JP2001107701A - Gas turbine moving blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine moving blade with improved reliability by improving cooling performance at a front portion which receives extremely high load at a gas surface and is hardly cooled. SOLUTION: Cooling passages 8 through which cooling medium 7 flows are formed inside the moving blade 2. Heat transfer promoting projections 17 are formed on a cooing surface of an inversion part 11 of each cooling passage 8 at a front of the moving blade 2. Cooling performance at the front of the moving blade 2 is thus improved.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine rotor blade in which a cooling passage for flowing a cooling medium is formed.

[0002]

2. Description of the Related Art FIG. 6 shows a structure of a turbine section and a cooling air system in a conventional gas turbine. As can be seen in the figure, the turbine section includes a rotating section composed of a rotor body 1 and a turbine rotor blade 2, a casing 3, a stationary blade 4 and a holding component 5.
And the like. Reference numeral 6 denotes a combustor. The high-temperature, high-pressure combustion gas from the combustor 6 is converted into a high-speed flow by the stationary blade 4, and the high-speed flow rotates the moving blade 2 to generate power.

[0003] Each component of the rotating part and the stationary part close to the combustion gas must be cooled by heat input from the combustion gas so that the temperature does not rise above an allowable temperature. In many cases, a cooling medium 7 indicated by an arrow flows therein. As the cooling medium 7, air extracted from a compressor or discharged air is often used. In some cases, the fluid may be once introduced into a cooler and cooled to an appropriate temperature before use. Further, steam or the like from outside the system may be applied to the fluid.

[0004] The cooling medium 7 flowing through the rotating body passes through the inner surface of the rotor main body 1, cools the inside of the turbine blade 2, and then flows along a path joining the high-temperature gas passage. In some cases, the cooling medium 7 that has been heat-exchanged by the turbine blades 2 or the like may be recovered, and the heat energy may be used outside the system.

FIG. 7 is a cross-sectional view showing a main part of a conventional cooling structure of the turbine moving blade 2. In the front stage of the turbine moving blade 2 which is particularly exposed to a high-temperature gas, a cooling medium is required to withstand a high heat load. In many cases, a cooling passage 8 is provided to pass through the inside 7 to cool the inside by convection. The cooling passage 8 is often formed of a serpentine flow path that repeats reciprocating several times from the design request, and the flow path is formed by a reversing portion 11 provided near the turbine blade tip 9 and the turbine blade root 10. Is inverted.

The cooling medium 7 cools the turbine blade 2 from the inner surface while passing through the cooling passage 8. At this time, in the case of a turbine blade having a higher heat load, a film cooling hole 12 is provided on the blade surface, and a part of the cooling medium is blown out to the gas flow path side along the blade surface, as if by a low-temperature curtain. In some cases, film cooling is performed so as to cover the wing surface and reduce the heat load from the outer surface.

The turbine rotor blade tip 9 is a region where cooling is difficult to be performed. The cooling is mainly performed by increasing the heat transfer coefficient by the inner surface cooling in the reversing portion 11 or the film cooling hole 1.
2, a method of blowing out the film cooling air 13 to form a region having a relatively low temperature around the turbine blade 2 to reduce the heat load has been adopted. However, the heat load on the gas surface of the turbine rotor blade tip 9 is extremely high, exceeding the allowable design temperature, and there is a concern that the reliability of the turbine rotor blade may be impaired.

Further, as shown in FIGS. 8 and 9, the tip end portion 9 of the turbine rotor blade has a reduced gap 19 so as to reduce leakage from the gap with the components of the stationary system 18 in order to ensure the performance of the gas turbine. In many cases, leakage air is reduced by devising the shape of the turbine blade tip 9. As a method of reducing the leakage air in terms of shape, a turbine blade tip 9
Has a frame-shaped protrusion 20 on both end surfaces or one end surface to reduce pressure leak and reduce leaked air, while reducing the damage to the moving blade when the moving blade comes into contact with the stationary system 18. (Figure 9). However, in this cooling means, there is no means for convection cooling the projection 20 at the tip.
As an advanced cooling method, there is film cooling. However, even if the film cooling is applied to the blade tip, the film cooling medium 21 repeats rapid contraction and rapid expansion, and the film cooling medium 21 is immediately diffused. Cooling does not work effectively. For these reasons, the above-mentioned protruding tip portion is often burned by the high-temperature gas, which is one of the problems of the blade cooling technology of the high-temperature gas turbine.

