Interstage seal device between compressor and turbine rotor of small gas turbine

Abstract

The present invention discloses an interstage sealing device for a small gas turbine compressor and a turbine rotor, wherein the diffuser, sealing ring and combustion chamber nozzle ring are all in annular shape; the diffuser, sealing ring and combustion chamber nozzle ring are coaxially assembled with the compressor impeller and the turbine rotor; the front end wall, distal ring wall, proximal ring wall and thermal insulation tile of the sealing ring together constitute an interstage sealing inner cavity; the front end wall is fixed to the proximal axial center of the diffuser; the rear end of the compressor impeller is distal to the axial center and is formed with a circle of inner isolation rings protruding backwards; the proximal side of the proximal ring wall fits the distal side of the isolation ring; the proximal end of the diffuser is on the distal side of the compressor impeller and has a radial gap; the distal ring wall of the sealing ring and the distal end of the thermal insulation tile are sealed with the combustion chamber nozzle ring. The interstage sealing device for a small gas turbine compressor and a turbine rotor can effectively reduce the compressor temperature and ensure the pressure distribution of the flow path, has a simple and compact structure, is effective and reliable in sealing, and is easy to assemble.

Description

Technical Field

[0001] The invention relates to a gas turbine, in particular to an interstage sealing device between a compressor and a turbine rotor of a small gas turbine. Background Art

[0002] A gas turbine is a rotary power machine that uses a continuously flowing gas as a working fluid and converts the thermal energy of the working fluid into mechanical work.

[0003] At present, most small gas turbines (power ≤ 10MW) are modified from aerospace engines, and axial flow channels are mostly used between the compressor, combustion chamber and turbine rotor. Since high-temperature gas may leak through unplanned paths, resulting in pressure drop, it is expected that the interstage channels will not affect each other, the gas will be confined within the planned path, and the related parts will be guaranteed not to be affected by the high-temperature gas. However, the sealing device of aerospace engines is often not suitable for the axial flow channel of such small gas turbines. Therefore, an interstage sealing device suitable for the compressor impeller and turbine rotor of small gas turbines is urgently needed at this stage. Summary of the invention

[0004] The technical problem to be solved by the present invention is to provide an interstage sealing device for a small gas turbine compressor and a turbine rotor, which can effectively reduce the compressor temperature and ensure the pressure distribution of the flow path. It has a simple and compact structure, effective and reliable sealing, and is easy to assemble.

[0005] In order to solve the above technical problems, the present invention provides an interstage sealing device for a small gas turbine compressor and a turbine rotor, wherein the diffuser 3, the sealing ring 8, and the combustion chamber nozzle ring 4 are all in annular shape;

[0006] 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;

[0007] 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;

[0008] A coaxial distal ring wall 81 and a proximal ring wall 82 are formed on the rear side of the front end wall 80 of the sealing ring 8;

[0009] A circle of heat-insulating tiles 13 is installed between the rear ends of the distal annular wall 81 and the proximal annular wall 82;

[0010] The front end wall 80, the distal ring wall 81, the proximal ring wall 82 and the heat insulation tile 13 of the sealing ring 8 together constitute the interstage sealing cavity 1-20;

[0011] The front end wall 80 of the sealing ring 8 is fixed to the near-axial portion of the diffuser 3;

[0012] The rear end proximal to the axial center of the compressor impeller 5 is butted against the front end proximal to the axial center of the turbine rotor 6;

[0013] A spacing area is formed between the rear end far from the axial center of the compressor impeller 5 and the front end far from the axial center of the turbine rotor 6;

[0014] The rear end of the compressor impeller 5 is far from the axial center and is formed with an inner isolation ring 51 extending backward;

[0015] The proximal side of the proximal ring wall 82 is in contact with the distal side of the inner isolation ring 51;

[0016] The proximal end of the diffuser 3 is on the distal side of the compressor impeller 5 and has a radial gap;

[0017] The distal ring wall 81 of the sealing ring 8 and the distal end of the heat insulating tile 13 are in sealing cooperation with the combustion chamber nozzle ring 4;

[0018] There is an axial gap between the heat insulation tile 13 and the front end of the turbine rotor 6 far from the axial center;

[0019] There is an axial gap between the front end wall 80 of the sealing ring 8 and the rear end far from the axial center of the compressor impeller 5.

