Gas turbine moving blade

Abstract: Gas turbine moving blade is improved so as to prevent occurrence of cracks caused by thermal stresses due to temperature differences between blade and platform when gas turbine is stopped. In steady operation time of moving blade (20), cooling air (40 to 43) enters cooling passages (23 to 26) to flow through cooling passages (23a, 24a to 24c, 25a to 25c) for cooling the blade (20) to then flow out of the blade (20). Recessed portion (1) having smooth curved surface is provided in platform (22) near blade fitting portion on blade trailing edge side. Fillet (R) of the blade fitting portion on the blade trailing edge side is formed with curved surface having curvature larger than conventional case. Hub slot below the fillet (R) for blowing air is formed having slot cross sectional area larger than other slots of blade trailing edge. TBC is applied to blade (20) surface. By these improvements, thermal stresses due to temperature differences between the blade (20) and the platform (22) in gas turbine stop time are made smaller and occurrence of cracks is prevented.

Description: BACKGROUND OF THE INVENTION Field of the Invention The present invention relates generally to a gas turbine moving blade and more particularly -to a gas turbine moving blade which is improved in a cooling structure of blade and platform so as to prevent occurrence of cracks due to thermal stresses caused by temperature changes in gas turbine starts and stops or in high temperature combustion gas. Description of the Prior Art In Fig. 7, which is a cross sectional view of a representative first stage moving blade of a prior art gas turbine, numeral 20 designates the moving blade, numeral 21 designates a blade root portion and numeral 22 designates a platform. In the blade root portion 21, there are provided cooling passages 23, 24, 25, 26, which are independent of each other. The cooling passage 23 is a passage on a blade leading edge side to communicate with a cooling passage 23a provided in a blade leading edge portion. Cooling air 40 flows into the cooling passage 23 from a turbine rotor side to flow through the cooling passage 23a and to flow out of a blade tip portion for cooling the blade leading edge portion and, at the same time, to flow out of cooling holes 29 for effecting a shower head film cooling of the blade leading edge portion. Cooling air 41 flows into the cooling passage 24 to flow through a cooling passage 24a provided in the blade and then turns at the blade tip portion to flow through a cooling passage 24b and turns again at a blade base portion to flow through a cooling passage 24c and to flow out of the blade tip portion. In this process of the flow, the cooling air 41 cools a blade interior and, at the same time, flows out of cooling holes, to be described later with respect to Fig. 8, onto a blade surface for effecting a film cooling thereof. Cooling air 42 entering the cooling passage 25 and cooling air 43 entering the cooling passage 26 join together to flow through a cooling passage 25a and then turn at the blade tip portion to flow through a cooling passage 25b and turn again at the blade base portion to flow through a cooling passage 25c. In this process of the flow, the cooling air 42, 43 cools the blade interior and, at the same time, flows out of cooling holes, to be described later with respect to Fig. 8, onto the blade surface for effecting the film cooling thereof and a portion still remaining of the cooling air 42, 43 flows out of cooling holes 28 of a blade trailing edge 27 for effecting a pin fin cooling of a blade trailing edge portion. In Fig. 8, which is a cross sectional view taken on line B-B of Fig. 7, a portion of the cooling air flowing through the cooling passage 23a in the blade leading edge portion flows out of the blade through the cooling holes 29 for effecting the shower head film cooling of the blade leading edge portion. Also, a portion of the cooling air flowing through the cooling passage 24c flows outside obliquely through cooling holes 30 for effecting the film cooling of the blade surface. Likewise, a portion of the cooling air flowing through the cooling passage 25c flows outside obliquely through cooling holes 31 for effecting the film cooling of the blade trailing edge portion. It is to be noted that although the cooling holes 29, 30, 31 only are illustrated, there are actually provided a multiplicity of cooling holes other than the mentioned three kinds of the cooling holes 29, 30, 31. In Fig. 9, which is an explanatory plan view of a cooling structure of the platform 22, Fig. 9(a) shows an example to cool a front portion, or a blade leading edge side portion, of the platform 22 as well as to cool