WO1998058158A1 - Dispositif d'etancheite pour aubes de stator de turbine a gaz - Google Patents
Dispositif d'etancheite pour aubes de stator de turbine a gaz Download PDFInfo
- Publication number
- WO1998058158A1 WO1998058158A1 PCT/JP1998/002565 JP9802565W WO9858158A1 WO 1998058158 A1 WO1998058158 A1 WO 1998058158A1 JP 9802565 W JP9802565 W JP 9802565W WO 9858158 A1 WO9858158 A1 WO 9858158A1
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- WO
- WIPO (PCT)
- Prior art keywords
- air
- seal
- blade
- ring
- sealing
- Prior art date
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
Definitions
- the present invention relates to a gas turbine vane, in which the supply of sealing air is improved, the leakage of the ⁇ is reduced, the ⁇ is efficiently supplied to the inner shroud, and the rating is achieved by cooling the sealing air.
- the clearance between the rotor side and the stationary side during operation can be reduced.
- Figure 14 is a general block diagram of a gas turbine.
- the evening bin 1501 since the evening bin 1501 is exposed to the high-temperature combustion gas, the air from the compressor 150 is extracted once and supplied to the evening bin 1501, where it is supplied to the stationary blade, the rotor blade, and the rotor blade. Guiding and cooling them.
- FIG. 15 is a cross-sectional view showing a typical air supply path for sealing to stationary vanes of a conventional gas turbine, and shows the configuration of the blades in the evening bin 151 in FIG.
- 21 is a rotor blade
- 22 is its platform
- 23 is a seal plate at the bottom of the platform
- 24 and 25 are both ends of the platform
- 26 is a blade root
- the rotor blades 21 composed of these members are provided in multiple m in the circumferential direction of the mouth.
- Reference numeral 31 denotes a stationary blade, which is disposed adjacent to the rotor blade 21; 32, an outer shroud thereof; and 33, an inner shroud.
- 34, 35 are both ⁇ of the inner shroud
- -36 is the cavity at the lower part of the inner shroud 32
- 37 is a seal ring holding ring, which has a labyrinth seal 37a at its end, Wing root part 2 6 times ⁇ ⁇ minutes and sliding ing.
- Reference numeral 38 denotes an air hole, which is provided to penetrate the seal ring retaining ring 37, and defines a space between the cavity 36 and the blade root 26 of the adjacent rotor blade 21.
- Reference numerals 40a and 40b denote scene portions between the adjacent platform 22 and the inner shroud 33, respectively, and seal members are provided in the gaps between ⁇ 24 and 34, 25, and 35, respectively. Is provided.
- Reference numeral 50 denotes a blade ring, inside of which an outer shroud 32 of the stationary blade 31 is fixed via heat shield rings 32a and 32b.
- Numeral 51 denotes a space provided in the blade ring 50, which corresponds to a space 53 formed by the blade ring 50, the heat shield rings 32a, 32b, and the outer shroud 32, the tip of which is shown.
- Reference numeral 52 denotes a seal tube, which passes through the inner shroud 33 from the outer shroud 32 to the inside of the stator vane 31.
- the cooling air 54 from the compressor is supplied from the air 51 of 50, flows into the space 53, and on the other hand, passes through the seal tube 52 and the lower part of the inner shroud 33.
- the cooling air from the cavity 36 blows out from the air hole 38 into the rear space of the adjacent rotor blade 2 1 as shown by S 1, and further passes through the labyrinth seal 37 a to the rotor blade 2 1
- S 1 flows out of the seal part 40 a and S 2 flows out of the seal part 4 Ob, and the combustion gas flows inside the inner shroud 33 of the stator vane 31.
- the air entering the space 53 cools the surface of the outer shroud 32 and enters the cooling passage in the stationary blade as described in FIG. It is blown out from the hole and discharged into the combustion gas passage.
- the diameter of the hollow 1 is 2 to 50, and the inner diameter of the seal tube 52 is limited to be larger depending on the thickness of the blade and the degree of the warp of the blade. Therefore, the inflowing air is subjected to pressure loss and the pressure is reduced. In addition, the cooling air flowing into the space 53 is further reduced by the outer shroud 32 and the heat shield rings 32 a, 3
- the pressure of the cooling air 54 flowing into the air hole 51 of the blade ring 50 is about 6 kgZcrf, and about 5 kg / cnf due to pressure loss in the space 53
- the pressure inside the cavity 36 is further reduced by 3.5 kgZcnf due to the pressure loss, and is almost the same as the pressure 3.5 kg / cnf between the moving blade 21 and the stationary blade 31 adjacent to each other. The effect is reduced.
- the first problem with the conventional seal structure of the conventional gas vane bin stationary blade described above is that cooling air is supplied from the air holes 51 of the blade ring 50, and the cooling air is supplied between the blade ring 50 and the outer shroud 32. Flows into the space 53, and flows into the cavity 36 below the inner shroud 3 3 from the seal tube 52, but as described above, the outer shroud 32 and the heat shield rings 3 2a, 3 2b It leaks out of the gap.
- the seal tube 52 also receives a pressure loss, and when flowing into the cavity 36 of the inner shroud, the pressure force decreases. As a result, there is no difference with the pressure of the combustion gas, and it is difficult to obtain a sufficient pressure as the seal air.
