Turbine blade as well as turbine and gas turbine equipped with it

Abstract

Turbine blade, which includes the following: a flow profile (2510, 3510, 4510) with a leading edge (2511, 3511, 4511) and a trailing edge (2512, 3512, 4512); an inner lining (2520, 3520, 4520) arranged at one end of the flow profile (2510, 3510, 4510) to support the flow profile (2510, 3510, 4510); an outer cladding (2530, 3530, 4530) which is arranged at the other end of the flow profile (2510, 3510, 4510) to support the flow profile (2510, 3510, 4510) and is opposite the inner cladding (2520, 3520, 4520); a first cooling passage (C21, C31, C41) and a second cooling passage (C22, C32, C42) extending in one vertical direction; and a first barrier (2561, 3561, 4561) that divides the first cooling passage (C21, C31, C41) and the second cooling passage (C22, C32, C42) and extends in a vertical direction of the flow profile (2510, 3510, 4510), wherein a longitudinal end of the first barrier (2561, 3561, 4561) is located further outward than an end of the flow profile (2510, 3510, 4510); and a first curved passage section (B21, B31, B41) connecting the first cooling passage (C21, C31, C41) and the second cooling passage (C22, C32, C42), wherein the first curved passage section (B21, B31, B41) is arranged within the inner lining (2520, 3520, 4520) or the outer lining (2530, 3530, 4530), and where: the first barrier (2561) has a plurality of inlet holes (2571) that penetrate the first barrier (2561) in its thickness direction, and some inlet holes (2571) are arranged further outwards than the end of the flow profile (2510) with respect to a center of the turbine blade (2500); or several porous plates (4570) are designed to project from the first barrier (4561), the porous plates (4570) being arranged on the first curved passage section (B41); or the first barrier (3561) comprises a first passage (3571) extending in a thickness direction of the first barrier (3561), and a second passage (3572) connected to the first passage (3571) and extending to an end of the first barrier (3561), the first passage (3571) being located inside the flow profile (3510), and the second passage (3572) extending from an interior of the flow profile (3510) to an interior of the inner lining (3520).

Description

BACKGROUND area

[0001] Devices and methods consistent with exemplary embodiments relate to a turbine blade and to a turbine and a gas turbine equipped therewith. Description of the state of the art

[0002] A gas turbine is a power plant that mixes fuel with compressed air in a compressor, combusts the mixture of fuel and compressed air, and rotates a turbine using the high-temperature gas produced by the combustion. The gas turbine is used to power a generator, aircraft, ship, train, or similar engines.

[0003] The gas turbine comprises a compressor, a combustor, and a turbine. The compressor draws in and compresses outside air and transfers the compressed air to the combustion chamber. The combustion chamber mixes fuel with the compressed air supplied by the compressor and burns the mixture of fuel and compressed air to produce a high-pressure, high-temperature combustion gas. The combustion gas generated by the combustion is discharged to the turbine. A turbine blade provided in the turbine is rotated by the combustion gas, generating power. The generated power can be used in various fields, such as power generation and driving a mechanical device.

[0004] In order to increase turbine efficiency, recently the temperature of the gas flowing into the turbine (i.e., turbine inlet temperature (TIT)) has a trend of continuous increase and consequently the importance of heat resistance treatment and cooling of the turbine blade has been emphasized.

[0005] In particular, among the internal passages of the blades, the structural durability of the curved passage section, where one of the passage directions is changed, is rated as low due to the concentrated heat load. If the structural durability of the curved passage section is rated as low, the overall service life of the turbine blade is reduced, which increases maintenance costs.

[0006] DE 10 2015 203 175 A1 shows a guide or rotor blade device and casting core.

[0007] US 8 628 294 B1 shows a guide vane for an industrial turbine.

[0008] US 2009 0081 029 A1 shows a gas turbine component with reduced cooling air requirement.

[0009] US 8 221 055 B1 shows a turbine stator blade with end wall cooling. SUMMARY

[0010] Aspects of one or more exemplary embodiments provide a turbine blade, a turbine, and a gas turbine capable of reducing thermal stress.

[0011] Additional aspects will be set forth in part in the description which follows and in part will be obvious from the description, or may be learned by practice of the example embodiments.

[0012] According to one aspect of an exemplary embodiment, a turbine blade is provided, comprising: an airfoil having a leading edge and a trailing edge, an inner fairing disposed at one end of the airfoil to support the airfoil, an outer fairing disposed at the other end of the airfoil to support the airfoil and configured to face the inner fairing, a first cooling passage and a second cooling passage configured to extend in a height direction thereof, and a first curved passage portion configured to connect the first cooling passage and the second cooling passage, wherein the first curved passage portion is disposed within one of the inner fairing and the outer fairing.

[0013] The first curved passage portion may comprise a first curved surface curved in an arc shape about a first point located within the inner panel or the outer panel.

[0014] The turbine blade further includes a first barrier configured to divide the first cooling passage and the second cooling passage and extending in a height direction of the airfoil, and a longitudinal end of the first barrier is disposed further outward than an end of the airfoil.

[0015] The inner fairing may include an inner platform connected to an inner end of the airfoil and an inner hook projecting from the inner platform, and the inner platform may include an inner projection projecting to form a space therein, and the first curved passage portion may be disposed within the inner projection.

[0016] According to a first aspect of the present invention, the first barrier includes a plurality of introduction holes penetrating the first barrier in its thickness direction, and some introduction holes are arranged farther outward from a center of the turbine blade than the end of the airfoil.

[0017] Some injection holes may be located at an interface where the airfoil and the inner platform are connected.

[0018] According to a second aspect of the present invention, the first barrier includes a first passage extending in a thickness direction of the first barrier and a second passage connected to the first passage and extending to an end of the first barrier, and the first passage is disposed within the airfoil, and the second passage extends from an interior of the airfoil to an interior of the inner panel.

[0019] According to a third aspect of the present invention, a plurality of porous plates are formed to protrude from the first barrier, the porous plates being arranged at the first curved passage portion.

[0020] The turbine blade may further include a third cooling passage configured to extend in a height direction thereof and a second curved passage portion configured to connect the second cooling passage and the third cooling passage, and the second curved passage portion may include a second curved surface curved in an arc shape around a second point located inside the outer shroud.

[0021] The outer fairing may include an outer platform connected to an outer end of the airfoil and an outer hook projecting from the outer platform, and the outer platform may include an outer projection projecting outwardly and forming a space therein, and the second curved passage portion may be disposed within the outer projection.

[0022] The outer projection may include a transverse cross-section shaped like the airfoil.

[0023] The turbine blade may further include a third cooling passage configured to extend in a height direction thereof, and a second barrier configured to partition the second cooling passage and the third cooling passage and extending in a height direction of the airfoil, and a longitudinal end of the second barrier may be disposed within the outer shroud.

[0024] According to one aspect of another exemplary embodiment, a turbine is provided, comprising: a rotor disk configured to be rotatable; and a plurality of turbine blades and turbine buckets mounted on the rotor disk. The turbine bucket is configured according to one of the above three aspects.

