Airfoil and gas turbine comprising it

Abstract

An airfoil according to an embodiment of the present invention comprises an intake surface, a pressure surface, a leading edge, a trailing edge, and a through hole. The intake surface protrudes outwardly, forming a convex curved surface. The pressure surface is formed on the opposite side of the intake surface and forms a concave curved surface. The leading edge is positioned on the upstream side of the airflow and is formed by the contact of the leading edge of the pressure surface and the intake surface. The trailing edge is positioned on the downstream side of the airflow and is formed by the contact of the trailing edge of the pressure surface and the intake surface. The through hole is formed by penetrating from the pressure surface toward the intake surface.

Description

Technology Field

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

[0002] A turbine is a mechanical device that generates rotational force through impulse or reaction force by utilizing the flow of a compressible fluid, such as steam or gas; examples include steam turbines that use steam and gas turbines that use high-temperature combustion gases.

[0003] Among these, the gas turbine is largely composed of a compressor, a combustor, and a turbine. The compressor is equipped with an air inlet for introducing air, and a plurality of compressor vanes and compressor blades are alternately arranged within the compressor housing.

[0004] The combustion device generates high-temperature, high-pressure combustion gas by supplying fuel to the compressed air compressed by the compressor and igniting it with a burner.

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

[0006] The above rotor is rotatably supported at both ends by bearings. Additionally, a plurality of discs are fixed to the rotor, and each blade is connected to them, while a drive shaft, such as that of a generator, is connected to the end on the exhaust chamber side.

[0007] Since these gas turbines do not have a reciprocating mechanism like the piston of a four-stroke engine, there are no mutual friction parts like piston-cylinder, so the consumption of lubricating oil is extremely low, and the amplitude characteristic of reciprocating machines is greatly reduced, and high-speed motion is possible.

[0008] To briefly explain the operation of a gas turbine, air compressed by a compressor is mixed with fuel and combusted to produce high-temperature combustion gases, which are then injected toward the turbine. As the injected combustion gases pass through the turbine vanes and blades, they generate rotational force, causing the rotor to rotate. Prior art literature

[0009] Published Patent Application No. 10-2024-0095996 The problem to be solved

[0010] The present invention aims to provide an airfoil capable of improving the compression efficiency of a compressor and a gas turbine including the same. means of solving the problem

[0011] An airfoil according to an embodiment of the present invention comprises an intake surface, a pressure surface, a leading edge, a trailing edge, and a through hole. The intake surface protrudes outwardly, forming a convex curved surface. The pressure surface is formed on the opposite side of the intake surface and forms a concave curved surface. The leading edge is positioned on the upstream side of the airflow and is formed by the contact of the leading edge of the pressure surface and the intake surface. The trailing edge is positioned on the downstream side of the airflow and is formed by the contact of the trailing edge of the pressure surface and the intake surface. The through hole is formed by penetrating from the pressure surface toward the intake surface.

[0012] In an airfoil according to an embodiment of the present invention, the through hole may be a tip hole formed on the upper part of the pressure surface or a hub hole formed on the lower part of the pressure surface.

[0013] In an airfoil according to an embodiment of the present invention, the through hole may be formed in a straight or curved shape.

[0014] An airfoil according to an embodiment of the present invention, wherein the through hole is formed in a diagonal direction having a predetermined angle with respect to the tangential direction of the pressure surface.

[0015] In an airfoil according to an embodiment of the present invention, the preset angle may be in the range of 10 to 60 degrees.

[0016] In an airfoil according to an embodiment of the present invention, the through-hole may be formed such that the cross-sectional area decreases from the pressure surface toward the suction surface.

[0017] In an airfoil according to an embodiment of the present invention, the through hole may include a first path that starts at the entrance of the through hole and ends before reaching the exit, and a second path formed toward the exit of the through hole at the end of the first path.

[0018] In an airfoil according to an embodiment of the present invention, a first path may be formed in a diagonal direction having a first angle predetermined with respect to a pressure surface, and a second path may be formed in a diagonal direction having a second angle smaller than the first angle.

[0019] In an airfoil according to an embodiment of the present invention, the through hole may include a third path connecting the inlet of the through hole and a first outlet, and a fourth path formed to connect a second outlet at one point of the third path.

[0020] In an airfoil according to an embodiment of the present invention, the second outlet may be formed closer to the trailing edge side than the first outlet.

