JP2001271603A - Gas turbine moving blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To improve the cooling effect of a platform by improving strength by reducing the stress of a rear edge hub part, tip tinning, and a fillet R, concerning a gas turbine moving blade. SOLUTION: A serpentine passage, composed of a cooling passage 5a in a front edge part, cooling passages 6a, 6b and 6c in which cooling air flows from the center to the front edge side, and cooling passages 7a, 7b and 7c in which cooling air flows to the rear edge side is formed on a single stage moving blade 1 to cool a blade. Fillets R4a, 4b' on the front and rear edge sides are made larger than conventional fillets or the other fillets R for improving the strength of both ends, and recessed parts 3a, 3b are provided on the front and rear edge sides of a platform 2. In the platform 2, cooling passages are provided on both sides on the back side and the front side to allow cooling air to flow from the front edge to the rear edge side, and many cooling holes are provided on the side surface on the back side to allow air to flow from the cooling passages. The fillet R is made large, the strength is improved by the recessed parts 4a, 4b, and the cooling effect of the platform is 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 blade, in which the cooling structure of blades and platforms is improved so as to avoid cracks and damage due to high-temperature combustion gas, thereby improving the cooling effect. It is.

[0002]

2. Description of the Related Art FIG. 8 is a typical sectional view of a one-stage moving blade of a conventional gas turbine. In the figure, 60 is a moving blade, 61 is a blade root, and 62 is a platform. Cooling passages 63, 64, 65, and 66 are provided independently of each other at the blade root portion. The cooling passage 63 is a passage on the leading edge side, communicates with a cooling passage 63a in the blade, and cools air 40 from the rotor side. It flows in and flows out of the cooling hole 69 while cooling the front edge from the cooling passage 63a, and the front edge is cooled by the shower head film. The cooling air 41 flows from the cooling passage 64, passes through the cooling passage 64a in the blade,
, Flows toward the base from 64b, flows into 64c, flows out from the tip, and in this process flows out from the cooling hole to the blade surface as described later with reference to FIG.

The cooling air 42 flows from the cooling passage 65, and the cooling air 43 flows from the passage 66, merges and flows through the cooling passage 65a, enters the passage 65b from the distal end, flows toward the base, and flows into the 65c. In the course of flowing through the passage 65c, as described later with reference to FIG. 9, the film flows out from the cooling hole to the surface to perform film cooling, and the remaining air flows out from the cooling hole 68 in the trailing edge 67 to perform pin fin cooling.

FIG. 9 is a sectional view taken along line EE in FIG.
As shown in the drawing, a part of the cooling air in the cooling passage 63a at the leading edge flows out of the blade through the cooling hole 69, performs showerhead film cooling, and cools the leading edge. Also, cooling passage 6
Part of the cooling air flowing through the inside of the cooling hole 4c obliquely flows out of the cooling hole 70 in the flowing process to cool the surface of the film, and similarly in the process of flowing through the passage 65c.
From the wing surface, and the trailing edge is film-cooled. In the example shown, the cooling holes are 69, 70,
Although only 71 locations are shown, a number of cooling holes are provided in addition to this.

FIGS. 10A and 10B are plan views of a conventional platform, in which FIG. 10A shows both sides of the platform, and FIG. 9A, passages 50 a and 50 b communicating with the cooling passage 63 at the leading edge of the moving blade 60 are formed in both sides of the platform 62.
a and 72b pass through the passages 50a and 50b, flow on both sides of the platform 62, cool both sides, and flow out to the trailing edge side as 72c and 72d.

(B) shows another type of platform cooling system, in which a plurality of cooling holes 51a and 51b opening on the upper surface of the platform 62 are provided in addition to the passages 50a and 50b of (a). ing. Although not shown, these cooling holes communicate with one of the cooling passages communicating with the blades, allow cooling air to flow, flow out to the upper surface of the platform 62, and cool the ventral and rear sides of the upper surface, respectively. I have to. In the moving blade of the gas turbine, as shown in FIGS. 10A and 10B, by cooling the moving blade 60 and also cooling the platform 62, the influence of heat by the high-temperature combustion gas is reduced.

