10c, thereby ensuring a seal and thus acting as a barrier to block the flow of combustion gases.

[0104] Therefore, by providing a bent extension 12 on the damping pin 10 installed on the turbine blade 190 in order to reduce the vibration energy caused by friction, another effect can be obtained, such as reducing the amount of combustion gas leakage to the turbine blade 190.

[0105] Figure 7This is a perspective view showing the damping pin and auxiliary damping pin involved in the embodiments of the present invention. Figure 8 This is a cross-sectional view showing the state in which an auxiliary damping pin is inserted into the damping pin according to an embodiment of the present invention.

[0106] like Figure 7 As shown, in the hollow portion 11a of the main body portion 11 of the second damping pin 10b and the third damping pin 10c involved in the embodiment of the present invention, an auxiliary damping pin 20 may be inserted in order to minimize the risk of leakage of combustion gas when at least one of the second damping pin 10b and the third damping pin 10c is damaged.

[0107] like Figure 8 As shown, when the auxiliary damping pin 20 is inserted into at least one of the second damping pin 10b and the third damping pin 10c, if the damping pin is damaged due to design error or physical impact, in order to prevent the chain of damage impacts from being transmitted to the auxiliary damping pin 20, the auxiliary damping pin 20 may be formed to be relatively smaller than the inner diameter of the hollow portion 11a of the second damping pin 10b and the third damping pin 10c.

[0108] That is, when there is a free space between the circumferential surface of the auxiliary damping pin 20 and the inner diameter of the hollow portion 11a of the second damping pin 10b and the third damping pin 10c, the risk of damage to the auxiliary damping pin 20 will be reduced even if at least one of the second damping pin 10b and the third damping pin 10c is damaged, and the installation of the auxiliary damping pin 20 can minimize the leakage of combustion gases.

[0109] The auxiliary damping pin 20 can be formed in the same shape as the first damping pin 10a, but is not limited thereto, and can be in various shapes that can minimize combustion gas leakage.

[0110] The above description of the present invention is exemplary, and those skilled in the art should understand that it can be easily modified into other specific forms without changing the technical concept or essential features of the invention. Therefore, it should be understood that the above embodiments are exemplary in all respects and not restrictive. The scope of the present invention is defined by the scope of the claims described below, and all modifications or variations derived from the meaning, scope, and equivalent concepts of the claims should be interpreted as being included within the scope of the present invention.

Claims

1. A device for preventing combustion gas leakage from turbine blades, in, include: The first damping pin is arranged between the platform and the root of the turbine blade, extending axially and inserted into the pin groove. The second and third damping pins are respectively arranged on the lower parts of both sides of the first damping pin and inserted into the pin groove; and A retaining plate is arranged below the second damping pin and the third damping pin; The second damping pin and the third damping pin include: The main body has a hollow section; and The curved extension extends outward from the circumferential surfaces at both ends of the main body; The curved extension includes a notched groove for inserting the contact end of the first damping pin on one side and the contact end of the retaining plate on the other side.

2. The turbine blade combustion gas leakage prevention device according to claim 1, wherein, The main body includes at least one of the following cross-sectional shapes: an annular cross-sectional shape perpendicular to the length direction; and a "C"-shaped cross-sectional shape having a cut portion along the length direction and perpendicular to the length direction.

3. The turbine blade combustion gas leakage prevention device according to claim 2, wherein, The main body includes: The upper curved portion contacts the turbine blades; and The lower curved portion is formed on one side of the upper curved portion and inserted into the pin groove.

4. The turbine blade combustion gas leakage prevention device according to claim 1, wherein, The curved extension has an arc-shaped cross-section.

5. The turbine blade combustion gas leakage prevention device according to claim 1, wherein, The notch groove includes: The first cut portion contacts the vertical ends of the first damping pin and the retaining plate, and is cut along a direction corresponding to the length direction of the main body; and The second cut contacts the horizontal ends of the first damping pin and the retaining plate, and is cut along a direction that intersects with the direction corresponding to the length direction of the main body.

6. The turbine blade combustion gas leakage prevention device according to claim 1, wherein, It also includes an auxiliary damping pin, which is inserted into the hollow portion of at least one of the second and third damping pins. The auxiliary damping pin is formed with an inner diameter that is relatively smaller than that of at least one of the second damping pin and the third damping pin.

7. The turbine blade combustion gas leakage prevention device according to claim 1, wherein, The second damping pin, the first damping pin, and the third damping pin are continuously arranged along the periphery of the blade root. The first damping pin is arranged to extend axially between the platform and the blade root, and the second and third damping pins are arranged on the lower sides of the first damping pin in a direction that intersects with each other.

8. A gas turbine, wherein, The gas turbine includes: The compressor draws in and compresses outside air; The combustion chamber mixes and burns the air compressed by the compressor with fuel; A turbine, in which turbine blades and turbine stator are mounted within a turbine housing, and the turbine blades are rotated by combustion gases discharged from the combustion chamber; The gas turbine also includes: A first damping pin is arranged between the platform and the root of the turbine blade, extending axially and inserted into a pin groove. The second and third damping pins are arranged on the lower sides of the first damping pin and inserted into the pin grooves; and A retaining plate is arranged below the second damping pin and the third damping pin; The second damping pin and the third damping pin include: The main body has a hollow section; and The curved extension extends outward from the circumferential surfaces at both ends of the main body; The curved extension includes a notched groove for inserting the contact end of the first damping pin on one side and the contact end of the retaining plate on the other side.

9. The gas turbine according to claim 8, wherein, The main body includes at least one of the following cross-sectional shapes: an annular cross-sectional shape perpendicular to the length direction; and a "C"-shaped cross-sectional shape having a cut portion along the length direction and perpendicular to the length direction.

10. The gas turbine according to claim 9, wherein, The main body includes: The upper curved portion contacts the turbine blades; and The lower curved portion is formed on one side of the upper curved portion and inserted into the pin groove.

11. The gas turbine according to claim 8, wherein, The curved extension has an arc-shaped cross-section.

12. The gas turbine according to claim 8, wherein, The notch groove includes: The first cut portion contacts the vertical ends of the first damping pin and the retaining plate, and is cut along a direction corresponding to the length direction of the main body; and The second cut contacts the horizontal ends of the first damping pin and the retaining plate, and is cut along a direction that intersects with the direction corresponding to the length direction of the main body.

13. The gas turbine according to claim 8, wherein, It also includes an auxiliary damping pin, which is inserted into the hollow portion of at least one of the second and third damping pins. The auxiliary damping pin is formed with an inner diameter that is relatively smaller than that of at least one of the second damping pin and the third damping pin.

14. The gas turbine according to claim 8, wherein, The second damping pin, the first damping pin, and the third damping pin are continuously arranged along the periphery of the blade root. The first damping pin is arranged to extend axially between the platform and the blade root. The second damping pin and the third damping pin are arranged on the lower sides of the first damping pin in a direction that intersects with each other.

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