Swept compressor blades and compressors and gas turbines including them

Abstract

The present invention provides a swept compressor blade with excellent aerodynamic performance and a gas turbine incorporating the same. [Solution] The compressor blade is constructed by laminating cross-sections of a predetermined thickness, which have a closed curve shape connecting a concave curve and a convex curve, and includes an airfoil that extends radially between the hub and the shroud, extends in the cord line direction between the leading edge and the trailing edge, and extends in the thickness direction between the pressure surface and the suction surface. The airfoil has a swept profile, which is formed by laminating cross-sections of a predetermined thickness at an inclination in the radial direction, and the leading edge and trailing edge are formed at an inclination in front of or behind the cord line. The swept profile is formed by a first region to which a rear sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

Description

Technical Field

[0001] The present invention relates to a swept compressor blade, a compressor and a gas turbine including the same, and more particularly, to a compressor blade having a swept profile forward or backward, a compressor 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. Further, a rotor is arranged so as to penetrate the centers of the compressor, the combustor, the turbine, and the exhaust chamber.

[0006] Both ends of the rotor are rotatably supported by bearings. A plurality of disks are fixed to the rotor, and respective blades are 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 greatly reduced, and it has the merit of being capable of high-speed movement.

[0008] In simple terms, a gas turbine works as follows: compressed air is mixed with fuel and burned to produce high-temperature combustion gases, which are then injected into the turbine. As these gases pass through the turbine vanes and blades, they generate rotational force, causing the rotor to rotate.

[0009] U.S. Patent Publication No. 9790797 discloses a swept fan blade in the subsonic region.

[0010] The disclosed conventional fan blades apply a rearward sweep over a span of 0-50%, a transitional region over a span of 50-60%, and a forward sweep over a span of 60-100%, with a rearward sweep angle of 5-10 degrees.

[0011] However, conventional technology defines the shape of the fan blades of the propulsor in a gas turbine engine, and not only can it not be directly applied to compressor blades, but even if applied to compressor blades, it cannot achieve superior aerodynamic performance. [Prior art documents] [Patent Documents]

[0012] [Patent Document 1] U.S. Patent Publication No. 9790797 (Registered October 17, 2017) [Overview of the Initiative] [Problems that the invention aims to solve]

[0013] The present invention aims to provide a swept compressor blade and a gas turbine including the same, which have excellent aerodynamic performance by defining the sweep angle according to the span region of the compressor blade. [Means for solving the problem]

[0014] A compressor blade for a gas turbine according to one embodiment of the present invention is constructed by laminating cross-sections of a predetermined thickness, each having a closed curve shape connecting a concave curve and a convex curve, and includes an airfoil that extends radially between the hub and the shroud, extends in the direction of the chord line between the leading edge and the trailing edge, and extends in the thickness direction between the pressure surface and the intake surface. The airfoil has a swept profile, which is formed by laminating cross-sections of a predetermined thickness with an inclination in the radial direction, and the leading edge and trailing edge are formed with an inclination in front of or behind the chord line.

[0015] The swept profile is formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

[0016] In a compressor blade according to one embodiment of the present invention, the first region is preferably located at a radial span position of 30 to 50% from the hub.

[0017] In a compressor blade according to one embodiment of the present invention, the second region is preferably from a radial span position of 70-50% to the shroud.

[0018] In a compressor blade according to one embodiment of the present invention, the rear sweep angle is preferably such that the leading edge and trailing edge are inclined 20 to 30 degrees behind the code line.

[0019] In a compressor blade according to one embodiment of the present invention, the forward sweep angle is preferably such that the leading edge and trailing edge are inclined 20 to 30 degrees forward of the code line.

[0020] In the compressor blade according to an embodiment of the present invention, the airfoil may be formed such that a cord line, which is a straight line connecting between a leading edge and a trailing edge, is disposed outside a pressure surface.

