KR102358606B1 - Turbine blade and gas turbine including turbine blade - Google Patents

Abstract

The present invention relates to a turbine blade and a gas turbine, and according to one aspect of the present invention, a platform unit; a root portion formed at an inner end of the platform portion in a radial direction; and an airfoil portion formed at a radially outer end of the platform portion, wherein the airfoil portion has a turbine blade cooling passage formed therein, and an outlet of the turbine blade cooling passage is formed at the radially outer end of the airfoil portion It provides a turbine blade comprising a blade tip, the blade tip is characterized in that the projection is formed.

Description

Turbine blades and a gas turbine comprising the turbine blades

The present invention relates to turbine blades used in gas turbines. More particularly, it relates to a blade tip structure for protecting a turbine blade cooling passage from hot combustion gases.

In general, a turbine is a machine that converts the energy of a fluid such as water, gas, steam, etc. into mechanical work, usually by planting several blades or blades on the circumference of a rotating body and blowing steam or gas thereto at high speed with impulse or reaction force. A turbo-type machine that rotates is called a turbine.

Examples of such turbines include a hydro turbine using the energy of high water, a steam turbine using the energy of steam, an air turbine using the energy of high-pressure compressed air, and a gas using the energy of high-temperature and high-pressure gas. turbines, etc.

Among them, the gas turbine includes a compressor, a combustor, a turbine and a rotor as shown in FIG. 1 .

The compressor includes a plurality of compressor vanes and a plurality of compressor blades arranged alternately with each other.

The combustor generates high-temperature and high-pressure combustion gas by supplying fuel to the compressed air compressed by the compressor and igniting it with a burner.

The turbine includes a plurality of turbine vanes and a plurality of turbine blades arranged alternately with each other.

The rotor is formed to pass through the central portions of the compressor, the combustor, and the turbine, both ends are rotatably supported by bearings, and one end is connected to a drive shaft of the generator.

In addition, the rotor includes a plurality of compressor disks coupled to the compressor blades, a plurality of turbine disks coupled to the turbine blades, and a torque tube for transmitting rotational force from the turbine disks to the compressor disks.

In the gas turbine according to this configuration, the compressed air in the compressor is mixed with fuel in the combustion chamber and is converted into high-temperature combustion gas by combustion, the combustion gas thus made is injected toward the turbine, and the injected combustion gas is the turbine blade. A rotational force is generated while passing, and the rotor rotates.

Since these gas turbines do not have a reciprocating mechanism such as a piston of a four-stroke engine, there is no mutual friction part such as a piston-cylinder, so the consumption of lubricant is extremely low. There are advantages.

On the other hand, in a turbine blade, a method of cooling the turbine blade by forming a cooling passage through which a cooling fluid flows inside the turbine blade is known because the combustion gas of high temperature directly contacts the turbine blade.

However, in such a conventional turbine blade, there is a problem that a part of the high-temperature combustion gas penetrates into the cooling passage outlet.

Accordingly, an object of the present invention is to provide a turbine blade capable of preventing infiltration of combustion gas into a turbine blade cooling passage, and a gas turbine including the turbine blade.

According to an aspect of the present invention for achieving the above technical problem, the platform unit; a root portion formed at an inner end of the platform portion in a radial direction; and an airfoil portion formed at a radially outer end of the platform portion, wherein the airfoil portion has a turbine blade cooling passage formed therein, and an outlet of the turbine blade cooling passage is formed at the radially outer end of the airfoil portion A turbine blade is provided, comprising a blade tip, wherein the blade tip has a protrusion formed therein.

According to another aspect of the present invention, the disk; A rotor is provided, comprising a plurality of turbine blades radially mounted on an outer circumferential surface of the disk, the turbine blades according to any one of claims 1 to 13.

According to another aspect of the present invention, the housing; a compressor for compressing the air introduced into the housing; a combustor that mixes fuel with the air compressed in the compressor and ignites to generate combustion gas; and a turbine for rotating the compressor by obtaining rotational force from the combustion gas generated from the combustor, wherein the turbine includes the turbine disk according to claim 14 .

According to the aspects of the present invention having the above configuration, it is possible to minimize the combustion gas flowing into the cooling passage by the protrusion, thereby minimizing the deterioration of the turbine-side blade end. Accordingly, the lifespan of the product can be improved and the maintenance cycle can be extended.

