JPH0953407A - Gas turbine moving blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To improve the rigidity of a cavity section and the cooling efficiency of a hub section by providing the cavity in the hub section of a moving blade and the inside of a blade root section, feeding cooling air into it, protruding pin fins on the inner wall of the cavity, or integrally forming pillar-like fins connected with both ends to the opposite inner walls. SOLUTION: A long and thin moving blade 1 is provided with a cavity 4 having core support ribs 14 in the hub section 18 of a blade 12 up to 25% of the blade axial length toward a blade end 16 from a blade root section 19 and a hub 11, and multiple holes 15 aligned in the axial length direction of the blade 12 are bored in the cavity 4 from the outer periphery section to the blade end 16 to feed cooling air. Many pin fins 5 protruded from the inner wall are integrally formed in the aligned state in the blade axial direction of the cavity 4. The rigidity of the thin hub section 18 can be increased, the cooling efficiency of the hub section 18 can be improved, and the setting of a core can be facilitated when the moving blade 1 is manufactured.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine blade provided with a cooling structure inside, which is applied to a thin large-sized rotor blade provided on a rear side of a gas turbine blade row.

[0002]

2. Description of the Related Art With the high temperature and high output of gas turbines,
The lengthening of the moving blades in the rear stage of the gas turbine cascade has become remarkable. In such a lengthened moving blade, the weight of the moving blade itself increases and the peripheral speed also increases.Therefore, the stress generated in the moving blade due to the centrifugal force during rotation of the moving blade is It is much higher than
Therefore, in such a moving blade, the thickness of the blade cross section is thinned from the hub portion, which is the base of the blade adjacent to the blade root portion, to the blade tip to make the blade thin and lightweight, and the width of the blade is reduced. Tapered blades that have become smaller toward the tip end have come into use.

Further, in such a lengthened rotor blade,
The shroud provided at the blade tip of the moving blade is an integral shroud blade (ISB) integrally formed with the blade, and the weight on the blade tip side, which is greatly affected by the centrifugal force, is reduced and adjacent to it. Shrouds are also used to suppress blade vibrations to improve vibration strength. Further, in such a large and thin rotor blade used in the rear stage of such a gas turbine blade cascade, in addition to the aforementioned thinning and tapering in order to avoid an increase in stress due to centrifugal force, a gas turbine As a result of the higher temperature and higher power output, the strength of the moving blades deteriorates and the strength decreases.

FIG. 4 shows a cross section of the moving blade 10 at a central portion in the blade thickness direction in which a cooling structure is provided inside the moving blade 10 in order to avoid such a decrease in the strength of the moving blade due to the high temperature. It is a figure. As shown in the drawing, the blade root portion 19 of the moving blade 10 and the hub 11 which is the boundary between the blade 12 and the blade root portion 19 to the blade tip 1
A cavity 13 is formed by using a ceramic core as a core at the time of casting inside the blade 12 in the six directions up to 25% of the blade axial length (hereinafter referred to as a hub portion 18 for convenience of description).

Further, a plurality of multi-holes 15 are bored in the blade axial direction from the outer peripheral portion of the cavity 13 to the blade tip 16, and the blade 12 and the hub portion 1 are convectively cooled.
In order to cool 8 and the root portion 16 of the blade, cooling air was sent from a turbine rotor (not shown) as shown by an arrow, passed through the cavity 13 and the multi-hole 15, and provided on the blade tip 16 or the shroud 17. The gas is discharged from the opening into the turbine passage. In addition, 14 is a moving blade 1.
At the time of casting 0, a ceramic core used as a core is formed in the hub portion 18 in order to form the cavity 13.
It is a core support rib provided to support the forming portion.

