CN109139122B - Internal cooling system of 2-stage turbine rotor of gas turbine - Google Patents
Internal cooling system of 2-stage turbine rotor of gas turbine Download PDFInfo
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- CN109139122B CN109139122B CN201811320228.3A CN201811320228A CN109139122B CN 109139122 B CN109139122 B CN 109139122B CN 201811320228 A CN201811320228 A CN 201811320228A CN 109139122 B CN109139122 B CN 109139122B
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- wheel disc
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- turbine
- stage turbine
- turbine wheel
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
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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
Abstract
An internal cooling structure of a 2-stage turbine rotor of a gas turbine belongs to the field of gas turbine structures. The invention aims to solve the problem that the existing rotor structure of the two-stage turbine gas turbine with different shafts and different rotating speeds has no applicable internal cooling structure. The invention comprises a cooling air guiding device, a first-stage turbine wheel disc, a second-stage turbine wheel disc, a turbine front end shaft and a turbine rear end shaft, wherein the turbine front end shaft is a hollow shaft, the turbine rear end shaft is arranged in the turbine front end shaft, the turbine front end shaft and the turbine rear end shaft are coaxially arranged, a first-stage turbine wheel disc connecting ring is arranged in the circumferential direction of the turbine front end shaft, a second-stage turbine wheel disc connecting ring is arranged in the circumferential direction of the turbine rear end shaft, the first-stage turbine wheel disc is fixedly arranged on the first-stage turbine wheel disc connecting ring, and the second-stage turbine wheel disc is fixedly arranged on the second-stage turbine wheel disc connecting ring. The invention maximally reduces the consumption of high-pressure cooling air, improves the efficiency of the gas turbine, and realizes an internal cooling system of turbine rotors with different axes and different rotation speeds.
Description
Technical Field
The invention relates to an internal cooling system of a gas turbine, and belongs to the field of gas turbine structures.
Background
At present, for a gas turbine unit, one effective method for improving the circulation efficiency is to improve the gas inlet temperature of a gas turbine, and one effective method for improving the initial temperature is to cool and protect a high-temperature part of the gas turbine through reasonable design of a cooling air system, and as the temperature which can be born by high-temperature materials used by the current turbine blades is limited, reliable cooling measures must be adopted for reducing the working temperature of the high-temperature part of the gas turbine, especially the turbine blades, so that the safe operation of the gas turbine is ensured, and the cooling effect of the turbine working blades is a necessary requirement for developing a high-efficiency gas turbine.
The cooling system schemes of the gas turbine discussed in the published patent at present are all aimed at the condition of integrally connecting the turbine rotors with the same coaxial rotation speed through means such as bolts or welding, however, the internal cooling system scheme of the turbine rotor is not applicable to the turbine rotors of the gas turbine with two-stage turbines, wherein the two-stage turbines are different in coaxial rotation speed.
Disclosure of Invention
The invention aims to solve the problem that the existing two-stage turbine gas turbine rotor structure with different rotation speeds has no applicable internal cooling structure, and further provides an internal cooling structure of a 2-stage turbine rotor of a gas turbine.
The technical scheme of the invention is as follows:
the internal cooling structure of the 2-stage turbine rotor of the gas turbine comprises a cooling air guiding device, a first-stage turbine wheel disc, a second-stage turbine wheel disc, a turbine front end shaft and a turbine rear end shaft, wherein the turbine front end shaft is a hollow shaft, the turbine rear end shaft is arranged in the turbine front end shaft, the turbine front end shaft and the turbine rear end shaft are coaxially arranged, a first-stage turbine wheel disc connecting ring is arranged in the circumferential direction of the turbine front end shaft, a second-stage turbine wheel disc connecting ring is arranged in the circumferential direction of the turbine rear end shaft, the first-stage turbine wheel disc is fixedly arranged on the first-stage turbine wheel disc connecting ring, and the second-stage turbine wheel disc is fixedly arranged on the second-stage turbine wheel disc connecting ring;
the cooling air guiding device is used for introducing cooling air flow of the gas turbine from the exhaust of the compressor, the cooling air guiding device and the first-stage turbine wheel disc form a blade bottom cavity, the cooling air guiding device and the shell form a first-stage turbine wheel disc front cavity, the cooling air flow from the high-pressure stage bleed air flows through the first-stage turbine wheel disc front cavity and then is mostly introduced into the wheel disc rear cavity, and the cooling air flow from the exhaust of the compressor are converged after the blade bottom cavity and then are introduced into the first-stage blades from the root of the first-stage blades;
the cooling air flow flows through the first-stage turbine wheel disc rear cavity and the second-stage turbine wheel disc front cavity, and cools the rear side of the root of the first-stage blade and the root of the second-stage blade respectively after cooling the side of the corresponding wheel disc to supply air to the second-stage blade.