[0021] A compressor of a gas turbine according to an embodiment of the present invention includes a plurality of compressor rotor disks mounted on a rotating shaft and a plurality of compressor blades fastened to each of the plurality of compressor rotor disks. The compressor blade is configured by laminating cross-sections of a predetermined thickness having a closed curve shape connecting a concave curve and a convex curve, extends in a radial span between a hub and a shroud, extends in a cord line direction between a leading edge and a trailing edge, and includes an airfoil extending in a thickness direction between a pressure surface and a suction surface.

[0022] The airfoil has a swept profile in which cross-sections of a predetermined thickness are laminated while being inclined in the radial direction, and the leading edge and the trailing edge are inclined in front of or behind the cord line.

[0023] The swept profile is formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position facing the first region to the shroud, and a third region located between the first region and the second region to which no sweep is applied.

[0024] In the compressor according to an embodiment of the present invention, the first region is preferably from a hub to a radial span position of 30 to 50%.

[0025] In the compressor according to an embodiment of the present invention, the second region is preferably from a radial span position of 70 to 50% to the shroud.

[0026] In the compressor according to an embodiment of the present invention, it is preferable that the rear sweep angle is formed such that the leading edge and the trailing edge are inclined 20 to 30 degrees backward from the cord line.

[0027] In the compressor according to an embodiment of the present invention, it is preferable that the front sweep angle is formed such that the leading edge and the trailing edge are inclined 20 to 30 degrees forward from the cord line.

[0028] In the compressor according to an embodiment of the present invention, the airfoil may be formed such that the cord line, which is a straight line connecting between the leading edge and the trailing edge, is disposed outside the pressure surface.

[0029] A gas turbine according to an embodiment of the present invention includes a compressor that sucks and compresses external air by a plurality of rotating compressor blades, a combustor that mixes and burns fuel with the air compressed by the compressor, and a turbine in which turbine blades are mounted inside a turbine casing and the turbine blades are rotated by combustion gas discharged from the combustor.

[0030] The compressor blade is configured by laminating cross-sections of a predetermined thickness having a closed curve shape connecting a concave curve and a convex curve, extending in the radial span between the hub and the shroud, extending in the cord line direction between the leading edge and the trailing edge, and including an airfoil extending in the thickness direction between the pressure surface and the suction surface.

[0031] The airfoil has a swept profile in which cross-sections of a predetermined thickness are laminated while being inclined in the radial direction, and the leading edge and the trailing edge are inclined forward or backward from the cord line.

[0032] The swept profile is formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

[0033] In a gas turbine according to one embodiment of the present invention, the first region is preferably located at a radial span position of 30 to 50% from the hub.

[0034] In a gas turbine according to one embodiment of the present invention, the second region is preferably from the 70-50% radial span position to the shroud.

[0035] In a gas turbine according to one embodiment of the present invention, the rear sweep angle is preferably such that the leading edge and trailing edge are inclined 20 to 30 degrees behind the chord line.

[0036] In a gas turbine according to one embodiment of the present invention, the forward sweep angle is preferably such that the leading edge and trailing edge are inclined 20 to 30 degrees forward of the chord line.

[0037] In a gas turbine according to one embodiment of the present invention, the airfoil may be formed such that a chord line, which is a straight line connecting the leading edge and the trailing edge, is located on the outside of the pressure surface. [Effects of the Invention]

[0038] According to the swept compressor blade of the present invention described above, by defining the sweep angle according to the span area of ​​the compressor blade, it is possible to design a swept compressor blade shape that has excellent aerodynamic performance. [Brief explanation of the drawing]

[0039] [Figure 1]This is a partially cut perspective view of a gas turbine according to one embodiment of the present invention. [Figure 2] This is a cross-sectional view showing the schematic structure of a gas turbine according to one embodiment of the present invention. [Figure 3] This figure shows a cross-section of a typical airfoil of a compressor blade according to one embodiment of the present invention. [Figure 4] This diagram illustrates the definition of sweep in the airfoil of a compressor blade. [Figure 5] This graph shows the sweep angle for each span region in the airfoil of the compressor blade according to the present invention. [Figure 6] This figure shows the sweep angle for each span region in the airfoil of the compressor blade according to the present invention in three dimensions. [Figure 7] This table shows a comparison of conventional technology and experimental examples, in which the aerodynamic performance of a compressor blade airfoil was tested while varying the sweep angle for each span region. [Figure 8] Figure 7 is a graph showing the efficiency for each case. [Modes for carrying out the invention]

[0040] 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.