- FIG. 16 is a cross-sectional view of a conventional gas vane vane having the same structure as the vane shown in FIG. 15, but mainly illustrating cooling of the vane.
- reference numeral 31 denotes a stationary blade, in which air passages 8 OA, 80 B, and 80 C are formed by fiber, forming a serpentine flow path, and 80 D is a trailing edge of the blade.
- Reference numeral 52 denotes a seal tube penetrating the inside of the stationary blade 31 vertically.
- Reference numeral 33 denotes an inner shroud, in which a cavity 36 is formed, and the lower end of the seal tube 52 is open.
- Reference numeral 37 denotes a seal ring holding ring, which holds the flange of the inner shroud 33 and the labyrinth seal 37a.
- Reference numeral 38 denotes an air hole provided in the retaining ring 37, which defines a space 72 between the cavity 36 and an adjacent moving blade.
- Numeral 32 denotes an outer channel, which is provided with a hole 62 for supplying cooling air. 2 1 is an adjacent bucket.
- cooling ⁇ 70 is supplied from the hole 6 2 of the outer shroud 3 2 to the air passage 8 OA on the leading edge side of the stationary blade 3 1, and flows inside to the next air passage 8 0 B, then flows outward, enters the adjacent air passageway 82C, continues to the inside, flows inward, cools the vanes 31 in turn, and passes through the air holes 60 at the trailing edge 80D. It flows along the outer surface of the trailing edge and cools the film.
- cooling air for sealing 71 was introduced, and as shown in FIG. Flows into the cavity 36, and the “ ⁇ ” flows out from the 3 ⁇ 4m hole 38 provided in the cavity 36 into the space 72 adjacent to the moving blade, and passes through the labyrinth seal 37 a to the space 7 ahead.
- cooling flows from the seal tube 52 into the cavity 36, and the inside of the cavity 36 is maintained at a higher pressure than the external combustion gas passage, so that the high-temperature combustion gas enters the inside.
- the air from the compressor passes through the disk cavity and enters the inside of the platform 22 from the radial hole provided on the blade root 26, and the air moves from here. Guided by wing 21, cooling blade 2 1.
- an air passage is provided inside to cool the vane, and this air passage is usually a Saintine passage, and the cooling air from the outer shroud is The air flows into the air passage to cool the inside of the stator vanes, and then discharges to the outside from the trailing edge.
- a seal tube is passed through the stator vane to supply the outer shroud force and the cooling air to the inside of the inner shroud cavity for sealing, and this is supplied to the outside combustion gas passage at a higher pressure.
- a large amount of cooling air is supplied to the stationary blade of the above configuration for cooling and for sealing.
- the cooling air 3 ⁇ 4m is discharged from the trailing edge to the combustion gas fiber after cooling the stationary blade, and for sealing.
- As the air a part of the cooling air is extracted, supplied to the cavity via the seal tube, and discharged to the space between the cavity and the adjacent front and rear rotor blades. Therefore, the second problem of the sealing device for gas turbine vanes is that, in addition to the pressure loss of the first problem, a large amount of air is consumed for cooling and sealing, and the This is due to the large capacity and the large power burden on the performance of the gas bin.
- FIG. 17 is a cross-sectional view showing the general arrangement of blades of a gas turbine, and shows the entire arrangement of the stationary blades shown in FIGS.
- 81 C, 82 C, 83 C, and 84 C are stationary vanes, each of which is arranged radially around the mouth and mounted on the stationary side.
- 8 1 S, 8 2 S, 8 3 S, and 8 4 S are rotor blades, each of which is mounted around the mouth through the blade root portion, and is disposed in the axial direction with the stator blade ⁇ !: It rotates with the rotor.
- 1 1 1 C, 1 1 2 C, 1 1 3 C, 1 1 4 C are static
- the inner shrouds of the wings 81C to 84C, 111S, 112S, 113S, and 114S are platform forms of the rotor blades 81S-84S, respectively.
- 3 7-1, 3 7-2, 3 7-3 are seal ring retaining rings, each of which is a stationary blade 8
- Fig. 17 Up to 84 C inner shroud 11 11 C to l 14 C It is fixed to the flange and has an annular shape placed around the rotor. Inside it, a labyrinth seal (close to the rotor) In this way, the example shown in Fig. 17 is a gas turbine in which the stationary blades and the moving blades are each composed of four stages. It rotates in the evening and defines a generator.
- the stationary blades, the moving blades and the inlet section of the gas turbine have an inlet combustion gas temperature of 800 ° C to 100 ° C or a recent gas.
- a 150 ° C class has been developed, and since it is exposed to this high-temperature gas, cooling 3 ⁇ 4m from the compressor is extracted and guided to cool them.
- the clearance CR ' is held on the surface facing the rotor.
- the clearance between the turbine stages is minimized from startup to the rated speed, and then increased to the rated speed by heating the combustion gas. At that point, the clearance CR 'becomes larger than the minimum clearance. It is preferable that the clearance CR 'be as small as possible, because the sealing performance is improved, and it is preferable.
- the design value becomes the minimum value after the start-up. Fiber, Pf Kureshi, etc. cannot be made too small. Therefore, the third problem is that when the rated rotation is reached after operation, a large clearance is created and the sealing performance is degraded.