1 is a cross-sectional view of a gas turbine;
2 is a perspective view showing a gas turbine blade in a gas turbine according to an embodiment of the present invention;
Figure 3 is a perspective view showing the blade tip portion of Figure 2;
Figure 4 is a perspective view showing the blade tip portion of Figure 2 from another angle;
5A and 5B are cross-sectional views showing whether hot gas penetrates according to the presence or absence of a protrusion;
Figure 6a is a cross-sectional view taken along line I-I of Figure 4;
Figure 6b is a sectional view taken along line II-II of Figure 4;
7A to 7C are cross-sectional views illustrating a blade tip of a gas turbine blade according to another embodiment of the present invention.

Hereinafter, a gas turbine according to the present invention will be described in detail with reference to the accompanying drawings.

Figure 1 is a cross-sectional view showing a gas turbine, Figure 2 is a perspective view showing a gas turbine blade in a gas turbine according to an embodiment of the present invention, Figure 3 is a perspective view showing the blade tip portion of Figure 2, Figure 4 is a perspective view showing the blade tip portion of FIG. 2 at a different angle, FIGS. 5A and 5B are cross-sectional views showing whether hot gas penetrates according to the presence or absence of a protrusion, and FIG. 6A is a cross-sectional view taken along line I-I of FIG. , FIG. 6B is a cross-sectional view taken along line II-II of FIG. 4, and FIGS. 7A to 7C are cross-sectional views illustrating a blade tip of a gas turbine blade according to another embodiment of the present invention.

1 , the gas turbine according to an embodiment of the present invention includes a housing 100 , a rotor 600 rotatably provided in the housing 100 , and rotational force from the rotor 600 . Compressor 200 that receives and compresses air flowing into the housing 100, a combustor 400 that mixes fuel with the compressed air in the compressor 200 and ignites to generate combustion gas, the combustor 400 ) to obtain rotational force from the combustion gas generated from the turbine 500 for rotating the rotor 600 , a generator interlocked with the rotor 600 for power generation, and a diffuser for discharging the combustion gas passing through the turbine 500 . may include

The housing 100 includes a compressor housing 110 in which the compressor 200 is accommodated, a combustor housing 120 in which the combustor 400 is accommodated, and a turbine housing 130 in which the turbine 500 is accommodated. can do.

Here, the compressor housing 110 , the combustor housing 120 , and the turbine housing 130 may be sequentially arranged from an upstream side to a downstream side in a fluid flow direction.

The rotor 600 includes a compressor disk 610 accommodated in the compressor housing 110 , a turbine disk 630 accommodated in the turbine housing 130 , and the compressor disk 630 accommodated in the combustor housing 120 . A torque tube 620 connecting 610 and the turbine disk 630, the compressor disk 610, a tie rod 640 and a fixing nut for fastening the torque tube 620 and the turbine disk 630 ( 650) may be included.

The compressor disk 610 may be formed in plurality, and the plurality of the compressor disks 610 may be arranged along the axial direction of the rotor. That is, the compressor disk 610 may be formed in multiple stages.

In addition, each compressor disk 610 may be formed in a substantially disk shape, and a compressor disk slot coupled to a compressor blade 210 to be described later may be formed on an outer periphery thereof.

The compressor disk slot may be formed in a fir-tree shape to prevent the compressor blade 210, which will be described later, from being separated from the compressor disk slot in the rotational radial direction of the rotor.

Here, the compressor disk 610 and the compressor blade 210 to be described later are typically coupled to each other in a tangential type or an axial type, but in this embodiment, they are formed to be coupled in an axial type. Accordingly, the plurality of compressor disk slots according to the present embodiment may be formed, and the plurality of compressor disk slots may be radially arranged along the circumferential direction of the compressor disk 610 .

The turbine disk 630 may be formed similarly to the compressor disk 610 . That is, the turbine disk 630 may be formed in plurality, and the plurality of turbine disks 630 may be arranged along the axial direction of the rotor. That is, the turbine disk 630 may be formed in multiple stages.

In addition, each turbine disk is formed in a substantially disk shape, and a turbine disk slot coupled to a turbine blade 700 to be described later may be formed on the outer periphery.

The turbine disk slot may be formed in the shape of a fir tree to prevent the turbine blade 700, which will be described later, from being separated from the turbine disk slot in the rotational radial direction of the rotor.

Here, the turbine disk 630 and the turbine blade 700 to be described later are typically coupled to each other in a tangential type or an axial type, but in this embodiment, they are formed to be coupled in an axial type. Accordingly, the turbine disk slot according to this embodiment is formed in plurality, the plurality of the turbine disk slot may be radially arranged along the circumferential direction of the turbine disk (630).