both side portions, or blade ventral and dorsal side portions, of the platform 22 and Fig. 9(b) shows another example to cool upper surface portions of both of the side portions of the platform 22 in addition to the cooled portions of Fig. 9(a). In Fig. 9(a), there are bored cooling passages 50a, 50b in the front portion and both of the side end portions of the platform 22 so as to communicate with the cooling passage 23 of the leading edge portion of the moving blade 20. Cooling air 72a, 72b flows through the cooling passages 50b, 50a, respectively, for cooling the front portion and both of the side portions of the platform 22 and flows out through a rear portion, or a blade trailing edge side portion, of the platform 22 as air 72c, 72d. In Fig. 9(b), in addition to the cooling passages 50a, 50b of Fig. 9(a), there are provided a plurality of cooling holes 51a, 51b, respectively, in both of the side portions of the platform 22 so as to open at an upper surface of the platform 22. These cooling holes 51a, 51b communicate with one or more of the cooling passages leading to the interior of the moving blade 20, so that cooling air flows through the cooling holes 51a, 51b to flow out onto the upper surface of the platform 22 and cools both of the side portions of the platform 22. Thus, in the gas turbine moving blade, the moving blade 20 as well as the platform 22 are cooled as described with respect to Figs. 7 to 9, so that thermal influences given by the high temperature combustion gas are mitigated. In Fig. 10, which shows an example of a second stage moving blade in the prior art, Fig. 10(a) is a cross sectional view thereof, Fig. 10(b) is a cross sectional view taken on line F-F of Fig. 10(a) and Fig. 10(c) is a cross sectional view taken on line G-G of Fig. 10(a). In Figs. 10(a) and (b), numeral 180 designates the second stage moving blade, numeral 181 designates a blade root portion and numeral 182 designates a platform. In the blade foot portion 181, there are provided cooling passages 183, 184, 185, which are independent of each other. The cooling passage 183 is a passage on a blade leading edge side to communicate with a cooling passage 183a provided in a blade leading edge portion. Cooling air 190 flows into the cooling passage 183 from a turbine rotor side to flow through the cooling passage 183a for cooling the blade leading edge portion and to flow outside through a blade tip portion. Cooling air 191 flows into the cooling passage 184 to flow through a cooling passage 184a provided in the blade and then turns at the blade tip portion to flow through a cooling passage 184b and turns again toward inside at a blade base portion. In the blade base portion, the cooling air 191 and cooling air 192 flowing through the cooling passage 185 join together and flow into a cooling passage 184c. In the cooling passage 184c, the cooling air 191, 192 flows between pin fins 185 for enhancing the cooling effect and flows outside through slots 186 provided in a blade trailing edge as well as through a hole of the blade tip portion. In this process of the cooling air flow, the blade is cooled. In Fig. 10(c), there is provided a blade tip thinned portion 187 along each of blade tip edge portions of the moving blade 180 so as to function as a seal of air leaking toward blade rear stages from the blade tip. Numeral 188 designates a plug, which plugs up openings provided for working purposes when the moving blade 180 is being manufactured. In the second stage moving blade 180 as so constructed also, the cooling air is led into the interior of the blade, so that thermal influences given by the high temperature combustion gas are mitigated. As mentioned above, in the gas turbine moving blade, the blade and the platform are cooled by flowing the cooling air and elevation of metal temperature due to the high temperature combustion gas is suppressed. While there is a large difference in the mass between the platform and a blade profile portion of the gas turbine moving blade, the platform and the blade profile portion are cooled by the cooling air during a gas turbine steady operation time and there occurs no large temperature difference between them, so that thermal stress influences caused by the temperature difference are also small. However, in an unsteady time to stop the gas turbine, while the blade profile portion, which is of a thin shape, is cooled earlier, the platform, which is of a larger mass, is cooled slowly and this causes a large temperature difference between them, which results in causing large thermal stresses. If large thermal stresses occur between the blade profile portion and the platform, as mentioned above, cracks may arise easily, especially at a portion