However, in the moving blade 10 provided with the cooling structure inside in this way, the cavity 13 is produced at the time of manufacture.
There is a problem in that it is difficult to manufacture the core for forming the core and it is difficult to set the core inside the moving blade 10 in which the cavity 13 is provided. Further, in a moving blade used in a gas turbine whose turbine inlet temperature becomes 1500 ° C. due to high temperature and high pressure performed to improve gas turbine efficiency, a cavity 13 is provided in the hub portion 18 described above, and the inside thereof is provided. There is a problem that the cooling is insufficient and the creep strength becomes a problem only by cooling with the cooling air introduced into the.

[0007]

SUMMARY OF THE INVENTION In order to solve the problems of the conventional large and thin gas turbine moving blades as described above, the present invention provides a core for forming a cavity inside the moving blades. It is easy to manufacture, especially the core in consideration of setting, and the core of the ceramic core can be easily set in the cavity, making it easier to reduce the weight of the rotor blade and to form the cavity for cooling. In addition, the rigidity of the cavity, especially the torsional rigidity is improved, the cooling efficiency of the hub is significantly improved, the problem of creep strength is solved, and the gas that can be used for a gas turbine with a higher inlet temperature is also used. An object is to provide a turbine rotor blade.

[0008]

For this reason, the gas turbine moving blade of the present invention has the following means. In order to cool the rotor blades from the inside, pin fins are provided so as to protrude from the cavity inner wall in the cavity provided inside the hub portion of the rotor blade and the blade root portion, or the columnar fins have cavity inner walls whose opposite ends face each other. It was arranged to be connected to each. In addition,
The pin fins or the columnar fins are preferably provided at least in the cavity provided inside the hub portion of the moving blade.

As a result, in the gas turbine rotor blade of the present invention, a cavity for cooling is formed to improve the strength of the hub portion and the blade root portion where the strength becomes critical, and particularly, the cooling of the hub portion is performed. By installing pin fins or columnar fins, the efficiency is remarkably improved, the temperature rise in this part can be reduced, and the creep strength is increased, so that it can be applied to a higher temperature gas turbine and the creep life can be extended. In addition, even if the blade is thinned by providing pin fins or columnar fins,
Fabrication of a ceramic core (core) to form a cavity,
And the setting becomes easy. Further, the pin fin or the columnar fin serves as a structural material particularly in the hub portion, which is significantly thinned by providing the cavity,
The strength of this portion rises, and the rigidity against torsion, which is a characteristic of the ISB blade, is improved.

[0010]

BEST MODE FOR CARRYING OUT THE INVENTION An embodiment of a gas turbine rotor blade of the present invention will be described below with reference to the drawings. FIG.
FIG. 1 is a cross-sectional view at a blade thickness direction central portion of a moving blade showing an embodiment of a gas turbine moving blade of the present invention. In the drawings showing the present embodiment, the same reference numerals as those shown in FIG. 3 are the same as those explained in FIG. 3, and their explanations are omitted.

As shown in the figure, a large and thin moving blade 1
A core support rib 14 is provided inside the blade root portion 19 of the blade and the hub portion 18 of the blade 12 up to 25% of the blade axial length in the direction of the blade tip 16 from the hub 11 which is the boundary between the blade root portion 19 and the blade 12. A cavity 4 is provided. A pin fin 5, which will be described later, is provided in this cavity 4. 2 is a cross-sectional view taken along the line AA of FIG. 1 arranged in the axial length direction of the blade 12 inside the cavity 4 from the outer peripheral portion to the blade tip 16.
A plurality of multi-holes 15 as shown in (A) are formed in the same manner as the above-mentioned conventional example.

2 is a partial sectional view taken along the line BB in FIG. 1 showing the first preferred embodiment of the present invention.
As shown in (B), the pin fin 5 is provided so as to project from the inner wall forming the cavity 4. This pin fin 5
Have a pin diameter of 2 mm and a pitch of 8 to 10 mm, and are provided in 11 rows in the blade axis direction of the cavity 4. As a result, the hollow 4 is provided, and the rigidity of the thinned hub portion 18 is increased. At the same time, the hollow portion 4 is introduced into the hollow 4 from a passage provided in a turbine rotor (not shown), passes through the multi-hole 15, and passes through the blade tip 1
The cooling efficiency of the hub portion 18 due to the cooling air flowing to the 6 side is improved, and the creep strength is increased.

