Blade cooling air supplying system for gas turbine

Abstract: In the present invention, an air pipe (13) extends through a stationary blade (10) between outer (11) and inner (12) shrouds. Further, an air passage (19A) is directed to a lower portion of the stationary blade and is communicated with the air pipe so that a serpentine cooling passage is formed. The air enters a cavity (14) from the air passage (16a, 16b) and is discharged to a gas passage through an air hole, a passage and a seal (12a, 12b). Thus, the cavity is sealed at a high pressure. Cooling air is supplied from the air passage (13) to a rotating blade (20) through a cooling air hole (18), a cooling air chamber (24), a radial hole (25) and a lower portion of a platform. The stationary blade (10) is cooled by the air through the air passage (19A, 19B, 19C, 19D, 19E, 19F). The cooling air can be supplied to the rotating blade at a low temperature and a high pressure as they are. Accordingly, the air can be also supplied to the rotating blade when a rotor is cooled by vapor.

Description: FIELD OF THE INVENTION AND RELATED ART STATEMENT The present invention relates to a blade cooling air supplying system for effectively cooling a blade of a gas turbine by the air, and particularly to a system which makes it a possible to cool rotating blade (moving blade) by the air when a rotor is cooled by vapor. Fig. 2 is a cross-sectional view of the interior of a conventional general gas turbine showing a flow of cooling air to a rotating blade. In Fig. 2, reference numerals 50, 51 and 52 respectively designate a stationary blade, an outer shroud and an inner shroud. Reference numeral 60 designates a rotating blade constructed such that this rotating blade 60 is attached to a rotor disk blade root portion 62 of a turbine disk 61 and is rotated between stationary blades 50. In the gas turbine constructed by the stationary blade 50 and the rotating blade 60 mentioned above, the rotating blade 60 is cooled by the air and is adapted to be cooled by using one portion of the rotor cooling air. Namely, a radial hole 65 is formed in the rotor disk blade root portion 62 and the rotor cooling air 100 is guided to each disk cavity 64. The rotor cooling air 100 is guided through the radial hole 65 to a lower portion of a platform 63, and is supplied to the rotating blade 60. Fig. 3 is a detailed view of the stationary and rotating blades in the gas turbine of the above construction. In Fig. 3, the stationary blade 50 has the outer shroud 51 and the inner shroud 52. An air pipe 53 axially extends through the interior of the stationary blade 50. Namely, in this stationary blade 50, air 110 for seal is guided from a side of the outer shroud 51 to a cavity 54 and flows out to a passage 56 through a hole 57. A pressure within the passage 56 is increased in comparison with that in a combustion gas passage and one portion of this pressure flows into the combustion gas passage so as to prevent the invasion of a high temperature gas. Reference numeral 55 designates a labyrinth seal similarly used to seal the high temperature gas. As mentioned above, the cooling air supplied to the rotating blade 60 guides the rotor cooling air 100 into the disk cavity 64 and also guides the rotor cooling air 100 to a shank portion 61 surrounded by a seal plate 66 in a lower portion of the platform 63 through the radial hole 65 extending through the interior of the rotor disk blade root portion 62. The rotor cooling air 100 is then supplied from this shank portion 61 to a passage for cooling the rotating blade 60. The air from a compressor may be also cooled through a cooler instead of usage of one portion of the rotor cooling air and may be guided to the disk cavity 64. As mentioned above, the blades of the conventional gas turbine are cooled by the air and the rotating blade 60 is particularly cooled by guiding one portion of the rotor cooling air. In recent years, a cooling system using vapor instead of the air has been researched. When a rotor system is cooled by the vapor, no air for cooling can be obtained from the rotor so that no rotating blade can be cooled by the air in the conventional structure. With respect to the stationary blade 50, as explained with reference to Fig. 3, the air 110 for seal is blown out to the cavity 54 of the stationary blade 50 from the air pipe 53 extending through the interior of the stationary blade. Thus, the interior of the cavity 54 is held at a high pressure and the pressure of the passage 56 is set to be higher than the pressure of the combustion gas passage so that the invasion of a high temperature gas into the interior of the stationary blade is prevented. Namely, the air 110 for seal blown out to the cavity 54 partially flows out to the high temperature combustion gas passage through the hole 57 and the passage 56. When an amount of this flowing-out air is increased, efficiency of the gas turbine is reduced. OBJECT AND SUMMARY OF THE INVENTION An object of the present invention is to provide an improved blade cooling air supplying system of a gas turbine in which the air for cooling a rotating blade is supplied from a stationary blade to the rotating blade and in which means for supplying the air for sealing the stationary blade is also provided. The present invention provides a blade cooling air supplying system of a gas turbine which has plural rotating blades each attached to a rotor through a blade root portion and also has plural stationary blades arranged alternately with the rotating blades such that each of the stationary blades has outer and inner shrouds, a