where there is the severest thermal influence, that is, at blade hub portions where the blade and the platform join together on the blade leading edge and trailing edge sides and also cracks are likely to arise at other portions where there are thermal stress influences, that is, at the cooling holes of the blade trailing edge, the blade tip thinned portion and the like. The cracks of the mentioned portions are caused by combination of creep ruptures caused by high temperature and high stress repeated by long time operations and fatigue failures caused by repeated stresses due to operation starts and stops and, in order to avoid such cracks, it is necessary to reduce the temperature and thermal stresses as much as possible at portions where stress concentrations are caused (blade and platform fitting portions of the blade leading edge and trailing edge portions). US-A-5 857 837 discloses a coolable air foil for a gas turbine engine which comprises a platform and a blade fitted to a blade fitting portion of the platform. A blade cooling passage is provided in the blade and cooling air blow holes are providing in and around the blade so that the blade may be cooled by cooling air flowing through the blade cooling passage and out of the blade through the cooling air blow holes. Of the cooling air blow holes at the blade trailing edge a lowermost air blow hole is formed larger than the cooling air blow holes provided at the blade trailing edge above the lowermost hole. The blade fitting portion is formed having a fillet exterior with a curved surface at the downstream side and a recess portion with a curved surface and extending in the direction orthogonal to a turbine axial direction is provided in an end face portion of a rear side of the platform near the blade fitting portion on the blade trailing edge side. EP-A-0 954 594 discloses a cooled moving blade for a gas turbine in which the base portion of the blade at the connection portion to the platform is formed in an elliptically curved surface. SUMMARY OF THE INVENTION In view of the problems in the prior art, therefore, it is an object of the present invention to provide a gas turbine moving blade which is improved in structural portions of blade and platform which are prone to be influenced by thermal stresses, especially blade and platform fitting portions and blade trailing edge cooling holes, as well as improved in cooling structures of a blade tip portion and platform front and rear both end portions so that cracks caused by thermal stresses due to temperature differences may be suppressed and life and reliability of the blade may be enhanced. In order to achieve the mentioned object, the present invention provides means of the following: (1) A gas turbine moving blade comprising a platform and a blade fitting portion where the blade is fitted to the platform as well as comprising a blade cooling passage provided in the blade, a platform cooling passage provided in the platform and cooling air blow holes provided in and around the blade so that the blade may be cooled by cooling air flowing through the blade cooling passage, flowing through the platform cooling passage and flowing out of the blade through the cooling air blow holes, characterized in that there is provided a recessed portion, having a smooth curved surface and extending in a direction orthogonal to a turbine axial direction, in an end face portion of a rear side portion of the platform near the blade fitting portion on a blade trailing edge side; the blade fitting portion is formed having a fillet exterior with a curved surface; and the cooling air blow holes provided in a blade trailing edge includes a hole provided in a blade hub portion positioned at a lowermost end of the cooling air blow holes provided in the blade trailing edge, the hole having a hole cross sectional area larger than that of each of the cooling air blow holes provided in the blade trailing edge above the hole. In the gas turbine moving blade as mentioned above, there is applied a coating of a heat resistant material to the blade and platform so that the blade fitting portions of blade leading edge and trailing edge portions may be applied to with the coating thicker than other portions of the blade and portions of the platform near and around the blade leading edge and trailing edge portions may be applied to with the coating thinner than other portions of the platform. (2) A gas turbine moving blade as mentioned in (1) above, characterized in that the curved surface of the fillet exterior of the blade fitting portion is formed to an elliptical curve (3) A gas turbine moving blade as mentioned in (1) above, characterized in that the platform cooling passage is