[0025] According to one aspect of another exemplary embodiment, a gas turbine is provided, including: a compressor configured to compress air drawn therein from an outside; a combustor configured to mix fuel with air compressed by the compressor and combust a mixture of the fuel and the compressed air; and a turbine having a plurality of turbine blades configured to be rotated by combustion gas discharged from the combustor. The turbine may include a rotor disk configured to be rotatable, and a plurality of turbine blades and turbine blades installed on the rotor disk. The turbine blade is configured according to any one of the above three aspects. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] The above and other aspects will become more apparent from the following description of exemplary embodiments with reference to the accompanying drawings, in which: Fig. 1 is a diagram illustrating an internal structure of a gas turbine according to an exemplary embodiment; Fig. 2 a longitudinal cross-sectional diagram showing part of the gas turbine of Fig. 1 represents; Fig. 3 is a perspective diagram illustrating a turbine blade according to an exemplary embodiment; Fig. 4 is a longitudinal cross-sectional diagram illustrating the turbine blade according to an exemplary embodiment; Fig. 5 is a transverse cross-sectional diagram illustrating the turbine blade according to an example not covered by the present invention; Fig. 6 is a longitudinal cross-sectional diagram illustrating a turbine blade according to an example not covered by the present invention; Fig. 7 is a transverse cross-sectional diagram illustrating the turbine blade according to an exemplary embodiment; Fig. 8 is a longitudinal cross-sectional diagram illustrating a turbine blade according to another exemplary embodiment; and Fig. 9 is a longitudinal cross-sectional diagram illustrating a turbine blade according to another exemplary embodiment. DETAILED DESCRIPTION

[0027] Various modifications and various embodiments will be described in detail with reference to the drawings so that those skilled in the art can easily practice the disclosure. However, it should be understood that the various embodiments are not intended to limit the scope of the disclosure to the specific embodiment, but should be interpreted to include all modifications, equivalents, and alternatives of the embodiments included within the spirit and technical scope disclosed herein.

[0028] The terminology used herein is for the purpose of describing specific embodiments only and is not intended to limit the scope of the disclosure. The singular terms "a," "an," and "the" may also include the plural terms unless the context clearly indicates otherwise. Throughout the disclosure, terms such as "comprising," "including," and "having / comprising" should be construed to mean the presence of such features, integers, steps, operations, components, parts, and / or combinations thereof, and not to preclude the presence or possibility of adding one or more other features, integers, steps, operations, components, parts, and / or combinations thereof.

[0029] Furthermore, terms such as "first," "second," and so on may be used to describe a variety of elements, but the elements should not be limited by these terms. The terms are simply used to distinguish one element from other elements. The use of such atomic numbers should not be construed as limiting the meaning of the term. The components assigned to such an atomic number should not be limited, for example, in order of use, order of arrangement, or the like. If necessary, any atomic number may be used interchangeably.

[0030] Hereinafter, exemplary embodiments will be described in detail with reference to the accompanying drawings. Reference should now be made to the drawings, in which the same reference numerals are used throughout the various drawings to designate the same or similar components. Details of well-known configurations and functions may be omitted to avoid unnecessarily obscuring the gist of the present disclosure. For the same reason, some components are exaggerated, omitted, or schematically illustrated in the accompanying drawings.

[0031] Fig. 1 is a diagram illustrating an internal structure of a gas turbine according to an exemplary embodiment, and Fig. Figure 2 is a longitudinal cross-sectional diagram showing part of the gas turbine of Fig. 1 represents.

[0032] For example, a thermodynamic cycle of a gas turbine 1000 according to the exemplary embodiment may ideally correspond to a Brayton cycle. The Brighton cycle may consist of four processes including isentropic compression (i.e., adiabatic compression), rapid heating at constant pressure, isentropic expansion (i.e., adiabatic expansion), and heat dissipation at constant pressure. In other words, the gas turbine may draw in atmospheric air, compress the air to a high pressure, burn fuel in a constant-pressure environment to release heat energy, expand the high-temperature combustion gas to convert the heat energy of the combustion gas into kinetic energy, and discharge exhaust gas containing residual energy to the atmosphere. That is, the Brayton cycle may be performed in four processes, including compression, heating, expansion, and heat dissipation.

[0033] With reference to Fig. 1 and Fig. 2, the gas turbine 1000 embodying the Brayton cycle may include a compressor 1100, a combustor 1200, and a turbine 1300.

[0034] Compressor 1100 can draw in air from the outside and compress it. Compressor 1100 can supply the compressed air, which is compressed by a compressor blade 1130, to the combustion chamber 1200 and also supply the compressed air to a high-temperature region that needs to be cooled in gas turbine 1000. Since the drawn-in air undergoes an adiabatic compression process in compressor 1100, the pressure and temperature of the air flowing through compressor 1100 are increased.

[0035] The compressor 1100 is designed in the form of a centrifugal compressor or an axial compressor. The centrifugal compressor is used in a small gas turbine, whereas a multi-stage axial compressor 1100 is used in a large gas turbine such as the gas turbine 1000 used in Fig. 1, is used to compress a large amount of air. In the multi-stage axial compressor 1100, a compressor blade 1130 moves the compressed air to a compressor vane 1140 located in a subsequent stage, while the introduced air is compressed by rotation along with the rotation of a central tie rod 1120 and a rotor disk. The air is gradually compressed to a high pressure as it passes through the compressor blade 1130, which is formed in a multi-stage structure.

[0036] The compressor blade 1140 is mounted within a housing 1150 such that a plurality of compressor blades 1140 form each stage. The compressor blade 1140 guides the compressed air moved by the compressor blade 1130 located in a preceding stage toward the compressor blade 1130 located in the following stage. In an exemplary embodiment, at least some of the plurality of compressor blades 1140 may be mounted to be rotatable to adjust the amount of air introduced within a predetermined range.

[0037] The compressor 1100 may be driven using some of the power output from the turbine 1300. To this end, a rotating shaft of the compressor 1100 and a rotating shaft of the turbine 1300 may be directly connected through a torque tube 1170. In the case of the large gas turbine 1000, almost half of the output generated by the turbine 1300 may be consumed to drive the compressor 1100.

[0038] Combustion chamber 1200 can generate high-energy combustion gas by mixing and burning the compressed air supplied from compressor 1100 with the fuel at a constant pressure. Combustion chamber 1200 generates high-temperature, high-pressure combustion gas with high energy by mixing and burning the introduced compressed air with the fuel, and raises the temperature of the combustion gas to a heat-resistant limit temperature that the combustion chamber and turbine can withstand through the constant-pressure combustion process.

[0039] A plurality of combustion chambers forming the combustion chamber 1200 may be arranged within the housing in a cell-like configuration. Each of the combustion chambers includes a combustor including a fuel injector, a combustor liner forming a combustion chamber, and a transition piece that becomes a connecting part between the combustor and the turbine.

[0040] The high-temperature, high-pressure combustion gas discharged from the combustion chamber 1200 is supplied to the turbine 1300. The supplied high-temperature, high-pressure combustion gas expands and exerts an impact or reaction force on a turbine blade 1400 of the turbine 1300 to generate rotational torque. Part of the rotational torque is supplied to the compressor 1100 through the torque pipe 1170, and the remaining part, which is the excess torque, is used to drive a generator or the like.

[0041] The turbine 1300 includes a rotor disk 1310 and a plurality of turbine blades 1400 and turbine vanes 1500 arranged radially on the rotor disk 1310.