[0041] 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.

[0042] 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.

[0043] 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.

[0044] 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 compressor blades 1110 arranged radially. The compressor 1100 is connected to the turbine 1300 by a rotating shaft, causing the plurality of compressor blades 1110 to rotate. Air is compressed and moved by the rotation of the compressor blades 1110. The size and installation angle of the compressor 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 compressor blades 1110.

[0045] 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.

[0046] 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.

[0047] 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.

[0048] The compressor section 1100 is equipped with a plurality (for example, 14) of compressor rotor discs 1120, each of which is fastened by tie rods 1600 so as not to be separated axially.

[0049] Specifically, each compressor rotor disc 1120 is aligned axially with respect to each other, with a tie rod 1600, which constitutes the axis of rotation, 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.

[0050] Multiple compressor blades 1110 are radially connected to the outer circumferential surface of the compressor rotor disk 1120. Each compressor blade 1110 is fastened to the compressor rotor disk 1120 with a dovetail portion 1112.

[0051] Between each compressor rotor disc 1120 are vanes (not shown) fixed to the housing. Unlike the rotor discs, the vanes are fixed in place and do not rotate. Their role is to align the flow of compressed air that has passed through the blades of the compressor rotor discs and guide the air to the blades of the rotor disc located downstream.

[0052] The fastening method for the dovetail section 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, other fastening devices other than those described above, such as fasteners such as keys or bolts, can be used to fasten the blades to the rotor disk.

[0053] The tie rod 1600 is positioned to penetrate the centers of multiple 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.

[0054] The configuration of the tie rod 1600 can consist of various structures depending on the gas turbine, and is not necessarily limited to the configuration shown in Figure 2. That is, as shown in the figure, it may have a configuration in which a single tie rod penetrates the center of the rotor disc, or it may have a configuration in which multiple tie rods are arranged around the circumference, or a combination of these is also possible.

[0055] 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.

[0056] In the combustor 1200, the incoming compressed air is mixed with fuel and burned to create high-energy, high-temperature, high-pressure combustion gases. During the isobaric combustion process, the temperature of the combustion gases is raised to the heat resistance limit that the combustor and turbine components can withstand.

[0057] 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.

[0058] 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.

[0059] 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.

[0060] 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.

[0061] Cooling air, which has cooled the aforementioned transition piece, flows through the annular space of the liner, and compressed air from outside the flow sleeve can be supplied as cooling air through cooling holes provided in the flow sleeve and collide with the outer wall of the liner.

[0062] Meanwhile, the high-temperature, high-pressure combustion gases from the combustor are supplied to the turbine 1300 mentioned 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 mentioned above, and any power exceeding the power required to drive the compressor is used to drive a generator or the like.

[0063] The turbine 1300 is fundamentally similar in structure to that of a compressor. That is, the turbine 1300 is also equipped with multiple turbine rotor discs 1320, similar to the compressor rotor discs of a compressor. Therefore, the turbine rotor discs 1320 also include multiple turbine blades 1340 arranged radially. The turbine blades 1340 can also be coupled to the turbine rotor discs 1320 by means of dovetails or the like. 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 passing through the blades.

[0064] Figure 3 is a diagram showing a typical cross-section of an airfoil of a compressor blade according to one embodiment of the present invention, and Figure 4 is a diagram illustrating the definition of sweep in the airfoil of a compressor blade. Figure 5 is a graph showing the sweep angle for each span region in the airfoil of a compressor blade according to the present invention, and Figure 6 is a diagram showing the sweep angle for each span region in the airfoil of a compressor blade according to the present invention in three dimensions.