connected with a platform cooling air supply system and there are provided in the platform cooling air supply system an opening/closing valve for opening and closing the platform cooling air supply system and a control unit for controlling the opening/closing valve so as to be closed while a gas turbine is operated and to be opened for a predetermined time when the gas turbine is stopped. (4) A gas turbine moving blade as mentioned in (1) above, characterized in comprising a shank portion for fixing the platform, the shank portion being formed in an elongated shape having a height (H) of the shank portion in a. turbine radial direction larger than a width (W) of the shank portion in a turbine rotational direction (H > W). In the invention (1), there is provided the recessed portion, or cut-out portion, having the smooth curved surface, in the rear end face portion of the platform near the blade fitting portion on the blade trailing edge side and a thick portion of the platform near this blade fitting portion is thinned by the recessed portion. Thus, there is eliminated a sharp thickness change between the thin blade portion and the thick platform portion and also the mass of the platform right under the thin blade portion is reduced by the recessed portion t'o make the thermal capacity there smaller and thus the thermal capacity difference also can be made smaller. Thereby, the temperature difference caused by the difference in the cooling velocity at the time of gas turbine stop or the like becomes also smaller and occurrence of the cracks as have been caused by the thermal stresses at the blade fitting portion can be prevented. Further, the fillet of the blade fitting portion is made in the curved surface which has partially the linear portion, so that the fillet R is made larger than the conventional case in the curvature and the rigidity of this portion is strengthened. Moreover, the lowermost hole of the cooling air blow holes provided in the blade trailing edge is made to have a hole cross sectional area larger than that of the other cooling air blow holes provided in the blade trailing edge and thereby the cooling effect of this portion is enhanced, the temperature difference in the blade fitting portion becomes smaller to suppress occurrence of the thermal stresses and occurrence of the cracks can be avoided securely. In the invention, the thermal barrier coating (TBC) of the heat resistant material is applied to the blade, so that temperature lowering of the blade after the stop of the gas turbine becomes slower and thereby the temperature difference between the blade fitting portion and the platform becomes smaller and the thermal stresses are made smaller. Also, temperature lowering of the blade portion where the thicker TBC is applied becomes further slower and the temperature difference between the blade and the platform becomes further smaller. Moreover, by the platform portion where the thinner TBC is applied, temperature lowering of the platform at and around this portion is comparatively fast, so that the temperature difference between the blade fitting portion and the platform becomes further smaller and thus the thermal stresses caused there are made further smaller. Also, in the embodiment of the invention (2), the fillet exterior of the blade fitting portion is formed to the elliptical curve so that the curvature there becomes large and the stress concentration in this portion can be mitigated. In the other embodiment of the invention (3), when the gas turbine is stopped, the control unit opens the opening/closing valve for the predetermined time so that cooling air from the platform cooling air supply system may be led actively into the cooling passage of the platform and the platform is cooled even in the stop of the gas turbine. Hence, cooling of the platform which is slower in the temperature lowering than the thin moving blade is accelerated, the temperature difference between the blade and the platform is made smaller to suppress occurrence of the thermal stresses and occurrence of the cracks is prevented. In the further embodiment of the invention (4), the shank portion which fixes the platform is elongated in the height as compared with the conventional one, so that deformation caused by the thermal stresses at the connection portion of the blade and the platform is absorbed by the damping effect which is given by the elongation of the shank portion to mitigate the influences of the thermal stresses and thereby occurrence of the cracks is prevented. BRIEF DESCRIPTION OF THE DRAWINGS Fig. 1 is a cross sectional view of a gas turbine moving blade of a first embodiment according to the present invention. Fig. 2 shows a blade fitting portion