[0021] A gas turbine according to an embodiment of the present invention comprises a compressor, a combustor, and a turbine. The compressor compresses incoming air and includes multi-stage compressor blades installed radially and compressor vanes installed between the multi-stage compressor blades. The combustor mixes the air compressed by the compressor with fuel and combusts it. The turbine generates power from the gas combusted by the combustor. At least one of the compressor blades or the compressor vanes includes an airfoil. The airfoil includes an intake surface, a pressure surface, a leading edge, a trailing edge, and a through hole. The intake surface protrudes outwardly, forming a convex curved surface. The pressure surface is formed on the opposite side of the intake surface and forms a concave curved surface. The leading edge is positioned on the upstream side of the air flow and is formed by the contact of the leading edge of the pressure surface and the intake surface. The trailing edge is positioned on the downstream side of the air flow and is formed by the contact of the trailing edge of the pressure surface and the intake surface. The through hole is formed by penetrating from the pressure surface toward the intake surface.

[0022] In a gas turbine according to an embodiment of the present invention, the through hole may be a tip hole formed on the upper part of the pressure surface or a hub hole formed on the lower part of the pressure surface.

[0023] In a gas turbine according to an embodiment of the present invention, the through-hole may be formed in a straight or curved shape.

[0024] In a gas turbine according to an embodiment of the present invention, the through-hole may be formed in an oblique direction having a predetermined angle with respect to the tangential direction of the pressure surface.

[0025] In a gas turbine according to an embodiment of the present invention, a preset angle may be in the range of 10 to 60 degrees.

[0026] In a gas turbine according to an embodiment of the present invention, the through-hole may be formed such that the cross-sectional area decreases from the pressure surface toward the suction surface.

[0027] In a gas turbine according to an embodiment of the present invention, the through hole may include a first path that starts at the inlet of the through hole and ends before reaching the outlet, and a second path formed toward the outlet of the through hole at the end of the first path.

[0028] In a gas turbine according to an embodiment of the present invention, a first path may be formed in a diagonal direction having a first angle predetermined with respect to a pressure surface, and a second path may be formed in a diagonal direction having a second angle smaller than the first angle.

[0029] In a gas turbine according to an embodiment of the present invention, the through hole may include a third path connecting the inlet of the through hole and a first outlet, and a fourth path formed to connect a second outlet at one point of the third path.

[0030] In a gas turbine according to an embodiment of the present invention, the second outlet may be formed closer to the trailing edge side than the first outlet. Effects of the invention

[0031] When an airfoil according to an embodiment of the present invention is applied to a compressor blade or compressor vane, the surge margin and polytropic efficiency of the compressor can be increased, thereby improving the compression efficiency of the compressor. Brief explanation of the drawing

[0032] FIG. 1 is a perspective view showing a gas turbine according to an embodiment of the present invention with a portion cut away. FIG. 2 is a cross-sectional view showing the schematic structure of a gas turbine according to an embodiment of the present invention. FIG. 3 is a perspective view showing an airfoil according to a first embodiment of the present invention, which is shown applied to a compressor blade. FIG. 4 is a side view showing an airfoil according to the first embodiment of the present invention, which is shown applied to a compressor vane. Figure 5 is a top view of the A-A' section of Figure 3. Figure 6 is a top view of the B-B' section of Figure 3. Figure 7 is a diagram illustrating the flow separation phenomenon caused by an airfoil according to the prior art. FIG. 8 is a drawing illustrating the flow separation phenomenon caused by an airfoil according to the first embodiment of the present invention. Figure 9 is a diagram showing the compression ratio relative to the flow rate. Figure 10 is a diagram showing polytropic efficiency relative to flow rate. FIG. 11 is a drawing showing an airfoil according to a second embodiment of the present invention. FIG. 12 is a drawing showing an airfoil according to a third embodiment of the present invention. FIG. 13 is a drawing showing an airfoil according to a fourth embodiment of the present invention. Specific details for implementing the invention

[0033] The present invention is capable of various modifications and may have various embodiments, and specific embodiments are illustrated and described in detail in the detailed description. However, this is not intended to limit the present invention to specific embodiments, and it should be understood that it includes all modifications, equivalents, and substitutions that fall within the spirit and scope of the invention.

[0034] The terms used in this invention are used merely 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 "comprising" or "having" are intended to indicate the existence of the features, numbers, steps, actions, components, parts, or combinations thereof described in the specification, and should be understood as not precluding the existence or addition of one or more other features, numbers, steps, actions, components, parts, or combinations thereof.