[0042] The rotor disk 1310 has a substantially disk shape, and a plurality of grooves are formed in an outer peripheral portion thereof. The groove is formed to have a curved surface, and the turbine blade 1400 and the turbine vane 1500 are inserted into the groove. The turbine blade 1400 can be coupled to the rotor disk 1310 in a dovetail manner. The turbine vane 1500, which is fixed to the casing, is provided between the turbine blades 1400 to guide a flow direction of the combustion gas passing through the turbine blade 1400.

[0043] Fig. 3 is a perspective diagram illustrating a turbine blade according to an exemplary embodiment, Fig. 4 is a longitudinal cross-sectional diagram illustrating the turbine blade according to an example not covered by the present invention, and Fig. 5 is a transverse cross-sectional diagram illustrating a blade according to the example.

[0044] With reference to Fig. 3 to 5, the turbine blade 1500 includes an inner fairing 1520, an outer fairing 1530, and an airfoil 1510 disposed between the inner fairing 1520 and the outer fairing 1530.

[0045] The airfoil 1510 may be a curved plate with a vane shape and have an optimized airfoil according to the specification of the gas turbine 1000. The airfoil 1510 may include a leading edge 1511 located on an upstream side and a trailing edge 1512 located on a downstream side with respect to a flow direction of the combustion gas.

[0046] A front surface of the airfoil 1510, onto which the combustion gas is introduced, is formed with a suction surface that protrudes outward to have an outwardly convex curved surface, and a rear surface of the airfoil 1510 is formed with a pressure surface with a curved surface that is concavely recessed toward the suction surface. A pressure difference occurs between the suction surface and the pressure surface of the airfoil 1510 to rotate the turbine 1300.

[0047] A plurality of cooling holes 1513 are formed in a surface of the airfoil 1510. The cooling holes 1513 communicate with a cooling passage formed within the airfoil 1510 to supply the cooling air to the surface of the airfoil 1510.

[0048] The inner fairing 1520 is coupled to the rotor disk 1310 and is disposed at an inner end of the airfoil 1510 to support the airfoil 1510. The inner fairing 1520 includes an inner platform 1522 coupled to the interior of the airfoil 1510 and an inner hook 1524 projecting downwardly from the inner platform 1522 and coupled to the rotor disk 1310.

[0049] The inner platform 1522 is formed in a substantially rectangular plate shape and is formed with an inner protrusion 1528 protruding to form a space therein. The inner protrusion 1528 protrudes inward, which is a direction toward the rotor disk 1310, and has a transverse cross section shaped like the airfoil. That is, the transverse cross section of the inner protrusion 1528 includes a convex surface and a concave surface, and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof.

[0050] The outer fairing 1530 is coupled to a blade carrier (not shown) installed outward in a radial direction thereof and is disposed at an outer end of the airfoil 1510 to support the airfoil 1510. The outer fairing 1530 includes an outer platform 1532 coupled to the outer end of the airfoil 1510 and an outer hook 1534 projecting upward from the outer platform 1532 and coupled to the blade carrier.

[0051] The outer platform 1532 is formed in a substantially rectangular plate shape, and the outer platform 1532 is formed with an outer protrusion 1538 projecting to form a space therein. The outer protrusion 1538 projects outward and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the outer protrusion 1538 includes a convex surface and a concave surface, and is formed such that an interval between the convex surface and the concave surface decreases toward the side end thereof. The outer protrusion 1538 may be formed with an inlet (E11) through which the cooling medium is introduced and an outlet (O11) through which the cooling medium is discharged. The cooling medium may include, for example, the air compressed by the compressor. It goes without saying that the cooling medium is not limited to this.

[0052] A first cooling passage (C11), a second cooling passage (C12), a third cooling passage (C13), a fourth cooling passage (C14), a first curved passage portion (B11), a second curved passage portion (B12), a first barrier 1561, a second barrier 1562, and a third barrier 1563 are formed within the turbine blade 1500. The turbine blade 1500 may be formed by casting.

[0053] The first cooling passage (C11) is connected to the outlet (O11) and is configured to extend from the outer fairing 1530 to the interior of the inner fairing 1520 through the airfoil 1510. The first cooling passage (C11) is formed by the first barrier 1561 and the leading edge 1511 and is configured to penetrate the outer protrusion 1538, the outer platform 1532, and the airfoil 1510 in a height direction thereof.

[0054] The second cooling passage (C12) is connected to the inlet (E11) and is configured to extend from the outer panel 1530 to the interior of the inner panel 1520 through the airfoil 1510. The second cooling passage (C12) is formed by the first barrier 1561 and the second barrier 1562. The air introduced through the second cooling passage (C12) can be supplied to the first cooling passage (C11).

[0055] The first barrier 1561 is disposed between the first cooling passage (C11) and the second cooling passage (C12) and divides the first cooling passage (C11) and the second cooling passage (C12). The longitudinal outer end of the first barrier 1561 is attached to the outer shroud 1530, and the longitudinal inner end of the first barrier 1561 is positioned spaced from the end of the airfoil 1510. That is, the first barrier 1561 is configured to penetrate the airfoil 1510 and extend to the interior of the inner shroud 1520.

[0056] The first cooling passage (C11) and the second cooling passage (C12) are connected by the first curved passage section (B11), which is spaced from the end of the airfoil 1510 and disposed within the inner shroud 1520. The first curved passage section (B11) may, for example, be disposed within the inner protrusion 1528. Furthermore, the first curved passage section (B11) includes a first curved surface 1516 curved in an arc shape and having an arcuate cross-section with a first radius (R11) around a first point (P11). Here, the first point (P11) is spaced from the end of the airfoil 1510 and disposed within the inner shroud 1520. That is, the first point (P11) may also be disposed within the first barrier 1561. Consequently, the first barrier 1561 can serve as a heat dissipation plate that dissipates heat.

[0057] Although the turbine blade 1500 is completely cooled by the cooling passage, it is confirmed that a heat stress is concentrated in the first bent passage portion (B11), which is a portion where the cooling passage is switched, thereby reducing the durability of the structure.

[0058] The hot gas passes through only one section where the airfoil 1510 is formed, and the inner liner 1520 and the outer liner 1530 are inserted into other elements, thus preventing them from contacting the hot gas. Consequently, by disposing the first curved passage section (B11) within the inner liner 1520, the heat load applied to the first curved passage section (B11) can be minimized.

[0059] Furthermore, the heat load is maximized at the center of the curved surface curved in the first curved passage section (B11), so that when the first point (P11) is located within the inner liner 1520, the durability of the structure can be significantly improved. Even if part of the first curved passage section (B11) is located within the inner liner 1520, the degree of heat load reduction appears to be insignificant when the first point (P11) is located within the airfoil 1510.

[0060] It should be understood that although the first curved passage portion (B11) is formed inside by way of example, it is not limited thereto. For example, if the first curved passage portion is formed outside, the first curved passage portion and the first point may also be arranged inside the outer panel.

[0061] In the prior art structure, it was impossible to form the curved passage portion within the inner panel because the inner platform is thin. However, by forming the inner protrusion 1528, the first curved passage portion (B11) can be arranged within the inner protrusion 1528, thereby minimizing heat stress and improving the durability of the structure.

[0062] The third cooling passage (C13), located adjacent to the second cooling passage (C12), is formed to extend from the outer panel 1530 to the interior of the inner panel 1520 through the airfoil 1510. The third cooling passage (C13) is formed by the second barrier 1562 and the third barrier 1563 and is formed to penetrate the outer protrusion 1538, the outer platform 1532, and the airfoil 1510 in a height direction thereof.

