CN111237017A - A gas turbine interstage sealed heat dissipation structure - Google Patents

Abstract

The invention discloses an interstage sealing heat dissipation structure of a gas turbine, which mainly comprises: a sealing ring, an inner shroud of a stationary blade, upstream and downstream disk cavities, an air supply cavity, an axial hole and a sealing end; the air supply cavity Surrounded by the sealing ring and the inner shroud of the stator blade, the cold air enters the upstream and downstream disc cavities through the axial holes distributed on the upper and downstream end faces of the air supply cavity, and cools the wheel disc , and then the external gas is flowed through the sealing end to prevent the high-temperature gas from entering the disc cavity; the invention has the advantages that the structure is simple and easy to implement, and the wheel disc is cooled by a combination of impact cooling and convection cooling, and alleviates the The problem of frictional resistance and temperature rise of the cold air is solved, the control of the cooling air volume is also simpler, and the deviation is smaller. Influence of sealing structure between downstream disc cavities on cooling air control.

Description

A gas turbine interstage sealed heat dissipation structure

technical field

The invention relates to the field of high-temperature component turbines of gas turbines, in particular to an interstage sealing and heat dissipation structure of a gas turbine.

Background technique

In a gas turbine, the working temperature of the rotor disc is relatively high. If its temperature is not controlled, its excessively high temperature will lead to a decrease in strength, thereby affecting the safe operation of the unit. The usual practice is to introduce cooling air to cool the wheel disc. The cooling air is generally supplied from the air supply cavity at the outer band of the stator. After the gas flows through the inner channel of the stator, it enters the inner shroud of the stator and is sealed between the stages. The air supply chamber formed by the ring finally cools the adjacent disks through the sealing teeth at the entrance of the disk cavity. Finally, the gas after cooling the disks will flow into the mainstream through the dynamic and static gaps, sealing the dynamic and static gaps. Prevent high-temperature gas from intruding into the disc cavity and cause ablation to the roulette.

The solution in the prior art is that the supply of cooling air is realized through a single-sided radial hole on the interstage sealing ring, the air flow flows from the air supply chamber through the radial hole into one side of the disk cavity, and a part of the air flow flows through the primary seal. The upstream disk cavity, the other part flows into the downstream disk cavity through the sealing teeth between the sealing ring and the rotating shaft, and the last two parts of the cold air flow into the mainstream through the sealing structure. Therefore, the radial clearance of the seals at all levels is very important for the cooling and sealing of the disc cavity. However, the existing technical solution has the problem that due to the start and stop of the unit and the influence of vibration, thermal expansion and other factors during the operation of the unit, the sealing part will rub against the rotor, resulting in changes in the radial clearance of the seal, which cannot be guaranteed in the design. Therefore, the control deviation of the cooling air volume is relatively large. In addition, this method is mainly convective cooling for the cooling of the roulette, and the cooling efficiency is low. Influence, the airflow temperature will increase additionally, which is not conducive to the cooling of the roulette, and more cooling air needs to be consumed under the same cooling requirement of the roulette.

SUMMARY OF THE INVENTION

In view of this, the present invention provides a gas turbine interstage sealing heat dissipation structure, which can solve the problem of the existing gas turbine interstage sealing heat dissipation.

For this purpose, the present invention is implemented by the following technical solutions.

An interstage sealed heat dissipation structure of a gas turbine, comprising: an air supply cavity, an upstream disk cavity, a downstream disk 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 the sealing ring and the inner shroud of the stator blade, which communicates with the upstream disk through the upstream axial hole and the downstream axial hole on the sealing ring respectively. a cavity and the downstream disc cavity;

The upstream disk cavity is an annular cavity surrounded by the previous stage wheel disk, the sealing ring and the inner shroud of the stator blade;

The downstream disk cavity is an annular cavity surrounded by the rear-stage wheel disk, the sealing ring and the inner shroud of the stator blade;

The sealing end is installed on the inner end face of the sealing ring close to the rotating shaft;

Cold air enters the air supply cavity, flows into the upstream disk cavity and the downstream disk cavity through the upstream axial hole and the downstream axial hole respectively, and flows into the external high-temperature gas through the sealed end.

Further, there are a plurality of the upstream axial holes and the downstream axial holes, which are respectively uniformly distributed along the circumferential direction of the two end faces of the sealing ring.

Further, both the upstream axial hole and the downstream axial hole are tapered holes, and the hole diameter on the outer side is smaller than the hole diameter on the side close to the air supply cavity.

Furthermore, the upstream axial hole and the downstream axial hole are both pre-rotation holes, and the pre-rotation 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 sealing structure, a wear-bearing sealing structure or a compound straight-through sealing structure.

Further, the sealing end is a rotary sealing structure of the rotating stationary system formed by two sides of the inner shroud of the stator and the adjacent wheel discs respectively.

Further, the sealing end is either a radial seal or an axial seal.

Furthermore, the radial sealing of the sealing end is that the annular flanges on both sides of the inner shroud of the stator vane form a dislocation stacking structure with a gap between the annular flanges of the adjacent disks.

The present invention has the following advantages:

1. The structure of the present invention is simple and easy to implement. The wheel disc is cooled by a combination of impact cooling and convection cooling, and the sealing structure on the cold air flow path is reduced, which solves the problem of frictional resistance and temperature rise caused by the air passing through the seal. The quantity control is also simpler and the deviation is smaller.

