CN212454634U - Interstage sealing device for compressor and turbine rotor of small gas turbine - Google Patents
Interstage sealing device for compressor and turbine rotor of small gas turbine Download PDFInfo
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- CN212454634U CN212454634U CN202021913154.7U CN202021913154U CN212454634U CN 212454634 U CN212454634 U CN 212454634U CN 202021913154 U CN202021913154 U CN 202021913154U CN 212454634 U CN212454634 U CN 212454634U
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Abstract
The utility model discloses an interstage sealing device of a compressor and a turbine rotor of a small-sized gas turbine, wherein a diffuser, a sealing ring and a nozzle ring of a combustion chamber are all in a ring shape; the diffuser, the sealing ring, the nozzle ring of the combustion chamber, the impeller of the gas compressor and the turbine rotor are coaxially assembled together; the front end wall, the far shaft ring wall, the near shaft ring wall and the heat insulation tile of the sealing ring form an interstage sealing inner cavity together; the front end wall is fixed at the near-axis part of the diffuser; the rear end of the compressor impeller is far away from the shaft center part and is provided with a circle of inner isolation ring which extends backwards; the near-axis side of the near-axis ring wall is attached to the far-axis side of the isolating ring; the near shaft end of the diffuser is arranged at the far shaft side of the compressor impeller and is provided with a radial gap; the far shaft ring wall of the sealing ring and the far shaft end of the heat insulation tile are in sealing fit with the nozzle ring of the combustion chamber. The interstage sealing device for the gas compressor and the turbine rotor of the small gas turbine has the advantages of effectively reducing the temperature of the gas compressor, ensuring the pressure distribution of a flow path, along with simple and compact structure, effective and reliable sealing and easiness in assembly.
Description
Technical Field
The utility model relates to a gas turbine, in particular to small-size gas turbine compressor and turbine rotor's interstage sealing device.
Background
A Gas Turbine (Gas Turbine) is a rotary power machine which uses continuously flowing Gas as a working medium and converts the heat energy of the working medium into mechanical work.
At present, most small gas turbines (the power is less than or equal to 10MW) are formed by modifying an aerospace engine, axial flow channels are adopted among a gas compressor, a combustion chamber and a turbine rotor, and because high-temperature gas can cause pressure drop through leakage flow of an unplanned path, the interstage channels are expected not to influence each other, the gas is limited in the planned path, and related parts are guaranteed not to be influenced by the high-temperature gas. The sealing device of the aerospace engine is not suitable for the axial flow channel of the small-sized gas turbine, so that an interstage sealing device suitable for a compressor impeller and a turbine rotor of the small-sized gas turbine is urgently needed at the present stage.
SUMMERY OF THE UTILITY MODEL
The to-be-solved technical problem of the utility model is to provide an interstage seal device of small-size gas turbine compressor and turbine rotor, existing effect reduces the compressor temperature, can guarantee the pressure distribution of flow path again, and simple structure is compact, and is sealed effective reliable, easily assembly.