[0063] The third cooling passage (C13) is connected to the inlet (E11), and the air introduced through the inlet (E11) moves by being divided into the second cooling passage (C12) and the third cooling passage (C13), respectively. The air moving through the third cooling passage (C13) can be supplied to the fourth cooling passage (C14).

[0064] The second barrier 1562 is disposed between the second cooling passage (C12) and the third cooling passage (C13), dividing the second cooling passage (C12) and the third cooling passage (C13). The longitudinal inner end of the second barrier 1562 is fixed to the inner liner 1520, and the longitudinal outer end thereof is disposed farther outward than the end of the airfoil 1510. That is, the second barrier 1562 penetrates the airfoil 1510 and extends to the interior of the outer liner 1530, but is spaced from the inlet (E11). Consequently, it is possible to prevent the heat load from being concentrated at the end of the second barrier 1562.

[0065] The fourth cooling passage (C14), located between the third cooling passage (C13) and the trailing edge 1512, is configured to extend toward the airfoil 1510 and the interior of the inner fairing 1520. The fourth cooling passage (C14) is formed by the third barrier 1563 and the trailing edge 1512. The fourth cooling passage (C14) receives air from the third cooling passage (C13), and the air introduced into the fourth cooling passage (C14) is exhausted through the trailing edge 1512. A rear-end cooling slot 1519 is formed on the trailing edge 1512, a split protrusion 1568 is formed on the rear-end cooling slot 1519, and the air cools the trailing edge 1512 while being exhausted through the rear-end cooling slot 1519.

[0066] The third barrier 1563 is disposed between the third cooling passage (C13) and the fourth cooling passage (C14) and divides the third cooling passage (C13) and the fourth cooling passage (C14). The longitudinal outer end of the third barrier 1563 is attached to the outer shroud 1530, and the longitudinal inner end of the third barrier 1563 is spaced from the end of the airfoil 1510 and positioned farther outward than the center of the turbine blade 1500. That is, the third barrier 1563 is configured to penetrate the airfoil 1510 and extend to the interior of the inner shroud 1520.

[0067] The third cooling passage (C13) and the fourth cooling passage (C14) are connected by the second curved passage section (B12), which is spaced from the end of the airfoil 1510 and disposed within the inner shroud 1520. The second curved passage section (B12) may, for example, be disposed within the inner platform 1522 and the inner boss 1528. Furthermore, the second curved passage section (B12) includes a second curved surface 1517 curved in an arcuate shape and having an arcuate cross-section with a second radius (R12) about a second point (P12). Here, the second point (P12) is spaced from the end of the airfoil 1510 and disposed within the inner shroud 1520. Consequently, the heat load generated in the second curved passage portion (B12) can be minimized and the durability of the structure in the second curved passage portion (B12) can be improved.

[0068] Fig. 6 is a longitudinal cross-sectional diagram illustrating a blade according to an exemplary embodiment, and Fig. 7 is a transverse cross-sectional diagram illustrating the blade according to the exemplary embodiment.

[0069] With reference to Fig. 6 and Fig. 7, a turbine blade 2500 includes an inner fairing 2520, an outer fairing 2530, and an airfoil 2510 disposed between the inner fairing 2520 and the outer fairing 2530.

[0070] The airfoil 2510 may be a curved plate with a vane shape and have an optimized airfoil according to the specification of a gas turbine. The airfoil 2510 may include a leading edge 2511 located on an upstream side and a trailing edge 2512 located on a downstream side with respect to a flow direction of the combustion gas.

[0071] The inner fairing 2520 is coupled to the rotor disk and disposed at the inner end of the airfoil 2510 to support the airfoil 2510. The inner fairing 2520 includes an inner platform 2522 coupled to the interior of the airfoil 2510 and an inner hook 2524 projecting downwardly from the inner platform 2522 and coupled to the rotor disk.

[0072] The inner platform 2522 is formed in a substantially rectangular plate shape and is formed with an inner protrusion 2528 protruding to form a space therein. The inner protrusion 2528 protrudes inward, which is a direction toward the rotor disk, and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the inner protrusion 2528 includes a convex surface and a concave surface, and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof.

[0073] The outer fairing 2530 is coupled to a blade carrier (not shown) installed outwardly in a radial direction thereof and disposed at the outer end of the airfoil 2510 to support the airfoil 2510. The outer fairing 2530 includes an outer platform 2532 coupled to the outer end of the airfoil 2510 and an outer hook 2534 projecting upward from the outer platform 2532 and coupled to the blade carrier.

[0074] The outer platform 2532 is formed in a substantially rectangular plate shape and is formed with an outer protrusion 2538 protruding to form a space therein. The outer protrusion 2538 protrudes outward and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the outer protrusion 2538 includes a convex surface and a concave surface and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof. The outer protrusion 2538 may be formed with an inlet (E21) through which the cooling medium is introduced and an outlet (O21) through which the cooling medium is discharged. The cooling medium may include, for example, the air compressed by the compressor. It goes without saying that the cooling medium is not limited to this.

[0075] A first cooling passage (C21), a second cooling passage (C22), a third cooling passage (C23), a fourth cooling passage (C24), a first curved passage portion (B21), a second curved passage portion (B22), a first barrier 2561, a second barrier 2562, and a third barrier 2563 are formed within the turbine blade 2500. The turbine blade 2500 may be formed by casting.

[0076] The first cooling passage (C21) is connected to the outlet (O21) and is configured to extend from the outer fairing 2530 to the interior of the inner fairing 2520 through the airfoil 2510. The first cooling passage (C21) is formed by the first barrier 2561 and the leading edge 2511 and is configured to penetrate the outer protrusion 2538, the outer platform 2532, and the airfoil 2510 in a height direction thereof.

[0077] The second cooling passage (C22) is connected to the inlet (E21) and is configured to extend from the outer panel 2530 to the interior of the inner panel 2520 through the airfoil 2510. The second cooling passage (C22) is formed by the first barrier 2561 and the second barrier 2562. The air introduced through the second cooling passage (C22) can be supplied to the first cooling passage (C21).

[0078] The first barrier 2561 is disposed between the first cooling passage (C21) and the second cooling passage (C22) and divides the first cooling passage (C21) and the second cooling passage (C22). The longitudinal outer end of the first barrier 2561 is attached to the outer shroud 2530, and the longitudinal inner end of the first barrier 2561 is positioned spaced from the end of the airfoil 2510. That is, the first barrier 2561 is configured to penetrate the airfoil 2510 and extend to the interior of the inner shroud 2520.

[0079] The first cooling passage (C21) and the second cooling passage (C22) are connected by the first curved passage section (B21), which is spaced from the end of the airfoil 2510 and disposed within the inner shroud 2520. The first curved passage section (B21) may, for example, be disposed within the inner protrusion 2528. Furthermore, the first curved passage section (B21) includes a first curved surface 2516 curved in an arcuate shape and having an arcuate cross-section with a first radius (R21) about a first point (P21). Here, the first point (P21) is spaced from the end of the airfoil 2510 and disposed within the inner shroud 2520.

[0080] Although the turbine blade 2500 is completely cooled by the cooling passage, it is confirmed that the heat stress is concentrated in the first bent passage portion (B21), which is a portion where the cooling passage is switched, thereby reducing the durability of the structure.