- the clearance as a factor to improve the reduction of the sealing pressure due to the pressure loss of the first problem and the consumption of a large amount of air of the second problem is optimized, and this clearance is reduced during operation. It is desired. Disclosure of the invention
- the present invention is directed to solving the first problem by providing a seal space supplied from the blade ring into the stator vane.
- the air supply system is devised to reduce the amount of air leaking from the outer shroud and to supply sealing air with sufficient pressure to the inner shroud to increase the sealing effect of the gas turbine turbine vane sealing device. That's the challenge.
- the present invention relates to a sealing device having such an enhanced cooling effect, in which, in addition to the above-mentioned improvement of the sealing effect, the present invention employs a structure in which a seal tube for supplying a seal can be attached and detached. It is an object of the present invention to provide a gas turbine vane sealing device that can be easily removed for time and maintenance and has a simple structure to achieve a sealing effect.
- the present invention improves the supply system of the cooling air for the blade and the air for sealing in the stationary blade of the gas bottle, and uses both the cooling air and the sealing air.
- the task was to improve cooling of the gas turbine and contribute to the improvement of gas turbine performance.
- the rated speed was to improve the sealing performance by cooling the seal ring with air for sealing so that the clearance at the time of reaching is smaller than before.
- the present invention provides the following means (1) to (4) in order to solve the above respective problems.
- the air passes through the space formed by the g3 ⁇ 4, the heat shield ring and the outer shroud, and the air is guided from the space to the inner shroud through the seal tube in the stator vane, and the cavity of the inner shroud
- the air hole of the self-supporting wing ring communicates with the air hole communicating with the seal tube and the space.
- a sealing device for a gas turbine vane comprising an air hole.
- the sealing device for a gas vane bin stationary vane wherein the seal tube is detachably connected to a g ⁇ ⁇ air hole communicating with the seal tube.
- the air from the compressor is extracted and passed through the space formed by the outer shroud, guided into the stator vane from the same space, passed through the stator vane, and formed by the inner shroud and the seal ring retaining ring.
- the air extracted from the self-compressor is cooled by a cooler to the sealing device of the gas turbine vane, which guides the inside of the formed cavity and sets the inside of the inside shroud to a higher pressure than the combustion gas passage to seal the inside of the inside shroud.
- the air turbine cooled by the cooler cools the holding ring of the self-sealing ring with air cooled by the cooler.
- the space formed by the blade ring, the heat shield ring, and the outer shroud is supplied from the air communicating with this space to cool the surface of the outer shroud and to cool the inner shroud.
- the cooling air is guided to the cooling passage, and is discharged into the combustion gas passage at the trailing edge of the blade while cooling the inside of the blade.
- the seal tube is supplied with air from the air that communicates with the seal tube, and since this seal tube is a space, the amount of air that leaks from the space through the gap between the outer shroud and the heat shield ring connection is reduced. It is not affected and can supply air to the inner shroud with almost no pressure loss. Therefore, the cavity in the inner shroud can be maintained higher than the pressure in the combustion gas passage, and the sealing effect can be enhanced.
- the tip of the seal tube since the tip of the seal tube is detachably connected to the air hole, the tip of the seal tube inserted into the stator vane can be easily inserted into the air at the time of the stationary vane or maintenance. It can be fixed and can be easily removed.
- the air for cooling the vanes passes through the air passage inside the vanes and is guided to the cavity as a seal for the stationary vanes, so that no sealing tube force is required as in the conventional case.
- the cooling air from the outer shroud was guided directly to the cavity from the seal tube as sealing air, and the cooling air was wasted to the combustion gas passage from the trailing edge after cooling the blades.
- the cooling air passes through each of the insides of the stationary blades, cools the blades, and then passes through the holes in the inner shroud. It flows into the cavity for sealing, and the inside of the cavity is kept at a high pressure from the outside to prevent high-temperature gas from entering the inside. Therefore, the need for a seal tube is eliminated, which contributes to cost reduction, effectively utilizes cooling air, reduces cooling air, and improves the performance of the gas turbine.
- the clearance between the stationary side and the roll side during the rated operation can be further reduced, in addition to the improvement of the sealing effect by the above configurations (1) and (2).
- the seal ring retaining ring of the gas bottle is an annular shape attached to the flange of the inner shroud of the stationary blade, and holds the seal ring inside thereof.
- the seal ring that is, the mouth rotates while maintaining a predetermined clearance between the stationary side and the rotor side.
- the smaller the clearance the more the sealing performance improves.
- the clearance will be much larger than the initially set clearance, and the sealing performance will decrease.
- this clearance is determined by the difference between the dimensions for which the stationary side and the mouth and evening side are thanked, and the characteristics 14 are different between the stationary side and the mouth and evening side, and the stationary side is larger than the rotor side.
- the difference becomes larger than the initially set clearance.
- the seal ⁇ is cooled by a cooler, and the seal ring holding ring is cooled by this air. Therefore, the seal ring holding ring, that is, the stationary side is activated (cold) and operated (hot). ) And the difference of ⁇ Jg becomes smaller than before without cooling. As a result, the translation on the stationary side during the rated operation is smaller than before, and the clearance when the rated number of revolutions is reached, that is, the difference between the claims on the stationary side and the rotor side is also smaller than before. As a result, the air for sealing is reduced and the sealing performance is improved.