The lower parts of the first-stage turbine wheel disc and the second-stage turbine wheel disc form a two-stage wheel disc interstage lower chamber, the inside of the front end shaft of the turbine and the rear end shaft of the turbine form an interaxial circulation chamber, and the interaxial circulation chamber is communicated with the two-stage wheel disc interstage lower chamber;
the second-stage turbine wheel disc and the second-stage turbine wheel disc connecting ring circumferentially arranged on the turbine rear end shaft form a second-stage turbine wheel disc lower cooling cavity, a plurality of first through holes are radially formed in the installation position of the second-stage turbine wheel disc and the second-stage turbine wheel disc connecting ring, a second-stage turbine wheel disc rear cavity is formed between the second-stage turbine wheel disc and the turbine cylinder body, cooling air flows through the two-stage wheel disc inter-stage lower cavity into the second-stage turbine wheel disc lower cooling cavity, and the cooling air flows into the second-stage turbine wheel disc rear cavity through the first through holes to cool the rear sides of the roots of the second-stage blades.
The upper end of the cooling air guiding device is matched with the top end of the first-stage turbine wheel disc by adopting a comb tooth sealing structure, the lower part of the cooling air guiding device is matched with the middle part of the first-stage turbine wheel disc by adopting a first comb tooth sealing structure, and the cooling air guiding device and the top of the first-stage turbine wheel disc form a blade bottom cavity; the lower shell of the cooling air guiding device is provided with an air sealing sleeve, the air sealing sleeve is matched with the front end shaft of the turbine by adopting a second comb tooth sealing structure, and the cooling air guiding device, the shell, the first-stage turbine wheel disc and the air sealing sleeve form a first-stage turbine wheel disc front cavity.
Further, the lower end of the baffle plate member arranged on the stationary blade inner ring is matched and arranged with the lower end of the first-stage turbine wheel disc by adopting a third comb tooth sealing structure.
Further, the left end and the right end of the stationary blade inner ring are respectively provided with a fourth comb tooth sealing structure and a fifth comb tooth sealing structure which are matched with the first-stage turbine wheel disc and the second-stage turbine wheel disc.
The turbine rear end shaft is provided with a shaft sleeve, the shaft sleeve and the turbine front end shaft are connected through a sixth comb tooth sealing structure, an axial annular gap is formed between the shaft sleeve and the bottom of the second-stage turbine wheel disc, the compressor low-pressure stage bleed air channel is communicated with the lower-stage turbine wheel disc cavity through the sixth comb tooth sealing structure, and meanwhile, the compressor low-pressure stage bleed air channel is communicated with the lower-stage turbine wheel disc cooling cavity through the axial annular gap. And further, a first-stage turbine wheel disc bottom cooling chamber is formed between the turbine front end shaft and the first-stage turbine wheel disc, a plurality of second through holes are radially formed in the turbine front end shaft, the secondary high-pressure stage air-entraining channel of the compressor is communicated with the first-stage turbine wheel disc bottom cooling chamber through the second through holes, and the first-stage turbine wheel disc bottom cooling chamber is communicated with the two-stage wheel disc interstage lower chamber.
The cooling air flow of the exhaust gas of the compressor enters the cavity at the bottom of the blade through the third through hole of the cooling air guiding device and then is introduced into the root of the blade of the first-stage turbine movable blade through the upper hole of the first-stage wheel disc, and cooling air is provided for the first-stage movable blade; the middle position of the first-stage turbine wheel disc is provided with a plurality of fourth through holes, and the front cavity of the first-stage turbine wheel disc is communicated with the rear cavity of the first-stage turbine wheel disc through the fourth through holes.
Further, a plurality of second-stage wheel disc upper holes are formed in the upper portion of the second-stage turbine wheel disc, the second-stage wheel disc upper holes are communicated with the second-stage turbine wheel disc front cavity, and cooling air flow entering the second-stage turbine wheel disc front cavity is introduced into the roots of the second-stage turbine movable blades through the second-stage wheel disc upper holes and provides cooling air for the second-stage movable blades.