[0020] Preferably, a piston ring mounting groove is formed along the circumferential direction at the rear end of the distal side of the distal ring wall 81 of the sealing ring 8;

[0021] The piston ring 11 is assembled into the piston ring mounting groove;

[0022] The combustion chamber nozzle ring 4 includes an outer nozzle ring 41 and an inner nozzle ring 42;

[0023] The front portion of the outer nozzle ring 41 is bent radially inward, and the rear portion is bent axially backward;

[0024] The front portion of the inner nozzle ring 42 is bent radially outward, and the rear portion is bent axially backward;

[0025] The inner nozzle ring 42 and the outer nozzle ring 41 form a hot gas injection channel;

[0026] The front end of the outer nozzle ring 41 is formed with an outer isolation ring 45 extending forward;

[0027] The front end of the outer isolation ring 45 is sleeved outside the rear end of the distal ring wall 81 of the sealing ring 8;

[0028] The outer isolating ring 45 is sealed with the distal ring wall 81 of the sealing ring 8 via the piston ring 11 .

[0029] Preferably, the heat-insulating tile 13 is fan-shaped;

[0030] The two wings of the heat-insulating tile 13 are respectively provided with grooves and convex ribs;

[0031] The plurality of heat-insulating tiles 13 are matched with the grooves and the convex ribs, and are connected end to end to form a complete sealing ring.

[0032] Preferably, the front-to-rear thickness of the distal end of the heat-insulating tile 13 is greater than the axial gap between the rear end of the distal ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41;

[0033] The distal end of the heat-insulating tile 13 is squeezed and sealed by the rear end of the distal ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41 .

[0034] Preferably, a heat insulation tile installation groove is formed along the circumferential direction at the rear end of the distal axis side of the proximal ring wall 82 of the sealing ring 8;

[0035] The proximal end of the heat insulation tile 13 is inserted into the heat insulation tile installation groove and is sealed with the proximal ring wall 82 of the sealing ring 8 .

[0036] Preferably, the rear end of the proximal ring wall 82 of the sealing ring 8 is thicker than the front end.

[0037] Preferably, a comb tooth structure 510 is formed along the circumferential direction on the distal axis side of the inner isolation ring 51 extending backward from the distal axis center portion of the rear end of the compressor impeller 5;

[0038] The proximal side of the proximal ring wall 82 of the sealing ring 8 is in contact with the comb tooth structure 510 on the distal side of the inner isolation ring 51 .

[0039] Preferably, the radial clearance between the proximal end of the diffuser 3 and the distal end of the compressor impeller 5 is between 2 mm and 4 mm;

[0040] The axial clearance between the heat-insulating tile 13 and the front end of the turbine rotor 6 is between 2 mm and 8 mm.

[0041] The axial clearance between the front end wall 80 of the sealing ring 8 and the rear end far from the axial center of the compressor impeller 5 is between 3 mm and 9 mm.

[0042] The interstage sealing device of the compressor and turbine rotor of a small gas turbine of the present invention forms three recirculation areas between the compressor impeller 5, the turbine rotor 6 and the nozzle ring 4, which does not affect the main flow channel and can effectively prevent leakage of the interstage fluid, blocking the high-temperature combustion gas in the main flow channel of the turbine from entering the compressor and the interstage cavity, effectively reducing the compressor temperature and ensuring the pressure distribution of the flow path. The structure is simple and compact, the sealing is effective and reliable, and it is easy to assemble. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] In order to more clearly illustrate the technical solution of the present invention, the following briefly introduces the drawings required for use in the present invention. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0044] Figure 1 It is a cross-sectional schematic diagram of an embodiment of an interstage sealing device between a compressor and a turbine rotor of a small gas turbine according to the present invention;

[0045] Figure 2 A schematic diagram of a recirculation zone formed by an embodiment of an interstage sealing device for a small gas turbine compressor and a turbine rotor of the present invention

[0046] Figure 3 It is a partially enlarged schematic diagram of an embodiment of an interstage sealing device between a compressor and a turbine rotor of a small gas turbine according to the present invention;

[0047] Figure 4 The figure is a schematic diagram of the thermal insulation tile assembly of an embodiment of the interstage sealing device of a small gas turbine compressor and a turbine rotor of the present invention.