- the seal ring holding ring since the seal ring holding ring is cooled, the seal ring holding ring, that is, the thermal expansion on the stationary side becomes gentler from the start-up than before, so that the initial setting clearance can be set larger than before. Because it is possible, it is advantageous in design and manufacturing.
- FIG. 1 is an overall configuration diagram of a gas evening bin plant to which the gas evening bin stationary blade of the present invention is applied.
- FIG. 2 is an overall cross-sectional view of the gas turbine stationary blade sealing device according to the first embodiment of the present invention.
- FIG. 3 is a perspective view showing an assembled state of the gas turbine vane sealing device according to the first embodiment of the present invention.
- FIG. 4 is a cross-sectional view of the seal tube tip ⁇ of the sealing device for a gas turbine bin stationary blade according to the first embodiment of the present invention.
- FIG. 5 is a cross-sectional view showing another example of the distal end portion of the seal tube of the sealing device for a gas turbine bin according to the first embodiment of the present invention.
- FIG. 6 is a cross-sectional view of a tip portion of a seal tube of a seal device for a gas turbine stationary blade according to a second embodiment of the present invention.
- FIG. 7 is a cross-sectional view showing another example of the seal tube tip ⁇ of the sealing device for a gas jet bin stationary blade according to the second embodiment of the present invention.
- FIG. 8 is a cross-sectional view showing an application example of FIG. 6 at a tip portion of a seal tube of a seal device for a gas turbine stationary blade according to a second embodiment of the present invention.
- FIG. 9 is a cross-sectional view showing an application example of FIG. 7 at a seal tube tip ⁇ of a seal device for a gas turbine stationary blade according to a second embodiment of the present invention.
- FIG. 10 is a cross-sectional view of the vicinity of the seal ring retaining ring of the gas turbine stationary blade according to the first and second embodiments of the present invention.
- FIG. 11 is a characteristic diagram showing a comparison between the conventional example of the low side and the stationary side due to cooling of the gas ring bottle seal ring retaining ring according to the first and second embodiments of the present invention. You.
- FIG. 12 is a cross-sectional view of a sealing device for a gas turbine stationary blade according to a third embodiment of the present invention.
- FIG. 13 is a sectional view taken along the line A-II in FIG.
- Fig. 14 is a general configuration diagram of a conventional gas bin.
- FIG. 15 is a cross-sectional view of a conventional seal structure for a gas turbine bottle.
- Figure 16 is a cross-sectional view of a conventional gas turbine stationary blade, showing the flow of sealing air and cooling air.
- FIG. 17 is a diagram showing a blade arrangement and a seal ring retaining ring of a conventional gas turbine.
- Figure 1 shows the overall configuration of the gas turbine.
- the gas turbine consists of a JE compressor 150, a turbine 151, and a combustor 152, and burns with air from the compressor 150.
- the fuel is burned in the unit 15 2 to generate hot combustion gas, which is sent to the turbine 15 1.
- the air from the compressor 150 is extracted by the extraction line 153, sent to the turbine 151, guided to the inside of the nozzle, and into the moving and stationary blades.
- a cooler 154 is provided in the middle of the bleed line 153 according to il to cool the cooling air from the compressor 150.
- FIG. 2 is a sectional view of a sealing device for a stationary blade in the turbine 15 1 of the gas turbine described above, and shows a first embodiment of the present invention.
- FIG. 2 the structure of the moving blades indicated by reference numerals 21 to 26, the structure of the stationary blades indicated by reference numerals 31 to 38, and the entirety of the stationary blades and the moving blades indicated by 40, 50, 51, 53 Since the structure has the same function as the conventional one shown in FIG. 15, the description thereof is omitted, and the force to be quoted and described as it is ⁇
- the characteristic parts of the present invention are those indicated by 1 to 6, and will be described in detail below.
- 1 is an empty space, which is provided separately from the empty space 1 in E 050, and is 1 in the space 53.
- Reference numeral 2 denotes a seal tube, which passes from the outer shroud 3 2 to the inside of the stator vane 3 1, penetrates the inner shroud 3 3, and is inserted up to the cavity 36, and a seal tube distal end 3 is provided at a distal end of the seal tube 3. It extends inside and is inserted into the sky 1.
- Reference numeral 4 denotes a projection fixed around the tip 3 of the seal tube 2
- reference numeral 5 denotes a locking portion provided on the outlet side of the cavity 51 of the blade ring 50 and having an enlarged inner diameter
- Reference numeral 6 denotes a coil-shaped panel, which is provided around the distal end portion 3 of the seal tube, one end of which enters the empty space 51 with an elastic force applied, and is locked to the step portion of the locking portion 5, The other end is held down by the wei part 4, and the elastic force presses the protrusion part 4, fixing the seal tube tip ⁇ 3.
- FIG. 3 is a perspective view showing the stationary state of the vane.