[0065] As described above with reference to Figures 1 and 2, the gas turbine 1000 includes a compressor 1100, a combustor 1200, and a turbine 1300. In the compressor 1100, multiple compressor blades 1100 can be coupled to a compressor rotor disk 1120. The compressor blades are indicated by the drawing reference numerals "1110" or "100".

[0066] The airfoil of the compressor blade 100 is constructed by laminating cross-sections of a predetermined thickness, each consisting of a closed curve shape connecting a concave curve and a convex curve. Figure 3 shows an example of the cross-section of the airfoil of the compressor blade 100.

[0067] As shown in Figure 3, the cross-section of the airfoil of the compressor blade 100 may have a closed curve shape connecting the leading edge 110 and trailing edge 120, the suction surface 130 and the pressure surface 140. The suction surface 130 is formed as a substantially convex curved surface, and the pressure surface 140 is formed as a substantially concave curved surface.

[0068] The airfoil can extend radially between the hub and the shroud, in the cord line direction between the leading edge 110 and the trailing edge 120, and in the thickness direction between the pressure surface 140 and the intake surface 130.

[0069] As shown in Figure 2, the compressor blade 100 extends from the hub, which is the radially inner end, to the shroud, which is the radially outer end. Therefore, the airfoil of the compressor blade 100 can form a radial span between the hub and the shroud, with a 0% span position at the hub and a 100% span position at the shroud.

[0070] As shown in Figure 3, the camber line 150 passing through the center of the airfoil's thickness is formed in a curve, and the chord line 160, which is a straight line connecting the leading edge 110 and the trailing edge 120, may be positioned at a predetermined angle with respect to the axis of the compressor 1100. One side point of the suction surface 130 can serve as the reference point (origin) of the stacked axial coordinate system.

[0071] The airfoil of the compressor blade 100 may be formed such that a code line 160, which is a straight line connecting the leading edge 110 and the trailing edge 120, is located outside the pressure surface 140.

[0072] The pressure surface 140 of the airfoil may be formed in a mostly concave curved shape, except for a portion near the leading edge 110 and the trailing edge 120. Therefore, the code line 160 connecting the leading edge 110 and the trailing edge 120 in a straight line may be positioned mostly outside the pressure surface 140.

[0073] The airfoil is formed by laminating layers of a predetermined thickness inclined radially, and has a swept profile in which the leading edge 110 and trailing edge 120 are formed inclined forward or backward of the code line.

[0074] Figure 4 shows a diagram illustrating the definition of sweep.

[0075] The right side of Figure 4 shows a cross-section of the baseline of the airfoil and a cross-section of the swept layer laid on top of it. The left side of Figure 4 shows the case of a rear sweep, where the airfoil cross-section is shifted backward at a predetermined angle along the code line direction and then laid, and the case of a forward sweep, where the airfoil cross-section is shifted forward at a predetermined angle along the code direction and then laid.

[0076] The airfoils of the compressor blades may be positioned with their chord lines inclined at a predetermined angle with respect to the axial direction of the gas turbine, and the cross-section of the airfoils may be positioned parallel to the tangential direction of the gas turbine.

[0077] The swept profile of the airfoil may be formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

[0078] The span refers to a length or height that extends radially outward from 0 to 1 (=100%), and may be formed by stacking unit cross-sections of airfoil so that they are swept backward or forward in each span region.

[0079] As shown in Figures 5 and 6, the swept profile of the airfoil may be formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

[0080] The swept profile of the airfoil may be formed by sweeping backward at a predetermined angle α in a first region from the 0% span position on the hub side to the 30-50% radial span position. For example, the swept profile of the airfoil may be swept backward from the hub to the 30% radial span position.