FIGS. 11A and 11B show an example of a conventional two-stage rotor blade. FIG. 11A is a sectional view, FIG. 11B is a sectional view taken along line FF in FIG. It is -G sectional drawing. 11A and 11B, reference numeral 80 denotes a two-stage bucket.
81 is a blade root part and 82 is a platform. Cooling passages 83, 84, 85 are provided independently of each other at the blade root portion 81. The cooling passage 83 is a leading edge side passage, communicates with a passage 83a in the blade, and cooling air 90 flows in from the rotor side. , While cooling the leading edge, flows out from the tip. Cooling air 91 flows in from the cooling passage 84, passes through the cooling passage 84a in the blade, flows into 84b from the tip side, and
4b, flows inward and flows into 84c, where the cooling passage 8
5 and the cooling air 92 flowing from the
Through the space, the cooling effect is enhanced, and flows out of the slot 86 at the trailing edge to the outside, and also flows out of the tip portion to the outside. The blades are cooled in the course of the flow of the cooling air.

In FIG. 11 (c), a tip thinning 87 is provided around the tip of the moving blade 80 and functions as a seal for air leaking from the tip of the moving blade to a subsequent stage.
Reference numeral 88 denotes a plug which closes a hole provided for processing at the time of manufacturing the rotor blade. The two-stage bucket 80 having such a configuration is used.
In this case, the cooling air is guided inside and cooled to prevent the influence of the high temperature due to the high temperature combustion gas.

[0009]

As described above, in the moving blade of a gas turbine, the blade and the platform are cooled by flowing cooling air to suppress a rise in metal temperature due to a high-temperature combustion gas. In the moving blade of a gas turbine, there is a large difference in mass between the platform and the blade profile. If a large temperature difference occurs between the two, a large thermal stress is generated. If a large thermal stress is generated between the wing profile and the platform, cracks are likely to occur, especially at the most severely severe points on the wing and the trailing edge hub at the root of the platform. Damage due to high temperature occurs at the affected portion, that is, the tip thinning at the tip. Such cracks and damages are caused by a combination of creep rupture based on high temperature and high stress due to operation for many months and fatigue rupture due to repeated start / stop stress. For that purpose, it is necessary to reduce the temperature and (thermal) stress of the portion where the stress concentration occurs (the root portion of the leading edge and trailing edge of the platform) as much as possible.

Therefore, according to the present invention, in the blade and the platform, the cooling structure at the root portion, the trailing edge, the tip end portion and the both ends of the platform, which are particularly susceptible to thermal stress, is improved. An object of the present invention is to provide a gas turbine rotor blade capable of suppressing the occurrence of cracks and thermal damage, extending the life of the blade, and improving reliability.

[0011]

The present invention provides the following means (1) to (6) to solve the above-mentioned problems.

(1) A tip thinning is provided at the tip, and a serpentine cooling passage through which cooling air flows is provided inside the blade. Cooling air flows into the serpentine cooling passage, and a cooling air blowing hole provided in the blade. In a gas turbine blade that cools by flowing more out to the surface and also flowing to the platform, the serpentine cooling passage is composed of two flow paths in which cooling air flows in from the center of the blade root and flows to the leading edge and trailing edge. The outer shape of the root portion of the wing forms a curved surface, and a concave portion is formed on each of both end surfaces of the root portion at the leading edge side and the trailing edge side of the platform along a direction orthogonal to the rotation axis direction. A cooling passage for flowing cooling air from the leading edge side to the trailing edge side is provided inside both ends of the side and the ventral side, and a plurality of cooling passages are arranged along the cooling path on the back side. There the gas turbine blade and the other end communicates to the cooling passage comprises a cooling hole that opens to the back end surface.

(2) The gas turbine rotor blade according to (1), wherein the shape of the leading edge portion and the trailing edge portion of the root portion of the blade is a combination of a straight line and a curved line.

(3) The gas turbine rotor blade according to (1), wherein the tip thinning is provided only on the back side, and a plug provided on the tip portion has a circular shape.

(4) The height (H) of the shank portion for fixing the platform is longer than the width (W) of the shank portion in the circumferential direction to lengthen the shank portion. The gas turbine rotor blade according to any one of (3).