- the seal tube 2 When assembling the stator vane 31 as shown in the figure, the seal tube 2 is inserted into the air hole 51 of the blade ring 50, and the seal tube 2 is attached to the outer shroud 32 with the spring 6 attached. Insert 2 into hole 5 1 and cover with 0. Next, the heat shield rings 32a and 32b are fed one by one in the circumferential direction (R direction), and the outer shroud 32 is attached. Removal does the reverse. When assembled in this manner, the spring 6 extends between the projection 4 of the seal tube tip 3 and the enlarged locking portion 5 of the hole 51 by the elastic force thereof, and the seal tube 2 is moved to the blade ring 5. Can be fixed to 0.
- FIG. 4 is a cross-sectional view showing a state in which the seal tube tip 3 is attached to the air hole 51.As shown in the figure, the seal tube tip 3 slightly enters the air hole 51, and the periphery thereof Panel 6 has been.
- the spring 6 is inserted with an elastic force applied between the protrusion 4 provided on the seal tube 2 and the engaging portion 5 having an enlarged inner diameter in the air hole 51 of the blade ring 50, and is inserted.
- the seal tube 2 is fixed with the elastic force inserted into the hole of the blade ring 50.
- FIG. 5 is a cross-sectional view showing another example of the first embodiment in which the seal tube tip is attached to the air hole.
- a detachable adapter 7 is inserted into the tip of the seal tube 2, and The tip is inserted into the hole 51 of the wing ring 50.
- the adapter 7 is provided with a part 8, and is inserted between the part 8 and the enlarged locking part 5 provided in the air hole 51 while giving a panel 6 elasticity, The seal tube is fixed by the elastic force.
- the seal tube 2 is fixed via the adapter 7.
- a part of the cooling air 54 extracted from the compressor flows through the hole 1, and the blade ring 50, the outer shroud 32 and It flows into the space 53 formed by the heat shield rings 32a and 32b.
- the air that has entered the space 53 cools the surface of the outer shroud 32 as before, and enters the cooling passage inside the stationary vane 31 (not shown), as described in the conventional example of FIG. It is blown out from the trailing edge while cooling the inside of the blade, and is discharged into the combustion gas passage.
- ⁇ 3 ⁇ 4 of the air entering this space 53 is the force that leaks into the combustion gas passage as shown by S3 and S4 from the gap between the outer shroud 32 and the heat shield rings 32a and 32b. of Since it is independent of the system of the air hole 51 and the seal tube 2, it does not affect the pressure of the sealing air.
- the cooling air 54 flows into the seal tube 2 from the hole 51, enters the lower cavity 36, and then passes through the hole 38 provided in the seal ring retaining ring 37. As shown by the middle S1, the gas flows out into the space between the adjacent moving blades 21 and the stationary blades 31, and from there flows through the seal portion 40a into the combustion gas passage.
- the air in the cavity 36 passes through the air hole 38, passes through the labyrinth seal 37a, enters the space with the adjacent rotor blade 21 on the downstream side, and passes through the seal portion 4Ob to form S.
- the cooling air 54 is 6 kg / crfSS, and when it flows into the holes 1 and 51, the space 53 has a pressure loss due to 3 ⁇ 4 ⁇ from the sky 1.
- ⁇ m which is about 5 kgZcrf as in the past
- ⁇ ⁇ m flowing from the space 51 is independent of the space 53, and flows into the cavity 36 from the seal tube 2 with almost no pressure loss. And maintain about 5 kgZcm 2 in the cavity 36.
- the inside of the cavity 36 has been reduced to about 3.5 kg / cm 2 due to the pressure loss of the air flowing from the seal tube 52.
- 5 kg / cm 2 And can maintain high pressure. Due to the high pressure in the cavity 36, as described above, air flows out of the seal portions 40a and 40b into the combustion gas passages as indicated by S1 and S2, and the combustion gas passages are approximately 3. Since the pressure is about 5 kgZcni 2 , the sealing pressure increases, and a sufficient sealing effect can be obtained.
- the spring 6 is also compressed during assembly and removal of the stator vanes, and the tip of the seal tube 2 is slightly inserted into the air hole 51 of the blade ring 50, and the panel 6 is easily extended and fixed. If the structure shown in FIG. 5 is adopted, mounting and listening can be easily performed by operating the adapter 1 and the panel 6. Furthermore, in the case of fixing with panel 6, the expansion of the seal tube can be performed even if it is extended due to the force of the seal tube. can do. 6 to 9 show a seal device for a gas turbine stationary blade according to a second embodiment of the present invention. The difference from the first embodiment of 3 ⁇ 4M is that a blade 9 at a seal tube tip 3 using a bellows 9. It is at the place where the insertion part of the ring 50 into the air hole 51 is sealed. The other configuration is the same as that of the first embodiment, and the description is omitted.
- FIG. 6 a bellows 9 is provided in the configuration in FIG. 4, and the seal tube tip 3 is sealed.
- the upper end of the bellows 9 is fixed around the air hole 51 of the wing ring 50 by brazing, and the lower end is fixed around the seal tube 2 by brazing.
- Fig. 7 shows the configuration of Fig. 5 with a rose 9 attached.
- the lower end of the rose 9 is fixed around the adapter 7. The rest is the same as in Figure 6 above.
- FIG. 8 shows the configuration of FIG. 6, in which the spring 6 and the projection 4 are eliminated, and only the bellows 9 is attached in the same manner.