of the first embodiment of Fig. 1, wherein Fig. 2(a) is a side view of the blade fitting portion, Fig. 2(b) is a rear view seen from line A-A of Fig. 2(a) and Fig. 2(c) is a view showing a fillet R of Fig. 2(a). Fig. 3 is a rear view of a blade trailing edge showing a modified form of a hub slot of Fig. 2(b). Fig. 4 is a perspective view of a gas turbine moving blade including a shank portion thereof, wherein Fig. 4(a) shows a prior art one and Fig. 4(b) shows a second embodiment according to the present invention. Fig. 5 is a cooling system diagram of a gas turbine moving blade of a third embodiment according to the present invention. Fig. 6 is a plan view of a platform of the third embodiment according to the present invention, including a cooling system diagram thereof. Fig. 7 is a cross sectional view of a representative first stage moving blade of a prior art gas turbine. Fig. 8 is a cross sectional view taken on line B-B of Fig. 7. Fig. 9 is an explanatory plan view of a cooling structure of a platform of the prior art moving blade of Fig. 7, wherein Fig. 9(a) shows an example to cool a front portion and both side portions of the platform and Fig. 9(b) shows an example to cool upper face portions of the platform in addition to the cooled portions of Fig. 9(a). Fig. 10 shows an example of a second stage moving blade of a prior art gas turbine, wherein Fig. 10(a) is a cross sectional view thereof, Fig. 10(b) is a cross sectional view taken on line F-F of Fig. 10(a) and Fig. 10(c) is a cross sectional view taken on line G-G of Fig. 10(a). DESCRIPTION OF THE PREFERRED EMBODIMENTS Herebelow, embodiments according to the present invention will be described concretely with reference to figures. In Fig. 1, which is a cross sectional view of a gas turbine moving blade of a first embodiment according to the present invention, there is provided a recessed groove or cut-out portion 1, which is grooved in or cut out of a thick portion and has a rounded smooth curved surface, at a blade fitting portion where a moving blade 20 and a platform 22 join together to be fitted to each other on a blade trailing edge side. The recessed groove 1 is provided in an end face portion of a rear portion, or a blade trailing edge side portion, of the platform 22, extending in a direction orthogonal to a turbine rotor axial direction and having such a groove depth as not affecting lines of load force of the blade. In Fig. 2 showing the blade fitting portion of the first embodiment of Fig. 1, Fig. 2(a) is a side view thereof, Fig. 2(b) is a rear view seen from line A-A of Fig. 2(a) and Fig. 2(c) is a view showing a fillet R of Fig. 2(a). As shown by the shape of the fillet R of Fig. 2(c) provided at the blade fitting portion on the blade trailing edge side, while fillets of other portions than the blade trailing edge side portion have a smaller curvature, 6 mm for example, in the present first embodiment, the fillet R is made to an elliptical curve of 20 mm x 40 mm. By so making the fillet R larger, the stress concentration can be suppressed. Also, in Fig. 2, while there are provided cooling holes 28 in a blade trailing edge portion and slots 33 in a blade trailing edge, one slot 2 nearest to the platform 22 of the slots 33 (that is, the slot of the lowermost end, which is near a blade hub portion and is called a hub slot) is made to have a slot cross sectional area larger than that of other slots 33. For example, the hub slot 2 is of a 1.6 mm diameter while other slots 33 are of a 1 mm diameter. Thus, the construction is made so as to enhance the cooling effect of this portion. In Fig. 3, which is a rear view of the blade trailing edge showing a modified form of the hub slot of Fig. 2(b), while the slots 33 are formed by pedestals 34 provided between each of the slots 33, one pedestal 34a nearest to the platform 22 of the pedestals 34 is cut off so as to connect two slots to each other to thereby form a hub slot 3. Thus, the hub slot 3 which is nearest to the platform 22 is made to have a slot cross sectional area larger than that of other slots 33. Other structures of the moving blade 20 and the platform 22 are same as those shown in Figs. 1 and 2 and description thereon will be omitted. By the construction of the slots as described above, heat transfer area in the slot portions of the blade fitting portion on the blade trailing edge side is increased and cooling air flowing therethrough is increased in volume and temperature of the portion where the stress concentration occurs easily in operation can be reduced. Thus, the thermal stress influences in this portion are mitigated and occurrence of cracks can be prevented. Further, in the moving blade 20 of the present embodiment, a TBC (thermal barrier coating) is applied to the entire surface of the moving blade 20 including the recessed groove 1 and the hub slot 2, 3. Moreover, in so applying the TBC, (1) the blade fitting portions to the platform 22 on the blade leading edge and trailing edge sides are applied to with a thicker TBC as compared with other portions of the blade 20 and also (2) the platform 22 on the blade leading edge and trailing edge sides is applied to with a thinner TBC as compared with other portions of the platform 22. By the TBC so applied, when the gas turbine is stopped, cooling velocity of the blade is lowered as a whole, so that the temperature is lowered slowly, the temperature difference between the blade fitting portion and the platform becomes smaller and the thermal stress caused in this portion is reduced. Also, according to (1) above, in the portions of the blade where the TBC is applied thicker, the temperature lowering becomes slower and the temperature difference between those portions of the blade and the platform becomes further smaller. Hence, the thermal stress caused in this portion is further reduced. Furthermore, according to (2) above, in other portions than the portion of the platform where the TBC is applied thinner, the temperature lowering becomes slower and the temperature difference between those portions of the platform becomes further smaller. Hence, the thermal stress caused in the platform is further reduced. According to the gas turbine moving blade of the first embodiment as described above, cooling air flows in the same way as in the conventional case of Figs. 7 to 9, that is, the cooling air 40 to 43 enters the interior of the moving blade 20 from inside of the platform 22 for cooling the moving blade 20 to then flow out into the gas path through the blade tip portion on the blade leading edge side and through the cooling holes 29 to 31 and the blade trailing edge portion and; at the same time, enters the cooling passages 50a, 50b on both side end portions, or blade ventral and dorsal side end portions, of the platform 22 for cooling the platform 22 to then flow out toward the rear portion, or the blade trailing edge side portion, of the platform 22. In this cooling process as well as at the time of gas turbine stop, while, in the conventional case, the temperature difference between the blade profile portion and the platform 22 becomes large due to mass difference between them to thereby cause thermal stresses, in the present invention, there is provided the recessed groove or the cut-out portion 1 in the rear portion, or the blade trailing edge side portion, of the platform 22 and thereby the following effect can be obtained. That is, by the recessed groove 1, there is eliminated a sharp thickness change between a thin portion of the blade fitting portion of the moving blade 20 and a thick portion of the platform 22 as well as a thickness right under the thin portion of the blade fitting portion is recessed, so that thermal capacity there is reduced and also thermal capacity difference therearound is made smaller. Thus, the cracks as have been so far caused by thermal stresses at the fitting portion of the moving blade 20 and the platform 22 can be prevented. Also, the fillet R at the blade fitting portion is made larger than in the conventional case so that rigidity at this curved surface portion is increased and occurrence of cracks at this portion can be suppressed. Moreover, there is provided the hub slot 2, 3 at the portion of the fillet R and the hub slot 2, 3 has a slot cross sectional area larger than that of the other slots 33. Hence, heat transfer area in the thickness changing portion of the blade fitting portion is increased and also the cooling air is increased in volume so as to enhance the cooling effect. Thereby, in addition to the effect to reduce the thermal capacity by the recessed groove 1 right under the hub slot 2, 3, a large temperature difference therearound is suppressed synergically and occurrence of cracks can be prevented. Also, there is applied the TBC and yet it is applied thicker to the blade fitting portion and thinner to the platform 22 of that portion, so that, by this coating also, the thermal influences can be made smaller. In Fig. 4, which is a perspective view of a gas turbine moving blade comprising a shank portion thereof, Fig. 4(a) shows a prior art one and Fig. 4(b) shows a second embodiment according to the present invention comprising the recessed groove of the first embodiment of Fig. 1 and an improvement in the shank portion. In the shank portion of the present second embodiment, the shank portion to support fixedly the platform 22 is elongated in the height and thinned in the width. That