The torque tube 620 is a torque transmission member that transmits the rotational force of the turbine disk 630 to the compressor disk 610 , and has one end at the most downstream end in the air flow direction among the plurality of compressor disks 610 . It may be engaged with the located compressor disk, and the other end may be engaged with the turbine disk located at the most upstream end in the flow direction of the combustion gas among the plurality of turbine disks 630 . Here, a protrusion is formed on each of one end and the other end of the torque tube 620 , and a groove meshing with the protrusion is formed in each of the compressor disk 610 and the turbine disk 630 , the torque tube 620 . ) may be prevented from rotating relative to the compressor disk 610 and the turbine disk 630 .

In addition, the torque tube 620 may be formed in a hollow cylinder shape so that air supplied from the compressor 200 may flow to the turbine 500 through the torque tube.

In addition, the torque tube 620 is strongly formed in deformation and distortion due to the characteristics of a gas turbine continuously operated for a long period of time, and can be easily assembled and disassembled for easy maintenance.

The tie rod 640 is formed to pass through the plurality of compressor disks 610 , the torque tube 620 and the plurality of turbine disks 630 , and one end of the plurality of compressor disks 610 of air It is fastened in the compressor disk located at the most upstream end in the flow direction, and the other end is opposite to the compressor 200 with respect to the turbine disk located at the most downstream end in the flow direction of the combustion gas among the plurality of turbine disks 630 . may protrude and be fastened to the fixing nut 650 .

Here, the fixing nut 650 presses the turbine disk located at the most downstream end toward the compressor 200, and the interval between the compressor disk located at the most upstream end and the turbine disk located at the most downstream end is reduced. Accordingly, the plurality of compressor disks 610 , the torque tube 620 , and the plurality of turbine disks 630 may be compressed in the axial direction of the rotor. Accordingly, axial movement and relative rotation of the plurality of compressor disks 610 , the torque tube 620 and the plurality of turbine disks 630 may be prevented.

Meanwhile, in the present embodiment, one tie rod 640 is formed to pass through the central portions of the plurality of compressor disks 610 , the torque tube 620 and the plurality of turbine disks 630 , but is limited thereto. it is not That is, separate tie rods may be provided on the compressor 200 side and the turbine 500 side, respectively, and a plurality of tie rods may be radially disposed along the circumferential direction, and a mixture thereof may be used.

Both ends of the rotor according to this configuration may be rotatably supported by bearings, and one end may be connected to the drive shaft of the generator.

The compressor 200 includes a compressor blade 210 that rotates together with the rotor and a compressor vane 220 that is fixedly installed in the housing 100 to align the flow of air flowing into the compressor blade 210 . can do.

The compressor blades 210 are formed in plurality, the plurality of compressor blades 210 are formed in a plurality of stages along the axial direction of the rotor, and the plurality of compressor blades 210 rotate at each stage of the rotor. It may be formed radially along the direction.

In addition, each compressor blade 210 includes a plate-shaped compressor blade platform portion, a compressor blade root portion extending from the compressor blade platform portion to a centripetal side in a rotational radial direction of the rotor, and a rotation radius of the rotor from the compressor blade platform portion It may include a compressor blade airfoil portion extending in a direction distal to the centrifugal side.

The compressor blade platform unit may be in contact with a neighboring compressor blade platform unit and may serve to maintain a gap between the compressor blade airfoil units.

As described above, the compressor blade root portion may be formed in a so-called axial type to be inserted into the compressor disk slot along the axial direction of the rotor.

In addition, the compressor blade root portion may be formed in a fir tree shape to correspond to the compressor disk slot.

Here, in the present embodiment, the compressor blade root and the compressor disk slot are formed in a fir tree shape, but are not limited thereto and may be formed in a dovetail shape or the like. Alternatively, the compressor blade 210 may be fastened to the compressor disk 610 using another fastening device other than the above type, for example, a fastener such as a key or bolt.

The compressor blade root portion and the compressor disk slot are formed to be larger than the compressor blade root portion so that the compressor blade root portion and the compressor disk slot can be easily coupled to each other. A turbine tip clearance may be formed between the compressor blade root and the compressor disk slot.

In addition, although not shown separately, the compressor blade root portion and the compressor disk slot are fixed by separate pins, so that the compressor blade root portion is prevented from being separated from the compressor disk slot in the axial direction of the rotor.

The compressor blade airfoil part is formed to have an airfoil optimized according to the gas turbine specifications, and is located on the upstream side in the air flow direction to enter the compressor blade airfoil part leading edge (leading edge) and the air flow direction The compressor blade airfoil portion may include a trailing edge positioned on the upstream and downstream side from which air is emitted.