- FIG. 9 shows the configuration of FIG. 7 in which not only the bellows 9 but also the blank are added.
- the lower end of the bellows 9 is fixed by attaching or the like.
- the lower end may be fixed or fastened by elastic force or the like.
- FIG. 10 is an enlarged cross-sectional view of the vicinity of the seal ring holding ring in the first and second embodiments of the present invention, showing a portion of a second-stage stationary blade.
- the air extracted from the compressor 150 is cooled by the cooler 154 from the seal tube 2 of the stator vane 3 1 using the cooler 54 shown in FIG. 1, and the cooling air cooled by the cooler 154 is supplied to the cavity. It is guided into 36 and keeps the cavity at a high pressure as before while cooling the seal ring retaining ring 37.
- ⁇ ⁇ and ⁇ 2 are the radial diameters of the front and rear stages opposite to the seal ring attached to the seal ring holding ring 37, respectively, and ⁇ is the seal ring holding ring 22.
- 9 shows the thermal elongation of the end face of the held seal ring on the stationary side.
- delta beta has by Uni AA i which will be described later, and different Shasaru beauty JP I 1 production and [Delta] [alpha] 2 .
- FIG. 11 shows the thermal expansion in the radial direction between the stationary side and the evening side.
- FIG. 7 is a diagram showing a comparison between a conventional device without cooling 37 and a device of the present invention where cooling is performed.
- ⁇ ⁇ ' is the thermal elongation of the end face of the labyrinth seal 37 a (seal ring) supported by the seal ring retaining ring 37. It shows that the raw material is "I,” which is gradually saturated after the operation from the initial clearance S, and is almost constant and saturated at 4 mmliUi at the rated speed.
- ⁇ ,, ⁇ 2 are observations of the surface facing the seal ring on the rotor side, and than 10 minutes after the operation, thanks were given more rapidly than ⁇ ⁇ , and then gradually increased.
- the motor is saturated at a speed lower than ⁇ ⁇ ⁇ ⁇ at the rated speed.
- ⁇ ,, ⁇ 2 show almost the same thermal elongation, but ⁇ A! Is the former side, and the thermal elongation is somewhat higher, but both are characterized by almost the same thermal elongation.
- ⁇ ⁇ indicates the seal ring end face of the present invention when ⁇ is cooled by the cooler 154 and the seal ring retaining ring 37 is cooled, and the initial value is S 2 , and the conventional S, Is set to be rather large, and its characteristics are more gradual than the conventional ⁇ ⁇ ', and are saturated at low speeds at rated speed.
- the mi monkey beauty delta beta of the present invention the initial clearance is an S 2, Yes is set larger than the conventional S 1, the air cooled by the cooler 5 4 as described above, shea one Luling Cooling the retaining ring 37, the difference in ffi between the startup (cold) and the operation (hot) is smaller than that of the conventional one, and its claim B is also slower than ⁇ 'and the minimum clearance MC R is also about 10 minutes later, which is slower than before.
- the air for sealing to the vanes at each stage is cooled by the cooler 154 and the cavity 3 is cooled. 6 and cools the seal ring retaining ring 3 7 so that the iUS difference between startup and operation is smaller than that of the conventional case where the seal ⁇ is not cooled.
- the clearance with the stationary seal can be reduced.
- FIG. 12 is a cross-sectional view of a gas turbine vane according to a third embodiment of the present invention
- FIG. 13 is a cross-sectional view taken along line AA of FIG.
- reference numeral 31 denotes a stationary vane, in which air passages 80 A, 80 B, and 80 C are sequentially communicated in the same manner as in the prior art to form a serpentine flow path.
- 80 D is the trailing edge, and is provided with a large number of film cooling air holes 60. 6 4 is the evening sun, the air:
- the fins on the inner wall of each passage of 8 OA, 80 B, 80 C each disturb the flow of the cooling air flowing in This is for improving the heat transfer efficiency.
- 3 3 is an inner shroud, inside of which a cavity 36 is formed.
- Reference numeral 37 denotes a seal ring retaining ring, which holds the inner ffi! (The flange of the shroud 33 and the labyrinth seal 37a.
- Reference numeral 38 denotes an air hole provided in the seal ring retaining ring 37. A space 72 between the cavity 36 and the adjacent rotor blade is provided.
- Numeral 6 3 denotes an empty space, which is provided in the inner shroud 33 and flows into the cavity 36 after cooling the stationary vanes 31 by flowing through the air passageway 80 C to form an empty space for sealing. It is a hole to be used as a mind.
- 3 2 is an outer shroud provided with a hole 6 2 for supplying cooling air, and the hole 6 2 is a stationary blade 3 1 0 A. 2 1 is an adjacent bucket.
- Reference numerals 72 and 73 denote spaces between adjacent blades.
- the cooling blade 70 is supplied to the air passage 8 OA on the leading edge side of the stationary blade 31, and the cooling air 70 is supplied to the blade 62 of the third embodiment as described above. It flows inward and flows into the next air passage 80B, flows outward and enters the next adjacent air passage 80C, and flows inward while flowing out from the trailing edge sky ⁇ : 60 As the edge is film cooled, the remaining cooling flows into cavity 36 through holes 63 provided in inner shroud 33.