[0009]

In view of the above-mentioned problems of the conventional gas turbine rotor blade, the present invention provides a gas turbine rotor blade having a cooling passage formed therein. It is an object of the present invention to provide a gas turbine rotor blade with an extremely high load, improved cooling performance at a tip portion where cooling is difficult to perform, and improved reliability.

[0010]

SUMMARY OF THE INVENTION In order to solve the above-mentioned problems, the present invention relates to a gas turbine rotor blade having a cooling passage formed therein. Provided is a gas turbine rotor blade having a surface formed with a projection for promoting heat transfer.

In the gas turbine rotor blade of the present invention, as the projection for promoting heat transfer formed on the cooling surface of the inverted portion of the cooling passage, one having an appropriate shape such as an orthogonal fin, a pin-shaped projection, or a dimple is employed. can do.

According to the gas turbine rotor blade of the present invention, the cooling surface of the reversal part of the cooling passage at the tip of the turbine rotor blade is:
The protrusion for promoting heat transfer provided thereon acts as a radiating fin to increase the amount of heat absorption, thereby sufficiently cooling the heat entering the blade tip from the gas side.

In the gas turbine rotor blade of the present invention, in order to form the above-mentioned heat transfer promoting projection on the cooling surface of the inverted portion of the cooling passage at the tip of the rotor blade, the core of the gas turbine rotor blade is cast at the time of casting. The plug can be easily manufactured by forming a projection on the inner surface of the plug used to close the hole at the tip of the gas turbine rotor blade used for support. When the projections are formed on the inner surface of the plug in this manner, the metal temperature of the plug itself is reduced, so that the joint formed by brazing or welding is well protected from high temperatures.

In order to solve the above-mentioned problems of the present invention, the present invention further provides a gas turbine blade having a radius R of 0.2 to 2t, where t is a blade thickness at at least one side of the blade tip. Provided is a gas turbine rotor blade having a multiple arc (curved portion) and suppressing high temperature gas separation and high turbulence at the tip of the rotor blade.

By forming the curved portion at the blade tip in this way, regardless of the magnitude of the tip clearance at the blade tip, the high-temperature gas due to the pressure difference generated from the blade abdominal side to the back side is obtained. Since the leakage of the leakage can be prevented, the heat transfer from the high-temperature gas side to the blade tip can be suppressed.

When this curved portion (R or multiple arcs) is provided on the ventral side of the rotor blade, high heat transfer due to the rapid contraction of the gas flow path generated near the tip of the rotor blade ventral side is suppressed. When the portion is provided on the back side of the moving blade, it is possible to suppress the high heat transfer generated due to the separation vortex of the high-temperature gas due to the rapid expansion of the gas flow path generated near the front end of the moving blade.

As described above, according to the present invention, at least one side of the blade tip is provided with an R or multiple arc (curved portion) of 0.2 to 2 times the blade thickness t of the blade tip. Separation and turbulence of the high-temperature gas at the blade tip can be suppressed, heat transfer from the high-temperature gas to the blade tip can be suppressed, and the reliability of the gas turbine blade can be increased.

In the gas turbine rotor blade of the present invention, when a film cooling hole is formed at the tip of the gas turbine rotor blade and cooling air is blown at the tip of the rotor blade or the top of the rotor blade, as described above, Since the separation of the leakage flow is prevented, the diffusion of the film cooling medium can be prevented, and the film can be cooled more effectively, which is preferable.

Further, in the gas turbine rotor blade according to the present invention, when the protrusion for promoting heat transfer on the cooling surface of the cooling passage reversal portion and the formation of the curved portion are performed together, the rotor blade is removed from the high-temperature gas. While suppressing heat transfer to the tip,
Cooling from the inside of the moving blade tip is promoted, and a more reliable gas turbine moving blade can be provided.

[0020]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Hereinafter, a gas turbine rotor blade according to the present invention will be described in detail with reference to the embodiments shown in FIGS. In the following embodiment, FIGS.
Parts that are substantially the same as those of the conventional turbine blade shown in FIG. 9 are denoted by the same reference numerals for simplicity of description, and redundant description thereof will be omitted.