[0081] Furthermore, in a third region from the 30-50% radial span position to the 70-50% radial span position, stacking is possible without sweeping. For example, the swept profile of the airfoil may be formed with a sweep angle of 0 degrees from the 30% span position to the 70% span position.

[0082] Furthermore, the swept profile of the airfoil may be formed by sweeping forward at a predetermined angle α in a second region from the 70-50% radial span position to the 100% radial span position on the shroud side. For example, the swept profile of the airfoil may be swept forward from the 70% span position to the 100% span position.

[0083] Figure 7 is a table showing the conventional technology, comparative examples, and experimental examples in which the aerodynamic performance was tested while varying the sweep angle for each span region in the airfoil of the compressor blades, and Figure 8 is a graph showing the efficiency for each case in Figure 7.

[0084] As shown in Figure 7, aerodynamic tests were conducted on the airfoil of the compressor blades by applying a rearward sweep to the first region and a forward sweep to the second region in the radial span from 0 to 100%, while varying the sweep angle.

[0085] Case 0 represents the case where the sweep angle is 0 degrees across the entire span from 0 to 100%, and therefore no sweeping occurs.

[0086] Cases 1, 2, and 3 show the results when a backward sweep is applied to the 0-10% span region and a forward sweep is applied to the 90-100% span region. The sweep angles applied were 10 degrees, 20 degrees, and 30 degrees, respectively.

[0087] Cases 4, 5, and 6 show the results when a backward sweep is applied to the 0-30% span region and a forward sweep is applied to the 70-100% span region. The sweep angles applied were 10 degrees, 20 degrees, and 30 degrees, respectively.

[0088] Cases 7, 8, and 9 show the results when a backward sweep is applied to the 0-50% span region and a forward sweep is applied to the 50-100% span region. The sweep angles applied were 10 degrees, 20 degrees, and 30 degrees, respectively.

[0089] In Figure 8, the efficiency of the compressor blade airfoil refers to the ratio of the change in air pressure to the change in air temperature flowing inside the compressor. In other words, it is the ratio of the change in air pressure to the energy input to rotate the compressor blade.

[0090] First, in case 0, where no sweeping is applied at all, the airfoil efficiency showed an extremely low value of less than 0.9460. This can be interpreted as a decrease in the airfoil's compression efficiency due to the characteristics of the multistage axial flow compressor, particularly the sharp increase in losses due to wall flow caused by boundary layer growth at the trailing end.

[0091] As shown in Figure 8, the efficiency of the airfoil was high, exceeding 0.9473 in cases 5, 6, 8, and 9. It can be seen that the efficiency generally increases as the amount of the radial span region that is swept increases.

[0092] From here on, it is preferable that the airfoil is formed such that a rearward sweep is applied from the 0-30% span region to the 0-50% span region, and a forward sweep is applied from the 70-100% span region to the 50-100% span region.

[0093] Furthermore, it is preferable that the rear sweep angle is formed with the leading edge and trailing edge inclined 20 to 30 degrees behind the chord line, and the front sweep angle is formed with the leading edge and trailing edge inclined 20 to 30 degrees in front of the chord line.

[0094] This is because if the sweep angle is less than 20 degrees, the compression efficiency will be lower than the target value, and if the sweep angle exceeds 30 degrees, the lamination angle will be excessively large, making it difficult to form the airfoil.

[0095] According to the swept compressor blade of the present invention, by appropriately defining the rearward and forward sweep angles according to the span area of ​​the compressor blade, it is possible to design a swept compressor blade shape that has excellent compression efficiency.