In order to solve the technical problem, in the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine, the diffuser 3, the sealing ring 8 and the nozzle ring 4 of the combustion chamber are all in a ring shape;
the diffuser 3, the sealing ring 8, the combustion chamber nozzle ring 4, the compressor impeller 5 and the turbine rotor 6 are coaxially assembled together;
the diffuser 3 is fixed in the diffuser casing 2, and a radial cold air flow channel is formed between the front side of the diffuser 3 and the diffuser casing 2;
a far shaft ring wall 81 and a near shaft ring wall 82 which are coaxial are formed on the rear side of the front end wall 80 of the sealing ring 8;
a circle of heat insulation tiles 13 are arranged between the rear ends of the far collar wall 81 and the near collar wall 82;
the front end wall 80, the far shaft ring wall 81, the near shaft ring wall 82 and the heat insulation tile 13 of the sealing ring 8 jointly form an interstage seal inner cavity 1-20;
the front end wall 80 of the sealing ring 8 is fixed at the near axial center part of the diffuser 3;
the back end near-shaft center part of the compressor impeller 5 is butted with the front end near-shaft center part of the turbine rotor 6;
a spacing region is formed between the far shaft center part at the rear end of the compressor impeller 5 and the far shaft center part at the front end of the turbine rotor 6;
the rear end of the compressor impeller 5 is far away from the shaft center part and is provided with a circle of inner isolation ring 51 protruding backwards;
the proximal side of the proximal collar wall 82 abuts the distal side of the inner spacer ring 51;
the near shaft end of the diffuser 3 is arranged at the far shaft side of the compressor impeller 5 and has a radial gap;
the far shaft ring wall 81 of the sealing ring 8 and the far shaft end of the heat insulation tile 13 are in sealing fit with the combustion chamber nozzle ring 4;
the heat insulation tile 13 and the front end far shaft center part of the turbine rotor 6 have an axial gap;
the front end wall 80 of the sealing ring 8 has an axial clearance with the rear end far shaft center part of the compressor impeller 5.
Preferably, a piston ring mounting groove is formed in the circumferential direction at the rear end of the far shaft side of the far shaft ring wall 81 of the sealing ring 8;
a piston ring 11 is fitted to the piston ring mounting groove;
the combustion chamber nozzle ring 4 comprises an outer nozzle ring 41 and an inner nozzle ring 42;
the front part of the outer nozzle ring 41 is bent inwards along the radial direction, and the rear part is bent backwards along the axial direction;
the front part of the inner nozzle ring 42 is bent outwards along the radial direction, and the rear part is bent backwards along the axial direction;
the inner nozzle ring 42 and the outer nozzle ring 41 form a hot gas injection channel;
a circle of outer isolating rings 45 protruding forwards are formed at the front end of the outer nozzle ring 41;
the front end of the outer isolating ring 45 is sleeved outside the rear end of the far shaft ring wall 81 of the sealing ring 8;
the outer isolating ring 45 is in sealing fit with the far shaft ring wall 81 of the sealing ring 8 through a piston ring 11.
Preferably, the heat insulation tile 13 is in a fan ring shape;
two wings of the heat insulation tile 13 are respectively provided with a groove and a convex rib;
the heat insulation tiles 13 are matched with the convex ribs through the grooves and are connected end to form a complete sealing ring.
Preferably, the front-rear thickness of the far axial end of the heat insulation tile 13 is larger than the axial gap between the rear end of the far axial ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41;
the far shaft end of the heat insulation tile 13 is pressed, attached and sealed by the back end of the far shaft ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41.
Preferably, a heat insulation tile installation groove is formed in the circumferential direction at the rear end of the far shaft side of the shaft ring wall 82 of the sealing ring 8;
the near shaft end of the heat insulation tile 13 is inserted into the heat insulation tile installation groove and is in sealing fit with the near shaft ring wall 82 of the sealing ring 8.
Preferably, the seal ring 8 is thicker at the rear end proximal to the collar wall 82 than at the front.
Preferably, a labyrinth structure 510 is formed on the far axial side of the inner separating ring 51, which protrudes backwards from the far axial center part of the back end of the compressor impeller 5, along the circumferential direction;
the proximal side of the proximal annular wall 82 of the seal ring 8 abuts the labyrinth 510 of the distal side of the inner spacer ring 51.
Preferably, the radial clearance between the near shaft end of the diffuser 3 and the far shaft side of the compressor impeller 5 is between 2mm and 4 mm;
the axial clearance between the heat insulation tile 13 and the front end far shaft center part of the turbine rotor 6 is between 2mm and 8 mm;
the axial clearance between the front end wall 80 of the sealing ring 8 and the rear end far shaft center part of the compressor impeller 5 is between 3mm and 9 mm.