The compressor vanes 220 may be formed in plurality, and the plurality of compressor vanes 220 may be formed in a plurality of stages along the axial direction of the rotor. Here, the compressor vane 220 and the compressor blade 210 may be alternately arranged along the air flow direction.

In addition, the plurality of compressor vanes 220 may be formed radially along the rotation direction of the rotor at each stage.

In addition, each compressor vane 220 may include a compressor vane platform part formed in an annular shape along the rotational direction of the rotor and a compressor vane airfoil part extending from the compressor vane platform part in the rotational radial direction of the rotor. have.

The compressor vane platform part is formed in the proximal part of the compressor vane airfoil part and is formed in the root side compressor vane platform part which is fastened to the compressor housing 110 and the compressor vane airfoil part is formed in the tip part and faces the rotor. It may include a tip side compressor vane platform part.

Here, the compressor vane platform unit according to the present embodiment supports the root side compressor vane platform unit and the compressor vane airfoil unit in order to more stably support the compressor vane airfoil unit by supporting not only the blade part but also the tip part of the compressor vane airfoil part. including, but not limited to, a tip side compressor vane platform portion. That is, the compressor vane platform portion may be formed to support only the blade portion of the compressor vane airfoil portion including the root-side compressor vane platform portion.

Meanwhile, each compressor vane 220 may further include a compressor vane root portion coupling the root-side compressor vane platform portion and the compressor housing.

The compressor vane airfoil part is formed to have an airfoil optimized according to the gas turbine specifications, and is located on the upstream side in the air flow direction, and the leading edge of the compressor vane airfoil part into which air is incident and the air flow direction upstream side The compressor vane airfoil portion may include a trailing edge positioned and from which air exits.

The combustor 400 mixes and combusts the air flowing in from the compressor 200 with fuel to produce high-energy, high-temperature, high-pressure combustion gas, and the combustor 400 and the turbine 500 withstand the isostatic combustion process. It can be configured to raise the combustion gas temperature to the limit of possible heat resistance.

Specifically, the combustor 400 may be formed in plurality, and the plurality of combustors 400 may be arranged in the combustor housing 120 along the rotational direction of the rotor.

In addition, each combustor 400 includes a liner into which the air compressed by the compressor 200 flows, a burner that injects fuel into the air flowing into the liner and burns it, and a combustion gas generated from the burner into the turbine 500 . ) may include a transition piece to guide.

The liner may include a flame barrel forming a combustion chamber and a flow sleeve enclosing the flame barrel to form an annular space.

The burner may include a fuel injection nozzle formed on a front side of the liner to inject fuel into the air flowing into the combustion chamber, and a spark plug formed on a wall portion of the liner to ignite the fuel and air mixed in the combustion chamber. can

The transition piece may be formed so that an outer wall of the transition piece is cooled by the air supplied from the compressor 200 so as not to be damaged by the high temperature of the combustion gas.

That is, a cooling hole for injecting air into the transition piece is formed, and the air can cool the main body therein through the cooling hole.

On the other hand, the air that has cooled the transition piece flows into the annular space of the liner, and on the outer wall of the liner, air from the outside of the flow sleeve is provided as cooling air through a cooling hole provided in the flow sleeve to collide have.

Here, although not shown separately, between the compressor 200 and the combustor 400, there is a deworler serving as a guide blade to adjust the flow angle of the air flowing into the combustor 400 to the design flow angle. can be formed.

The turbine 500 may be formed similarly to the compressor 200 .

That is, the turbine 500 is a turbine blade 700 rotated together with the rotor and a turbine vane 520 fixedly installed in the housing 100 to align the flow of air introduced into the turbine blade 700 . may include

The turbine blades 700 are formed in plurality, the plurality of the turbine blades 700 are formed in a plurality of stages along the axial direction of the rotor, and the plurality of the turbine blades 700 rotates the rotor at each stage. It may be formed radially along the direction.

In addition, each turbine blade 700 includes a plate-shaped turbine blade platform portion, a turbine blade root portion extending from the turbine blade platform portion to a centripetal side in a rotational radial direction of the rotor, and a rotation radius of the rotor from the turbine blade platform portion It may include a turbine blade airfoil portion extending distally in a direction.

The turbine blade platform unit may be in contact with the neighboring turbine blade platform unit and serve to maintain a gap between the turbine blade airfoil units.

As described above, the turbine blade root portion may be formed in a so-called axial type to be inserted into the turbine disk slot along the axial direction of the rotor.