[0096] Although one embodiment of the present invention has been described above, any person with ordinary skill in the art can modify and change the present invention in various ways, such as 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]

[0097] 1000: Gas turbine, 1010: Housing 1100: Compressor, 1110: Compressor 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: Compressor blade 110: Leading edge, 120: Trailing edge 130: Intake surface, 140: Pressure surface 150: Camber line, 160: Cord line

Claims

1. It is constructed by laminating cross-sections of a predetermined thickness, each consisting of a closed curve shape connecting a concave curve and a convex curve, and includes an airfoil that extends radially between the hub and the shroud, extends in the cord line direction between the leading edge and the trailing edge, and extends in the thickness direction between the pressure surface and the intake surface. The airfoil is formed by laminating sections of a predetermined thickness inclined radially, and has a swept profile in which the leading edge and trailing edge are inclined forward or backward of the code line. The swept profile is formed by a compressor blade comprising a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

2. The compressor blade according to claim 1, wherein the first region extends to a radial span position of 30 to 50% from the hub.

3. The compressor blade according to claim 2, wherein the second region is from a radial span position of 70-50% to the shroud.

4. The compressor blade according to claim 1, wherein the angle of the rear sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees behind the cord line.

5. The compressor blade according to claim 1, wherein the angle of the forward sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees forward of the cord line.

6. The compressor blade according to any one of claims 1 to 5, wherein the airfoil is formed such that a code line, which is a straight line connecting the leading edge and the trailing edge, is located on the outside of the pressure surface.

7. Multiple compressor rotor discs mounted on the rotating shaft, It includes a plurality of compressor blades fastened to each of the plurality of compressor rotor disks, The compressor blade is constructed by stacking sections of a predetermined thickness, each having a closed curve shape connecting a concave curve and a convex curve, and includes an airfoil extending radially between the hub and the shroud, extending in the cord line direction between the leading edge and the trailing edge, and extending in the thickness direction between the pressure surface and the suction surface. The airfoil is formed by laminating sections of a predetermined thickness inclined radially, and has a swept profile in which the leading edge and trailing edge are inclined forward or backward of the code line. A compressor in which the swept profile is formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

8. The compressor according to claim 7, wherein the first region extends to a position 30-50% radially spanned from the hub.

9. The compressor according to claim 8, wherein the second region is from a radial span position of 70-50% to the shroud.

10. The compressor according to claim 7, wherein the angle of the rear sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees behind the code line.

11. The compressor according to claim 7, wherein the angle of the forward sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees forward of the code line.

12. The compressor according to any one of claims 7 to 11, wherein the airfoil is formed such that a code line, which is a straight line connecting the leading edge and the trailing edge, is located outside the pressure surface.

13. A compressor that draws in and compresses external air using multiple rotating compressor blades, A combustor that mixes fuel with the air compressed by the aforementioned compressor and burns it, The turbine includes a turbine in which turbine blades are mounted inside the turbine casing, and the turbine blades are rotated by combustion gases discharged from the combustor. The compressor blade is constructed by stacking sections of a predetermined thickness, each having a closed curve shape connecting a concave curve and a convex curve, and includes an airfoil extending radially between the hub and the shroud, extending in the cord line direction between the leading edge and the trailing edge, and extending in the thickness direction between the pressure surface and the suction surface. The airfoil is formed by laminating sections of a predetermined thickness inclined radially, and has a swept profile in which the leading edge and trailing edge are inclined forward or backward of the code line. A gas turbine in which the swept profile is formed by a first region to which a rearward sweep is applied from the hub, a second region to which a forward sweep is applied from a position opposite the first region to the shroud, and a third region located between the first and second regions to which no sweep is applied.

14. The gas turbine according to claim 13, wherein the first region extends to a radial span position of 30 to 50% from the hub.

15. The gas turbine according to claim 14, wherein the second region is from a radial span position of 70-50% to the shroud.

16. The gas turbine according to claim 13, wherein the angle of the rear sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees behind the chord line.

17. The gas turbine according to claim 13, wherein the angle of the forward sweep is such that the leading edge and trailing edge are inclined 20 to 30 degrees forward of the chord line.

18. The gas turbine according to any one of claims 13 to 17, wherein the airfoil is formed such that a code line, which is a straight line connecting the leading edge and the trailing edge, is located outside the pressure surface.

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