[0035] Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the attached drawings. In this case, identical components in the attached drawings are indicated by the same reference numerals whenever possible. Furthermore, detailed descriptions of known functions and configurations that may obscure the essence of the present invention will be omitted. For the same reason, some components in the attached drawings have been exaggerated, omitted, or schematically illustrated.

[0037] FIG. 1 is a perspective view showing a gas turbine according to an embodiment of the present invention with a portion cut away, and FIG. 2 is a cross-sectional view showing the internal structure of a gas turbine according to an embodiment of the present invention.

[0038] As illustrated in FIG. 1, a gas turbine (1000) according to an embodiment of the present invention includes a compressor (1100), a combustor (1200), and a turbine (1300). The compressor (1100) has a plurality of compressor blades (1110) installed radially. The compressor (1100) rotates the compressor blades (1110), and air moves as it is compressed by the rotation of the compressor blades (1110). The size and installation angle of the compressor blades (1110) may vary depending on the installation location. In the first embodiment, the compressor (1100) is connected directly or indirectly to the turbine (1300) so that a portion of the power generated from the turbine (1300) can be transferred and used to rotate the compressor blades (1110).

[0039] 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 shape and a fuel nozzle module (1220).

[0040] As illustrated in FIG. 2, a gas turbine (1000) according to an embodiment of the present invention is provided with a housing (1010), and a diffuser (1400) is provided at the rear of the housing (1010) through which combustion gas passing through the turbine is discharged. Then, a combustor (1200) is disposed in front of the diffuser (1400) to receive compressed air and combust it.

[0041] When described based on the direction of air flow, a compressor (1100) is located on the upstream side of the housing (1010), and a turbine (1300) is positioned on the downstream side. Additionally, a torque tube (1500) is positioned between the compressor (1100) and the turbine (1300) as a torque transmission member that transmits the rotational torque generated by the turbine (1300) to the compressor (1100).

[0042] The compressor (1100) is equipped with a plurality (e.g., 14) of compressor rotor disks (1120), and each of the compressor rotor disks (1120) is connected by a tie rod (1600) so as not to be separated in the axial direction.

[0043] Specifically, each compressor rotor disk (1120) is aligned along the axial direction with the tie rod (1600) forming the rotation axis passing approximately through the center. Here, each adjacent compressor rotor disk (1120) is positioned so that the opposing surfaces are compressed by the tie rod (1600), making relative rotation impossible.

[0044] A plurality of compressor blades (1110) are radially connected to the outer surface of the compressor rotor disk (1120). Each compressor blade (1110) is equipped with a dovetail portion (1112) and is connected to the compressor rotor disk (1120).

[0045] Between each rotor disk (1120), a compressor vane (1130) is positioned and fixed to the housing. Unlike the rotor disks, the compressor vane (1130) is fixed so as not to rotate and serves to align the flow of compressed air passing through the compressor blade (1110) of the compressor rotor disk (1120) and guide the air to the compressor blade (1110) of the rotor disk (1120) located downstream.

[0046] The fastening method of the dovetail section (1112) is of the tangential type and the axial type. This can be selected according to the required structure of the commercially available gas turbine and can have a commonly known dovetail or fir-tree shape. In some cases, the blade can be fastened to the rotor disk using a fastening device other than the above shapes, such as a key or a bolt.

[0047] A tie rod (1600) is positioned to penetrate the center of a plurality of compressor rotor disks (1120) and turbine rotor disks (1320), and the tie rod (1600) may consist of one or more tie rods. One end of the tie rod (1600) is fastened within the compressor rotor disk located at the uppermost side, and the other end of the tie rod (1600) is fastened by a fixing nut (1450).

[0048] Since the shape of the tie rod (1600) can be formed in various structures depending on the gas turbine, it is not necessarily limited to the shape shown in FIG. 2. That is, as illustrated, it may have a shape where one tie rod penetrates the center of the rotor disk, or a shape where multiple tie rods are arranged circumferentially, and a combination of these is also possible.

[0049] Although not shown in the drawings, a vane acting as a guide vane may be installed at the next position after the diffuser in the compressor of a gas turbine to match the flow angle of the fluid entering the combustor inlet to the design flow angle after increasing the fluid pressure, and this is called a deswirler.

[0050] In the combustor (1200), the introduced compressed air is mixed with fuel and combusted to produce high-energy, high-temperature, high-pressure combustion gas, and the combustion gas temperature is raised to a heat resistance limit that the combustor and turbine parts can withstand through an isostatic combustion process.