And, the turbine blade root portion may be formed in the shape of a fir tree to correspond to the turbine disk slot.

Here, in the present embodiment, the turbine blade root portion and the turbine disk slot are formed in a fir tree shape, but are not limited thereto and may be formed in a dovetail shape or the like. Alternatively, the turbine blade 700 may be fastened to the turbine disk 630 using a fastener other than the above type, for example, a key or a fastener such as a bolt.

And, the turbine blade root portion and the turbine disk slot, the turbine disk slot is formed larger than the turbine blade root portion so that the turbine blade root portion and the turbine disk slot can be easily fastened, in a coupled state A gap may be formed between the turbine blade root and the turbine disk slot.

And, although not shown separately, the turbine blade root portion and the turbine disk slot are fixed by separate pins, so that the turbine blade root portion is prevented from being separated from the turbine disk slot in the axial direction of the rotor.

The turbine blade airfoil part is formed to have an airfoil optimized according to the gas turbine specification, and is located on the upstream side in the flow direction of the combustion gas so that the leading edge of the turbine blade airfoil part into which the combustion gas is incident and the combustion gas flow direction It may include a turbine blade airfoil portion trailing edge positioned on the downstream side from which combustion gases are emitted.

The turbine vane 520 may be formed in plurality, and the plurality of turbine vanes 520 may be formed in a plurality of stages along the axial direction of the rotor. Here, the turbine vane 520 and the turbine blade 700 may be alternately arranged with each other along the air flow direction.

In addition, the plurality of turbine vanes 520 may be radially formed along the rotational direction of the rotor at each stage.

In addition, each turbine vane 520 may include a turbine vane platform portion formed in an annular shape along the rotational direction of the rotor and a turbine vane airfoil portion extending from the turbine vane platform portion in the rotational radial direction of the rotor. have.

The turbine vane platform part is formed at the tip of the turbine vane airfoil part and is formed at the tip end of the turbine vane airfoil part and the root-side turbine vane platform part which is fastened to the turbine housing 130 and faces the rotor. It may include a tip side turbine vane platform part.

Here, the turbine vane platform part according to this embodiment supports the root side turbine vane platform part and the turbine vane airfoil part to more stably support the turbine vane airfoil part by supporting not only the blade part but also the tip part of the turbine vane airfoil part. including, but not limited to, a tip side turbine vane platform portion. That is, the turbine vane platform portion may be formed to support only the blade portion of the turbine vane airfoil portion including the root-side turbine vane platform portion.

Meanwhile, each turbine vane 520 may further include a turbine vane root for coupling the root-side turbine vane platform and the turbine housing 130 .

The turbine vane airfoil part is formed to have an airfoil optimized according to gas turbine specifications, and is located on the upstream side in the flow direction of the combustion gas, so that the leading edge of the turbine vane airfoil part into which the combustion gas is incident and on the flow direction of the combustion gas It may include a turbine vane airfoil portion trailing edge located on the downstream side from which combustion gases are emitted.

Here, the turbine 500, unlike the compressor 200, comes in contact with high-temperature and high-pressure combustion gas, and thus requires a cooling means for preventing damage such as deterioration.

Accordingly, the gas turbine according to the present embodiment may further include a cooling passage for extracting compressed air from a portion of the compressor 200 and supplying it to the turbine 500 . Hereinafter, the air in the cooling passage will be referred to as a cooling fluid.

The cooling flow path may extend from the outside of the housing 100 (external flow path), extend through the inside of the rotor (internal flow path), or both an external flow path and an internal flow path may be used.

In addition, the cooling passage communicates with a turbine blade cooling passage formed inside the turbine blade 700 , so that the turbine blade 700 may be cooled by the cooling fluid.

In addition, the turbine blade cooling passage communicates with a turbine blade film cooling hole formed on the surface of the turbine blade 700 , and a cooling fluid is supplied to the surface of the turbine blade 700 , so that the turbine blade 700 is heated. It can be so-called film cooling by cooling air.

In addition, the turbine vane 520 may also be formed to be cooled by receiving a cooling fluid from the cooling passage similarly to the turbine blade 700 .

Meanwhile, the turbine 500 requires a turbine tip clearance between the tip of the turbine blade 700 and the inner peripheral surface of the turbine housing 130 so that the turbine blade 700 can rotate smoothly.