[0081] The hot gases pass through only one section where the airfoil 2510 is formed, and the inner liner 2520 and the outer liner 2530 are inserted into other elements, preventing them from contacting the hot gas. Consequently, by disposing the first curved passage section (B21) within the inner liner 2520, the heat load applied to the first curved passage section (B21) can be minimized.

[0082] The first barrier 2561 may include a plurality of introduction holes 2571 penetrating the first barrier 2561 in a thickness direction to introduce the flow of air. The introduction holes 2571 may be arranged to be spaced apart from each other in a height direction of the first barrier 2561, as well as to be spaced apart from each other in a width direction of the first barrier 2561. Some introduction holes 2571 may be arranged farther outward from the center of the turbine blade 2500 than the end of the airfoil 2510. Some introduction holes 2571 are formed inside the inner liner 2520, and the introduction hole 2571 may pass through the portion where the first point (P21) is located.Consequently, the first barrier 2561 is cooled through the introduction hole 2571, and a portion with a large heat load in the turbine blade 2500 can be cooled by the first barrier 2561.

[0083] In addition, some introduction holes 2571 may be arranged on a boundary line (L21) where the airfoil 2510 and the inner platform 2522 are connected. That is, a part of one introduction hole 2571 may be arranged within the airfoil 2510, and a part of it may be arranged within the inner platform 2522. Consequently, the heat load at the boundary portion can be minimized.

[0084] The third cooling passage (C23), located adjacent to the second cooling passage (C22), is formed to extend from the outer panel 2530 to the interior of the inner panel 2520 through the airfoil 2510. The third cooling passage (C23) is formed by the second barrier 2562 and the third barrier 2563 and is formed to penetrate the outer protrusion 2538, the outer platform 2532, and the airfoil 2510 in a height direction thereof.

[0085] The third cooling passage (C23) is connected to the inlet (E21), and the air introduced through the inlet (E21) moves by being divided into the second cooling passage (C22) and the third cooling passage (C23), respectively. The air moving through the third cooling passage (C23) can be supplied to the fourth cooling passage (C24).

[0086] The second barrier 2562 is disposed between the second cooling passage (C22) and the third cooling passage (C23), dividing the second cooling passage (C22) and the third cooling passage (C23). The longitudinal inner end of the second barrier 2562 is fixed to the inner liner 2520, and the longitudinal outer end thereof is disposed farther outward than the end of the airfoil 2510. That is, the second barrier 2562 penetrates the airfoil 2510 and extends to the inside of the outer liner 2530, spaced from the inlet (E21). Consequently, it is possible to prevent the heat load from being concentrated at the end of the second barrier 2562.

[0087] The second barrier 2562 may include a plurality of introduction holes 2572 penetrating the second barrier 2562 in its thickness direction. The introduction holes 2572 may be arranged to be spaced apart from each other in a width direction of the second barrier 2562. The introduction hole 2572 may be arranged farther outward from the center of the turbine blade 2500 than the end of the airfoil 2510. That is, the introduction hole 2572 may be arranged inside the outer liner 2530. Consequently, the second barrier 2562 is cooled by the introduction hole 2572, and a portion with a large heat load in the turbine blade 2500 may be cooled by the second barrier 2562.

[0088] The fourth cooling passage (C24), located between the third cooling passage (C23) and the trailing edge 2512, is formed to extend toward the airfoil 2510 and the interior of the inner fairing 2520. The fourth cooling passage (C24) is formed by the third barrier 2563 and the trailing edge 2512. The fourth cooling passage (C24) receives air from the third cooling passage (C23), and the air introduced into the fourth cooling passage (C24) is exhausted through the trailing edge 2512. A rear-end cooling slot 2519 is formed on the trailing edge 2512, and a split protrusion 2568 is formed on the rear-end cooling slot 2519, and the air cools the trailing edge 2512 while being exhausted through the rear-end cooling slot 2519.

[0089] The third barrier 2563 is disposed between the third cooling passage (C23) and the fourth cooling passage (C24) and divides the third cooling passage (C23) and the fourth cooling passage (C24). The longitudinal outer end of the third barrier 2563 is attached to the outer shroud 2530, and the longitudinal inner end of the third barrier 2563 is spaced from the end of the airfoil 2510 and positioned farther outward than the center of the turbine blade 2500. That is, the third barrier 2563 is configured to penetrate the airfoil 2510 and extend toward the interior of the inner shroud 2520.

[0090] The third cooling passage (C23) and the fourth cooling passage (C24) are connected by the second curved passage section (B22), which is spaced from the end of the airfoil 2510 and disposed within the inner shroud 2520. The second curved passage section (B22) may, for example, be disposed within the inner platform 2522 and the inner boss 2528. Furthermore, the second curved passage section (B22) includes a second curved surface 2517 curved in an arcuate shape and having an arcuate cross-section with a second radius (R22) around a second point (P22). Here, the second point (P22) is spaced from the end of the airfoil 2510 and disposed within the inner shroud 2520. Consequently, the heat load generated in the second curved passage portion (B22) can be minimized and the durability of the structure in the second curved passage portion (B22) can be improved.

[0091] The third barrier 2563 may include a plurality of introduction holes 2573 penetrating the third barrier 2563 in its thickness direction. The introduction holes 2573 may be arranged to be spaced apart from each other in a height direction of the third barrier 2563, as well as to be spaced apart from each other in a width direction of the third barrier 2563. Some introduction holes 2573 may be arranged farther outward from the center of the turbine blade 2500 than the end of the airfoil 2510. Some introduction holes 2573 are formed inside the inner liner 2520, and the introduction hole 2573 may pass through the portion where the second point (P22) is located. Consequently, the third barrier 2563 is cooled through the introduction hole 2573, and a portion with a large heat load in the turbine blade 2500 can be cooled by the third barrier 2563.

[0092] In addition, some introduction holes 2573 may be arranged at the interface where the airfoil 2510 and the inner platform 2522 are connected. That is, a part of one introduction hole 2573 may be arranged within the airfoil 2510, and a part of it may be arranged within the inner platform 2522. Consequently, the heat load at the interface can be minimized.

[0093] Fig. 8 is a longitudinal cross-sectional diagram illustrating a blade according to another exemplary embodiment.

[0094] With reference to Fig. 8, a turbine blade 3500 includes an inner fairing 3520, an outer fairing 3530, and an airfoil 3510 disposed between the inner fairing 3520 and the outer fairing 3530.

[0095] The airfoil 3510 may be a curved plate with a vane shape and have an optimized airfoil according to the specification of a gas turbine. The airfoil 3510 may include a leading edge 3511 located on an upstream side and a trailing edge 3512 located on a downstream side with respect to a flow direction of the combustion gas.

[0096] The inner fairing 3520 is coupled to a rotor disk and disposed at the inner end of the airfoil 3510 to support the airfoil 3510. The inner fairing 3520 includes an inner platform 3522 coupled to the interior of the airfoil 3510 and an inner hook 3524 projecting downwardly from the inner platform 3522 and coupled to the rotor disk.

[0097] The inner platform 3522 is formed in a substantially rectangular plate shape and is formed with an inner protrusion 3528 protruding to form a space therein. The inner protrusion 3528 protrudes inward, which is a direction toward the rotor disk, and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the inner protrusion 3528 includes a convex surface and a concave surface, and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof.