The utility model discloses a small-size gas turbine compressor and turbine rotor's interstage sealing device is formed with three backward flow districts between compressor impeller 5, turbine rotor 6 and the nozzle cascade 4, does not influence the sprue and can effectually prevent that interstage fluid from revealing, and the high temperature gas that has obstructed the turbine sprue gets into compressor and interstage inner chamber, and existing effect reduces the compressor temperature, can guarantee the pressure distribution of flow path again, and simple structure is compact, and is sealed effective reliable, easily assembles.
Drawings
In order to more clearly illustrate the technical solution of the present invention, the drawings required for the present invention are briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
Fig. 1 is a schematic cross-sectional view of an embodiment of an interstage seal device of a compressor and a turbine rotor of a small gas turbine according to the present invention;
FIG. 2 is a schematic view of a backflow region formed in an embodiment of an interstage seal device for a compressor and a turbine rotor of a small gas turbine according to the present invention
Fig. 3 is a partially enlarged schematic view of an embodiment of an interstage seal device between a compressor and a turbine rotor of a small gas turbine according to the present invention;
fig. 4 is a schematic view of a heat insulation tile assembly of an embodiment of the interstage seal device of the compressor and the turbine rotor of the small gas turbine of the present invention.
Description of reference numerals:
2 a diffuser casing; 3, a diffuser; 4 a nozzle ring; 5, an air compressor impeller; 6 a turbine rotor; 8, sealing rings; 80 a seal ring front end wall; 81 distal collar wall; 82 a proximal collar wall; 13 heat insulation tile; 1-20 interstage seal cavity; 51 an inner spacer ring; 510 a comb tooth structure; 1-10 of a first reflux zone; 1-30 reflux zone two; 1-40 reflux zone III; 11 a piston ring; 41 an outer nozzle ring; 42 an inner nozzle ring; 45 outer spacer ring.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings, and it should be understood that the described embodiments are only some embodiments, but not all embodiments, of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by a person of ordinary skill in the art without creative efforts belong to the protection scope of the present invention.
Example one
As shown in fig. 1 and 2, a diffuser 3, a sealing ring 8 and a nozzle ring 4 of a combustion chamber of an interstage sealing device of a compressor and a turbine rotor of a small-sized gas turbine are all annular;
the diffuser 3, the sealing ring 8, the combustion chamber nozzle ring 4, the compressor impeller 5 and the turbine rotor 6 are coaxially assembled together;
the diffuser 3 is fixed in the diffuser casing 2, and a radial cold air flow channel is formed between the front side of the diffuser 3 and the diffuser casing 2;
as shown in fig. 3, a front end wall 80 of the seal ring 8 is formed with a distal collar wall 81 and a proximal collar wall 82 coaxially at a rear side thereof;
a circle of heat insulation tiles 13 are arranged between the rear ends of the far collar wall 81 and the near collar wall 82;
the front end wall 80, the far shaft ring wall 81, the near shaft ring wall 82 and the heat insulation tile 13 of the sealing ring 8 jointly form an interstage seal inner cavity 1-20;
the front end wall 80 of the sealing ring 8 is fixed at the near axial center part of the diffuser 3;
the back end near-shaft center part of the compressor impeller 5 is butted with the front end near-shaft center part of the turbine rotor 6;
a spacing region is formed between the far shaft center part at the rear end of the compressor impeller 5 and the far shaft center part at the front end of the turbine rotor 6;
the rear end of the compressor impeller 5 is far away from the shaft center part and is provided with a circle of inner isolation ring 51 protruding backwards;
the proximal side of the proximal collar wall 82 abuts the distal side of the inner spacer ring 51;
the near shaft end of the diffuser 3 is arranged at the far shaft side of the compressor impeller 5 and has a radial gap;
the far shaft ring wall 81 of the sealing ring 8 and the far shaft end of the heat insulation tile 13 are in sealing fit with the combustion chamber nozzle ring 4;
the heat insulation tile 13 and the front end far shaft center part of the turbine rotor 6 have an axial gap;
the front end wall 80 of the sealing ring 8 has an axial clearance with the rear end far shaft center part of the compressor impeller 5.