[0051] Combustors constituting the combustion system of a gas turbine may be arranged in multiple numbers within a housing formed in a cell shape and are composed of a burner including a fuel injection nozzle, a combustor liner forming a combustion chamber, and a transition piece that serves as a connection part between the combustor and the turbine.

[0052] Specifically, the liner provides a combustion space where fuel injected by the fuel nozzle is mixed with the compressed air of the compressor and combusted. This liner may include a flame chamber that provides a combustion space where the fuel mixed with air is combusted, and a flow sleeve that surrounds the flame chamber and forms an annular space. Additionally, a fuel nozzle is connected to the front end of the liner, and a spark plug is connected to the side wall.

[0053] Meanwhile, a transition piece is connected to the rear end of the liner to allow combustion gases, combusted by a spark plug, to be sent to the turbine side. The outer wall of this transition piece is cooled by compressed air supplied from the compressor to prevent damage caused by the high temperature of the combustion gases.

[0054] To this end, the transition piece is provided with cooling holes to allow air to be injected into the interior, and the compressed air cools the main body inside through the holes before flowing toward the liner.

[0055] Cooling air that has cooled the aforementioned transition piece flows through the annular space of the liner, and compressed air from the outside of the flow sleeve can collide with the outer wall of the liner by being supplied as cooling air through cooling holes provided in the flow sleeve.

[0056] Meanwhile, the combustion gas from the combustion chamber is supplied to the turbine (1300). As the combustion gas expands, it collides with the rotating blades of the turbine, generating a reaction force and causing rotational torque. This rotational torque is transmitted to the compressor via the torque tube (1500), and any power exceeding the power required to drive the compressor is used to drive a generator, etc.

[0057] The turbine (1300) is basically similar in structure to a compressor. That is, the turbine (1300) is also equipped with a plurality of turbine rotor disks (1320) similar to the compressor rotor disks of a compressor. Accordingly, the turbine rotor disks (1320) also include a plurality of turbine blades (1310) arranged radially. The turbine blades (1310) can also be connected to the turbine rotor disks (1320) in a manner such as a dovetail. In addition, between the blades (1310) of the turbine rotor disks (1320), turbine vanes (1330) are provided, which are fixed to the turbine casing (1350) via a turbine vane carrier (1340), thereby guiding the flow direction of the combustion gas passing through the blades. The turbine vane carrier (1340) is formed to be connected to the turbine casing (1350) and is equipped with at least one turbine vane (1330).

[0059] FIGS. 3 and FIGS. 4 are drawings illustrating an airfoil according to a first embodiment of the present invention, where FIG. 3 is a perspective view showing it applied to a compressor blade and FIG. 4 is a side view showing it applied to a compressor vane. FIG. 5 is a top view of the A-A' section of FIG. 3, and FIG. 6 is a top view of the B-B' section of FIG. 3.

[0060] Referring to FIGS. 3 and 4, the airfoil (2000_1) according to the first embodiment of the present invention may be made of a curved plate with a wing shape and may be formed to have an optimized airfoil shape according to the specifications of the gas turbine (1000).

[0061] The airfoil (2000_1) can be applied to at least one of the compressor blade (1110) and the compressor vane (1130). Meanwhile, the compressor blade (1110) and the turbine blade (1310), and the compressor vane (1130) and the turbine vane (1330) have similar external appearances, but their internal structures are different. That is, the airfoil applied to the turbine blade (1310) and the turbine vane (1330) has a cavity and a cooling passage formed inside to prevent overheating by combustion gas, but the airfoil (2000_1) applied to the compressor blade (1110) and the compressor vane (1130) is not overheated by combustion gas, so a separate cooling passage is not formed inside the airfoil (2000_1).

[0062] The airfoil (2000_1) includes a suction surface (2400) that protrudes outwardly and forms a convex curved surface, and a pressure surface (2300) that forms a concave curved surface toward the suction surface (2400).

[0063] The airfoil (2000_1) includes a leading edge (2100) and a trailing edge (2200). The leading edge (2100) is positioned on the upstream side of the airflow and is formed by the contact of the leading edge of the pressure surface (2300) and the suction surface (2400). The trailing edge (2200) is positioned on the downstream side of the airflow and is formed by the contact of the trailing edge of the pressure surface (2300) and the suction surface (2400).