(First Embodiment) First, a gas turbine blade according to a first embodiment shown in FIG. 1 will be described. FIG.
In the figure, reference numeral 17 denotes a heat transfer promoting projection formed intentionally on the cooling surface of the reversing portion 11 of the cooling passage. The projections 17 may have an appropriate shape, such as orthogonal fins provided in parallel with the surface of the high-temperature gas flow so as to be orthogonal to the surface, a large number of pin-shaped projections having a circular cross section, and dimple-shaped projections. It is preferable to adopt a shape that promotes turbulence with respect to the cooling medium flowing through the cooling passage.

As described above, in the gas turbine blade of the first embodiment, in which the projection 17 for promoting heat transfer is formed on the cooling surface of the reversing portion 11 of the cooling passage 8 at the turbine blade tip 9, The cooling surface of the portion 11 performs forced convection heat transfer with the cooling medium 7 and the reversing portion 11 is cooled by heat input from the gas side. As a result, the projections 17 act as radiation fins, and increase the amount of heat absorbed.

(Second Embodiment) Next, a gas turbine blade according to a second embodiment shown in FIGS. 2 and 3 will be described. 2 and 3, reference numeral 14 denotes a hole used to support the core during blade casting. The actual cooling passage of the turbine blade 2 is often molded by a technique such as precision casting, and after the molding, the holes 14 supporting the core at the time of casting are formed.
Are often present in the vicinity of the cooling passage reversal portion 11 as shown in FIG.

The support hole 14 is often closed by joining a plug 15 to the turbine blade 2 by welding, brazing, or the like in the process of processing the blade. This second
In the gas turbine blade according to the embodiment, the plug 15
A projection 17 is provided on the cooling surface 16 to promote heat transfer to enhance the cooling of the blade tip 9.

FIG. 3 shows an example of the projection 17 formed on the cooling surface 16 side of the plug 15. FIG. 3A shows a configuration in which a number of fins 17-1 in which a number of fins are juxtaposed in a direction perpendicular to the flow of the cooling medium are provided, and FIG. 3B shows a number of pin-shaped projections 17-2 having a circular cross section. (C) shows a plurality of dimple-shaped projections 17-3 formed therein. In each of (a) to (c), the upper part is a side view and the lower part is a plan view.

FIG. 3 shows three types of projections (a) to (c). Regarding the shape of the projections 17, it is expected that a simple fin group capable of improving the heat radiation efficiency will promote turbulence. Various shapes such as possible shapes may be appropriately adopted.

In the gas turbine rotor blade of the second embodiment, the projection 17 is attached to the cooling surface of the plug 15 that closes the hole 14 at the tip of the turbine rotor blade used for supporting the core during casting. By acting as a radiation fin to increase the amount of heat absorbed, it is possible to reduce the metal temperature of the plug 15 itself,
Cooling due to the heat conduction effect can also be expected in the vicinity such as brazing and welding joints for attaching the solder.

(Third Embodiment) Next, a gas turbine rotor blade according to a third embodiment shown in FIGS. 4 and 5 will be described. 4 and 5, reference numeral 22 denotes a rotor blade root, and reference numeral 23 denotes a blade platform. Reference numeral 24 denotes a film cooling hole formed at the tip of the blade.

The cooling medium 7 for cooling the turbine blade is supplied from a gas turbine compressor or from the outside, and is supplied from the turbine blade root portion 22 or the like by various methods. Among the turbine rotor blades, a cooling passage 8 for a cooling medium 7 is provided inside a turbine blade that needs to be cooled according to various design conditions. The cooling medium 7 cools the blade surface of the turbine blade and, in some cases, the platform 23, and uses a part of the cooling medium for cooling the blade tip 9. Blade tip 9
Film cooling holes 2 as shown
4 or the like may be applied.

Among the turbine blade tips 9 having a high heat load,
As shown by reference numeral 30 in the drawing, R or multiple arcs (curved portions) of 0.2 to 2 t with respect to the blade thickness t of the blade tip are provided at both ends or one of them. By forming the curved portion in this way, regardless of the size of the gap (tip clearance) 19, the leakage flow 29 of the high-temperature gas 28 due to the pressure difference generated from the ventral side 26 to the dorsal side 27 of the moving blade 2
Can be prevented from peeling off. As a result, heat transfer from the gas side to the tip of the blade is suppressed, and overheating of the blade tip 9 can be prevented.