(5) A tip thinning is provided at the tip, and a serpentine cooling passage through which cooling air flows is provided inside the blade. Cooling air flows into the serpentine cooling passage, and a cooling air blowing hole provided in the blade. In the gas turbine blade, which flows out to the surface and also flows to the platform for cooling, the serpentine cooling passage is a flow passage through which cooling air flows in from the center of the blade root and flows to the trailing edge, and the outer shape of the blade root A curved surface and a concave portion is formed on the end surface of the root portion on the trailing edge side of the platform along a direction perpendicular to the rotation axis direction. A cooling passage for flowing cooling air to the edge is provided, and a plurality of cooling passages are arranged along the cooling passage on the back side, one end of which is communicated with the cooling passage and the other end is open on the back end surface. Gas turbine blade characterized by comprising a cooling hole for.

(6) The gas turbine rotor blade according to (5), wherein the tip thinning is provided only on the back side.

(1) of the present invention is applied to a one-stage blade, and the serpentine cooling passage has two systems, a passage flowing to the leading edge side and a passage flowing to the trailing edge side, so that the inside of the blade can be effectively made. At the same time as cooling, a concave portion, that is, a slack portion is formed on the leading edge side and the trailing edge side of the wing root portion of the platform, and the concave portion reduces the thickness of the wing root portion of the platform. Accordingly, there is no sharp change in the wall thickness between the thin portion of the blade and the platform, and the mass just below the thin wall of the blade is reduced by the concave portion, the heat capacity is reduced, and the difference in heat capacity is reduced. For this reason, the temperature difference due to the difference in cooling rate is also reduced, and the occurrence of cracks due to the thermal stress generated at the connection between the blade and the platform in the related art is prevented. Furthermore, the platform is cooled at both ends on the back side and the abdomen side by the cooling passage,
Cooling air is blown to the end face from cooling holes arranged along the back side cooling passage, exposed to the hot combustion gas of the platform, and the thermally severe back side end face is effectively cooled.

In (2) of the present invention, the front edge and the rear edge
The fillet shape of the two places, for example, the upper part is inclined with a straight line,
By forming a curved surface at the base portion as a curved line and making the curved surface closer to a straight line to make the radius R larger than the other dorsal and ventral fillets R (radius), the steel property of this portion is increased, and the thermal stress is increased. Can be suppressed, and the occurrence of cracks can be prevented.

In (3) of the present invention, since the tip thinning is eliminated on the abdominal side, and especially on the back side which is affected by heat, the sealing performance of the tip is kept to a minimum and the tip thinning is performed at a high temperature. Damage can be reduced. In addition, since the plug is formed in a circular shape, the plug can be easily mounted, and damage due to high temperature is reduced.

In (4) of the present invention, since the shank portion is longer than before, the change that occurs at the blade-platform connection portion due to thermal stress is due to the longer shank portion fixing the platform. Absorbed by the damping effect, the effect of thermal stress is relieved, and the occurrence of cracks is prevented.

(5) of the present invention is applied to a two-stage bucket, in which the serpentine cooling passage is composed of a passage through which cooling air flows in from the center and flows to the trailing edge side, thereby effectively cooling the inside of the blade. At the same time, a concave portion, that is, a slack portion is formed on the end surface of the platform at the trailing edge side of the wing root portion, and the thick portion of the wing root portion of the platform is reduced by the concave portion. Accordingly, there is no sharp change in the wall thickness between the thin portion of the blade and the platform, and the mass just below the thin wall of the blade is reduced by the concave portion, the heat capacity is reduced, and the difference in heat capacity is reduced. For this reason, the temperature difference due to the difference in cooling rate is also reduced, and the occurrence of cracks in the root portion at the trailing edge due to the thermal stress generated at the connection portion between the conventional blade and the platform is prevented. Further, the platform is cooled at both ends on the back side and the abdomen side by the cooling passage, and cooling air is blown out from the cooling holes arranged along the back side cooling passage to the end face, and is exposed to the high-temperature combustion gas of the platform, The thermally severe dorsal end face is effectively cooled.

In (6) of the present invention, tip thinning is performed on the abdominal side, and especially on the dorsal side affected by heat.
The sealing performance of the tip can be kept to a minimum, and the damage of the tip thinning due to the high temperature can be reduced.