[0048] Description of reference numerals:

[0049] 2 diffuser casing; 3 diffuser; 4 nozzle ring; 5 compressor impeller; 6 turbine rotor; 8 sealing ring; 80 sealing ring front end wall; 81 distal ring wall; 82 proximal ring wall; 13 insulation tile; 1-20 interstage seal cavity; 51 inner isolation ring; 510 comb tooth structure; 1-10 recirculation zone one; 1-30 recirculation zone two; 1-40 recirculation zone three; 11 piston ring; 41 outer nozzle ring; 42 inner nozzle ring; 45 outer isolation ring. DETAILED DESCRIPTION

[0050] The following will be combined with the accompanying drawings to clearly and completely describe the technical solutions in the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0051] Embodiment 1

[0052] like Figure 1 , Figure 2 As shown, the diffuser 3, the sealing ring 8, and the combustion chamber nozzle ring 4 of the interstage sealing device of the small gas turbine compressor and the turbine rotor are all in the shape of a circular ring;

[0053] 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;

[0054] 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;

[0055] like Figure 3 As shown, a coaxial distal ring wall 81 and a proximal ring wall 82 are formed on the rear side of the front end wall 80 of the sealing ring 8;

[0056] A circle of heat-insulating tiles 13 is installed between the rear ends of the distal annular wall 81 and the proximal annular wall 82;

[0057] The front end wall 80, the distal ring wall 81, the proximal ring wall 82 and the heat insulation tile 13 of the sealing ring 8 together constitute the interstage sealing cavity 1-20;

[0058] The front end wall 80 of the sealing ring 8 is fixed to the near-axial portion of the diffuser 3;

[0059] The rear end proximal to the axial center of the compressor impeller 5 is butted against the front end proximal to the axial center of the turbine rotor 6;

[0060] A spacing area is formed between the rear end far from the axial center of the compressor impeller 5 and the front end far from the axial center of the turbine rotor 6;

[0061] The rear end of the compressor impeller 5 is far from the axial center and is formed with an inner isolation ring 51 extending backward;

[0062] The proximal side of the proximal ring wall 82 is in contact with the distal side of the inner isolation ring 51;

[0063] The proximal end of the diffuser 3 is on the distal side of the compressor impeller 5 and has a radial gap;

[0064] The distal ring wall 81 of the sealing ring 8 and the distal end of the heat insulating tile 13 are in sealing cooperation with the combustion chamber nozzle ring 4;

[0065] There is an axial gap between the heat insulation tile 13 and the front end of the turbine rotor 6 far from the axial center;

[0066] There is an axial gap between the front end wall 80 of the sealing ring 8 and the rear end far from the axial center of the compressor impeller 5.

[0067] The interstage sealing device between the compressor and the turbine rotor of the small gas turbine of the first embodiment is as follows: Figure 2As shown, the diffuser casing 2 is connected to the atmosphere. In the early stage, the compressor impeller 5 is driven to rotate by the output shaft of the gear box to compress the air into the radial cold air flow channel and enter the combustion chamber. After mixing with the fuel, the combustion expansion work drives the turbine rotor 6 to rotate. The diffuser 3, the sealing ring 8, and the combustion chamber nozzle ring 4 form a reflow zone 1-10, the diffuser 3, the sealing ring 8, and the compressor impeller 5 form a reflow zone 2 1-30, and the sealing ring 8, the compressor impeller 5, and the turbine rotor 6 form a reflow zone 3 1-40; the front end wall 80, the distal ring wall 81, the proximal ring wall 82, and the thermal insulation tile 13 of the sealing ring 8 together form the interstage sealing cavity 1-20.

[0068] In the interstage sealing device of the compressor and turbine rotor of a small gas turbine of the first embodiment, three recirculation areas are formed between the compressor impeller 5, the turbine rotor 6 and the nozzle ring 4, which do not affect the main flow channel and can effectively prevent leakage of the interstage fluid, and block the high-temperature combustion gas in the main flow channel of the turbine from entering the compressor and the interstage cavity, which can effectively reduce the compressor temperature and ensure the pressure distribution of the flow path. The structure is simple and compact, the sealing is effective and reliable, and it is easy to assemble.