The second-stage turbine wheel disc and the turbine cylinder body are installed in a matched mode through a seventh comb tooth sealing structure, and cooling airflow entering a rear cavity of the second-stage turbine wheel disc flows through the seventh comb tooth sealing structure and then cools the rear side of the root of the second-stage turbine movable blade.
Further, the first-stage turbine wheel disc is fixedly arranged on the first-stage turbine wheel disc connecting ring in a bolt connection mode, and the second-stage turbine wheel disc is fixedly arranged on the second-stage turbine wheel disc connecting ring.
The invention has the following beneficial effects: according to the invention, four strands of compressed air are led out from different parts of the compressor as cooling air, and part of exhaust gas of the high-pressure compressor flows into a cavity at the bottom of a blade of the first-stage turbine wheel disc through an air guide device, so that the cooling air is mainly provided for the first-stage turbine movable blade; the high-pressure interstage bleed air of the air compressor reaches the front cavity of the first-stage turbine wheel disc through the second comb tooth sealing structure, and then is introduced into the rear cavity of the first-stage turbine wheel disc at the rear side of the wheel disc through a plurality of fourth holes at the middle position of the first-stage turbine wheel disc, so that the two sides of the first-stage turbine wheel disc are mainly cooled; the low-pressure interstage bleed air of the air compressor is introduced into the lower chamber of the two-stage wheel disc interstage through the air compressor low-pressure stage bleed air channel of the circulation chamber between the central shafts in the front end shaft of the turbine; the secondary high-pressure interstage bleed air of the air compressor is introduced into the bottom of the first-stage turbine wheel disc through a secondary high-pressure bleed air channel of the air compressor, and after the bottom of the turbine wheel disc is cooled, the secondary high-pressure interstage bleed air of the air compressor is converged with the air flow of the low-pressure stage bleed air channel of the air compressor to jointly provide cooling air for the movable blades of the second-stage turbine wheel disc and cool the two sides of the second-stage turbine wheel disc; cooling each structural surface along the air flow path and finally discharging the air flow into a main gas channel;
the cooling medium adopts compressed air led out from the compressor, and follows the principle that cooling air of a high-temperature and high-pressure part of the turbine is led out from a high-pressure part of the compressor, and cooling air of a second-stage turbine disc part with lower temperature and pressure is mainly led out from a low-pressure stage at the front part of the compressor, so that the consumption of the high-pressure cooling air is reduced to the maximum extent, the efficiency of the gas turbine is improved, and an internal cooling system of turbine rotors with different axes and different rotation speeds is realized.
Drawings
FIG. 1 is a schematic cross-sectional view of a cooling system flow path structure within a turbine rotor of a gas turbine;
in the figure, a 1-cooling air guiding device, a 2-third through hole, a 3-comb seal structure, a 4-first-stage wheel disc upper hole, a 5-first-stage turbine wheel disc, a 6-fourth comb seal structure, a 7-stationary blade inner ring, an 8-partition member, a 9-first-stage turbine wheel disc rear cavity, a 10-fifth comb seal structure, a 11-second-stage wheel disc upper hole, a 12-second-stage turbine wheel disc, a 13-second-stage turbine wheel disc front cavity, a 14-seventh comb seal structure, a 15-turbine cylinder body, a 16-second-stage turbine wheel disc rear cavity, a 17-first through hole, a 18-second-stage turbine wheel disc lower cooling cavity, a 19-third comb seal structure and a 20-turbine rear end shaft are arranged, 21-axial annular gap, 22-shaft sleeve, 23-sixth comb tooth sealing structure, 24-two-stage wheel disc interstage lower chamber, 26-compressor low-pressure stage air-entraining channel, 27-air inlet pipe, 28-compressor secondary high-pressure stage air-entraining channel, 29-turbine front end shaft, 30-second through hole, 31-first-stage turbine wheel disc bottom cooling chamber, 32-second comb tooth sealing structure, 33-air envelope, 34-first stage turbine wheel disc front chamber, 35-first comb tooth sealing structure, 36-blade bottom chamber, 37-fourth through hole, 38-first stage turbine wheel disc connecting ring, 39-second stage turbine wheel disc connecting ring, 40-inter-axle circulation chamber, 41-shell.