[0024]

Embodiments of the present invention will be specifically described below with reference to the drawings. FIG. 1 is a cross-sectional view of a first-stage moving blade of a gas turbine according to a first embodiment of the present invention. FIGS. 2 to 5 are detailed views of each part in the first embodiment, and will be described in detail below with reference to these drawings.

In FIG. 1, reference numeral 1 denotes a one-stage rotor blade, and 2 denotes a platform. A slack portion 3a formed by a concave groove is provided on the leading edge side of the platform 2, and a smooth curved surface is also provided on the trailing edge side. Recessed portion 3b formed of
Are formed. Reference numerals 4a and 4b denote fillets R on the front edge side and the rear edge side, and the curved surfaces at both ends have a larger curvature than those on the back side and the abdomen side.

Reference numeral 17 denotes a blade root portion, in which cooling passages 5, 6, 7, and 7a are independently provided.
The cooling passage 5 communicates with the cooling passage 5a in the blade on the leading edge side, and cooling air 81 flows in from the rotor side, flows through the passage, cools, and flows out of the hole 10a at the tip. From the surface to a showerhead film cooling the leading edge. Cooling air 82 flows in from the cooling passage 6, passes through the cooling passage 6 a in the blade, returns from the tip to the inside of the passage 6 b, flows from 64 b toward the base, turns back, enters the passage 6 c, and enters the passage 6 c. And flows out of the hole 10b. In this process, as described later with reference to FIG. 2, the film flows out from the film cooling hole 8 to the blade surface to cool the film.

The cooling air 83 flows from the cooling passage 7,
The air flows in the cooling passage 7a in the blade toward the tip and turns back.
It enters the passage 7b, flows to the base, returns to the passage 7c, flows through the passage 7c toward the front end, and flows out from the hole 10e at the front end. In this process, as described later with reference to FIG. 2, the air flows out of the film cooling hole 11 to the blade surface to cool the blade surface, and further flows out of the trailing edge slot 12 to the trailing edge side. Reference numerals 13a and 13b denote knife edges, which have sharp leading and trailing edges, and are close to the sealing portion with the adjacent stationary blade to improve the sealing performance.

FIG. 2 is a sectional view taken along the line AA in FIG.
Although not shown in FIG. 1, each cooling passage 5a, 6
Turbulators are provided on inner wall surfaces on both sides of c, 6b, 6a, 7a, 7b, 7c. Cooling passage 5a on the leading edge side
Are provided with a large number of film cooling holes 9 at the top and bottom, and blow out cooling air to cool the surface of the film.
Further, a number of film cooling holes 8 are provided on the back side of the cooling passage 6c to blow the cooling air to cool the surface on the back side. Further, a large number of film cooling holes 11 are provided vertically on the ventral side of the cooling passage 7b, and the cooling air is blown out to cool the rear surface on the ventral side. Cooling air is blown from the slot 12 on the trailing edge side.

As described above, in the first embodiment, the cooling air flows in from the center of the blade at the root of the blade, and cools the leading edge side of the blade in the cooling passages 5a, 6a, 6b, 6c. The trailing edge side is cooled by the passages 7a, 7b, 7c to form a two-system serpentine flow path with the leading edge passage, and a long cooling path is formed inside the blade to improve the cooling effect. Further, a film cooling hole 9 is provided at the leading end on the leading edge side, a film cooling hole 8 is provided on the back side of the wing, a film cooling hole 11 is provided on the ventral side on the trailing edge side,
The cooling effect of the wings is enhanced by film cooling the surfaces of the wings.

FIG. 3 is a sectional view taken along the line BB in FIG.
In the drawing, the right side is shown in the figure arranged on the leading edge side and the left side is shown in the figure arranged on the trailing edge side. In the figure, as described in the conventional example of FIG.
And cooling air 72a, 7b from the leading edge side.
2b is guided, flows out to the trailing edge side like 72d and 72c, and cools both sides of the platform. In the first embodiment, a plurality of cooling holes 14 are further arranged along the cooling passage 50b on the back side, the cooling holes 14 communicate with the cooling passage 50b, and the cooling air is opened to the side surface to cool the cooling air. It blows out to the back side of the platform to enhance the cooling effect of this part.