However, the wider the turbine tip clearance is, the more advantageous it is in terms of preventing interference between the turbine blade 700 and the turbine housing 130, but it is disadvantageous in terms of combustion gas leakage, and vice versa. That is, the flow of combustion gas injected from the combustor 400 is a main flow flowing through the turbine blade 700 and a leakage flow passing through the turbine tip clearance between the turbine blade 700 and the turbine housing 130 . It can be distinguished, as the turbine tip clearance is wide, the leakage flow is increased and the gas turbine efficiency is lowered, but interference between the turbine blade 700 and the turbine housing 130 due to thermal deformation and the resulting damage are reduced. can be prevented. On the other hand, as the turbine tip clearance is narrower, the leakage flow is reduced and gas turbine efficiency is improved. have.

In the gas turbine according to this configuration, the air introduced into the housing 100 is compressed by the compressor 200 , and the air compressed by the compressor 200 is mixed with fuel by the combustor 400 . After combustion, it becomes combustion gas, the combustion gas generated in the combustor 400 flows into the turbine 500 , and the combustion gas introduced into the turbine 500 drives the rotor through the turbine blade 700 . After being rotated, it is discharged to the atmosphere through the diffuser, and the rotor rotated by combustion gas may drive the compressor 200 and the generator. That is, some of the mechanical energy obtained from the turbine 500 is supplied as energy required to compress the air in the compressor 200 , and the rest may be used to generate electric power by the generator.

On the other hand, the turbine blade (hereinafter, referred to as 'blade 700') according to this embodiment, the turbine blade platform unit (hereinafter referred to as 'platform unit 710'), as shown in FIG. It includes a turbine blade root portion (hereinafter referred to as 'root portion 720') and the turbine blade airfoil portion (hereinafter referred to as 'airfoil portion 730').

Hereinafter, the radial direction means the radial direction of the rotor 600 when the blade 700 is mounted on the rotor 600, and the axial direction is defined as meaning the longitudinal direction of the rotor 600 rotation shaft. do. These radial and axial directions are shown in FIG. 2 .

The root portion 720 is coupled to the radially inner side of the platform portion 710, and the airfoil portion 730 is coupled to the radially outer side. The root portion 720 is coupled to the rotor 600 .

The platform unit 710 may be formed in a plate structure in which a plurality of layers are overlapped. Meanwhile, in the present embodiment, the platform unit 710 is formed in a rectangular shape, but otherwise, all or part of the side surface may be formed in a curved shape, such as a C-shape or an S-shape. Also, when the blade 700 is coupled to the rotor, a groove may be formed in a side surface of the platform unit 710 for bonding between the adjacent platform units 710 .

The root portion 720 engages with the platform portion 710 at the radially outer side of the root portion 720 , and the radially inner side of the root portion 720 protrudes to engage the turbine disk 630 . made in the shape of That is, the blade 700 is coupled to the rotor 600 by the root portion 720 . In this case, the root portion 720 may include a coating layer for protecting the root portion 720 from the high-temperature gas (H).

In addition, the root portion 720 should be designed to withstand centrifugal stress well during rotation of the rotor 600, for example, the outer surface is protruded to have a fir-tree shape to be formed to be coupled with the turbine disk 630. can

The airfoil part 730 has the turbine blade cooling passage (hereinafter, referred to as a 'blade cooling passage 732') formed therein. As described above, a cooling fluid flows in the blade cooling passage 732 to prevent damage to the airfoil portion 730 due to the high-temperature combustion gas. The outlet of the blade cooling passage 732 is formed in the blade tip 738 which is the radially outer end of the airfoil portion 730 . In this case, a plurality of cooling passage outlets may be formed according to the structure of the blade cooling passage 732 .

Between the turbine blade airfoil unit leading edge (hereinafter referred to as 'leading edge 734a') and the turbine blade airfoil unit trailing edge (hereinafter referred to as 'trailing edge 734b'), hot gas There is a pressure surface 736a in the direction in which (H) collides. In addition, there is a suction surface 736b in a direction opposite to the pressure surface 736a. At this time, for rotation of the blade 700, the pressure surface 736a is formed to be concave, and the suction surface 736b is formed to be convex.

As shown in FIG. 5B, a shroud (S) surrounding the blade 700 is positioned on the radially outer side of the airfoil part 730, and the turbine tip is between the blade tip 738 and the shroud (S). clearance is formed.

As shown in FIG. 2 , the blade tip 738 is a radially outer cross-section of the airfoil portion 730 , and is formed in the shape of an airfoil. An outlet of the blade cooling passage 732 is formed at the blade tip 738 . At this time, a plurality of outlets of the blade cooling passage 732 are formed, and in this case, a partition wall may be provided between the outlets.