Further, since the pin fins 5 are provided so as to project inside the cavity 4, when the rotor blade 1 is manufactured, a ceramic, which is a core installed inside the rotor blade 1 to form the cavity 4, is formed. The core is supported only by the core supporting ribs 14, but the projecting portion of the pin fin 5 can support the core, which facilitates the manufacture of the core and facilitates the setting of the core. .

Further, the pin fin 5 may be replaced by a columnar fin 6 which is connected from the abdominal side to the dorsal side in the cavity 4, as shown in FIG. In the case where the columnar fins 6 are provided, the columnar fins 6 are provided in the hub portion as compared with the conventional hub portion 18 only having the cavity 13 and the pin fins 5 shown in FIG. It also has the effect of increasing the rigidity of 18.

The tip 1 of the gas turbine rotor blade of this embodiment is
The shroud 17 provided in 6 is more preferably an air-cooled shroud adapted to utilize the air from the multi-hole 15. That is, a shroud 17 as shown in FIG. 3A which is a plan view is fixed to a blade tip 16 which is an outer diameter end of the blade 12 shown in FIG. Reduced efficiency, rotor blade 1 during rotation
This shroud 1 is designed to reduce the vibration of
Further effects can be obtained by adopting the air-cooled shroud in which air passages connected to the multi-hole 15 are provided in the circumferential direction in the gas turbine rotor blade of the present embodiment.

[0016]

As described above, according to the gas turbine rotor blade of the present invention, the structure shown in the scope of claims
The following effects are obtained.

(1) Installed in the latter stage of the gas turbine,
The cooling efficiency of large and thin blades can be increased, and the creep life is improved.

(2) Further, by providing the pin fins, the manufacture and setting of the ceramic core can be facilitated even with a thin moving blade, and the rigidity of the moving blade can be increased. As described above, it is possible to enhance the reliability of the moving blade and the effect of dealing with high temperature and high output.

[Brief description of drawings]

FIG. 1 is a cross-sectional view of a long and thin moving blade at a central portion in a blade thickness direction showing an embodiment of a gas turbine moving blade of the present invention,

2 is a diagram showing the inside of the moving blade shown in FIG. 1, and FIG.
1 is a sectional view taken along the line AA of FIG. 1, and FIG. 2B is a view B of FIG.
-B sectional view, FIG. 2C is a sectional view taken along the line BB of FIG. 1 showing another embodiment different from FIG. 2B,

FIG. 3 is a diagram showing a shroud installed at the wing tip of FIG.
3A is a plan view taken along the line CC of FIG. 1, and FIG. 3B is a sectional view showing a cooling air passage provided so as to be connected to the blade and the shroud.

4 is a cross-sectional view of a conventional gas turbine rotor blade, and FIG.
FIG. 4A is a cross-sectional view of the rotor blade in the blade thickness direction central portion, and FIG.
4 is a cross-sectional view taken along the line DD of FIG.

[Explanation of symbols]

 4,13 Cavity 5 Pin fin 6 Columnar fin 10 Moving blade 11 Hub 12 Blade 14 Core support rib 15 Multi-hole 16 Blade tip 17 Shroud 18 Hub portion 19 Blade root portion

Claims (1)

[Claims] 1. A cavity is provided inside a hub portion and a blade root portion of a moving blade, and the cavity is communicated with the inside of the blade on the blade tip side of the hub portion and an opening provided at the blade tip portion, and the cavity is formed. In the gas turbine blade having a multi-hole for allowing cooling air to pass from the opening to the opening, pin fins protruding from the inner wall or columnar fins having opposite ends connected to opposite inner walls are provided in the cavity. And a gas turbine rotor blade.