In the platform according to the first embodiment of the present invention, the cooling holes 14 are arranged on the back side in addition to the cooling passages 50a and 50b provided on both the ventral side and the back side as described above. The cooling effect is improved, and as described with reference to FIG. 1, the leading and trailing edges are provided with the slack portions 3 a and 3 b, so that the leading and trailing edge wings which are most susceptible to thermal stress are attached. The heat capacity of the root platform can be reduced to balance with the wing, and the thermal stress in this portion can be averaged to reduce the influence of the thermal stress.

FIG. 4 corresponds to a section taken along the line CC in FIG.
2A and 2B are a cross-sectional view and a plan view showing a conventional structure, and FIGS. 3C and 3D are a cross-sectional view and a plan view showing a structure of the present invention. . In the conventional structure, a tip thinning 73 is erected around the tip of the wing 60, and a rectangular plug 74 for closing a hole at the time of manufacture is incorporated in the center of the tip. ), (C)
In (2), the tip thinning 15 is used only on the back side, and the tip thinning is removed on the ventral side. Further, the hole is round and the plug 16 is circular and welded from the upper surface to facilitate assembly.

As described above, the tip end portion in the first embodiment of the present invention eliminates the abdominal tip thinning and provides the tip thinning 15 only on the back side to minimize the deterioration of the sealing performance. The damage due to high temperature is avoided, the plug 16 is rounded and the hole is reduced, workability is improved, and damage due to high temperature is reduced.

FIG. 5 is a view showing the shape of the fillet R in the first embodiment. In the figure, before the shape of the fillet R of the root portion of the edge and the rear edge side of the platform 2, the prior art consists curve shown in dotted line in the figure of Y ₁ and solid line Y _2, the lateral distance 14mm wings high It was formed with a smooth curved surface up to about 13 mm. In contrast, in the present invention, the lateral distance of the fillet is 5 mm from the wing height of 20 mm.
Up to the point P in mm, make the surface inclined with a straight line, 5mm to 14mm
The range up to is the same as the conventional curved surface, and has a shape that smoothly joins at point P. The shape of the fillet is such that the vicinity of two places on the leading edge side and the trailing edge side is a large curved surface including a straight line as compared with the conventional case, and the other back side and ventral side are the same as the conventional shape.

As described above, by increasing the fillet R on the leading edge side and the trailing edge side, the thickness of the fillet in this portion is increased, and the bending strength of this portion is improved and the stress can be reduced. In addition, the effect of the above-mentioned slush portions 3a and 3b is added, the curvature of the leading edge and the trailing edge of the blade against thermal stress is improved, and the occurrence of cracks can be suppressed.

According to the first embodiment described above, the cooling passages 6a, 6b, and 6c which flow two times in the serpentine flow path inside the rotor blade toward the leading edge side, and the cooling passages which flow twice in the trailing edge side. 7a, 7b and 7c are provided to lengthen the flow path, and further, a film cooling hole 9 at the leading edge front end, a film cooling hole 8 at the back side of the front edge, and a film cooling hole 11 at the back of the ventral side are provided. The cooling structure is provided, and a cooling structure is provided in which a number of cooling holes 14 for blowing cooling air from the cooling passage 50b on the back side of the platform toward the side surface.

Further, the curvature of the fillet R at the roots of the leading and trailing edges of the wing is larger than that of the conventional one and larger than that of the dorsal and ventral sides.
a, 3b are provided, and the tip thinning 15 is provided only on the back side, and no tip thinning is provided on the ventral side.

By adopting the cooling structure described above, the cooling effect of the entire blade 1 is improved, the thermal stress at the root of the platform and the blade is reduced, the average stress of the fillet R is also reduced, and the bending strength is improved. However, damage to the chip thinning due to high temperatures can be avoided without impairing the sealing performance.

FIGS. 6A and 6B are perspective views showing a gas turbine rotor blade according to a second embodiment of the present invention. FIG. 6A shows a conventional rotor blade, and FIG. 6B shows a blade of the second embodiment. Show. In the second embodiment, the shank portion for supporting and fixing the platform 2 is long and thin. That is,
The shank portion has a height H which is larger than that of the conventional shank portion 90 shown in FIG.
> H ₀ , and the width W is thinned as W <W ₀ , and H> W as a whole. By making it longer and thinner in this way, it is possible to add an elastic force to changes due to thermal stress, disperse and absorb the thermal stress by the damping effect, and suppress the occurrence of cracks due to thermal stress. it can. Other configurations are the same as those of the first embodiment, and further enhance the effects of the first embodiment.