In the case of a conventional blade, a tip cooling hole was formed in the blade tip for cooling the blade tip, or a tip hole was formed in the shroud (S) to control the flow rate of the cooling fluid. However, in the turbine blade 700 according to an embodiment of the present invention, a tip cooling hole is not formed in the blade tip 738 and the flow rate is controlled by the throttle plate, so that the blade cooling passage 732 in the blade tip 738 ) is formed immediately.

In this case, as shown in FIG. 5A , the hot gas H can penetrate into the blade cooling passage through the turbine tip clearance between the blade tip 738 and the shroud S. As a result, since there is a possibility that the blade cooling passage may be damaged due to the penetration of the high temperature gas H, damage may be prevented by applying an oxidation-resistant coating to the inner wall of the blade cooling passage. However, when the oxidation-resistant coating is applied, there is a problem in that there is a risk of cracking.

In order to solve the above problem, the turbine blade 700 according to an embodiment of the present invention extends from the blade tip 738 in the axial direction toward the blade cooling passage 732 as shown in FIG. 3 . A protrusion 739 is formed. The protrusion 739 is formed on the pressure surface 736a to effectively achieve the purpose of preventing the penetration of hot gas.

The shape and size of the protrusion 739 may be variously implemented as will be described later.

As shown in FIGS. 7A to 7C , according to other embodiments of the present invention, the protrusion 739 may have a polygonal cross-section such as a trapezoid, or may be formed in a shape in which one or more sides are curved. However, this is not limited to the embodiments of the present invention, as long as the protrusion 739 can function to prevent penetration of the hot gas H In addition to this, it may be implemented in various shapes.

Meanwhile, the protrusion 739 may be chamfered or filled in a manufacturing or processing process. For example, a corner having an obtuse angle in the protrusion 739 having a trapezoidal cross-section may be processed by filleting. During such machining, detailed dimensions may be determined in consideration of the effect of the chamfer or fillet on the flow of the cooling fluid.

In addition, the size of the protrusion 739 is formed differently from the leading edge 734a to the trailing edge 734b. In the present embodiment, as shown in FIGS. 6A and 6B , the protrusion 739 is formed to become larger as it approaches the leading edge 734a, and to become smaller as it approaches the trailing edge 734b. However, on the contrary, the size of the protrusion 739 may be formed to increase as it approaches the trailing edge 734b and to become smaller as it approaches the leading edge 734a. In addition, the size of the protrusion 739 may be formed to be uniform anywhere from the leading edge 734a to the trailing edge 734b.

As shown in FIG. 5B , in the case of the blade tip 738 having the protrusion 739 formed therein, the blade cooling passage 732 of the hot gas H flowing into the turbine tip clearance by the protrusion 739 is inside the penetration is blocked.

Meanwhile, the protrusion 739 is integrally formed with the airfoil portion 730 . That is, the airfoil part 730 may be manufactured by casting including the protrusion part 739 , and thus, separate processing such as welding and attaching the protrusion part 739 is unnecessary.

Since the hot gas H flows along the pressure surface 736a in the direction from the leading edge 734a of the airfoil portion 730 to the trailing edge 734b, the projection 739 is the blade It is formed in the direction of the pressure surface (736a) of the tip (738).

The turbine blade 700 according to an embodiment of the present invention may be mounted on a three-stage turbine in a gas turbine having a multi-stage turbine. The turbine is generally equipped with a maximum of four stages. As the number of stages increases, the hot gas (H), which expands and increases in volume, flows in, so the size of the turbine blade must be increased accordingly, and the cooling method of the turbine blade must be appropriately determined. In consideration of these points, the present embodiment has been described as being used in a three-stage turbine, but otherwise, it is also possible to be used in a turbine of another stage.

Hereinafter, the effect of the turbine blade according to the present embodiment will be described.

The turbine blade 700 cools the blade tip 738 by discharging the cooling fluid by forming an outlet of the turbine blade cooling passage at the blade tip 738 . At this time, a portion of the hot gas (H) flows into the turbine tip clearance between the blade tip (738) and the shroud (S). In particular, in the starting section of the gas turbine, as shown in FIG. 5A , a part of the hot gas (H) may penetrate to the blade cooling passage. In this case, there is a possibility that the blade cooling passage is damaged by the hot gas (H). Accordingly, an oxidation-resistant coating is required to prevent damage to the blade cooling passage, but if the oxidation-resistant coating is applied, there is a risk of cracks occurring on the inner surface of the blade cooling passage.

