CN111237017A - Sealing and heat-dissipating structure for gas turbine interstage - Google Patents
Sealing and heat-dissipating structure for gas turbine interstage Download PDFInfo
- Publication number
- CN111237017A CN111237017A CN202010097147.2A CN202010097147A CN111237017A CN 111237017 A CN111237017 A CN 111237017A CN 202010097147 A CN202010097147 A CN 202010097147A CN 111237017 A CN111237017 A CN 111237017A
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- China
- Prior art keywords
- cavity
- seal
- downstream
- upstream
- disc
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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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
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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
The invention discloses a sealing heat dissipation structure between gas turbine stages, which mainly comprises: the sealing ring, the stationary blade inner shroud, the upstream disc cavity, the downstream disc cavity, the air supply cavity, the axial hole and the sealing end; the air supply cavity is enclosed by the sealing ring and the inner surrounding belt of the stationary blade, cold air enters the upstream disc cavity and the downstream disc cavity through the axial holes distributed on the upstream end surface and the downstream end surface of the air supply cavity, the wheel disc is cooled, and then external fuel gas is gathered through the sealing end to prevent high-temperature fuel gas from entering the disc cavity; the invention has the advantages that the structure is simple and easy to realize, the wheel disc is cooled by adopting a mode of combining impact cooling and convection cooling, the problem of friction resistance and temperature rise of cold air is relieved, the control of the cooling air quantity is simpler, the deviation is smaller, meanwhile, the invention realizes the synchronous cooling of the wheel disc at the upstream and the downstream through the axial hole, and the influence of the sealing structure between the disc cavities at the upstream and the downstream on the control of the cold air is reduced.
Description
Technical Field
The invention relates to the field of turbines of high-temperature components of gas turbines, in particular to a sealing and heat-dissipating structure between stages of a gas turbine.
Background
In a gas turbine, the working temperature of a rotor disc is high, and if the temperature of the rotor disc is not controlled, the strength of the rotor disc is reduced due to the high temperature, so that the safe operation of a unit is influenced. The common practice is to introduce cooling air and cool off the rim plate, cooling air generally can follow the air feed chamber air feed of quiet leaf peripheral band department, after gas flows through quiet leaf inner channel, get into the air feed chamber that quiet leaf inner band and stage seal ring constitute, cool off adjacent rim plate respectively after sealing the tooth through the dish chamber entry at last, finally, the gas after cooling the rim plate can be gathered into the mainstream through sound clearance, play the seal to sound clearance, prevent that high temperature gas from invading the dish chamber, cause the ablation to the rim plate.
The scheme of the prior art is that the air supply of cooling air is realized through a single-side radial hole on an inter-stage sealing ring, the air flow flows into a side disc cavity from an air supply chamber through the radial hole, one part of the air flow flows into an upstream disc cavity through a first-stage seal, the other part of the air flow flows into a downstream disc cavity through a seal tooth between the sealing ring and a rotating shaft, and finally two parts of cold air flow into a main flow through a sealing structure, wherein the air supply flow of the upstream disc cavity and the downstream disc cavity is controlled by each-stage seal, so that the radial clearance of each-stage seal is of great importance to the cooling and sealing of the. However, the problem that the prior art has is that, because of the influence of factors such as vibration, thermal expansion and the like in the process of starting and stopping the unit and running the unit, the sealing part can rub against the rotor, thereby causing the change of the sealing radial clearance, which can not be guaranteed at the design value, therefore, the control deviation of the cooling air amount is larger, in addition, the method mainly adopts a convection cooling mode for cooling the wheel disc, the cooling efficiency is lower, the cooling air passes through the sealing structure and then has heat exchange and temperature rise, because of the influence of friction resistance, the air flow temperature can be additionally increased, the cooling of the wheel disc is not facilitated, and more cooling air needs to be consumed under the same wheel disc cooling requirement.
Disclosure of Invention
In view of this, the invention provides a sealing and heat dissipating structure for an interstage of a gas turbine, which can solve the problem of sealing and heat dissipating of the interstage of the existing gas turbine.
For this purpose, the present invention is implemented by the following technical means.