[0064] The airfoil (2000_1) includes through holes (2500: 2510, 2520). The through holes (2500: 2510, 2520) may be formed in a straight or curved shape. Among the through holes (2500), the first through hole (2510) formed at the lower part of the pressure surface (2300) is called a hub hole, and the second through hole (2520) formed at the upper part of the pressure surface (2300) is called a tip hole. The first through hole (2510) and the second through hole (2520) are merely names based on their formed locations, and their shapes and functions are identical. Although the through holes (2500: 2510, 2520) are formed in multiple numbers, they are described as a single number for convenience.

[0065] When the airfoil (2000_1) is a component of the compressor blade (1110), the first through-hole (2510) may be formed adjacent to the platform (1111) coupled to the lower part of the airfoil (2000_1), and the second through-hole (2520) may be formed on the upper part of the airfoil (2000_1), on the side of the compressor housing (1150).

[0066] When the airfoil (2000_1) is a component of the compressor vane (1130), the first through-hole (2510) may be formed adjacent to the upper vane platform (1331) formed on the side of the compressor housing (1150), and the second through-hole (2520) may be formed adjacent to the lower vane platform (1332).

[0067] In the following description, the airfoil (2000_1) is described as an example applied to the compressor blade (1110). Of course, the airfoil (2000_1) described in the following description can be applied substantially the same way to the compressor vane (1130).

[0068] Referring to FIGS. 5 and 6, the through hole (2500) is formed by penetrating from the pressure surface (2300) toward the suction surface (2400). Since the interior of the airfoil applied to the turbine blade (1310) and turbine vane (1330) has cavities and cooling channels, a through hole cannot be formed. However, since the interior of the airfoil (2000_1) applied to the compressor blade (1110) and compressor vane (1130) is filled, a through hole (2500) penetrating from the pressure surface (2300) toward the suction surface (2400) of the airfoil (2000_1) can be formed.

[0069] The through hole (2500) can be formed in a diagonal direction having a predetermined angle (θ) with respect to the tangential direction of the pressure surface (2300). As a result, the inlet (2501) of the through hole (2500) is positioned close to the leading edge (2100) side, and the outlet (2502) of the through hole (2500) is positioned close to the trailing edge (2200) side.

[0070] The cross-sectional shape of the through hole (2500) is not specifically limited to circular, polygonal, or elliptical, but is preferably elliptical. If it is elliptical, the effect of reducing flow separation can be improved.

[0071] As the angle of inclination (θ) of the through hole (2500) moves further away from the main flow direction relative to the relative speed, the low-speed region expands, so the effect of reducing flow separation and increasing surge margin can be achieved.

[0072] Although the blade may be designed differently depending on the situation, it is preferable that the angle of inclination (θ) of the through hole (2500) be formed to be an angle of 10 to 60 degrees with respect to the tangential direction of the pressure surface (2300).

[0073] When the angle of inclination (θ) is less than 10 degrees, the exit (2502) of the through hole (2500) is too close to the trailing edge (2200), so the surge margin effect is small, and when the angle of inclination (θ) is greater than 60 degrees, the flow separation reduction effect is small, so the compression efficiency increase effect is relatively small.

[0075] FIG. 7 is a drawing illustrating a flow separation phenomenon caused by an airfoil according to the prior art, and FIG. 8 is a drawing illustrating a flow separation phenomenon caused by an airfoil according to the first embodiment of the present invention.

[0076] Flow separation is a phenomenon in which a fluid flows along the surface of an airfoil but separates at a viscous boundary layer, breaking off without contacting the airfoil. Flow separation has a negative impact on the performance of the airfoil and the efficiency of the gas turbine.

[0077] The airfoil (200) according to the prior art shown in FIG. 7 is an airfoil in which a through hole (2500) is not formed.

[0078] In the case of an airfoil (200) according to the prior art, as shown in FIG. 7, the air flow (F) branches off at the leading edge (210) and flow separation occurs at the separation point (SP1) of the suction surface (240), forming a vortex and consequently causing pressure loss and increased drag.

[0079] According to the airfoil (2000_1) of the first embodiment of the present invention, as shown in FIG. 8, the separation point (SP2) is formed closer to the trailing edge (2200) side by the air flow (F2) supplied from the pressure surface (2300) to the suction surface (2400) through the through hole (2500), thereby reducing the flow separation phenomenon compared to the conventional airfoil (200). As a result, pressure loss can be reduced and drag increase reduced compared to the conventional method, thereby improving compression efficiency.

[0081] Figure 9 is a diagram showing the compression ratio relative to the flow rate, and Figure 10 is a diagram showing the polytropic efficiency relative to the flow rate.