[0069] Embodiment 2

[0070] Based on the interstage sealing device of the small gas turbine compressor and the turbine rotor of the first embodiment, a piston ring mounting groove is formed along the circumferential direction at the rear end of the far-axis side of the far-axis ring wall 81 of the sealing ring 8;

[0071] The piston ring 11 is assembled into the piston ring mounting groove;

[0072] The combustion chamber nozzle ring 4 includes an outer nozzle ring 41 and an inner nozzle ring 42;

[0073] The front portion of the outer nozzle ring 41 is bent radially inward, and the rear portion is bent axially backward;

[0074] The front portion of the inner nozzle ring 42 is bent radially outward, and the rear portion is bent axially backward;

[0075] The inner nozzle ring 42 and the outer nozzle ring 41 form a hot gas injection channel;

[0076] The front end of the outer nozzle ring 41 is formed with an outer isolation ring 45 extending forward;

[0077] The front end of the outer isolation ring 45 is sleeved outside the rear end of the distal ring wall 81 of the sealing ring 8;

[0078] The outer isolating ring 45 is sealed with the distal ring wall 81 of the sealing ring 8 via the piston ring 11 .

[0079] Embodiment 3

[0080] Based on the interstage sealing device between the compressor and turbine rotor of a small gas turbine in the first embodiment, the heat insulation tile 13 is in the shape of a fan ring;

[0081] The two wings of the heat-insulating tile 13 are respectively provided with grooves and convex ribs;

[0082] The plurality of heat-insulating tiles 13 are matched with the grooves and the convex ribs, and are connected end to end to form a complete sealing ring.

[0083] Embodiment 4

[0084] Based on the interstage sealing device of the small gas turbine compressor and the turbine rotor of the third embodiment, the front and rear thickness of the far-axis end of the heat insulation tile 13 is slightly larger than the axial gap between the rear end of the far-axis ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41;

[0085] The distal end of the heat-insulating tile 13 is squeezed and sealed by the rear end of the distal ring wall 81 of the sealing ring 8 and the front part of the outer nozzle ring 41 .

[0086] Preferably, a heat insulation tile installation groove is formed along the circumferential direction at the rear end of the distal axis side of the proximal ring wall 82 of the sealing ring 8;

[0087] The proximal end of the heat insulation tile 13 is inserted into the heat insulation tile installation groove and is sealed with the proximal ring wall 82 of the sealing ring 8 .

[0088] Preferably, the rear end of the proximal ring wall 82 of the sealing ring 8 is thicker than the front end.

[0089] In the interstage sealing device of the small gas turbine compressor and the turbine rotor of the fourth embodiment, the heat insulation tile 13 forms an interstage sealing cavity 1-20 with the sealing ring 8 through the cooperation of the nozzle ring 4 and the sealing of the piston ring 11.

[0090] Embodiment 5

[0091] Based on the interstage sealing device between the compressor and the turbine rotor of a small gas turbine of the third embodiment, Figure 3 As shown, a comb tooth structure 510 is formed along the circumferential direction on the far-axis side of the inner isolation ring 51 extending backward from the far-axis center portion of the rear end of the compressor impeller 5;

[0092] The proximal side of the proximal ring wall 82 of the sealing ring 8 is in contact with the comb tooth structure 510 on the distal side of the inner isolation ring 51 .

[0093] In the interstage sealing device between the compressor and turbine rotor of the small gas turbine of the fifth embodiment, the proximal ring wall 82 of the sealing ring 8 cooperates with the compressor impeller 5 through the comb tooth structure to form a comb tooth throttling.

[0094] Embodiment 6

[0095] Based on the interstage sealing device of the small gas turbine compressor and the turbine rotor of the third embodiment, the radial clearance between the proximal end of the diffuser 3 and the distal end of the compressor impeller 5 is between 2 mm and 4 mm;

[0096] The axial clearance between the heat-insulating tile 13 and the front end of the turbine rotor 6 is between 2 mm and 8 mm.

[0097] The axial clearance between the front end wall 80 of the sealing ring 8 and the rear end far from the axial center of the compressor impeller 5 is between 3 mm and 9 mm.

[0098] The above are only preferred embodiments of the present application and are not intended to limit the present application. For those skilled in the art, the present application may have various changes and variations. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present application should all be included in the protection scope of the present application.