[0098] The outer fairing 3530 is coupled to a blade carrier (not shown) installed outwardly in a radial direction thereof and disposed at the outer end of the airfoil 3510 to support the airfoil 3510. The outer fairing 3530 includes an outer platform 3532 coupled to the outer end of the airfoil 3510 and an outer hook 3534 projecting upward from the outer platform 3532 and coupled to the blade carrier.

[0099] The outer platform 3532 is formed in a substantially rectangular plate shape and is formed with an outer protrusion 3538 protruding to form a space therein. The outer protrusion 3538 protrudes outward and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the outer protrusion 3538 includes a convex surface and a concave surface and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof. The outer protrusion 3538 is formed with an inlet (E31) through which the cooling medium is introduced and an outlet (O31) through which the cooling medium is discharged. The cooling medium may include, for example, the air compressed by the compressor. It goes without saying that the cooling medium is not limited to this.

[0100] A first cooling passage (C31), a second cooling passage (C32), a third cooling passage (C33), a fourth cooling passage (C34), a first curved passage portion (B31), a second curved passage portion (B32), a first barrier 3561, a second barrier 3562, and a third barrier 3563 are formed within the turbine blade 3500. The turbine blade 3500 may be formed by casting.

[0101] The first cooling passage (C31) is connected to the outlet (O31) and is configured to extend from the outer liner 3530 to the interior of the inner liner 3520 through the airfoil 3510. The first cooling passage (C31) is formed by the first barrier 3561 and the leading edge 3511 and is configured to penetrate the outer protrusion 3538, the outer platform 3532, and the airfoil 3510 in a height direction thereof.

[0102] The second cooling passage (C32) is connected to the inlet (E31) and is configured to extend from the outer panel 3530 to the interior of the inner panel 3520 through the airfoil 3510. The second cooling passage (C32) is formed by the first barrier 3561 and the second barrier 3562. The air introduced through the second cooling passage (C32) can be supplied to the first cooling passage (C31).

[0103] The first barrier 3561 is disposed between the first cooling passage (C31) and the second cooling passage (C32) and divides the first cooling passage (C31) and the second cooling passage (C32). The longitudinal outer end of the first barrier 3561 is attached to the outer cover 3530, and the longitudinal inner end of the first barrier 3561 is arranged to be spaced from the end of the airfoil 3510. That is, the first barrier 3561 is configured to penetrate the airfoil 3510 and extend to the interior of the inner cover 3520.

[0104] The first cooling passage (C31) and the second cooling passage (C32) are connected by the first curved passage section (B31), which is spaced from the end of the airfoil 3510 and disposed within the inner shroud 3520. The first curved passage section (B31) may, for example, be disposed within the inner protrusion 3528. Furthermore, the first curved passage section (B31) includes a first curved surface 3516 curved in an arcuate shape and having an arcuate cross-section with a first radius (R31) about a first point (P31). Here, the first point (P31) is spaced from the end of the airfoil 3510 and disposed within the inner shroud 3520.

[0105] It is confirmed that the turbine blade 3500 is completely cooled by the cooling passage, but the heat load is concentrated in the first bent passage portion (B31), which is a portion where the cooling passage is switched, thereby reducing the durability of the structure.

[0106] The hot gas passes only through the portion where the airfoil 3510 is formed, and the inner liner 3520 and the outer liner 3530 are inserted into other elements, preventing them from contacting the hot gas. Consequently, if the first curved passage portion (B31) is disposed within the inner liner 3520, the heat load applied to the first curved passage portion (B31) can be minimized.

[0107] The first barrier 3561 may include a first passage 3571 extending in a thickness direction of the first barrier 3561, and a second passage 3572 connected to the first passage 3571 and extending to the end of the first barrier 3561. The first passage 3571 is disposed within the airfoil 3510, and the second passage 3572 is formed to extend from the interior of the airfoil 3510 to the interior of the inner liner 3520 in a height direction of the first barrier 3561. The second passage 3572 may pass through the portion where the first point (P31) is located.

[0108] Air can be introduced into the first passage 3571 and discharged through the second passage 3572 to the end of the first barrier 3561. The air moves and cools the first barrier 3561, and a portion with a large heat load can be cooled by the first barrier 3561. Since the first point (P31) is located within the second passage 3572, the portion with the large heat load can be cooled efficiently.

[0109] The third cooling passage (C33), located adjacent to the second cooling passage (C32), is formed to extend from the outer panel 3530 to the interior of the inner panel 3520 through the airfoil 3510. The third cooling passage (C33) is formed by the second barrier 3562 and the third barrier 3563 and is formed to penetrate the outer protrusion 3538, the outer platform 3532, and the airfoil 3510 in a height direction thereof.

[0110] The third cooling passage (C33) is connected to the inlet (E31), and the air introduced through the inlet (E31) moves by being divided into the second cooling passage (C32) and the third cooling passage (C33). The air moving through the third cooling passage (C33) can be supplied to the fourth cooling passage (C34).

[0111] The second barrier 3562 is disposed between the second cooling passage (C32) and the third cooling passage (C33), and divides the second cooling passage (C32) and the third cooling passage (C33). The longitudinal inner end of the second barrier 3562 is fixed to the inner liner 3520, and the longitudinal outer end thereof is disposed farther outward than the end of the airfoil 3510. That is, the second barrier 3562 penetrates the airfoil 3510 and extends to the inside of the outer liner 3530, spaced from the inlet (E31). Consequently, it is possible to prevent the heat load from being concentrated at the end of the second barrier 3562.

[0112] The second barrier 3562 may include a first passage 3573 extending in a thickness direction of the second barrier 3562, and a second passage 3574 connected to the first passage 3573 and extending to the end of the second barrier 3562. The first passage 3573 is disposed within the airfoil 3510, and the second passage 3574 is configured to extend from the interior of the airfoil 3510 to the interior of the outer panel 3530 in a height direction of the second barrier 3562.

[0113] Air can be introduced into the second passage 3574 and discharged through the first passage 3573 to the end of the second barrier 3562. The air moves and cools the second barrier 3562, and a section with a large heat load can be cooled by the second barrier 3562.

[0114] The fourth cooling passage (C34), located between the third cooling passage (C33) and the trailing edge 3512, is configured to extend toward the airfoil 3510 and the interior of the inner fairing 3520. The fourth cooling passage (C34) is formed by the third barrier 3563 and the trailing edge 3512. The fourth cooling passage (C34) receives air from the third cooling passage (C33), and the air introduced into the fourth cooling passage (C34) is exhausted through the trailing edge 3512. A rear-end cooling slot is formed in the trailing edge 3512, a split protrusion 3568 is formed in the rear-end cooling slot, and the air is exhausted through the rear-end cooling slot to cool the trailing edge 3512.

[0115] The third barrier 3563 is disposed between the third cooling passage (C33) and the fourth cooling passage (C34) and divides the third cooling passage (C33) and the fourth cooling passage (C34). The longitudinal outer end of the third barrier 3563 is attached to the outer shroud 3530, and the longitudinal inner end of the third barrier 3563 is spaced from the end of the airfoil 3510 and is located farther outward than the center of the turbine blade 3500. That is, the third barrier 3563 is configured to penetrate the airfoil 3510 and extend to the interior of the inner shroud 3520.