In the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine according to the embodiment, as shown in fig. 2, the diffuser casing 2 is communicated with the atmosphere, the compressor impeller 5 is driven to rotate through the output shaft of the gearbox in the early stage, air is pressed into the radial cold air flow channel to enter the combustion chamber, and after the air is mixed with fuel, the air is combusted, expanded and worked to drive the turbine rotor 6 to rotate. The diffuser 3, the sealing ring 8 and the nozzle ring 4 of the combustion chamber form a first backflow region 1-10, the diffuser 3, the sealing ring 8 and the compressor impeller 5 form a second backflow region 1-30, and the sealing ring 8, the compressor impeller 5 and the turbine rotor 6 form a third backflow region 1-40; the front end wall 80, the far collar wall 81, the near collar wall 82 and the heat insulation tile 13 of the sealing ring 8 jointly form an interstage seal inner cavity 1-20.
In the interstage sealing device for the compressor and the turbine rotor of the small gas turbine, three backflow regions are formed among the compressor impeller 5, the turbine rotor 6 and the nozzle ring 4, so that a main flow channel is not influenced, interstage fluid can be effectively prevented from leaking, high-temperature gas in the main flow channel of the turbine is prevented from entering the compressor and an interstage inner cavity, the temperature of the compressor is effectively reduced, pressure distribution of a flow path can be guaranteed, the structure is simple and compact, sealing is effective and reliable, and assembly is easy.
Example two
Based on the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine in the first embodiment, a piston ring installation groove is formed in the circumferential direction at the rear end of the far shaft side of the far shaft ring wall 81 of the sealing ring 8;
a piston ring 11 is fitted to the piston ring mounting groove;
the combustion chamber nozzle ring 4 comprises an outer nozzle ring 41 and an inner nozzle ring 42;
the front part of the outer nozzle ring 41 is bent inwards along the radial direction, and the rear part is bent backwards along the axial direction;
the front part of the inner nozzle ring 42 is bent outwards along the radial direction, and the rear part is bent backwards along the axial direction;
the inner nozzle ring 42 and the outer nozzle ring 41 form a hot gas injection channel;
a circle of outer isolating rings 45 protruding forwards are formed at the front end of the outer nozzle ring 41;
the front end of the outer isolating ring 45 is sleeved outside the rear end of the far shaft ring wall 81 of the sealing ring 8;
the outer isolating ring 45 is in sealing fit with the far shaft ring wall 81 of the sealing ring 8 through a piston ring 11.
EXAMPLE III
Based on the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine of the first embodiment, the heat insulation tile 13 is in a sector ring shape;
two wings of the heat insulation tile 13 are respectively provided with a groove and a convex rib;
the heat insulation tiles 13 are matched with the convex ribs through the grooves and are connected end to form a complete sealing ring.
Example four
Based on the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine in the third embodiment, the front-back thickness of the far axial end of the heat insulation tile 13 is slightly larger than the axial gap between the back end of the far axial ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41;
the far shaft end of the heat insulation tile 13 is pressed, attached and sealed by the back end of the far shaft ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41.
Preferably, a heat insulation tile installation groove is formed in the circumferential direction at the rear end of the far shaft side of the shaft ring wall 82 of the sealing ring 8;
the near shaft end of the heat insulation tile 13 is inserted into the heat insulation tile installation groove and is in sealing fit with the near shaft ring wall 82 of the sealing ring 8.
Preferably, the seal ring 8 is thicker at the rear end proximal to the collar wall 82 than at the front.
In the fourth embodiment of the interstage sealing device of the compressor and the turbine rotor of the small gas turbine, the heat insulation tile 13 forms interstage sealing cavities 1-20 through the cooperation of the nozzle ring 4 and the sealing of the piston ring 11 and the sealing ring 8.