[0082] In FIG. 9, line L11 shows the compression ratio relative to the flow rate by a compressor equipped with a compressor blade (1110) equipped with an airfoil (2000_1) according to the first embodiment of the present invention, and line L21 shows the compression ratio relative to the flow rate by a compressor equipped with a compressor blade equipped with an airfoil (200) according to the prior art.

[0083] In Figure 9, the horizontal axis represents the flow rate that flowed during a reference time in a compressor of the same size, and the vertical axis represents the compression ratio at that flow rate.

[0084] Referring to FIG. 9, based on a flow rate of approximately 580 m³ / s (units omitted hereafter), when a flow rate of 580 or higher is applied, L11 and L21 show substantially the same trajectory, so there is no significant difference in compression ratio; however, when a flow rate of less than 580 is applied, it can be confirmed that there is a large difference between the compression ratios of L11 and L21. This means that when using the airfoil (2000_1) of the present invention, a high compression ratio can be achieved at a low flow rate.

[0085] In addition, since L21 starts its trajectory at a flow rate of 552 or higher, and L11 starts its trajectory at a flow rate of 536, which is smaller than that, it can be confirmed that the airfoil (2000_1) of the present invention shows a higher surge margin than the conventional one.

[0086] Meanwhile, a surge is a phenomenon in which the airflow within a compressor reverses or becomes unstable, which can lead to mechanical damage or performance degradation; therefore, a surge margin is established to prevent this. In other words, the surge margin serves as an important design and operating criterion that indicates the range in which the compressor can operate stably.

[0087] In FIG. 10, line L12 shows the polytropic efficiency relative to the flow rate by a compressor equipped with a compressor blade (1110) equipped with an airfoil (2000_1) according to the first embodiment of the present invention, and line L22 shows the polytropic efficiency relative to the flow rate by a compressor equipped with a compressor blade equipped with an airfoil (200) according to the prior art.

[0088] In Figure 10, the horizontal axis represents the flow rate that flowed during a reference time in a compressor of the same size, and the vertical axis represents the polytropic efficiency at that flow rate.

[0089] Polytropic efficiency is defined as in Equation (1) below.

[0091] Equation (1) :

[0093] An "ideal polytropic process" is an ideal quasi-static process in which temperature changes are very small, and "compression work used in an actual compressor" includes friction, specific heat loss, mechanical loss, etc., of the actual gas flow.

[0095] Polytropic efficiency can be calculated based on the pressure ratio inside the compressor using the following equation (2).

[0096] Equation (2) :

[0097] Here, n is the polytropic index, P1 is the compressor inlet pressure, P2 is the compressor outlet pressure, T1 is the compressor inlet temperature, and T2 is the compressor outlet temperature.

[0098] The higher the polytropic efficiency calculated in this way, the lower the energy loss of the compressor and the superior its performance.

[0099] Referring to FIG. 10, when a flow rate of 568 or higher is applied, L21 and L22 show substantially the same trajectory, so there is no significant difference in polytropic efficiency. However, when a flow rate of less than 568 is applied, it can be seen that a large difference between L21 and L22 is observed. This means that when using the airfoil (2000_1) of the present invention, high polytropic efficiency can be achieved at low flow rates.

[0100] Accordingly, when the airfoil (2000_1) according to the first embodiment of the present invention is applied to the compressor blade (1110) or compressor vane (1130), the surge margin and polytropic efficiency of the compressor (1100) can be increased to improve compression efficiency.

[0102] FIG. 11 is a drawing showing an airfoil according to a second embodiment of the present invention, FIG. 12 is a drawing showing an airfoil according to a third embodiment of the present invention, and FIG. 13 is a drawing showing an airfoil according to a fourth embodiment of the present invention.

[0103] Specifically, FIGS. 11 to 13 are top views of the A-A' cross-section or BB' cross-section of FIG. 3. Since the remaining configurations are substantially identical except for the shape of the through-hole (2500) in the airfoils (2000_2 to 2000_4) of FIGS. 11 to 13, a repeated description is omitted.

[0104] Referring to FIG. 11, in the airfoil (2000_2) according to the second embodiment of the present invention, the through hole (2500_2) is formed by penetrating from the pressure surface (2300) toward the suction surface (2400), and is formed such that the cross-sectional area decreases as it goes from the pressure surface (2300) toward the suction surface (2400).

[0105] That is, the area of ​​the inlet (2501) of the through hole (2500_2) formed on the pressure surface (2300) is the largest, and the area of ​​the outlet (2502) of the through hole (2500_2) formed on the suction surface (2400) is the smallest.