2. The present invention is aimed at the air supply method of the interstage sealing ring. It does not rely on the sealing structure between the upstream and downstream disc cavities to ventilate, and the air supply through the axial hole realizes the synchronous impact cooling of the adjacent discs, avoiding the sealing. The effect of clearance on the accuracy of cooling air volume control.

Description of drawings

It should be noted that, in the drawings of the present invention, for a plurality of identical structures with the same distribution features, partial views only show one of them. For the convenience of describing the present invention and simplifying the description, it is not intended to indicate or imply that the indicated device or element must have The particular orientation, construction and operation in the particular orientation are therefore not to be construed as limitations of the invention.

1 is a half-section schematic diagram in Embodiment 1 of the present invention;

Fig. 2 is the structural schematic diagram of the sealing ring of the present invention;

Fig. 3 is the front view of the downstream axial hole of the present invention;

Figure 4 is a side view of the downstream axial hole of the present invention;

FIG. 5 is a partial schematic diagram of the sealed end in Example 2 of the present invention.

In the picture:

1- Air supply cavity; 2- Upstream disc cavity; 3- Downstream disc cavity; 4- Upstream axial hole; 5- Downstream axial hole; 6- Sealed end; 7- Sealed end.

Detailed ways

In the description of the present invention, it should be noted that upstream-downstream refers to the left-right direction shown in the figure, which refers to the overall flow direction of gas during the operation of the gas turbine.

The present invention will be further described below with reference to the accompanying drawings.

Example 1

An interstage sealed heat dissipation structure of a gas turbine comprises: an air supply cavity 1 , an upstream disk cavity 2 , a downstream disk cavity 3 , an upstream axial hole 4 , a downstream axial hole 5 , a sealing end 6 and a sealing end 7 .

As shown in Figure 1, the air supply cavity 1 is a cavity formed by the sealing ring and the inner shroud of the stator. The side end faces are symmetrically machined with a plurality of axial holes, which are uniformly distributed along the circumferential direction of the surface, wherein the upstream axial hole 4 is located on the upstream side, and the downstream axial hole 5 is located on the downstream side. The air supply cavity 1 communicates with the upstream disk cavity 2 and the downstream disk cavity 3 through the upstream axial hole 4 and the downstream axial hole 5 respectively.

The upstream disc cavity 2 is an annular cavity formed by the former stage disc, the sealing ring and the inner shroud of the stator blade.

The downstream disc cavity 3 is an annular cavity formed by the rear-stage disc, the sealing ring and the inner shroud of the stator blade.

Preferably, as shown in FIG. 1 , both the upstream axial hole 4 and the downstream axial hole 5 are tapered holes, and the hole diameter on the outer side is smaller than the hole diameter on the side close to the air supply cavity 1 , and the tapered hole design is used to lift the hole into the air. The speed of the cold air flow in the disk cavity enhances the impact heat dissipation effect; further, the upstream axial hole 4 and the downstream axial hole 5 are both pre-rotation holes, and the pre-rotation direction is consistent with the rotation direction of the wheel disk, as shown in Figure 3-4, Assuming that the disc is clockwise relative to the stator vane and the sealing ring when viewed from the downstream-upstream direction of the gas The purpose is to provide a tangential speed of the cold air when it enters the downstream disk cavity 3, reduce the relative speed of the wheel disk and the cold air in the rotation direction, and improve the cooling effect.

An annular sealing end 6 is installed on the inner end face of the sealing ring close to the rotating shaft, which is located between the sealing ring and the rotating shaft, and the sealing end 6 is a labyrinth seal.

Preferably, the labyrinth seal is a wear-bearing sealing structure, which helps to design a smaller gap between the sealing ring and the rotating shaft, reduces the wear of the sealing mechanism, and provides a better sealing effect.

The sealing end 7 is preferably designed as a rotary sealing structure of the rotating stationary system formed by the two sides of the inner shroud of the stator and the adjacent wheel discs, and adopts a radial sealing structure, as shown in FIG. 1 , the inner shroud of the stator blade. The annular flanges on both sides respectively form a dislocation stacking structure with gaps with the annular flanges of the adjacent wheel discs.

During the working process, the cold air enters the air supply cavity 1 through the cold air passage in the stationary vane, and accelerates into the upstream disk cavity 2 and the downstream disk cavity 3 through the upstream axial hole 4 and the downstream axial hole 5 of the air supply cavity 1, respectively. After impacting the roulette, it continues to exchange heat with the roulette, and finally enters the external high-temperature gas through the sealing end 7, preventing the high-temperature gas from invading the disk 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 rotary sealing structure formed by the two sides of the inner shroud of the stationary blade and the adjacent wheel discs respectively, and adopts an axial sealing structure. The sealing structure, in particular, is that the annular flanges on both sides of the inner shroud of the stationary blade are abutted against the inner end face of the adjacent wheel disc, and a fixed gap is left.

The above embodiments are only used to illustrate the present invention, but not to limit the present invention. Although the present invention has been described in detail, those of ordinary skill in the art should be able to understand that various combinations, modifications or replacements of the technical solutions of the present invention will Without departing from the spirit and scope of the technical solutions of the present invention, all should be included in the scope of 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.

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