- the cooling air flowing into the cavity 36 is connected to the blades adjacent to the air holes 63 provided in the seal ring holding ring 37. It flows out into the space 72 and also flows out into the space 73 ahead through the labyrinth seal 37a.
- the cooling air 70 that has passed through the air passages 80 A, 80 B, and 80 C in the vane 31 and cooled the vanes 31 has a shroud 33 inside the air passage 80 C.
- the cavity 36 enters the cavity 36 through the air hole 63, and the inside of the cavity 36 is maintained at a higher pressure than the combustion gas passage in the tank portion, thereby preventing the outside high-temperature combustion gas from entering the inside.
- the sealing air flowing into the cavity 36 conventionally passes through the seal tube 52 into the stator vane 31 as shown in FIGS. 15 and 16.
- FIG. 15 and 16 In the first and second embodiments, as shown in FIG.
- a seal tube 2 is provided so as to penetrate inside the stationary blade 31, and a part of the cooling air is supplied from the outer shroud 32.
- cooling ⁇ m after cooling the stationary blade 31 is supplied into the cavity 36 from the empty m ⁇ 63 of the inner shroud 33.
- the conventional seal tube 13 and the first and second embodiments of the seal tube 2 are not required, and the air after cooling of the stationary blade 11 is used as the sealing air, so that cooling is effective. It can be used and cooling can be seen, contributing to the improvement of gas turbine performance. Available for H ⁇ 14
- the seal tube is detachably connected to a hole of a blade ring communicating with the seal tube.
- the tip of the seal tube can be easily inserted and fixed into the hole of the blade ring at the time of ffiir or maintenance of the stator blade, and the ear can be easily heard.
- the present invention also eliminates the need for a conventional seal tube, thereby contributing to cost reductions, and also uses the air after cooling of the stationary blades for sealing, so that the cooling air can be used effectively and the cooling air m mforce is reduced. And the performance of the gas turbine is improved.
- the seal ring retaining ring is cooled, and the mi side on the stationary side is smaller than before, so that the clearance between the stationary side and the mouth-evening side at the rated rotation is smaller than the conventional case without cooling. It can be made smaller and the sealing performance is improved.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
Claims
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP98924566A EP0919700B1 (en) | 1997-06-19 | 1998-06-10 | Device for sealing gas turbine stator blades |
US09/230,820 US6217279B1 (en) | 1997-06-19 | 1998-06-10 | Device for sealing gas turbine stator blades |
DE69825959T DE69825959T2 (de) | 1997-06-19 | 1998-06-10 | Vorrichtung zum dichten der leitschaufeln von gasturbinen |
CA002263508A CA2263508C (en) | 1997-06-19 | 1998-06-10 | Sealing device for gas turbine stator blades |
Applications Claiming Priority (6)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP16265197A JPH1113487A (ja) | 1997-06-19 | 1997-06-19 | ガスタービン静翼 |
JP9/162651 | 1997-06-19 | ||
JP17573497A JPH1122412A (ja) | 1997-07-07 | 1997-07-01 | ガスタービンシールリング保持環の冷却方法 |
JP9/175734 | 1997-07-01 | ||
JP18226397A JP3310909B2 (ja) | 1997-07-08 | 1997-07-08 | ガスタービン静翼のシール装置 |
JP9/182263 | 1997-07-08 |
Publications (1)
Publication Number | Publication Date |
---|---|
WO1998058158A1 true WO1998058158A1 (fr) | 1998-12-23 |
Family
ID=27322037
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/JP1998/002565 WO1998058158A1 (fr) | 1997-06-19 | 1998-06-10 | Dispositif d'etancheite pour aubes de stator de turbine a gaz |
Country Status (5)
Country | Link |
---|---|
US (1) | US6217279B1 (ja) |
EP (1) | EP0919700B1 (ja) |
CA (1) | CA2263508C (ja) |
DE (1) | DE69825959T2 (ja) |
WO (1) | WO1998058158A1 (ja) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP1057974A3 (en) * | 1999-05-31 | 2004-01-21 | Nuovo Pignone Holding S.P.A. | Stator nozzle for gas turbines |
Families Citing this family (66)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
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CN114412583A (zh) * | 2022-01-21 | 2022-04-29 | 中国联合重型燃气轮机技术有限公司 | 透平和具有其的燃气轮机 |