A gas turbine interstage seal heat dissipation structure, comprising: the sealing device comprises a gas supply cavity, an upstream disc cavity, a downstream disc cavity, an upstream axial hole, a downstream axial hole, a sealing end and a sealing end;
the air supply cavity is an annular cavity formed by a sealing ring and a stationary blade inner surrounding belt and is communicated with the upstream disc cavity and the downstream disc cavity through the upstream axial hole and the downstream axial hole in the sealing ring respectively;
the upstream disc cavity is an annular cavity which is formed by surrounding a previous stage wheel disc, the sealing ring and the inner surrounding belt of the stationary blade;
the downstream disc cavity is an annular cavity which is formed by enclosing a rear-stage wheel disc, the sealing ring and the inner surrounding belt of the stationary blade;
the sealing end is arranged on the end surface of the inner side of the sealing ring close to the rotating shaft;
and cold air enters the air supply cavity, flows into the upstream disc cavity and the downstream disc cavity through the upstream axial hole and the downstream axial hole respectively, and is collected into external high-temperature fuel gas through the sealing end.
Furthermore, a plurality of upstream axial holes and a plurality of downstream axial holes are uniformly distributed along the circumferential direction of the two end faces of the sealing ring respectively.
Further, the upstream axial hole and the downstream axial hole are taper holes, and the aperture close to the outer side is smaller than the aperture close to one side of the air supply cavity.
Furthermore, the upstream axial hole and the downstream axial hole are both prerotation holes, and the prerotation direction is consistent with the rotation direction of the wheel disc.
Further, the sealing end is a labyrinth seal.
Furthermore, the labyrinth seal is one of a honeycomb seal structure, a wear seal structure or a composite straight-through seal structure.
Furthermore, the sealing end is a rotating static system rotating sealing structure formed by two sides of the inner surrounding belt of the static blade and the adjacent wheel disc respectively.
Further, the sealed end is one of a radial seal or an axial seal.
Furthermore, the radial seal of the sealing end is that annular flanges on two sides of the inner peripheral belt of the stationary blade and annular flanges adjacent to the annular flange of the wheel disc form a staggered overlapping structure with gaps.
The invention has the following advantages:
1. the invention has simple and easy realization of the structure, cools the wheel disc by adopting a mode of combining impact cooling and convection cooling, reduces the sealing structure on the flow path of cold air, solves the problem of friction resistance and temperature rise of air flow through sealing, and has simpler control of the amount of the cold air and smaller deviation.
2. Aiming at the air supply mode of the interstage sealing ring, the synchronous impact cooling of the adjacent wheel discs is realized by supplying air through the axial holes without depending on ventilation of a sealing structure between an upstream disc cavity and a downstream disc cavity, and the influence of a sealing gap on the control precision of cooling air quantity is avoided.
Drawings
It is to be noted that while the drawings of the present invention show one of the several configurations of the present invention having the same distribution of features, some of the drawings are for the convenience of describing the present invention and simplifying the description, and are not intended to indicate or imply that the referenced device or element must have a particular orientation, be constructed and operated in a particular orientation, and therefore should not be considered limiting of the present invention.
FIG. 1 is a schematic half-sectional view in example 1 of the present invention;
FIG. 2 is a schematic view of a sealing ring structure of the present invention;
FIG. 3 is a downstream axial bore elevation view of the present invention;
FIG. 4 is a side view of a downstream axial bore of the present invention;
fig. 5 is a partial schematic view of a sealing end in embodiment 2 of the present invention.
In the figure:
1-a gas supply cavity; 2-an upstream disc chamber; 3-a downstream disc chamber; 4-an upstream axial bore; 5-downstream axial bore; 6-sealing the end; 7-sealing the end.
Detailed Description
In the description of the present invention, it should be noted that the upstream-downstream direction is the left-right direction shown in the figure, and refers to the overall gas flow direction in the operation of the gas turbine.
The invention will be further explained with reference to the drawings.
Example 1
A gas turbine interstage seal heat dissipation structure, comprising: the air supply cavity 1, the upstream disc cavity 2, the downstream disc cavity 3, the upstream axial hole 4, the downstream axial hole 5, the sealing end 6 and the sealing end 7.
As shown in fig. 1, the air supply cavity 1 is a cavity formed by the sealing ring and the inner peripheral band of the stationary blade, the whole cavity is annular, the air supply cavity is communicated with a cold air channel in the stationary blade outwards, a plurality of axial holes are symmetrically processed on the end faces of two sides of the sealing ring as shown in fig. 2, the axial holes are uniformly distributed along the circumferential direction of the surface of the sealing ring, wherein an upstream axial hole 4 is formed on one side close to the upstream, and a downstream axial hole 5 is formed on one side close to the downstream. The air supply cavity 1 is communicated with the upstream disc cavity 2 and the downstream disc cavity 3 through an upstream axial hole 4 and a downstream axial hole 5 respectively.
The upstream disc cavity 2 is an annular cavity formed by a front-stage wheel disc, a sealing ring and a stationary blade inner surrounding band.