When the curved portion 30 is applied to the abdominal side 26 of the moving blade, high heat transfer due to rapid contraction occurring near the moving blade tip 9 on the abdominal side 26 is provided. When applied, it is possible to suppress the high heat transfer from the high-temperature gas due to the separation vortex caused by the rapid expansion. Simultaneously, the separation of the leakage flow 29 is prevented by the film cooling air holes 24 blown out near the blade tip or at the blade top.
Diffusion of the film cooling medium 21 from the
Higher film cooling efficiency can be expected.

In the gas turbine rotor blade according to the third embodiment, a cooling medium 7 (cooling air) provided for the purpose of cooling cools the cooling passage 8 inside the rotor blade by various methods (serpentine convection cooling, impingement cooling, multi-cooling). The cooling medium 7 is cooled by hole cooling, pin fin cooling, or the like, and a part or all of the cooling medium 7 is supplied to the turbine blade tip 9.

On the other hand, the gas side surface separates high temperature gas from the leakage flow 29 mainly due to the pressure difference between the ventral side 26 and the back side 27 of the blade surface near the top of the blade.
In order to suppress high turbulence, the blade thickness t of the blade tip 9 is set to 0.
A curved portion 30 having a radius of 2 to 2 times or a multiple arc of the same degree is formed on one or both of the abdominal surface 26 and the dorsal surface 27 of the blade tip.

The curved portion 30 can suppress separation of the leakage flow 29 of the high-temperature gas, and can prevent separation of the high-temperature gas flow in the vicinity of the blade tip 9 (in the vicinity of the abdomen, at the blade tip,
The heat transfer from the high-temperature gas (near the back side) to the blade tip is suppressed, making it easier to design the blade cooling. It is possible to reduce the temperature up to.

As described above, the present invention has been specifically described based on the illustrated embodiments. However, the present invention is not limited to these embodiments, and specific structures within the scope of the present invention shown in the claims are set forth. Needless to say, various changes may be made to the configuration.

For example, in the above-described embodiment, the formation of the protrusion for promoting heat transfer on the cooling surface of the inverted portion of the cooling passage at the tip of the blade, the formation of the protrusion on both surfaces of the plug closing the hole at the tip of the blade, and R or multiple arcs are separately formed on the tip of the moving blade, but by combining and forming two or more of these, the cooling performance at the tip of the moving blade is improved to improve the reliability. It can be a turbine blade.

[0037]

As described above, according to the present invention, in a gas turbine rotor blade having a cooling passage formed therein, heat transfer is promoted to a cooling surface of a reversal portion of the cooling passage at a tip end portion of the gas turbine rotor blade. And a gas turbine rotor blade provided with a projection having an appropriate shape such as an orthogonal fin, a pin-shaped projection, and a dimple shape.

According to the gas turbine rotor blade of the present invention, the cooling surface of the reversal portion of the cooling passage at the tip of the turbine rotor blade is:
The protrusion for promoting heat transfer provided thereon acts as a radiating fin to increase the amount of heat absorption, thereby sufficiently cooling the heat entering the blade tip from the gas side.

In the gas turbine rotor blade of the present invention, the projection is formed on the inner surface of a plug used to close the hole at the tip of the gas turbine rotor blade used for supporting the core during casting of the gas turbine rotor blade. In such a case, the metal temperature of the plug itself is reduced, so that the joint by brazing or welding is well protected from high temperatures.

According to the present invention, the blade thickness t of the blade tip is preferably 0.2 to 2 t on at least one side of the blade tip of the gas turbine.
R or multiple arcs (curved portions) are provided to provide a gas turbine blade in which high temperature gas separation and high turbulence at the tip of the blade are suppressed.

As described above, according to the present invention, at least one side of the blade tip is provided with an R or multiple arc (curved portion) having a blade thickness t of 0.2 to 2 t at the blade tip. It is possible to suppress the separation and turbulence of the high-temperature gas in the section, suppress the heat transfer from the high-temperature gas to the tip of the blade, and improve the reliability of the gas turbine blade.