Detailed Description
The first embodiment is as follows: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine in this embodiment includes a cooling air guiding device 1, a first-stage turbine disk 5, a second-stage turbine disk 12, a turbine front end shaft 29 and a turbine rear end shaft 20, wherein the turbine front end shaft 29 is a hollow shaft, the turbine rear end shaft 20 is disposed in the turbine front end shaft 29, the turbine front end shaft 29 and the turbine rear end shaft 20 are coaxially disposed, a first-stage turbine disk connection ring 38 is disposed in the circumferential direction of the turbine front end shaft 29, a second-stage turbine disk connection ring 39 is disposed in the circumferential direction of the turbine rear end shaft 20, the first-stage turbine disk 5 is fixedly mounted on the first-stage turbine disk connection ring 38, and the second-stage turbine disk 12 is fixedly mounted on the second-stage turbine disk connection ring 39;
the cooling air guiding device 1 is used for introducing cooling air flow of the gas turbine, the cooling air guiding device 1 and the first-stage turbine wheel disc 5 form a blade bottom chamber 36 and a first-stage turbine wheel disc front chamber 34, and the cooling air flow flows through the first-stage turbine wheel disc front chamber 34 and the blade bottom chamber 36 to cool the root of the first-stage blade;
a stationary blade inner ring 7 is arranged between the first-stage turbine wheel disc 5 and the second-stage turbine wheel disc 12, a partition plate member 8 is arranged on the stationary blade inner ring 7, the partition plate member 8 partitions the first-stage turbine wheel disc 5 and the second-stage turbine wheel disc 12 to form a first-stage turbine wheel disc rear chamber 9 and a second-stage turbine wheel disc front chamber 13, and cooling air flows through the first-stage turbine wheel disc rear chamber 9 and the second-stage turbine wheel disc front chamber 13 to cool the rear sides of the roots of the first-stage blades and to be introduced into root cooling channels of the second-stage blades to provide cooling air for the second-stage blades;
the lower parts of the first-stage turbine wheel disc 5 and the second-stage turbine wheel disc 12 form a two-stage wheel disc interstage lower chamber 24, an inter-shaft circulation chamber 40 is formed between the inside of a turbine front end shaft 29 and a turbine rear end shaft, and the inter-shaft circulation chamber 40 is communicated with the two-stage wheel disc interstage lower chamber 24;
the second-stage turbine wheel disc 12 and the second-stage turbine wheel disc connecting ring 39 which are circumferentially arranged on the turbine rear end shaft 20 form a second-stage turbine wheel disc lower cooling chamber 18, a plurality of first through holes 17 are radially formed in the installation position of the second-stage turbine wheel disc 12 and the second-stage turbine wheel disc connecting ring 39, a second-stage turbine wheel disc rear chamber 16 is formed between the second-stage turbine wheel disc 12 and the turbine cylinder body 15, and cooling air flows through the two-stage wheel disc interstage lower chamber 24 into the second-stage turbine wheel disc lower cooling chamber 18 and enters the second-stage turbine wheel disc rear chamber 16 through the first through holes 17 to cool the rear sides of root parts of the second-stage blades. The internal cooling structure of the 2-stage turbine rotor of the gas turbine is a gas turbine internal cooling channel formed by a cylinder body and two-stage turbine wheel disc rotors, the rotating speed of a first-stage turbine wheel disc 5 and components thereof is different from that of a second-stage turbine wheel disc 12 and components thereof, the rotating speed of the first-stage turbine wheel disc 5 is greater than that of the second-stage turbine wheel disc 12, and structural elements such as a cavity, a hole, sealing and the like are adopted to control a flow path and distribute flow in the turbine rotor, so that the cooling system can meet the functions of providing moving blade cooling gas for two-stage turbines with different axes and different rotating speeds, cooling the inside of the turbine wheel disc and sealing and insulating a main gas channel, and the adopted cooling medium is compressed air led out from different stages of a compressor.