[0106] As in the second embodiment, when a through hole (2500_2) is formed, more air can be introduced through the inlet (2501) of the widened through hole (2500_2), and the introduced air is accelerated as it passes through the through hole (2500_2) with a gradually decreasing diameter and is discharged through the outlet (2502) of the through hole (2500_2), thereby minimizing air head loss and improving compression efficiency.

[0107] Referring to FIG. 12, in the airfoil (2000_3) according to the third embodiment of the present invention, the through hole (2500_3) is formed by penetrating from the pressure surface (2300) toward the suction surface (2400), and may include a first path (2530) and a second path (2540).

[0108] The first path (2530) starts at the entrance (2501) of the through hole (2500_3) and ends inside the airfoil (2000_3) before reaching the exit (2502), and the second path (2540) is formed toward the exit (2502) of the through hole (2500_3) from the end of the ended first path (2530). The exit (2502) of the through hole (2500_3) is formed closer to the trailing edge (2200) side compared to the first embodiment described above.

[0109] The inclination angle of the first path (2530) is the same as the inclination angle (θ) of the through hole (2500) in the first embodiment described above, and the inclination angle (θ2) of the second path (2540) is formed to be smaller than the inclination angle (θ) of the first path (2530).

[0110] The first path (2530) may be formed such that the cross-sectional area decreases as it goes from the pressure surface (2300) toward the suction surface (2400), similar to the second embodiment described above.

[0111] When a through hole (2500_3) is formed as in the third embodiment, air introduced through the inlet (2501) is discharged through the outlet (2502) via a second path (2540) having a smaller angle of inclination (θ2). At this time, since the outlet (2502) is formed closer to the trailing edge (2200), the separation point is formed closer to the trailing edge (2200) than in the first embodiment described above, thereby further reducing the flow separation phenomenon.

[0112] Referring to FIG. 13, in the airfoil (2000_4) according to the fourth embodiment of the present invention, the through hole (2500_4) is formed by penetrating from the pressure surface (2300) toward the suction surface (2400), and may include a third path (2550) and a fourth path (2560).

[0113] The third path (2550) is formed to connect the entrance (2501) of the through hole (2500_4) and the first exit (2502a), and the fourth path (2560) is formed to connect the second exit (2502b) at one point of the third path (2550).

[0114] The angle of inclination of the third path (2550) is the same as the angle of inclination (θ) of the through hole (2500) in the first embodiment described above, and the angle of inclination (θ3) of the fourth path (2560) is formed to be smaller than the angle of inclination (θ) of the third path (2550). The angle of inclination (θ3) of the fourth path (2560) may be substantially the same as the angle of inclination (θ2) of the second path (2540).

[0115] The second exit (2502b) of the through hole (2500_4) is formed closer to the trailing edge (2200) side compared to the first exit (2502a).

[0116] The third path (2550) can be formed such that the cross-sectional area decreases as it goes from the pressure surface (2300) toward the suction surface (2400), similar to the second embodiment described above.

[0117] When a through hole (2500_4) is formed as in the fourth embodiment, air introduced through the inlet (2501) branches off and flows at the connection point between the third path (2550) and the fourth path (2560). A portion of the branched air passes through the fourth path (2560), which has a smaller angle of inclination (θ3), and is discharged through the outlet (2502). At this time, since the outlet (2502) is formed closer to the trailing edge (2200), the separation point is formed closer to the trailing edge (2200) than in the first embodiment described above, thereby further reducing the flow separation phenomenon.

[0118] In the case of the third embodiment, after processing the first path (2530) on the pressure surface (2300) side so as to terminate inside the airfoil (2000_3), the second path (2540) is processed on the suction surface (2400) side, but there is an inconvenience in that the termination position of the first path (2530) must be processed precisely.

[0119] In the case of the fourth embodiment, it is sufficient to process the third path (2550) to penetrate the pressure surface (2300) and the suction surface (2400), so there is an advantage that processing is easier compared to the third embodiment.