EXAMPLE five
Based on the interstage sealing device of the compressor and the turbine rotor of the small gas turbine in the third embodiment, as shown in fig. 3, a labyrinth structure 510 is formed on the far shaft side of an inner isolating ring 51, which protrudes backwards from the far shaft center part at the back end of a compressor impeller 5;
the proximal side of the proximal annular wall 82 of the seal ring 8 abuts the labyrinth 510 of the distal side of the inner spacer ring 51.
In the fifth interstage sealing device of the gas compressor and the turbine rotor of the small-sized gas turbine, the ring wall 82 near the shaft of the sealing ring 8 is matched with the impeller 5 of the gas compressor through a labyrinth structure to form the labyrinth throttling.
EXAMPLE six
Based on the interstage sealing device of the compressor and the turbine rotor of the small-sized gas turbine in the third embodiment, the radial clearance between the near shaft end of the diffuser 3 and the far shaft side of the impeller 5 of the compressor is 2-4 mm;
the axial clearance between the heat insulation tile 13 and the front end far shaft center part of the turbine rotor 6 is between 2mm and 8 mm;
the axial clearance between the front end wall 80 of the sealing ring 8 and the rear end far shaft center part of the compressor impeller 5 is between 3mm and 9 mm.
The above are merely preferred embodiments of the present application and are not intended to limit the present application. Various modifications and changes may occur to those skilled in the art. Any modification, equivalent replacement, improvement and the like made within the spirit and principle of the present application shall be included in the protection scope of the present application.
Claims (8)
1. An interstage sealing device of a compressor and a turbine rotor of a small-sized gas turbine is characterized in that a diffuser (3), a sealing ring (8) and a nozzle ring (4) of a combustion chamber are all circular;
the diffuser (3), the sealing ring (8), the combustion chamber nozzle ring (4), the compressor impeller (5) and the turbine rotor (6) are coaxially assembled together;
the diffuser (3) is fixed in the diffuser casing (2), and a radial cold air flow channel is formed between the front side of the diffuser (3) and the diffuser casing (2);
a far shaft ring wall (81) and a near shaft ring wall (82) which are coaxial are formed on the rear side of the front end wall (80) of the sealing ring (8);
a circle of heat insulation tiles (13) are arranged between the rear ends of the far shaft ring wall (81) and the near shaft ring wall (82);
the front end wall (80) of the sealing ring (8), the far shaft ring wall (81), the near shaft ring wall (82) and the heat insulation tile (13) jointly form an interstage sealing inner cavity (1-20);
the front end wall (80) of the sealing ring (8) is fixed on the part, close to the axial center, of the diffuser (3);
the rear end near-shaft center part of the compressor impeller (5) is butted with the front end near-shaft center part of the turbine rotor (6);
a spacing region is formed between the far shaft center part at the rear end of the compressor impeller (5) and the far shaft center part at the front end of the turbine rotor (6);
the rear end of the compressor impeller (5) is far away from the shaft center part and is provided with a circle of inner isolating ring (51) which extends backwards;
the proximal side of the proximal ring wall (82) is attached to the distal side of the inner isolation ring (51);
the near shaft end of the diffuser (3) is arranged at the far shaft side of the compressor impeller (5) and has a radial gap;
the far shaft ring wall (81) of the sealing ring (8) and the far shaft end of the heat insulation tile (13) are in sealing fit with the nozzle ring (4) of the combustion chamber;
the heat insulation tile (13) and the front end far shaft center part of the turbine rotor (6) have an axial gap;
and an axial clearance is reserved between the front end wall (80) of the sealing ring (8) and the far shaft center part of the rear end of the compressor impeller (5).
2. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 1,
a piston ring mounting groove is formed in the rear end of the far shaft side of the far shaft ring wall (81) of the sealing ring (8) along the circumferential direction;
a piston ring (11) is fitted to the piston ring mounting groove;
the combustion chamber nozzle ring (4) comprises an outer nozzle ring (41) and an inner nozzle ring (42);
the front part of the outer nozzle ring (41) is bent inwards along the radial direction, and the rear part is bent backwards along the axial direction;
the front part of the inner nozzle ring (42) is bent outwards along the radial direction, and the rear part is bent backwards along the axial direction;
the inner nozzle ring (42) and the outer nozzle ring (41) form a hot gas injection channel;
a circle of outer spacing rings (45) protruding forwards are formed at the front ends of the outer nozzle rings (41);
the front end of the outer isolating ring (45) is sleeved outside the rear end of the far shaft ring wall (81) of the sealing ring (8);
the outer isolating ring (45) is in sealing fit with the far shaft ring wall (81) of the sealing ring (8) through a piston ring (11).
3. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 2,
the heat insulation tile (13) is in a fan ring shape;
two wings of the heat insulation tile (13) are respectively provided with a groove and a convex rib;
the heat insulation tiles (13) are matched with the convex ribs through the grooves and are connected end to form a complete sealing ring.
4. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 3,
the front-back thickness of the far shaft end of the heat insulation tile (13) is larger than the axial clearance between the rear end of the far shaft ring wall (81) of the sealing ring (8) and the front part of the outer nozzle ring (41);
the far shaft end of the heat insulation tile (13) is extruded, attached and sealed by the rear end of the far shaft ring wall (81) of the sealing ring (8) and the front part of the outer nozzle ring (41).
5. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 3,
a heat insulation tile installation groove is formed in the rear end of the shaft-far side of the shaft-near annular wall (82) of the sealing ring (8) along the circumferential direction;
the end of the heat insulation tile (13) close to the shaft is inserted into the heat insulation tile mounting groove and is in sealing fit with the wall (82) close to the shaft of the sealing ring (8).
6. The interstage seal device of the compressor and the turbine rotor of the small gas turbine as claimed in claim 5,
the seal ring (8) is thicker at the rear end near the collar wall (82) than at the front.
7. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 1,
a labyrinth structure (510) is formed on the far shaft side of an inner isolating ring (51) which protrudes backwards from the far shaft center part at the rear end of the compressor impeller (5) along the circumferential direction;
the shaft-proximal side of the shaft-proximal ring wall (82) of the sealing ring (8) is attached to the labyrinth structure (510) of the shaft-distal side of the inner isolation ring (51).
8. The interstage seal device of a compressor and a turbine rotor of a small gas turbine according to claim 1,
the radial clearance between the near shaft end of the diffuser (3) and the far shaft side of the compressor impeller (5) is between 2mm and 4 mm;
the axial clearance between the heat insulation tile (13) and the front end far shaft center part of the turbine rotor (6) is between 2mm and 8 mm;
the axial clearance between the front end wall (80) of the sealing ring (8) and the far shaft center part of the rear end of the compressor impeller (5) is between 3mm and 9 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202021913154.7U CN212454634U (en) | 2020-09-04 | 2020-09-04 | Interstage sealing device for compressor and turbine rotor of small gas turbine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202021913154.7U CN212454634U (en) | 2020-09-04 | 2020-09-04 | Interstage sealing device for compressor and turbine rotor of small gas turbine |
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| CN212454634U true CN212454634U (en) | 2021-02-02 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114876593A (en) * | 2022-07-06 | 2022-08-09 | 成都中科翼能科技有限公司 | Sealing structure for turbine rotor bearing cavity of core machine |
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2020
- 2020-09-04 CN CN202021913154.7U patent/CN212454634U/en active Active
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114876593A (en) * | 2022-07-06 | 2022-08-09 | 成都中科翼能科技有限公司 | Sealing structure for turbine rotor bearing cavity of core machine |
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