Claims

1. An interstage sealing device between a small gas turbine compressor and a turbine rotor, characterized in that: The diffuser (3), the sealing ring (8) and the combustion chamber nozzle ring (4) are all in annular 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 coaxial distal ring wall (81) and a proximal ring wall (82) are formed on the rear side of the front end wall (80) of the sealing ring (8); A circle of heat-insulating tiles (13) is installed between the rear ends of the distal annular wall (81) and the proximal annular wall (82); The front end wall (80), the distal ring wall (81), the proximal ring wall (82) and the heat insulation tile (13) of the sealing ring (8) together form an interstage sealing inner cavity (1-20); The front end wall (80) of the sealing ring (8) is fixed to the proximal axial center portion of the diffuser (3); The rear end proximal to the axial center of the compressor impeller (5) is butted against the front end proximal to the axial center of the turbine rotor (6); A spacing area is formed between a rear end distal to the axial center of the compressor impeller (5) and a front end distal to the axial center of the turbine rotor (6); The rear end of the compressor impeller (5) is far from the axial center and is formed with an inner isolation ring (51) extending backward; The proximal side of the proximal ring wall (82) is in contact with the distal side of the inner isolation ring (51); The proximal end of the diffuser (3) is on the distal side of the compressor impeller (5) and has a radial gap; The distal ring wall (81) of the sealing ring (8) and the distal end of the heat insulation tile (13) are in sealing cooperation with the combustion chamber nozzle ring (4); An axial gap is formed between the heat insulating tile (13) and the front end of the turbine rotor (6) and the far axial center. An axial gap is formed between the front end wall (80) of the sealing ring (8) and the rear end far from the axial center of the compressor impeller (5); A piston ring mounting groove is formed along the circumferential direction at the rear end of the distal axis side of the distal axis ring wall (81) of the sealing ring (8); The piston ring (11) is assembled into 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 portion of the outer nozzle ring (41) is bent radially inward, and the rear portion is bent axially backward; The front portion of the inner nozzle ring (42) is bent radially outward, and the rear portion is bent axially backward; The inner nozzle ring (42) and the outer nozzle ring (41) form a hot gas injection channel; The front end of the outer nozzle ring (41) is formed with an outer isolation ring (45) extending forward; The front end of the outer isolation ring (45) is sleeved outside the rear end of the distal ring wall (81) of the sealing ring (8); The outer isolation ring (45) is sealed with the distal ring wall (81) of the sealing ring (8) via a piston ring (11); A comb tooth structure (510) is formed along the circumferential direction on the distal axis side of an inner isolation ring (51) extending backward from the distal axis center of the rear end of the compressor impeller (5); The proximal side of the proximal ring wall (82) of the sealing ring (8) is in contact with the comb tooth structure (510) on the distal side of the inner isolation ring (51).

2. The interstage sealing device between the compressor and the turbine rotor of a small gas turbine according to claim 1, characterized in that: The heat-insulating tile (13) is in the shape of a fan ring; The two wings of the heat-insulating tile (13) are respectively provided with grooves and convex ribs; A plurality of heat-insulating tiles (13) are matched with the grooves and convex ribs and connected end to end to form a complete sealing ring.

3. The interstage sealing device between the compressor and the turbine rotor of a small gas turbine according to claim 2, characterized in that: The front and rear thickness of the distal end of the heat insulation tile (13) is greater than the axial gap between the rear end of the distal ring wall (81) of the sealing ring (8) and the front part of the outer nozzle ring (41); The distal end of the heat-insulating tile (13) is squeezed and sealed by the rear end of the distal ring wall (81) of the sealing ring (8) and the front part of the outer nozzle ring (41).

4. The interstage sealing device between the compressor and the turbine rotor of a small gas turbine according to claim 2, characterized in that: A heat insulation tile mounting groove is formed along the circumferential direction at the rear end of the distal axis side of the proximal axis ring wall (82) of the sealing ring (8); The proximal end of the heat insulation tile (13) is inserted into the heat insulation tile installation groove and is sealed with the proximal ring wall (82) of the sealing ring (8).

5. The interstage sealing device between the compressor and the turbine rotor of a small gas turbine according to claim 4, characterized in that: The rear end of the proximal ring wall (82) of the sealing ring (8) is thicker than the front end.

6. The interstage sealing device between the compressor and the turbine rotor of a small gas turbine according to claim 1, characterized in that: The radial clearance between the proximal end of the diffuser (3) and the distal end of the compressor impeller (5) is between 2 mm and 4 mm; The axial clearance between the heat insulation tile (13) and the front end of the turbine rotor (6) far from the axial center is between 2 mm and 8 mm; The axial clearance between the front end wall (80) of the sealing ring (8) and the rear end far from the axial center of the compressor impeller (5) is between 3 mm and 9 mm.

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