The second embodiment is as follows: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment is described, wherein the upper end of the cooling air guiding device 1 is mounted by adopting a comb seal structure 3 and matched with the top end of the first-stage turbine wheel disc 5, the middle part of the cooling air guiding device 1 and the middle part of the first-stage turbine wheel disc 5 are mounted by adopting a first comb seal structure 35, and the top end of the cooling air guiding device 1 and the top end of the first-stage turbine wheel disc 5 form a blade bottom cavity 36; the lower part of the cooling air guiding device 1 is connected with a shell, an air jacket 33 is arranged on the shell, the air jacket 33 is matched with the turbine front end shaft 29 by adopting a second comb tooth sealing structure 32, and the cooling air guiding device 1, the shell 41, the first-stage turbine wheel disc 5 and the air jacket 33 form a first-stage turbine wheel disc front cavity 34; the cooling air guiding device 1 is provided with a plurality of third through holes 2 in the axial direction, the upper part of the first-stage turbine wheel disc 5 is provided with a plurality of first-stage wheel disc upper holes 4, the third through holes 2 and the first-stage wheel disc upper holes 4 are respectively communicated with the blade bottom cavity 36, and cooling air flow of the exhaust gas of the compressor enters the blade bottom cavity 36 through the third through holes 2 of the cooling air guiding device 1 and then is introduced into the root of the first-stage turbine movable blade through the first-stage wheel disc upper holes 4; a plurality of fourth through holes 37 are formed in the middle of the first-stage turbine wheel disc 5, and the first-stage turbine wheel disc front cavity 34 is communicated with the first-stage turbine wheel disc rear cavity 9 through the fourth through holes 37. The arrangement is that one part of the exhaust gas of the compressor flows into a cavity 36 at the bottom of a movable vane blade of the first-stage turbine wheel disc 5 through a third through hole 2 on the front cooling air guiding device 1 of the first-stage turbine wheel disc, the other part of the exhaust gas is discharged into a main gas channel after passing through the front side of the root of the cooling vane of the comb teeth sealing structure 3, and the other part of the exhaust gas is introduced into the bottom of the movable vane of the first-stage turbine wheel disc through a hole 4 on the first-stage turbine wheel disc; the high-pressure stage bleed air of the compressor flows into a first-stage turbine wheel disc front cavity 34 through a second comb tooth sealing structure 32 formed by an air jacket 33 and a turbine front end shaft 29, is cooled to be divided into 2 air flows after passing through the front side of the first-stage turbine wheel disc, and one air flow is converged with the exhaust air flow of the compressor into a blade bottom cavity 36 after passing through a first comb tooth sealing structure 35 between the cooling air guide device 1 and the first-stage turbine wheel disc 5; the other part of the air flow is led into the second-stage turbine wheel disc front cavity 13 through a third comb tooth sealing structure 19 between the partition plate member 8 and the first-stage turbine wheel disc 5, and the other part of the air flow is led into the main gas channel through a fourth comb tooth sealing structure 6 formed by the first-stage turbine wheel disc 5 and the front part of the stator blade inner ring 7 after cooling the root rear side of the first-stage turbine blade and the front side of the stator blade inner ring.
And a third specific embodiment: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment is described, wherein the lower end of a partition member 8 mounted on a stationary blade inner ring 7 is mounted in cooperation with the lower end of a first-stage turbine disk 5 by a third comb-tooth sealing structure 19.
The specific embodiment IV is as follows: referring to fig. 1, the present embodiment is described as an internal cooling structure for a 2-stage turbine rotor of a gas turbine according to the present embodiment, wherein the left and right ends of the vane inner ring 7 are respectively mounted by adopting a fourth comb-teeth sealing structure 6 and a fifth comb-teeth sealing structure 10 in cooperation with the first stage turbine disk 5 and the second stage turbine disk 12.