Further, in the gas turbine rotor blade configured as described above, a film cooling hole is formed at the tip of the gas turbine rotor blade so that cooling air is blown to the rotor blade tip or the rotor blade top. As described above, since the leakage flow is prevented from being separated and the diffusion of the film cooling medium is prevented, the film can be cooled more effectively.

Further, the present invention provides a gas turbine blade in which the provision of the protrusion for promoting heat transfer on the cooling surface of the cooling passage reversal portion and the formation of the curved portion are performed. According to this, while suppressing the heat transfer from the high-temperature gas to the blade tip, cooling from inside the blade tip is promoted,
A gas turbine blade with even higher reliability is provided.

[Brief description of the drawings]

FIG. 1 is a cross-sectional view showing a reversal portion of a cooling passage at a tip portion of a gas turbine blade according to a first embodiment of the present invention.

FIG. 2 is a view showing a reversal part and a plug of a cooling passage in a gas turbine blade according to a second embodiment of the present invention;
(A) is a sectional view, (b) is a perspective view of the plug, (c)
FIG. 2A is a cross-sectional view along the line aa in FIG.

FIG. 3 is a view showing examples (a) to (c) of plugs used for a gas turbine blade according to a second embodiment of the present invention.

FIG. 4 is a perspective view of a gas turbine bucket according to a third embodiment of the present invention.

FIG. 5 is an enlarged sectional view taken along line BB of FIG. 4;

FIG. 6 is an explanatory diagram showing a structure of a rotor blade portion and a flow of a cooling medium in a conventional gas turbine.

FIG. 7 is a partially cutaway side view showing a structure of a conventional cooling passage of a gas turbine rotor blade.

FIG. 8 is a perspective view showing another example of the structure of a conventional gas turbine blade.

FIG. 9 is a sectional view taken along the line AA of FIG. 8;

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Rotor main body 2 Turbine rotor blade 3 Casing 4 Stator blade 5 Holding component 6 Combustor 7 Cooling medium 8 Cooling passage 9 Turbine rotor blade tip part 10 Turbine rotor blade root part 11 Inversion part 12 Film cooling hole 13 Film cooling air 14 Core support Hole 15 Plug 16 Cooling surface of plug 17 Projection for promoting heat transfer 18 Stationary system 19 Gap 20 Projection 21 Film cooling medium 22 Root of blade 23 Platform 24 Film cooling hole 26 Ventral side 27 Back side 28 Hot gas 29 Leakage flow 30 Curve (R or multiple arcs) t Blade thickness at tip

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor Kenichiro Takeishi 2-1-1 Shinama, Araimachi, Takasago City, Hyogo Prefecture Inside the Takasago Research Laboratory, Mitsubishi Heavy Industries, Ltd. (72) Inventor Yasushi Tomita 2-1, Shinama, Araimachi, Takasago City, Hyogo Prefecture No. 1 Mitsubishi Heavy Industries, Ltd. Takasago Works (72) Inventor Kiyoshi Suenaga 2-1-1, Araimachi Shinhama, Takasago City, Hyogo Prefecture Mitsubishi Heavy Industries, Ltd. Takasago Works F-term (reference) 3G002 CA05 CA06 CA08 CB05

Claims (6)

[Claims] 1. A gas turbine rotor blade having a cooling passage formed therein, wherein a projection for promoting heat transfer is formed on a cooling surface of a reversal portion of the cooling passage at a tip end portion of the gas turbine rotor blade. Gas turbine blades. 2. The gas turbine rotor blade according to claim 1, wherein the projection is formed on an inner surface of a plug used to support a core at the time of casting of the gas turbine rotor blade, the plug closing a hole at a tip end portion of the gas turbine rotor blade. Gas turbine blades. 3. A curved portion having a diameter of 0.2 to 2 times the blade thickness t of the moving blade tip portion is provided on at least one side of the tip portion of the gas turbine moving blade, so that the high temperature gas is separated at the moving blade tip portion. A gas turbine rotor blade characterized by suppressing turbulence. 4. The gas turbine moving blade according to claim 3, wherein a film cooling hole is formed at a tip portion of the gas turbine moving blade. 5. The gas turbine moving blade according to claim 1 or 2, wherein the curved portion according to claim 3 is provided at a tip end thereof. 6. The gas turbine blade according to claim 1, wherein the projection is at least one of an orthogonal fin, a pin-shaped projection, and a dimple shape.