FIGS. 7A and 7B show an example of a two-stage blade showing a gas turbine blade according to a third embodiment of the present invention. FIG. 7A is a longitudinal sectional view of the blade, and FIG. 7B is a rib inside the blade. Enlarged detail of the tip,
(C) is DD sectional drawing in (a). (A) In the figure, 21 is a two-stage bucket, and 22 is a platform. Cooling passages 23, 24, 25
Is provided, cooling air 50 flows into the cooling passage 23,
It enters the cooling passage 23a inside the blade, cools the leading edge, and flows out from the tip.

The cooling air 51 flows into the cooling passage 24,
It flows into the cooling passage 24a inside the wing, turns at the tip, returns to the passage 24b, flows toward the base, merges with the cooling air 52 that returns and flows into the cooling passage 25, flows into the passage 24c, and flows out to the tip. At the same time, it flows out of the trailing edge slot. In this process, the cooling passages 24a, 24a
4b also flows out from the tip. Reference numeral 26 denotes a slack portion, which forms a concave portion having a smooth curved surface on the trailing edge end surface of the platform 22. The fillet R28 at the root of the upper wing has a larger curvature than the other fillets. The shape of the fillet R is the same as that described with reference to FIG.

In the above-described blade structure, in addition to the cooling effect by the serpentine flow path composed of the cooling passages 23a, 24a, 24b, and 24c, the description has been given in the first embodiment by the fillet R28 and the slack portion 26 on the trailing edge side. Similarly,
The thermal stress at this portion is reduced, the strength of the fillet R28 is improved, and the occurrence of cracks can be prevented. In addition, since the same structure as that of the first embodiment is applied to cooling the platform, the description is omitted.

FIG. 7B is a sectional view taken along the line DD in FIG. 7A and shows chip thinning. As shown, the tip thinning 29 is provided only on the back side, and the abdomen side is omitted. Further, the plug 30 has a welded structure from the top to facilitate manufacture and assembly. By providing the chip thinning 29 only on the back side in this manner, the sealing performance is kept to a minimum, and damage to the chip thinning due to high temperatures is suppressed.

In the third embodiment described above, in addition to the cooling effect of the blades by the serpentine flow path by the cooling passages 23a, 24a, 24b and 24c, the weight is reduced by providing the groove 27a at the tip of the rib 27. At the same time, the flow of air is improved, the strength of the root portion at the trailing edge wing against thermal stress is improved by the fillet R28 and the squat 26, and damage to the chip thinning can be prevented. The cooling structure of other platforms is the first
The form can be applied.

[0045]

The gas turbine rotor blade of the present invention has the following features. (1) A tip thinning is provided at the tip, a serpentine cooling passage through which cooling air flows inside the blade, and cooling air is supplied to the serpentine cooling passage. In the gas turbine rotor blade which flows out to the surface from the cooling air blowout hole provided in the sink and also flows to the platform for cooling, the serpentine cooling passage has a cooling air flowing from a central portion of a blade root portion and a leading edge side. It is composed of two flow paths flowing to the trailing edge side, the outer shape of the root portion of the wing forms a curved surface, and both end surfaces of the platform at the leading edge side and the trailing edge side of the root portion,
A recess is formed along a direction orthogonal to the rotation axis direction, and the platform is provided with cooling passages through which cooling air flows from the leading edge side to the trailing edge side inside the back side and the inside of the abdominal side, and the cooling on the back side. A plurality of cooling holes are arranged along the passage, one end of which is connected to the cooling passage, and the other end is provided with a cooling hole which opens to the back end surface. With such a configuration, the inside of the blade is effectively cooled by the serpentine cooling passage, and a concave portion is formed on the leading edge side and the trailing edge side of the blade root portion of the platform, and the concave portion is formed by the concave portion. The thickness of the wings is reduced, the abrupt change in the thickness between the thin part of the wing and the platform disappears, the mass just below the thin wing is reduced by the concave part, the heat capacity is reduced, and the difference in heat capacity is reduced . For this reason, the temperature difference due to the difference in cooling rate is also reduced, and the occurrence of cracks due to the thermal stress generated at the connection between the blade and the platform in the related art is prevented. Further, the platform is cooled at both ends on the back side and the abdomen side by the cooling passage, and cooling air is blown out from the cooling holes arranged along the back side cooling passage to the end face, and is exposed to the high-temperature combustion gas of the platform, The thermally severe dorsal end face is effectively cooled.