[0120] Although embodiments of the present invention have been described above, those skilled in the art may modify and change the present invention in various ways by adding, changing, deleting, or adding components, etc., without departing from the spirit of the present invention as described in the claims, and such modifications and changes are also to be included within the scope of the rights of the present invention. Explanation of the symbols

[0121] 1000 : Gas turbine 1100 : Compressor 1110: Compressor blade 1120: Compressor rotor disc 1130 : Compressor vane 1150: Compressor casing 1200 : Combustion device 1300 : Turbine 1310: Turbine blade 1320: Turbine rotor disk 1330 : Turbine vane 1340 : Turbine vane carrier 1350 : Turbine casing 2000_1 ~ 2000_4 : Airfoil 2100: Leading Edge 2200: Trailing Edge 2300 : Pressure surface 2400 : Intake surface 2500, 2500_2, 2500_3, 2500_4 : Through hole

Claims

Claim 1 An airfoil comprising: a suction surface protruding outwardly with a convex curved surface; a pressure surface formed on the opposite side of the suction surface and having a concave curved surface; a leading edge positioned upstream of the airflow and formed by contact between the leading edge of the pressure surface and the suction surface; a trailing edge positioned downstream of the airflow and formed by contact between the trailing edge of the pressure surface and the suction surface; and a through hole formed penetrating from the pressure surface toward the suction surface. Claim 2 An airfoil according to claim 1, wherein the through hole is a tip hole formed on the upper part of the pressure surface or a hub hole formed on the lower part of the pressure surface. Claim 3 In claim 1, the airfoil, wherein the through hole is formed in a straight or curved shape. Claim 4 In claim 1, the airfoil is formed such that the through hole is formed in a diagonal direction having a predetermined angle with respect to the tangential direction of the pressure surface. Claim 5 In claim 4, the airfoil, wherein the preset angle is in the range of 1 to 60 degrees. Claim 6 In claim 1, the airfoil is formed such that the cross-sectional area of ​​the through-hole decreases from the pressure surface toward the suction surface. Claim 7 An airfoil according to claim 1, wherein the through hole comprises a first path that starts at the entrance of the through hole and ends before reaching the exit, and a second path formed toward the exit of the through hole at the end of the first path. Claim 8 An airfoil according to claim 7, wherein the first path is formed in a diagonal direction having a first angle predetermined with respect to the pressure surface, and the second path is formed in a diagonal direction having a second angle smaller than the first angle. Claim 9 An airfoil according to claim 1, wherein the through hole comprises a third path connecting the entrance of the through hole and a first exit, and a fourth path formed to connect a second exit at one point of the third path. Claim 10 In claim 9, the airfoil, wherein the second outlet is formed closer to the trailing edge side than the first outlet. Claim 11 A gas turbine comprising: a compressor that compresses incoming air and includes radially installed multi-stage compressor blades and compressor vanes installed between the multi-stage compressor blades; a combustor that mixes and burns the air compressed from the compressor with fuel; and a turbine that generates power from the gas burned from the combustor; wherein at least one of the compressor blades or the compressor vanes includes an airfoil, and the airfoil comprises: a suction surface protruding outwardly as a convex curved surface; a pressure surface formed on the opposite side of the suction surface as a concave curved surface; a leading edge formed by contact between the pressure surface and the suction surface, positioned upstream of the air flow; a trailing edge formed by contact between the pressure surface and the suction surface, positioned downstream of the air flow; and a through hole formed penetrating from the pressure surface toward the suction surface. Claim 12 A gas turbine according to claim 11, wherein the through hole is a tip hole formed on the upper part of the pressure surface or a hub hole formed on the lower part of the pressure surface. Claim 13 In claim 11, the gas turbine, wherein the through-hole is formed in a straight or curved shape. Claim 14 In claim 11, the gas turbine, wherein the through hole is formed in an oblique direction having a predetermined angle with respect to the tangential direction of the pressure surface. Claim 15 A gas turbine according to claim 14, wherein the aforementioned preset angle is in the range of 10 to 60 degrees. Claim 16 In claim 11, the gas turbine is formed such that the cross-sectional area of ​​the through-hole decreases from the pressure surface toward the suction surface. Claim 17 A gas turbine according to claim 11, wherein the through hole comprises a first path that starts at the entrance of the through hole and ends before reaching the exit, and a second path formed toward the exit of the through hole at the end of the first path. Claim 18 A gas turbine according to claim 17, wherein the first path is formed in a diagonal direction having a first angle predetermined with respect to the pressure surface, and the second path is formed in a diagonal direction having a second angle smaller than the first angle. Claim 19 A gas turbine according to claim 11, wherein the through hole comprises a third path connecting the inlet of the through hole and a first outlet, and a fourth path formed to connect a second outlet at one point of the third path. Claim 20 A gas turbine according to claim 19, wherein the second outlet is formed closer to the trailing edge side than the first outlet.

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