cavity for seal in a lower portion of the inner shroud, and a seal box in a lower portion of the cavity for seal, the blade cooling air supplying system comprising an air pipe extending through each of said stationary blades from the outer shroud to the inner shroud and inserted into said seal box, a rotating blade side cooling air introducing portion arranged in the blade root portion of each of said rotating blades and guide adapted so as to cooling air to each of said rotating blades, and a cooling air passage arranged in said seal box and communicating with said air pipe and opening toward an inlet of said rotating blade side cooling air introducing portion such that cooling air supplied to said air pipe is blown out from said cooling air passage to the inlet of said rotating blade side cooling air introducing portion and is sent from there to each rotating blade. GB 938,247 and US 3,945,758 disclose similar gar turbine blade cooling air supply systems. However, with the above object in view, the present invention is characterised in that the entirety of the cooling air supplied to said air pipe from an outer shroud side of each stationary blade is supplied to the rotating blade, while cooling air supplied to a leading edge portion passage of each stationary blade is sent afterwards as air for sealing to the cavity of each stationary blade. Accordingly, the cooling air can be directly supplied from each stationary blade to the rotating blade at a high pressure and a low temperature as they are. Accordingly, cooling effects of the rotating blade can be improved and the invention can be used as an air cooling system for the blades in a gas turbine in which the rotor is cooled by vapor. As already mentioned, the entirety of the cooling air from the air pipe is used to cool each rotating blade. The air for sealing each stationary blade is separately transmitted through a leading edge portion of the stationary blade and cools this leading edge portion. Thereafter, this air is used to pressurize the cavity. Accordingly, in the present invention, the cooling air is more effectively utilized than in the prior art. BRIEF DESCRIPTION OF THE DRAWINGS Fig. 1 is a cross-sectional view of root portions of stationary and rotating blades to which a blade cooling air supplying system in accordance with an embodiment of the present invention is applied; Fig. 2 is a cross-sectional view of a blade portion of a conventional gas turbine showing a flow of cooling air to the rotating blade; and Fig. 3 is a cross-sectional view of a rotating blade in which a cooling air supplying system to the rotating blade of the conventional gas turbine is applied. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS In Fig. 1, reference numeral 10 designates a stationary blade having an outside shroud 11 and an inner shroud 12. Reference numeral 13 designates an air pipe extending through the interior of the stationary blade and the air 100 for cooling is guided by this air pipe 13. Reference numeral 14 designates a cavity arranged in a lower portion of the inner shroud 12. A tube 13a connected to the air pipe 13 hermetically passes through the interior of the cavity 14. Reference numeral 15 designates a seal box for supporting a labyrinth seal 15a. Reference numerals 16a and 16b designate passages formed by seal portions 12a, 12b of the inner shroud 12 in both end portions thereof. Reference numeral 17 designates an air hole extending through the seal box 15 and communicating the cavity 14 with the passage 16a. Reference numeral 18 designates a cooling air passage arranged in the seal box 15. The cooling air passage 18 communicates the tube 13a continuously connected to the air pipe 13 with a cooling air chamber 24 on a rotating blade side. An air passage 19A for seal guides the air 101 from the outer shroud 11. Air passages 19B, 19C, 19D, 19E and 19F form a serpentine cooling flow passage. Reference numerals 20, 21 and 22 respectively designate an unillustrated rotating blade, a shank portion and a rotor disk blade root portion. This rotor disk blade root portion 22 has a projecting portion 22a. A seal portion 28 is formed between this projecting portion 22a and the seal box 15 of the stationary blade 10. Reference numerals 23 and 24 respectively designate a platform and a cooling air chamber in the blade root portion 22. The cooling air chamber 24 is formed by the projecting portion 22a, the seal chamber 28, the seal box 15 of the stationary blade 10 and the labyrinth seal 15a. The cooling air chamber 24 is communicated with the cooling air passage 18 arranged in the seal box 15 on a stationary blade side. Reference numeral 25 designates a radial hole formed in the rotor disk blade root portion 22. The radial hole 25 is communicated with the cooling air chamber 24 and an air reservoir 27 formed in the blade root portion 22 and the shank portion 21. Namely, an air introducing portion is constructed by the cooling air passage 24, the radial hole 25 and the air reservoir 27. Reference numeral 26 designates a seal plate in a lower portion of the platform 23. The passage 16b is formed by the seal plate 26 and the seal portion 12b on a stationary blade side. A turbulator 70 is arranged within the air passages 19A to 19F of the stationary blade 10 to provide turbulence to a cooling air flow and improve a heat transfer rate. In the above embodiment, the rotor is cooled by vapor and a vapor cavity 200 is arranged. The rotor is cooled by the vapor from the vapor cavity 200. The stationary blade 10 and the rotating blade 20 are cooled by the air. One portion of the air 101 