The downstream disc cavity 3 is an annular cavity formed by a rear-stage wheel disc, a sealing ring and a stationary blade inner surrounding band.
Preferably, as shown in fig. 1, the upstream axial hole 4 and the downstream axial hole 5 are both taper holes, the hole diameter near the outer side is smaller than the hole diameter near one side of the air supply cavity 1, the cold airflow velocity entering the disk cavity is increased through the taper hole design, the impact heat dissipation effect is enhanced, further, the upstream axial hole 4 and the downstream axial hole 5 are both pre-rotation holes, the pre-rotation direction is consistent with the rotation direction of the disk, in combination with fig. 3-4, it is assumed that, when viewed from the gas downstream-upstream direction, the disk is in the right-rotation direction relative to the stationary blade and the sealing ring, taking a downstream axial hole 5 in the horizontal direction as an example, the outlet hole of the downstream axial hole 5 is downwardly deviated along the tangential direction of the circumference, the center line of the downstream axial hole is β degrees with the normal phase direction, the purpose of providing a tangential velocity of the cold airflow when entering the downstream disk cavity 3, reducing the relative velocity of the disk and.
An annular sealing end 6 is arranged on the end face of the inner side, close to the rotating shaft, of the sealing ring and is located between the sealing ring and the rotating shaft, and the sealing end 6 is labyrinth seal.
Preferably, the labyrinth seal is a wear seal structure, which helps to design the clearance between the sealing ring and the rotating shaft to be smaller, reduces the abrasion of the sealing mechanism and provides better sealing effect.
The sealing end 7 is preferably designed to be a rotating static system rotating sealing structure formed by two sides of the inner surrounding belt of the stator blade and the adjacent wheel disc respectively, and adopts a radial sealing structure, as shown in fig. 1, and the annular flanges on two sides of the inner surrounding belt of the stator blade and the annular flanges of the adjacent wheel disc form a staggered stacking structure with gaps respectively.
In the working process, cold air enters the air supply cavity 1 through a cold air channel in the static vane, and respectively and acceleratedly rushes into the upstream disc cavity 2 and the downstream disc cavity 3 through the upstream axial hole 4 and the downstream axial hole 5 of the air supply cavity 1, continuously performs heat exchange with the wheel disc after impacting the wheel disc, and finally converges into external high-temperature gas through the sealing end 7 to prevent the high-temperature gas from invading the disc cavity.
Example 2
The specific difference between this embodiment and embodiment 1 is that, as shown in fig. 5, the sealing end 7 is a rotating-static system rotating sealing structure formed by two sides of the stationary blade inner shroud and the adjacent wheel disc respectively, and an axial sealing structure is adopted, specifically, annular flanges at two sides of the stationary blade inner shroud abut against the inner side end faces of the adjacent wheel discs and a fixed gap is left.
Although the present invention has been described in detail, those skilled in the art should understand that various combinations, modifications and substitutions can be made to the technical solution of the present invention without departing from the spirit and scope of the technical solution of the present invention, and the technical solution of the present invention is covered by the claims of the present invention.
Claims (9)
1. A gas turbine interstage seal heat dissipation structure, characterized by comprising: the device comprises a gas supply cavity (1), an upstream disc cavity (2), a downstream disc cavity (3), an upstream axial hole (4), a downstream axial hole (5), a sealing end (6) and a sealing end (7);
the air supply cavity (1) is an annular cavity formed by a sealing ring and a stationary blade inner peripheral belt, and is communicated with the upstream disc cavity (2) and the downstream disc cavity (3) through the upstream axial hole (4) and the downstream axial hole (5) on the sealing ring respectively;
the upstream disc cavity (2) is an annular cavity which is formed by a front-stage wheel disc, the sealing ring and the inner surrounding belt of the stationary blade in a surrounding way;
the downstream disc cavity (3) is an annular cavity which is formed by a rear-stage wheel disc, the sealing ring and the inner surrounding belt of the stationary blade in a surrounding way;
the sealing end (6) is arranged on the end surface of the inner side of the sealing ring close to the rotating shaft;
and cold air enters the air supply cavity (1), flows into the upstream disc cavity (2) and the downstream disc cavity (3) through the upstream axial hole (4) and the downstream axial hole (5) respectively, and is collected into external high-temperature fuel gas through the sealing end (7).
2. The interstage seal heat dissipation structure of a gas turbine as claimed in claim 1, wherein the upstream axial hole (4) and the downstream axial hole (5) are distributed uniformly in the circumferential direction of two end faces of the seal ring.