US12000308B2 (en) * | 2022-08-23 | 2024-06-04 | General Electric Company | Rotor blade assemblies for turbine engines |
Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS6229701A (ja) * | 1985-07-30 | 1987-02-07 | Toshiba Corp | ガスタ−ビン |
JPS648173B2 (ja) * | 1980-05-19 | 1989-02-13 | Asea Buraun Boeri Ag | |
JPH02233801A (ja) * | 1989-02-06 | 1990-09-17 | Westinghouse Electric Corp <We> | ガスタービン及びその羽根の冷却方法 |
JPH0552102A (ja) * | 1991-08-23 | 1993-03-02 | Toshiba Corp | ガスタービン |
JPH05288002A (ja) * | 1992-04-07 | 1993-11-02 | Mitsubishi Heavy Ind Ltd | 空冷セラミック静翼 |
JPH06257405A (ja) * | 1992-12-30 | 1994-09-13 | General Electric Co <Ge> | タービン |
JPH0754669A (ja) * | 1993-08-09 | 1995-02-28 | Mitsubishi Heavy Ind Ltd | ガスタービン冷却空気制御装置 |
JPH0960531A (ja) * | 1995-08-25 | 1997-03-04 | Mitsubishi Heavy Ind Ltd | ガスタービン |
Family Cites Families (18)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
NL70901C (ja) | 1945-01-23 | |||
BE514728A (ja) * | 1951-10-10 | |||
US2919891A (en) * | 1957-06-17 | 1960-01-05 | Gen Electric | Gas turbine diaphragm assembly |
GB938247A (en) * | 1962-03-26 | 1963-10-02 | Rolls Royce | Gas turbine engine having cooled turbine blading |
US3275294A (en) * | 1963-11-14 | 1966-09-27 | Westinghouse Electric Corp | Elastic fluid apparatus |
US3647311A (en) * | 1970-04-23 | 1972-03-07 | Westinghouse Electric Corp | Turbine interstage seal assembly |
US3945758A (en) * | 1974-02-28 | 1976-03-23 | Westinghouse Electric Corporation | Cooling system for a gas turbine |
US3892497A (en) * | 1974-05-14 | 1975-07-01 | Westinghouse Electric Corp | Axial flow turbine stationary blade and blade ring locking arrangement |
US4113406A (en) * | 1976-11-17 | 1978-09-12 | Westinghouse Electric Corp. | Cooling system for a gas turbine engine |
US4292008A (en) * | 1977-09-09 | 1981-09-29 | International Harvester Company | Gas turbine cooling systems |
JPS648173A (en) | 1987-06-29 | 1989-01-12 | Toshiba Corp | Image forming device |
US4930980A (en) * | 1989-02-15 | 1990-06-05 | Westinghouse Electric Corp. | Cooled turbine vane |
US5407237A (en) | 1993-02-26 | 1995-04-18 | The United States Of America As Represented By The Secretary Of The Air Force | Flexible coupling |
US5488825A (en) | 1994-10-31 | 1996-02-06 | Westinghouse Electric Corporation | Gas turbine vane with enhanced cooling |
DE69515502T2 (de) | 1994-11-10 | 2000-08-03 | Siemens Westinghouse Power Corp., Orlando | Gasturbinenschaufel mit einer gekühlten plattform |
JP3426902B2 (ja) * | 1997-03-11 | 2003-07-14 | 三菱重工業株式会社 | ガスタービン冷却静翼 |
JP3416447B2 (ja) * | 1997-03-11 | 2003-06-16 | 三菱重工業株式会社 | ガスタービンの翼冷却空気供給システム |
JP3495554B2 (ja) * | 1997-04-24 | 2004-02-09 | 三菱重工業株式会社 | ガスタービン静翼の冷却シュラウド |
-
1998
- 1998-06-10 EP EP98924566A patent/EP0919700B1/en not_active Expired - Lifetime
- 1998-06-10 DE DE69825959T patent/DE69825959T2/de not_active Expired - Lifetime
- 1998-06-10 CA CA002263508A patent/CA2263508C/en not_active Expired - Lifetime
- 1998-06-10 WO PCT/JP1998/002565 patent/WO1998058158A1/ja active IP Right Grant
- 1998-06-10 US US09/230,820 patent/US6217279B1/en not_active Expired - Lifetime
Patent Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS648173B2 (ja) * | 1980-05-19 | 1989-02-13 | Asea Buraun Boeri Ag | |
JPS6229701A (ja) * | 1985-07-30 | 1987-02-07 | Toshiba Corp | ガスタ−ビン |
JPH02233801A (ja) * | 1989-02-06 | 1990-09-17 | Westinghouse Electric Corp <We> | ガスタービン及びその羽根の冷却方法 |
JPH0552102A (ja) * | 1991-08-23 | 1993-03-02 | Toshiba Corp | ガスタービン |
JPH05288002A (ja) * | 1992-04-07 | 1993-11-02 | Mitsubishi Heavy Ind Ltd | 空冷セラミック静翼 |
JPH06257405A (ja) * | 1992-12-30 | 1994-09-13 | General Electric Co <Ge> | タービン |
JPH0754669A (ja) * | 1993-08-09 | 1995-02-28 | Mitsubishi Heavy Ind Ltd | ガスタービン冷却空気制御装置 |
JPH0960531A (ja) * | 1995-08-25 | 1997-03-04 | Mitsubishi Heavy Ind Ltd | ガスタービン |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP1057974A3 (en) * | 1999-05-31 | 2004-01-21 | Nuovo Pignone Holding S.P.A. | Stator nozzle for gas turbines |
Also Published As
Publication number | Publication date |
---|---|
EP0919700B1 (en) | 2004-09-01 |
EP0919700A4 (en) | 2000-12-13 |
DE69825959D1 (de) | 2004-10-07 |
US6217279B1 (en) | 2001-04-17 |
EP0919700A1 (en) | 1999-06-02 |
CA2263508A1 (en) | 1998-12-23 |
CA2263508C (en) | 2003-08-19 |
DE69825959T2 (de) | 2005-09-08 |
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