first flows into the interior of the stationary blade from the outside shroud 11 through the passage 19A on a leading edge side. This air cools the leading edge and is blown out to the cavity 14 and passes through the air hole 17 of the seal box 15 and also passes through the passage 16a at a pressure equal to or higher than a predetermined pressure. The air then passes through the seal portion 12a and partially flows out onto the side of a high temperature gas passage. Accordingly, a rotor side of the combustion gas passage is held at a pressure higher than the pressure of the combustion gas passage by this air 101 for seal so that the invasion of a high temperature gas onto the rotor side of the combustion gas passage is prevented. The remaining portion of the air 101 enters the passage 19B and is moved upward in the passage 19C from a lower portion of the passage 19B. Serpentine cooling is performed while the remaining portion of the air 101 sequentially passes through the passages 19D, 19E and 19F and is partially discharged from a trailing edge side. After this cooling, the air at a high temperature passes through the passage 16b and flows out to a gas flow passage on the trailing edge side from the seal portion 12b. In contrast to this, the cooling air 100 flows into the air pipe 13 from the outside shroud 11 and passes through the tube 13a continuously connected to a lower portion of the air pipe 13. The cooling air 100 further enters the cooling air chamber 24 through the cooling air passage 18 and stays as cooling air at a high pressure and a low temperature. The cooling air entering the cooling air chamber 24 further enters the air reservoir 27 through the radial hole 25 on the rotating blade side, and is guided from the platform 23 to an air passage for cooling arranged in an unillustrated rotating blade 20, and cools the rotating blade 20. In the above-mentioned embodiment, the air for cooling the rotating blade is supplied from only the air pipe 13 arranged in the stationary blade 10 and the tube 13a. The air pipe 13 and the tube 13a constitute an independent route. Accordingly, the air for cooling the rotating blade is directly supplied to the rotating blade 20 while the high pressure and the low temperature of the air are maintained. Therefore, the rotating blade 20 can be effectively cooled. The air 101 for seal within the cavity 14 is independently supplied from the passage 19A at a leading edge. The air 101 passing through this passage 19A cools a leading edge portion and is then used as a seal. Accordingly, the air 101 can be used for both seal and cooling so that the air can be effectively utilized. In the blade cooling air supplying system in the first embodiment having such features, the air can be also supplied to the blades, especially the rotating blade 20 in the case of a gas turbine for cooling the rotor by vapor. Accordingly, the blades can be cooled by the air.

Claims: A blade cooling air supplying system of a gas turbine which comprises plural rotating blades (20) each attached to a rotor through a blade root portion (22), and plural stationary blades (10) arranged alternately with the rotating blades such that each stationary blade has outer (11) and inner (12) shrouds, a cavity (14) for seal in a lower portion of the inner shroud, and a seal box (15) in a lower portion of the cavity for seal; the system comprising an air pipe (13) extending through each of said stationary blades from the outer shroud to the inner shroud and inserted into said seal box (15); a rotating blade side cooling air (100) introducing portion (24, 25, 27) arranged in the blade root portion of each rotating blade (10) and adapted so as to guide cooling air to each rotating blade; and a cooling air passage (18) arranged in the seal box and communicating with said air pipe (13) and opening toward an inlet of said rotating blade side cooling air introducing portion (25) such that cooling air (100) supplied to said air pipe (13) flows through said cooling air passage (18) of said seal box (15) to the inlet of said rotating blade side cooling air introducing portion (25) and is conducted from there to each rotating blade (20), characterised in that substantially all of the cooling air (100) supplied to the air pipe (13) from an outer shroud side of the stationary blade (10) is supplied to the rotating blade (20) while, cooling air (101) supplied to a leading edge portion passage (19A) of each stationary blade (10) is supplied as air for sealing to the cavity (14, 16) of each stationary blade after cooling a leading edge portion of said stationary blade. A blade cooling air supplying system according to claim 1 wherein the air pipe (13) is hermetically connected to said cooling air passage (18) by a conduit (13a). A blade cooling air supply system according to claim 2 wherein said conduit (13a) passes through said cavity (14). A blade cooling air supply system according to claims 1, 2 or 3 wherein the rotating blade side cooling air introducing portion (24, 25, 27) is at least partially formed (25) in a blade root portion (22) of the rotor.

Inventors: FUKUNO HIROKI, HASHIMOTO YUKIHIRO, SUENAGA KIYOSHI, TOMITA YASUOKI, FUKUNO, HIROKI, HASHIMOTO, YUKIHIRO, SUENAGA, KIYOSHI, TOMITA, YASUOKI, Fukuno, Hiroki, Mitsubishi Heavy Ind., Ltd., Hashimoto, Yukihiro, Mitsubishi Heavy Ind., Ltd., Suenaga, Kiyoshi, Mitsubishi Heavy Ind., Ltd., Tomita, Yasuoki, Mitsubishi Heavy Ind., Ltd.
Applicants: MITSUBISHI HEAVY IND LTD, MITSUBISHI HEAVY INDUSTRIES, LTD.
Application year: 1998
Main classification: F01D5
IPC codes: F01D9, F01D5, F02C7, F01D11