Fifth embodiment: referring to fig. 1, in the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment, an air intake pipe 27 is installed in an inter-shaft circulation chamber 40 formed by a turbine rear end shaft 20 and a turbine front end shaft 29, the inter-shaft circulation chamber 40 is partitioned by the air intake pipe 27 into a compressor sub-high pressure stage bleed air channel 28 and a compressor low pressure stage bleed air channel 26, a shaft sleeve 22 is installed on the turbine rear end shaft 20, an axial annular gap 21 is formed between the shaft sleeve 22 and the turbine front end shaft 29 through a sixth comb seal structure 23, the shaft sleeve 22 and the second stage turbine disk 12 form an axial annular gap 21, the compressor low pressure stage bleed air channel 26 is communicated with a two-stage disk inter-stage lower chamber 24 through the sixth comb seal structure 23, and the compressor low pressure stage bleed air channel 26 is communicated with the two-stage turbine disk lower cooling chamber 18 through the axial annular gap 21; a first-stage turbine wheel disk bottom cooling chamber 31 is formed between the turbine front end shaft 29 and the first-stage turbine wheel disk 5, a plurality of second through holes 30 are radially formed in the turbine front end shaft 29, the compressor secondary high-pressure stage bleed air channel 28 is communicated with the first-stage turbine wheel disk bottom cooling chamber 31 through the second through holes 30, and the first-stage turbine wheel disk bottom cooling chamber 31 is communicated with the two-stage wheel disk interstage lower chamber 24. So arranged, the low-pressure stage bleed air of the compressor passes through a sixth comb seal structure 23 and then flows into a lower chamber 24 between two stages of wheel disc stages through a low-pressure stage bleed air channel 26 of the compressor formed by a turbine rear end shaft 20 and an air inlet pipe 27; the secondary high-pressure stage bleed air of the compressor passes through an internal central secondary high-pressure stage bleed air channel 28 and then is introduced into a first-stage turbine wheel disc bottom cooling chamber 31 of the first-stage turbine wheel disc 5 through a plurality of second through holes 30 on a turbine front end shaft 29, and after the lower part of the first-stage turbine wheel disc is cooled, the secondary high-pressure stage bleed air reaches a two-stage wheel disc interstage lower chamber 24; the upper two-stage disc interstage lower chamber 24 merges and then splits into two streams:
one air flow flows into the front chamber 13 of the second-stage turbine wheel disc along the radial direction, merges with part of cooling air from the high-pressure stage, cools the front side of the second-stage turbine wheel disc 12 and the rear side of the partition plate member 8, then one part of the air flow is introduced into the bottom of the second-stage turbine movable blade through the second-stage wheel disc upper hole 11 on the second-stage turbine wheel disc, and the other part of the air flow flows into the main gas channel through the fifth comb tooth sealing structure 10 formed by the rear part of the second-stage stationary blade inner ring 7 and the front part of the second-stage turbine wheel disc 12, wherein the rear side of the stationary blade inner ring and the front side of the second-stage turbine wheel disc blade root are cooled;
the other strand of the air flows into a cooling cavity 18 at the lower part of the second-stage turbine wheel disc formed by the second-stage turbine wheel disc and a turbine rear end shaft 20 after cooling the bottom of the second-stage turbine wheel disc along an axial annular gap 21 between the bottom of the second-stage turbine wheel disc and the shaft sleeve 22, flows outwards to a rear cavity 16 of the second-stage turbine wheel disc along a plurality of first through holes 17 on the rear end structure of the second-stage turbine wheel disc, is discharged to a main gas channel through a seventh comb tooth sealing structure 14 after cooling the rear side of the second-stage turbine wheel disc along the path, and simultaneously cools the rear sides of the blade roots of the second-stage turbine wheel disc.
Specific embodiment six: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment is described, wherein a plurality of second-stage turbine disk upper holes 11 are formed in an upper portion of a second-stage turbine disk 12, the second-stage turbine disk upper holes 11 are communicated with a second-stage turbine disk front chamber 13, and cooling air flowing into the second-stage turbine disk front chamber 13 is led to second-stage turbine blades through the second-stage turbine disk upper holes 11.
Seventh embodiment: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment is described, wherein the second-stage turbine wheel disc 12 and the turbine cylinder 15 are mounted in a matched manner by a seventh comb seal structure 14, and the cooling air flow entering the rear chamber 16 of the second-stage turbine wheel disc passes through the seventh comb seal structure 14 and then cools the rear side of the root of the blade of the second-stage turbine blade.
Eighth embodiment: referring to fig. 1, the internal cooling structure of a 2-stage turbine rotor of a gas turbine according to the present embodiment is described, wherein the first-stage turbine wheel 5 is fixedly mounted on the first-stage turbine wheel connection ring 38 by a bolt connection method, and the second-stage turbine wheel 12 is fixedly mounted on the second-stage turbine wheel connection ring 39.
The present embodiment is only exemplary of the present patent, and does not limit the scope of protection thereof, and those skilled in the art may also change the part thereof, so long as the spirit of the present patent is not exceeded, and the present patent is within the scope of protection thereof.