is, as compared with a conventional shank portion 40a, having a height H 0 and a width W 0 , of a moving blade 20 shown in Fig. 4(a), a shank portion 40b shown in Fig. 4(b) has a height H and a width W, wherein H is larger than H 0 (H > H 0 ) and W is smaller than W 0 (W < W 0 ), and H is larger than W as a whole. By so making the shank portion 40b longer and thinner, the shank portion 40b is given a flexibility against thermal stress changes and because of a damping effect thereof, the thermal stresses are dispersed and absorbed. Thereby, occurrence of cracks due to the thermal stresses can be suppressed. In Fig. 5, which is a cooling system diagram of a gas turbine moving blade of a third embodiment according to the present invention, cooling air is led into a moving blade 20 for cooling thereof from a cooling air supply system 80 and then flows out through a blade trailing edge portion and, at the same time, a portion of the cooling air is led into a platform 22 for cooling thereof and then flows out through a rear portion, or a blade trailing edge side portion, of the platform 22. This cooling system is same as that of the conventional system described with respect to Figs. 7 to 9. In the present third embodiment, in addition to the cooling system mentioned above, there is provided a platform cooling air supply system 81, so that cooling air is led therefrom into cooling passages provided in the platform 22 via an opening/closing valve 11 and pipings 14a, 14b. Numeral 10 designates a control unit, and when the gas turbine is stopped, the control unit 10 is inputted with a gas turbine stop signal S to thereby control the opening/closing valve 11 so that cooling air may be supplied into the cooling passages of the platform 22 for a predetermined time after the stop of the gas turbine. Fig. 6 is a plan view of the platform 22 of the third embodiment, including a cooling system diagram thereof. Like in the prior art case, there are provided the cooling passages 50a, 50b in a front portion, or a blade leading edge side portion as well as in both side end portions, or blade ventral and dorsal side end portions, of the platform 22 so that cooling air flows therein for cooling the front portion and both of the side portions of the platform 22 and flows out through a rear portion of the platform 22. Further, in the front portion of the platform 22, there are provided passages 13a, 13b so as to communicate with the cooling passages 50a, 50b, respectively, of both of the side end portions of the platform 22. On the other hand, the passages 13a, 13b are connected with the pipings 14a, 14b, respectively, and the pipings 14a, 14b are connected to the platform cooling air supply system 81 via the opening/closing valve 11, as mentioned above. In the cooling system of the third embodiment constructed as described above, the opening/closing valve 11 is closed in the ordinary operation time of the gas turbine so that the ordinary cooling, as mentioned above, may be carried out. When the gas turbine is stopped, the gas turbine stop signal S is inputted into the control unit 10 and the control unit 10 controls to open the opening/closing valve 11 for the predetermined time. Thereby, cooling air from the platform cooling air supply system 81 is led into the cooling passages 50a, 50b of the platform 22, that is, even after the stop of the gas turbine, the cooling air is supplied into the platform 22 so that the platform 22 only may be cooled actively for the predetermined time, and when the platform 22 is so cooled for the predetermined time, the opening/closing valve 11 is closed by the control unit 10. In the conventional case, when the gas turbine is stopped, the platform 22, which has a mass larger than the moving blade 20, is slow in the temperature lowering to cause a large temperature difference between the thin blade 20 and the thick platform 22 and this causes large thermal stresses. But in the cooling system of the present invention, the platform 22 is cooled actively even after the stop of the gas turbine to accelerate the temperature lowering of the platform 22, so that no large temperature difference occurs between the moving blade 20 and the platform 22 and thereby occurrence of the thermal stresses is prevented and occurrence of the cracks can be suppressed. It is to be noted that, while the cooling system of the mentioned third embodiment has been described on the example where it is applied to a gas turbine moving blade in the prior art, this cooling system may be naturally applied to a gas turbine moving blade having constructions of the first and second embodiments and then the effect to prevent the occurrence of the cracks can be obtained further securely.