3. The gas turbine interstage seal heat dissipation structure according to claim 1, wherein the upstream axial hole (4) and the downstream axial hole (5) are both taper holes, and the hole diameter close to the outer side is smaller than that close to the gas supply cavity (1).
4. The gas turbine interstage seal heat dissipation structure according to any one of claims 1 to 3, wherein the upstream axial hole (4) and the downstream axial hole (5) are both prerotation holes, and the prerotation direction is consistent with the rotation direction of a wheel disc.
5. The gas turbine interstage seal heat dissipation structure according to claim 1, wherein the seal end (6) is a labyrinth seal.
6. The gas turbine interstage seal and heat dissipation structure of claim 5, wherein the labyrinth seal is one of a honeycomb seal structure, a wear seal structure or a composite straight through seal structure.
7. The interstage seal heat dissipation structure of the gas turbine as claimed in claim 1, wherein the seal end (7) is a rotating-static system rotating seal structure formed by two sides of the stator blade inner shroud and adjacent wheel discs respectively.
8. The gas turbine interstage seal heat dissipation structure according to claim 7, wherein the seal end (7) is one of a radial seal and an axial seal.
9. The interstage seal and heat dissipation structure of a gas turbine as claimed in claim 8, wherein the seal end (7) is in a staggered overlapping structure that annular flanges are arranged on two sides of the inner periphery of the stator blade and gaps are formed between the annular flanges and the annular flanges arranged on the side faces of the adjacent wheel discs respectively.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010097147.2A CN111237017A (en) | 2020-02-17 | 2020-02-17 | Sealing and heat-dissipating structure for gas turbine interstage |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010097147.2A CN111237017A (en) | 2020-02-17 | 2020-02-17 | Sealing and heat-dissipating structure for gas turbine interstage |
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| Publication Number | Publication Date |
|---|---|
| CN111237017A true CN111237017A (en) | 2020-06-05 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010097147.2A Pending CN111237017A (en) | 2020-02-17 | 2020-02-17 | Sealing and heat-dissipating structure for gas turbine interstage |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112610336A (en) * | 2020-12-21 | 2021-04-06 | 杭州汽轮动力集团有限公司 | Interstage seal ring sealing structure |
| CN112610335A (en) * | 2020-12-21 | 2021-04-06 | 杭州汽轮动力集团有限公司 | Sealing structure for turbine disk cavity of gas turbine |
| CN113047914A (en) * | 2021-04-22 | 2021-06-29 | 浙江燃创透平机械股份有限公司 | Sealing structure between turbine stages of gas turbine |
| CN113586251A (en) * | 2021-07-22 | 2021-11-02 | 西安交通大学 | Part cooling-wheel rim sealing structure for stepwise utilization of cooling airflow of gas turbine |
| CN114486222A (en) * | 2022-01-26 | 2022-05-13 | 沈阳航空航天大学 | Composite experimental device for movable blade heat engine of air seal type gas turbine |
-
2020
- 2020-02-17 CN CN202010097147.2A patent/CN111237017A/en active Pending
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112610336A (en) * | 2020-12-21 | 2021-04-06 | 杭州汽轮动力集团有限公司 | Interstage seal ring sealing structure |
| CN112610335A (en) * | 2020-12-21 | 2021-04-06 | 杭州汽轮动力集团有限公司 | Sealing structure for turbine disk cavity of gas turbine |
| CN112610336B (en) * | 2020-12-21 | 2021-11-12 | 杭州汽轮动力集团有限公司 | Interstage seal ring sealing structure |
| CN113047914A (en) * | 2021-04-22 | 2021-06-29 | 浙江燃创透平机械股份有限公司 | Sealing structure between turbine stages of gas turbine |
| CN113586251A (en) * | 2021-07-22 | 2021-11-02 | 西安交通大学 | Part cooling-wheel rim sealing structure for stepwise utilization of cooling airflow of gas turbine |
| CN113586251B (en) * | 2021-07-22 | 2023-03-14 | 西安交通大学 | Part cooling-wheel rim sealing structure for stepwise utilization of cooling airflow of gas turbine |
| CN114486222A (en) * | 2022-01-26 | 2022-05-13 | 沈阳航空航天大学 | Composite experimental device for movable blade heat engine of air seal type gas turbine |
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Address after: 310022 No. 357 Shiqiao Road, Zhejiang, Hangzhou Applicant after: Hangzhou Steam Turbine Holding Co.,Ltd. Address before: 310022 No. 357 Shiqiao Road, Zhejiang, Hangzhou Applicant before: HANGZHOU TURBINE POWER GROUP CO.,LTD. |