According to (2) of the present invention, the front edge and the rear edge
The fillet shape of the two places, for example, the upper part is inclined with a straight line,
By forming a curved surface at the base portion as a curved line and making the curved surface closer to a straight line to make the radius R larger than the other dorsal and ventral fillets R (radius), the steel property of this portion is increased, and the thermal stress is increased. Can be suppressed, and the occurrence of cracks can be prevented.

In (3) of the present invention, the tip thinning is eliminated on the abdominal side, especially on the back side which is affected by heat, so that the sealing performance of the tip is kept to a minimum and the tip thinning is performed at a high temperature. Damage can be reduced. In addition, since the plug is formed in a circular shape, it is easy to mount the plug, and damage due to high temperature is reduced.

In (4) of the present invention, since the shank is longer than before, the change in the connection between the blade and the platform due to thermal stress is due to the longer shank that fixes the platform. Cracking is prevented by absorbing the damping effect and relieving the influence of thermal stress.

According to a fifth aspect of the present invention, in the gas turbine rotor blade, the serpentine cooling passage is a passage through which cooling air flows from a central portion of the blade root portion and flows to a leading edge side and a trailing edge side. The outer shape forms a curved surface and a concave portion is formed on the end surface of the root portion on the trailing edge side of the platform along a direction orthogonal to the rotation axis direction. A cooling passage for flowing cooling air to the trailing edge side is provided, and a plurality of cooling holes are arranged along the back side cooling passage, one end of which is communicated with the cooling passage and the other end is opened to the back side end face. Features. With such a configuration, the inside of the blade is effectively cooled by the serpentine cooling passage, and a concave portion is formed on the end surface on the trailing edge side of the blade root portion of the platform. Is reduced, the abrupt change in the wall thickness between the thin portion of the blade and the platform is eliminated, and the heat capacity is reduced by reducing the mass immediately below the thin blade by the concave portion, and the difference in the heat capacity is reduced. For this reason, the temperature difference due to the difference in cooling rate is also reduced, and the occurrence of cracks in the root portion at the trailing edge due to the thermal stress generated at the connection portion between the conventional blade and the platform is prevented. Further, the platform is cooled at both ends on the back side and the abdomen side by the cooling passage, and cooling air is blown out from the cooling holes arranged along the back side cooling passage to the end face, and is exposed to the high-temperature combustion gas of the platform, The thermally severe dorsal end face is effectively cooled.

In (6) of the present invention, tip thinning is performed on the abdominal side, and especially on the dorsal side affected by heat.
The sealing performance of the tip can be kept to a minimum, and the damage of the tip thinning due to the high temperature can be reduced.

[Brief description of the drawings]

FIG. 1 is a longitudinal sectional view of a gas turbine bucket according to a first embodiment of the present invention.

FIG. 2 is a sectional view taken along line AA in FIG.

FIG. 3 is a sectional view taken along line BB in FIG.

FIGS. 4A and 4B are cross-sectional views taken along line CC in FIG.
2A and 2B are a sectional view and a plan view of a conventional chip thinning, and FIGS. 3C and 3D are a sectional view and a plan view of the chip thinning of the present invention.

FIG. 5 is a view showing a shape of a fillet R of the gas turbine rotor blade according to the first embodiment of the present invention.

FIGS. 6A and 6B are perspective views showing a gas turbine rotor blade according to a second embodiment of the present invention, wherein FIG. 6A shows a conventional blade and FIG. 6B shows the present invention.

7A and 7B show a gas turbine rotor blade according to a third embodiment of the present invention, wherein FIG. 7A is a longitudinal sectional view of a two-stage rotor blade, and FIG. 7B is a DD sectional view of FIG.