Claims: A gas turbine moving blade comprising: a platform (22) and a blade fitting portion where a blade (20) is fitted to said platform (22); a blade cooling passage (24a-24c,25a-25c) provided in the blade (20); a platform cooling passage (50a,50b) provided in said platform (22); cooling air blow holes (28,29,33) provided in and around the blade (20) so that the blade may be cooled by cooling air (40-43) flowing through said blade cooling passage (24a-24c,25a-25c), through said platform cooling passage (50a,50b) and out of the blade (20) through said cooling air blow holes (28,29,33); and a recessed portion (1) having a smooth curved surface and extending in a direction orthogonal to a turbine axial direction, provided in an end face portion of a rear side portion of said platform (22) near said blade fitting portion on the blade trailing edge side; wherein said blade fitting portion is formed having a fillet (R) exterior with a curved surface; and wherein said cooling air blow holes (33) provided in a blade trailing edge include a hole (2,3) provided in a blade hub portion positioned at a lowermost end of said cooling air blow holes (33) provided in the blade trailing edge, said hole (2,3) having a hole cross sectional area larger than that of each of said cooling air blow holes (33) provided in the blade trailing edge above said hole (2,3); and wherein there is applied a coating of a heat resistant material to said blade (20) and platform (22) such that said coating is applied to said blade fitting portions of blade leading edge and trailing edge portions thicker than to other portions of said blade, and such that said coating is applied to portions of said platform near and around the blade leading edge and trailing edge portions thinner than to other portions of said platform. A gas turbine moving blade as claimed in claim 1, wherein said curved surface of the fillet (R) exterior of said blade fitting portion is formed to an elliptical curve. A gas turbine moving blade as claimed in claim 1 or 2, wherein said platform cooling passage (50a, 50b) is connected with a platform cooling air supply system (81) and there are provided in said platform cooling air supply system (B1) an opening/closing valve (11) for opening and closing said platform cooling air supply system (81) and a control unit (10) for controlling said opening/closing valve so as to be closed while a gas turbine is operated and so as to be opened for a predetermined time when the gas turbine is stopped. A gas turbine moving blade as claimed in any one of claims 1 to 3, further comprising a shank portion (40b) for fixing said platform (22), said shank portion (40b) being formed in an elongated shape having a height (H) of said shank portion (40b) in a turbine radial direction larger than a width (W) of said shank portion in a turbine rotational direction.

Inventors: AOKI SUNAO, ARIMURA HISATO, HASHIMOTO YUKIHIRO, ISHIGURO TATSUO, KUBOTA JUN, SHIROTA AKIHIKO, SUENAGA KIYOSHI, TOMITA YASUOKI, TORII SHUNSUKE, AOKI, SUNAO, ARIMURA, HISATO, HASHIMOTO, YUKIHIRO, ISHIGURO, TATSUO, KUBOTA, JUN, SHIROTA, AKIHIKO, SUENAGA, KIYOSHI, TOMITA, YASUOKI, TORII, SHUNSUKE, Aoki, Sunao, Takasago Research &amp; Development Cent., Arimura, Hisato, Takasago Machinery Works, HASHIMOTO, Yukihiro, Takasago Machinery Works, Ishiguro, Tatsuo, Takasago Research &amp; Devel. Cent., Kubota, Jun, Takasago Machinery Works, Shirota, Akihiko, Takasago Machinery Works, SUENAGA, Kiyoshi, Takasago Machinery Works, TOMITA, Yasuoki, Takasago Machinery Works, Torii, Shunsuke, Takasago Machinery Works
Applicants: MITSUBISHI HEAVY IND LTD, MITSUBISHI HEAVY INDUSTRIES, LTD.
Application year: 2001
Main classification: F01D5
IPC codes: F01D5