[0116] The third cooling passage (C33) and the fourth cooling passage (C34) are connected by the second curved passage section (B32), which is spaced from the end of the airfoil 3510 and disposed within the inner shroud 3520. The second curved passage section (B32) may, for example, be disposed within the inner platform 3522 and the inner boss 3528. Furthermore, the second curved passage section (B32) includes a second curved surface 3517 curved in an arcuate shape and having an arcuate cross-section with a second radius (R32) around a second point (P32). Here, the second point (P32) is spaced from the end of the airfoil 3510 and disposed within the inner shroud 3520. Consequently, the heat stress generated in the second curved passage portion (B32) can be minimized and the durability of the structure in the second curved passage portion (B32) can be improved.

[0117] The third barrier 3563 may include a first passage 3575 extending in a thickness direction of the third barrier 3563, and a second passage 3576 connected to the first passage 3575 and extending to the end of the third barrier 3563. The first passage 3575 is disposed within the airfoil 3510, and the second passage 3576 is formed to extend from the interior of the airfoil 3510 to the interior of the inner panel 3520 in a height direction of the third barrier 3563. The second passage 3576 may pass through the portion where the second point (P32) is located.

[0118] Air can be introduced into the first passage 3575 and discharged to the end of the third barrier 3563 through the second passage 3576. The air moves and cools the third barrier 3563, and a section with a large heat load can be cooled by the third barrier 3563. Since the second point (P32) is located within the second passage 3576, the section with the large heat load can be cooled efficiently.

[0119] Fig. 9 is a longitudinal cross-sectional diagram illustrating a blade according to another exemplary embodiment.

[0120] With reference to Fig. 9, a turbine blade 4500 includes an inner fairing 4520, an outer fairing 4530, and an airfoil 4510 disposed between the inner fairing 4520 and the outer fairing 4530.

[0121] The airfoil 4510 may be a curved plate with a vane shape and have an optimized airfoil according to the specification of a gas turbine. The airfoil 4510 may include a leading edge 4511 located on an upstream side and a trailing edge 4512 located on a downstream side with respect to a flow direction of the combustion gas.

[0122] The inner fairing 4520 is coupled to a rotor disk and disposed at the inner end of the airfoil 4510 to support the airfoil 4510. The inner fairing 4520 includes an inner platform 4522 coupled to the interior of the airfoil 4510 and an inner hook 4524 projecting downwardly from the inner platform 4522 and coupled to the rotor disk.

[0123] The inner platform 4522 is formed in a substantially rectangular plate shape and is formed with an inner protrusion 4528 protruding to form a space therein. The inner protrusion 4528 protrudes inward, which is a direction toward the rotor disk, and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the inner protrusion 4528 includes a convex surface and a concave surface, and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof.

[0124] The outer fairing 4530 is coupled to a blade carrier (not shown) installed outward in a radial direction thereof and disposed at the outer end of the airfoil 4510 to support the airfoil 4510. The outer fairing 4530 includes an outer platform 4532 coupled to the outer end of the airfoil 4510 and an outer hook 4534 projecting upward from the outer platform 4532 and coupled to the blade carrier.

[0125] The outer platform 4532 is formed in a substantially rectangular plate shape and is formed with an outer protrusion 4538 protruding to form a space therein. The outer protrusion 4538 protrudes outward and has the transverse cross section shaped like the airfoil. That is, the transverse cross section of the outer protrusion 4538 includes a convex surface and a concave surface and is formed such that a gap between the convex surface and the concave surface decreases toward the side end thereof. The outer protrusion 4538 may be formed with an inlet (E41) through which the cooling medium is introduced. The cooling medium may include, for example, the air compressed by the compressor. It goes without saying that the cooling medium is not limited to this.

[0126] A first cooling passage (C41), a second cooling passage (C42), a third cooling passage (C43), a first curved passage portion (B41), a second curved passage portion (B42), a first barrier 4561, and a second barrier 4562 are formed within the turbine blade 4500. The turbine blade 4500 may be formed by casting.

[0127] The first cooling passage (C41) is connected to the inlet (E41) and is configured to extend from the outer liner 4530 to the interior of the inner liner 4520 through the airfoil 4510. The first cooling passage (C41) is formed by the first barrier 4561 and the leading edge 4511 and is configured to penetrate the outer protrusion 4538, the outer platform 4532, and the airfoil 4510 in a height direction thereof. The air introduced into the first cooling passage (C41) can be supplied to the second cooling passage (C42).

[0128] The second cooling passage (C42) is configured to extend from the outer panel 4530 to the interior of the inner panel 4520 through the airfoil 4510. The second cooling passage (C42) is formed by the first barrier 4561 and the second barrier 4562. The second cooling passage (C42) receives air from the first cooling passage (C41) and supplies the air to the third cooling passage (C43).

[0129] The first barrier 4561 is disposed between the first cooling passage (C41) and the second cooling passage (C42) and divides the first cooling passage (C41) and the second cooling passage (C42). The longitudinal outer end of the first barrier 4561 is attached to the outer cover 4530, and the longitudinal inner end of the first barrier 4561 is arranged to be spaced from the end of the airfoil 4510. That is, the first barrier 4561 is configured to penetrate the airfoil 4510 and extend to the interior of the inner cover 4520.

[0130] The first cooling passage (C41) and the second cooling passage (C42) are connected by the first curved passage section (B41), which is spaced from the end of the airfoil 4510 and disposed within the inner shroud 4520. The first curved passage section (B41) may, for example, be disposed within the inner protrusion 4528. Furthermore, the first curved passage section (B41) includes a first curved surface 4516 curved in an arcuate shape and having an arcuate cross-section with a first radius (R41) about a first point (P41). Here, the first point (P41) is spaced from the end of the airfoil 4510 and disposed within the inner shroud 4520.

[0131] It is confirmed that the turbine blade 4500 is completely cooled by the cooling passage, but a heat stress is concentrated in the first bent passage portion (B41), which is the portion where the cooling passage is switched, thereby reducing the durability of the structure.

[0132] The hot gases pass only through the portion where the airfoil 4510 is formed, and the inner liner 4520 and the outer liner 4530 are inserted into other elements, preventing them from contacting the hot gas. Consequently, by disposing the first curved passage portion (B41) within the inner liner 4520, the heat load applied to the first curved passage portion (B41) can be minimized.

[0133] A plurality of porous plates 4570 are formed to protrude from the first barrier 4561. Some porous plates 4570 may be disposed within the first curved passage section (B41), and some porous plates 4570 may be disposed within the airfoil 4510 adjacent to the first curved passage section (B41). As described above, when the plurality of porous plates 4570 are formed to protrude from the first barrier 4561, the first barrier 4561 and the turbine blade 4500 can be cooled by the porous plate 4570, thereby reducing thermal stress. The porous plate 4570 may be vertically attached to an outer surface of the first barrier 4561 or may be disposed to be inclined toward the end of the turbine blade 4500.

[0134] The third cooling passage (C43), located between the second cooling passage (C42) and the trailing edge 4512, is configured to extend from the outer fairing 4530 to the airfoil 4510. The third cooling passage (C43) is formed by the second barrier 4562 and the trailing edge 4512. The third cooling passage (C43) receives air from the second cooling passage (C42), and the air introduced into the third cooling passage (C43) is exhausted through the trailing edge 4512. A rear-end cooling slot is formed in the trailing edge 4512, a split protrusion 4568 is formed in the rear-end cooling slot, and the air is exhausted through the rear-end cooling slot to cool the trailing edge 4512.