Claims (10)
1. The utility model provides an inside cooling structure of gas turbine 2 level turbine rotor, includes cooling air guider (1), first level turbine rim plate (5), second level turbine rim plate (12), turbine front end axle (29), turbine rear end axle (20) and casing (41), turbine front end axle (29) be the hollow shaft, turbine rear end axle (20) set up in turbine front end axle (29), turbine front end axle (29) and turbine rear end axle (20) coaxial arrangement, the circumference of turbine front end axle (29) be provided with one-level turbine rim plate go-between (38), the circumference of turbine rear end axle (20) be provided with second level turbine rim plate go-between (39), first level turbine rim plate (5) fixed mounting is on one-level turbine rim plate go-between (38), second level turbine rim plate (12) fixed mounting is on second level turbine rim plate go-between (39), its characterized in that:
the cooling air guiding device (1) is used for introducing cooling air flow of the gas turbine from the exhaust of the compressor, the cooling air guiding device (1) and the first-stage turbine wheel disc (5) form a blade bottom chamber (36), the shell (41) and the first-stage turbine wheel disc (5) form a first-stage turbine wheel disc front chamber (34), and the cooling air flow flows through the first-stage turbine wheel disc front chamber (34) and the blade bottom chamber (36) to cool the root of the first-stage blade;
a stator blade inner ring (7) is arranged between the first-stage turbine wheel disc (5) and the second-stage turbine wheel disc (12), a partition plate member (8) is arranged on the stator blade inner ring (7), the partition plate member (8) partitions the first-stage turbine wheel disc (5) and the second-stage turbine wheel disc (12) to form a first-stage turbine wheel disc rear chamber (9) and a second-stage turbine wheel disc front chamber (13), and cooling air flows through the first-stage turbine wheel disc rear chamber (9) and the second-stage turbine wheel disc front chamber (13) to cool the rear sides of the root parts of the first-stage blades and the root parts of the second-stage blades;
the lower part of the first-stage turbine wheel disc (5) and the lower part of the second-stage turbine wheel disc (12) form a two-stage wheel disc interstage lower chamber (24), an interaxial circulation chamber (40) is formed between the interior of a turbine front end shaft (29) and a turbine rear end shaft, and the interaxial circulation chamber (40) is communicated with the two-stage wheel disc interstage lower chamber (24);
the second-stage turbine wheel disc (12) and the second-stage turbine wheel disc connecting ring (39) circumferentially arranged on the turbine rear end shaft (20) form a second-stage turbine wheel disc lower cooling cavity (18), a plurality of first through holes (17) are radially formed in the installation position of the second-stage turbine wheel disc (12) and the second-stage turbine wheel disc connecting ring (39), a second-stage turbine wheel disc rear cavity (16) is formed between the second-stage turbine wheel disc (12) and the turbine cylinder body (15), and cooling air flows into the second-stage turbine wheel disc lower cooling cavity (18) through cooling air flowing through the two-stage wheel disc inter-stage lower cavity (24) and enters the second-stage turbine wheel disc rear cavity (16) through the first through holes (17) to cool the rear sides of root parts of second-stage blades.
2. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the upper end of the cooling air guiding device (1) is provided with a comb tooth sealing structure (3) which is matched with the top end of the first-stage turbine wheel disc (5), the middle part of the cooling air guiding device (1) is matched with the middle part of the first-stage turbine wheel disc (5) by a first comb tooth sealing structure (35), and the top end of the cooling air guiding device (1) and the top of the first-stage turbine wheel disc (5) form a blade bottom cavity (36); the lower part of the cooling air guiding device (1) is connected with a shell (41), an air jacket (33) is arranged on the shell (41), the air jacket (33) is matched with the turbine front end shaft (29) by adopting a second comb tooth sealing structure (32), and the cooling air guiding device (1), the shell (41), the first-stage turbine wheel disc (5) and the air jacket (33) form a first-stage turbine wheel disc front cavity (34).
3. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the lower end of a baffle plate member (8) arranged on the stationary blade inner ring (7) is matched and arranged with the lower end of the first-stage turbine wheel disc (5) by adopting a third comb tooth sealing structure (19).
4. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the left end and the right end of the stationary blade inner ring (7) are respectively provided with a fourth comb tooth sealing structure (6) and a fifth comb tooth sealing structure (10) which are matched with the first-stage turbine wheel disc (5) and the second-stage turbine wheel disc (12).
5. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the turbine rear end shaft (20) and the turbine front end shaft (29) form an inter-shaft circulation chamber (40) internally provided with an air inlet pipe (27), the inter-shaft circulation chamber (40) is separated by the air inlet pipe (27) to form a compressor secondary high-pressure stage air-entraining channel (28) and a compressor low-pressure stage air-entraining channel (26), the turbine rear end shaft (20) is provided with a shaft sleeve (22), an axial annular gap (21) is formed between the shaft sleeve (22) and the turbine front end shaft (29) through a sixth comb tooth sealing structure (23), the compressor low-pressure stage air-entraining channel (26) is communicated with a two-stage turbine wheel disc interstage lower chamber (24) through the sixth comb tooth sealing structure (23), and meanwhile, the compressor low-pressure stage air-entraining channel (26) is communicated with a two-stage turbine wheel disc lower cooling chamber (18) through the axial annular gap (21).
6. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 5, wherein: a first-stage turbine wheel disc bottom cooling chamber (31) is formed between the turbine front end shaft (29) and the first-stage turbine wheel disc (5), a plurality of second through holes (30) are radially formed in the turbine front end shaft (29), the compressor secondary high-pressure stage air-entraining channel (28) is communicated with the first-stage turbine wheel disc bottom cooling chamber (31) through the second through holes (30), and the first-stage turbine wheel disc bottom cooling chamber (31) is communicated with the two-stage wheel disc interstage lower chamber (24).
7. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: a plurality of third through holes (2) are formed in the axial direction of the cooling air guiding device (1), a plurality of first-stage wheel disc upper holes (4) are formed in the upper portion of the first-stage turbine wheel disc (5), the third through holes (2) and the first-stage wheel disc upper holes (4) are respectively communicated with the blade bottom cavity (36), and cooling air flow of exhaust air of the compressor enters the blade bottom cavity (36) through the third through holes (2) of the cooling air guiding device (1) and then is introduced into the root of a first-stage turbine movable blade through the first-stage wheel disc upper holes (4);
a plurality of fourth through holes (37) are formed in the middle of the first-stage turbine wheel disc (5), and the first-stage turbine wheel disc front cavity (34) is communicated with the first-stage turbine wheel disc rear cavity (9) through the fourth through holes (37).
8. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the upper part of the second-stage turbine wheel disc (12) is provided with a plurality of second-stage wheel disc upper holes (11), the second-stage wheel disc upper holes (11) are communicated with the second-stage turbine wheel disc front chamber (13), and cooling air flow entering the second-stage turbine wheel disc front chamber (13) is introduced into the roots of the second-stage turbine movable blades through the second-stage wheel disc upper holes (11).
9. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the second-stage turbine wheel disc (12) and the turbine cylinder body (15) are installed in a matched mode through a seventh comb tooth sealing structure (14), and cooling airflow entering a rear cavity (16) of the second-stage turbine wheel disc flows through the seventh comb tooth sealing structure (14) and then cools the rear sides of the roots of the blades of the second-stage turbine movable blades.
10. The internal cooling structure of a gas turbine 2-stage turbine rotor according to claim 1, wherein: the first-stage turbine wheel disc (5) is fixedly arranged on the first-stage turbine wheel disc connecting ring (38) in a bolt connection mode, and the second-stage turbine wheel disc (12) is fixedly arranged on the second-stage turbine wheel disc connecting ring (39).
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| CN109798153B (en) * | 2019-03-28 | 2023-08-22 | 中国船舶重工集团公司第七0三研究所 | Cooling structure applied to turbine wheel disc of marine gas turbine |
| CN110656984B (en) * | 2019-07-29 | 2022-04-01 | 中国航发沈阳发动机研究所 | Device and method for measuring airflow at outlet of disk edge position of turbine disk |
| CN111894734A (en) * | 2020-08-12 | 2020-11-06 | 哈电发电设备国家工程研究中心有限公司 | Turbine of small and medium-sized gas turbine and working method |
| CN112444398A (en) * | 2020-11-23 | 2021-03-05 | 东方电气集团东方汽轮机有限公司 | Gas turbine rotor cold air conveying test piece and test parameter design method |
| CN112432793A (en) * | 2020-11-23 | 2021-03-02 | 东方电气集团东方汽轮机有限公司 | Gas turbine wheel disc air extraction test piece and modeling test parameter design method |
| CN114776402A (en) * | 2021-01-22 | 2022-07-22 | 中国航发商用航空发动机有限责任公司 | Bearing cavity sealing system of aircraft engine and control method thereof |
| CN113847280A (en) * | 2021-10-10 | 2021-12-28 | 中国航发沈阳发动机研究所 | Compressor rotor interstage bleed air structure |
| CN114215611B (en) * | 2021-12-01 | 2023-07-14 | 东方电气集团东方汽轮机有限公司 | Gas seal assembly for axial positioning of turbine movable blade of gas turbine |
| CN115788677A (en) * | 2022-11-18 | 2023-03-14 | 中国联合重型燃气轮机技术有限公司 | Gas turbine torque disc air supply flow path structure and gas turbine |
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