FIG. 8 is a longitudinal sectional view of a conventional one-stage bucket.

FIG. 9 is a sectional view taken along the line EE in FIG. 8;

FIG. 10 is a plan view of a platform of a conventional gas turbine blade, in which (a) is an example having cooling passages on both sides,
(B) shows an example in which a cooling hole is further added.

11A and 11B show a conventional two-stage bucket, in which FIG.
(B) is an FF cross-sectional view in (a), (b) is (a)
It is GG sectional drawing in.

[Explanation of symbols]

 Reference Signs List 1 1 stage rotor blade 2, 22 Platform 3a, 3b, 26 Thread portion 4a, 4b, 28 Fillet R 5, 6, 7 Cooling passage 5a, 6a, 6b, 6c Cooling passage 7a, 7b, 7c Cooling passage 8, 9, DESCRIPTION OF SYMBOLS 11 Film cooling hole 10a, 10b, 10c, 10d, 10e Hole 12 Slot 14 Cooling hole 15, 29 Chip thinning 16 Plug 18 Shank part 21 Two-stage bucket 23, 24, 25 Cooling passage 23a, 24a, 24b, 24c Cooling Passage 27 Rib 30 Plug

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor Yasushi Tomita 2-1-1 Shinhama, Arai-machi, Takasago City, Hyogo Prefecture Inside the Takasago Machinery Works, Mitsubishi Heavy Industries, Ltd. No. 1 Inside the Mitsubishi Heavy Industries, Ltd. Takasago Plant (72) Inventor Tatsuo Ishiguro 2-1-1, Araimachi Shinhama, Takasago City, Hyogo Prefecture Inside the Mitsubishi Heavy Industries, Ltd. Takasago Research Laboratory F-term (reference) 3G002 AA04 AA06 AB01 AB08 BA02 BB02 CA05 CA06 CA08 CB01 FB01

Claims (6)

[Claims] 1. A blade having a tip thinning at a tip thereof, a serpentine cooling passage through which cooling air flows inside the blade, and a cooling air flowing into the serpentine cooling passage through a cooling air blowing hole provided on the blade. In a gas turbine rotor blade for cooling to flow to the surface and also to the platform, the serpentine cooling passage includes two flow paths through which cooling air flows in from the center of the blade root and flows to the leading edge side and the trailing edge side, The outer shape of the root portion of the wing forms a curved surface, and a concave portion is formed on each of both end surfaces on the leading edge side and the trailing edge side of the root portion along the direction orthogonal to the rotation axis direction. A cooling passage for allowing cooling air to flow from the leading edge side to the trailing edge side inside both ends on the ventral side, and a plurality of cooling passages are arranged along the cooling path on the dorsal side, and one end is the same. A gas turbine rotor blade having a cooling hole communicating with a cooling passage and having the other end opened at a back end face. 2. The gas turbine moving blade according to claim 1, wherein the shape of the leading edge portion and the trailing edge portion of the blade root portion is formed by combining a straight line and a curved line. 3. The gas turbine blade according to claim 1, wherein the tip thinning is provided only on the back side, and a plug provided on the tip portion has a circular shape. 4. A shank portion for fixing the platform, wherein a height (H) of the shank portion is longer than a circumferential width (W) thereof to make the shank portion longer.
4. The gas turbine bucket according to any one of claims 1 to 3. 5. A blade having a tip thinning at a tip thereof, a serpentine cooling passage through which cooling air flows inside the blade, and a cooling air flowing through the serpentine cooling passage through a cooling air blowing hole provided in the blade. In the gas turbine rotor blade which flows out to the surface and also flows to the platform for cooling, the serpentine cooling passage is a flow passage through which cooling air flows in from the center of the blade root and flows to the trailing edge side. A curved surface is formed, and a concave portion is formed on the end surface of the root portion on the trailing edge side of the platform along a direction orthogonal to the rotation axis direction. And a plurality of cooling passages are arranged along the cooling passage on the back side, one end of which is connected to the cooling passage and the other end is open on the back end surface. A gas turbine blade having a cooling hole. 6. The gas turbine rotor blade according to claim 5, wherein the tip thinning is provided only on the back side.