[0135] The second barrier 4562 is disposed between the second cooling passage (C42) and the third cooling passage (C43) and divides the second cooling passage (C42) and the third cooling passage (C43). The longitudinal inner end of the second barrier 4562 is attached to the inner liner 4520, and the longitudinal outer end thereof is disposed farther outward than the end of the airfoil 4510. That is, the second barrier 4562 is configured to penetrate the airfoil 4510 and extend toward the interior of the outer liner 4530.

[0136] The second cooling passage (C42) and the third cooling passage (C43) are connected by the second curved passage section (B42), which is spaced from the end of the airfoil 4510 and disposed within the outer shroud 4530. The second curved passage section (B42) may, for example, be disposed within the outer platform 4532 and the outer projection 4538. Furthermore, the second curved passage section (B42) includes a second curved surface 4517 curved in an arcuate shape and having an arcuate cross-section with a second radius (R42) around a second point (P42). Here, the second point (P42) is spaced from the end of the airfoil 4510 and disposed within the outer shroud 4530. Consequently, the heat load generated in the second curved passage portion (B42) can be minimized and the durability of the structure in the second curved passage portion (B42) can be improved.

[0137] A plurality of porous plates 4570 are formed to protrude from the second barrier 4562. Some porous plates 4570 may be disposed within the second curved passage section (B42), and some porous plates 4570 may be disposed within the airfoil 4510 adjacent to the second curved passage section (B42). As described above, when the plurality of porous plates 4570 are formed to protrude from the second barrier 4562, the second barrier 4562 and the turbine blade 4500 may be cooled by the porous plate 4570, thereby reducing thermal stress.

[0138] Although exemplary embodiments have been described with reference to the accompanying drawings, it will be apparent to those skilled in the art that various modifications in form and detail may be made therein without departing from the spirit and scope as defined in the appended claims. Therefore, the description of the exemplary embodiments should be construed in a descriptive sense and not as limiting the scope of the claims, and many alternatives, modifications, and variations will be apparent to those skilled in the art.

Claims

[1] Turbine blade comprising: an airfoil (2510, 3510, 4510) having a leading edge (2511, 3511, 4511) and a trailing edge (2512, 3512, 4512); an inner fairing (2520, 3520, 4520) disposed at one end of the airfoil (2510, 3510, 4510) to support the airfoil (2510, 3510, 4510); an outer fairing (2530, 3530, 4530) arranged at the other end of the airfoil (2510, 3510, 4510) to support the airfoil (2510, 3510, 4510) and opposite the inner fairing (2520, 3520, 4520); a first cooling passage (C21, C31, C41) and a second cooling passage (C22, C32, C42) extending in a height direction thereof; and a first barrier (2561, 3561, 4561) dividing the first cooling passage (C21, C31, C41) and the second cooling passage (C22, C32, C42) and extending in a height direction of the airfoil (2510, 3510, 4510), wherein a longitudinal end of the first barrier (2561, 3561, 4561) is arranged further outward than an end of the airfoil (2510, 3510, 4510); and a first curved passage section (B21, B31, B41) connecting the first cooling passage (C21, C31, C41) and the second cooling passage (C22, C32, C42), wherein the first curved passage section (B21, B31, B41) is arranged within the inner panel (2520, 3520, 4520) or the outer panel (2530, 3530, 4530), and where: the first barrier (2561) has a plurality of introduction holes (2571) penetrating the first barrier (2561) in its thickness direction, and some introduction holes (2571) are arranged further outward than the end of the airfoil (2510) with respect to a center of the turbine blade (2500); or a plurality of porous plates (4570) are formed to protrude from the first barrier (4561), the porous plates (4570) being arranged on the first curved passage portion (B41); or the first barrier (3561) comprises a first passage (3571) extending in a thickness direction of the first barrier (3561), and a second passage (3572) connected to the first passage (3571) and extending to one end of the first barrier (3561), and the first passage (3571) is disposed within the airfoil (3510), and the second passage (3572) extends from an interior of the airfoil (3510) to an interior of the inner panel (3520). [2] The turbine blade of claim 1, wherein the first curved passage portion (B21, B31, B41) comprises a first curved surface (2516, 3516, 4516) curved in an arc shape about a first point (P21, P31, P41) located within the inner liner (2520, 3520, 4520) or the outer liner (2530, 3530, 4530). [3] Turbine blade according to one of claims 1 to 2, wherein the inner fairing (2520, 3520, 4520) comprises an inner platform (2522, 3522, 4522) connected to an inner end of the airfoil (2510, 3510, 4510) and an inner hook (2524, 3524, 4524) projecting from the inner platform (2522, 3522, 4522), and wherein the inner platform (2522, 3522, 4522) comprises an inner projection (2528, 3528, 4528) projecting to form a space therein, and the first curved passage portion (B21, B31, B41) is disposed within the inner projection (2528, 3528, 4528). [4] The turbine blade according to claim 1 or 2, wherein, when the first barrier (2561) has the plurality of introduction holes (2571) penetrating the first barrier (2561) in its thickness direction, and some introduction holes (2571) are arranged further outward than the end of the airfoil (2510) with respect to a center of the turbine blade (2500), some introduction holes (2571) are arranged at an interface in which the airfoil (2510) and the inner platform (2522) are connected. [5] A turbine blade according to claim 1 or claim 2, further comprising a third cooling passage (C23, C33, C43) extending in a height direction thereof, and a second curved passage portion (B42) connecting the second cooling passage (C22, C32, C42) and the third cooling passage (C23, C33, C43), wherein the second curved passage portion (B42) includes a second curved surface (2517, 3517, 4517) curved in an arc shape around a second point (P42) located inside the outer shroud (2530, 3530, 4530). [6] Turbine blade according to claim 5, wherein the outer fairing (2530, 3530, 4530) comprises an outer platform (2532, 3532, 4532) connected to an outer end of the airfoil (2510, 3510, 4510) and an outer hook (2534, 3534, 4534) projecting from the outer platform (2532, 3532, 4532), and wherein the outer platform (2532, 3532, 4532) comprises an outer projection (2538, 3538, 4538) projecting outwardly and forming a space therein, and the second curved passage portion (B42) is disposed within the outer projection (2538, 3538, 4538). [7] A turbine blade according to claim 5, wherein the outer projection (2538, 3538, 4538) comprises a transverse cross-section shaped like the airfoil. [8] Turbine blade according to one of claims 1 to 2, further comprising a third cooling passage (C23, C33, C43) extending in a height direction thereof, and a second barrier (2562, 3562, 4562) connecting the second cooling passage (C22, C32, C42) and the third cooling passage (C23, C33, C43) and extending in a height direction of the airfoil (2510, 3510, 4510), wherein a longitudinal end of the second barrier (2562, 3562, 4562) is disposed within the outer panel (2530, 3530, 4530). [9] The turbine blade according to claim 8, wherein the second barrier (2562) comprises a plurality of introduction holes (2572) penetrating the second barrier (2562) in its thickness direction, and some introduction holes (2572) are arranged inside the outer liner (2530). [10] Turbine with the turbine blade (2500, 3500, 4500) according to one of claims 1 to 9. [11] Gas turbine comprising the turbine (1300) according to claim 10.

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