However, when the protrusion 739 is formed on the blade tip 738 , it is possible to prevent the high temperature gas H from penetrating into the blade cooling passage 732 . This is because the protrusion 739 is located on the flow path of the hot gas H as shown in FIG. 5B . Accordingly, there is little possibility that the inner surface of the blade cooling passage 732 is damaged. In addition, the oxidation-resistant coating on the inner surface of the blade cooling passage 732 is unnecessary, so that the manufacturing process and cost do not increase, and the risk of cracking due to the oxidation-resistant coating is eliminated.

In addition, since the protrusion 739 is not attached separately, but is manufactured integrally with the airfoil portion 730, there is little cost increase factor and there is no particular difficulty in manufacturing. That is, the turbine blade 700 according to an embodiment of the present invention can achieve the purpose of preventing the penetration of hot gas (H) by processing the protrusion 739 having a simple shape on the blade tip 738 . .

100: housing 110: compressor housing
120: combustor housing 130: turbine housing
200: compressor 210: compressor blade
220: compressor vane 400: combustor
500: turbine 520: turbine vane
600: rotor 610: compressor disc
620: torque tube 630: turbine disk
640: tie rod 650: fixing nut
700: turbine blade 710: platform unit
720: root portion 730: airfoil portion
732: blade cooling passage 734a: leading edge
734b: trailing edge 736a: pressure side
736b: suction surface 738: blade tip
739, 739', 739'': overhang
H: hot gas
S: shroud

Claims (14)

platform unit;
a root portion formed at an inner end of the platform portion in a radial direction; and
Including; an airfoil portion formed at the radially outer end of the platform portion,
The airfoil part,
A turbine blade cooling passage is formed therein, and a blade tip having an outlet of the turbine blade cooling passage formed at an outer end thereof in a radial direction, wherein a protrusion is formed in the blade tip,
The turbine blade cooling passage is in an open form that communicates with the outside of the airfoil portion by forming an outlet at the blade tip,
Cooling air flowing in a direction from the root portion toward the blade tip flows in the turbine blade cooling passage,
The airfoil part,
a leading edge facing upstream of the flow of hot gas;
a trailing edge facing the opposite direction of the leading edge;
a pressure surface formed on one side between the leading edge and the trailing edge; and
Between the leading edge and the trailing edge, a suction surface formed in a direction opposite to the pressure surface; includes,
The protrusion is formed in the direction of the pressure surface,
The lower surface of the protrusion is formed such that air flowing along the lower surface is in contact with at least an end point in the vicinity of the shroud on the inner surface of the blade tip facing the protrusion,
The upper surface of the protrusion is formed at the same height as the blade tip,
The turbine blade, characterized in that the high-temperature gas is prevented from flowing into the turbine blade cooling passage by the flow of air flowing along the lower surface and the upper surface of the protrusion while passing through the turbine blade cooling passage. The method of claim 1,
The turbine blade, characterized in that the projection extends in a direction perpendicular to the radial direction. 3. The method of claim 2,
The protrusion is
Turbine blade, characterized in that formed only in a part of the section from the leading edge to the trailing edge. 3. The method of claim 2,
The protrusion is
A turbine blade, characterized in that the axial cross section is formed in a polygonal shape. 3. The method of claim 2,
The protrusion is
A turbine blade, characterized in that the axial cross section is formed to include a curve. 5. The method of claim 4,
The turbine blade, characterized in that the projection is chamfered. 3. The method of claim 2,
The protrusion is
From the leading edge to the trailing edge, the turbine blade, characterized in that the protruding thickness is formed differently. 3. The method of claim 2,
The protrusion is
Turbine blade, characterized in that from the leading edge to the trailing edge, the axial cross-sectional area is formed differently. According to claim 1,
A turbine blade, characterized in that a plurality of cooling passage outlets are formed at the blade tip. 10. The method of claim 9,
The protrusion is a turbine blade, characterized in that formed only in some of the outlets of the plurality of cooling passages. 3. The method of claim 2,
The airfoil part,
A turbine blade, characterized in that a plurality of film cooling holes are formed on the pressure surface. The method of claim 1,
The turbine blade is a turbine blade, characterized in that mounted on the three-stage turbine. disk;
Comprising a plurality of turbine blades radially mounted on the outer circumferential surface of the disk,
13. A turbine disk, characterized in that the turbine blades are according to any one of claims 1 to 12. housing;
a compressor for compressing the air introduced into the housing;
a combustor that mixes fuel with the air compressed in the compressor and ignites to generate combustion gas; and
A turbine for rotating the compressor by obtaining rotational force from the combustion gas generated from the combustor;
A gas turbine, characterized in that the turbine comprises a turbine disk according to claim 13 .

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