CN115289499B - Hollow support plate of gas inlet of combustion chamber of gas turbine - Google Patents
Hollow support plate of gas inlet of combustion chamber of gas turbine Download PDFInfo
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- CN115289499B CN115289499B CN202211219496.2A CN202211219496A CN115289499B CN 115289499 B CN115289499 B CN 115289499B CN 202211219496 A CN202211219496 A CN 202211219496A CN 115289499 B CN115289499 B CN 115289499B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/26—Controlling the air flow
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
Abstract
The invention belongs to the technical field of gas turbines, and particularly relates to a hollow support plate of a gas inlet of a combustion chamber of a gas turbine; comprises a support plate main body; a front diffusion channel is arranged at the air inlet of the combustion chamber of the gas turbine; the support plate main body is in a wing shape, and a first air entraining channel and a second air entraining channel are arranged in the support plate main body; the first bleed air channel is used for guiding air on the front side of the support plate main body to the outside of the front diffusion channel; the second air-entraining channel can guide the air at the position, close to the wall surface of the flow divider, of the outer flow-dividing channel to the rear side of the support plate main body and lead the air out towards the rear side of the inner flow-dividing channel. Reasonable technical measure is adopted in the scheme, the air entraining channels are arranged in the hollow support plate, air is entrained through the air entraining channels, the support plate main body is smoother to an air flow field entering a combustion chamber, the caused blockage can be reduced, the stability of the internal flow field of the multi-channel diffuser is improved, meanwhile, the flow field distortion effect in the circumferential direction is also hindered, and the performance of the diffuser is improved.
Description
Technical Field
The invention belongs to the technical field, and particularly relates to a hollow support plate of an air inlet of a combustion chamber of a gas turbine.
Background
A gas turbine, i.e., a gas turbine engine, is an internal combustion type power machine that converts energy of combustion gas into useful work, and is widely used in various fields, for example: the power generation device is applied to the field of civil power generation or used as a power device in an airplane or a large ship. The working process of the gas turbine is as follows: the compressor continuously sucks air from the atmosphere and compresses the air; the compressed air enters the combustion chamber, is mixed with the gas sprayed into the combustion chamber and then is combusted, so that high-temperature gas is formed, then the high-temperature gas flows into the gas turbine to expand and do work, and the high-temperature gas is used for pushing the turbine to drive the gas compressor to rotate together; the gas turbine is a device with good cleanness and high efficiency, and has the advantages of small volume, low weight and the like.
The pressure ratio of a high-performance gas turbine engine is usually high, the airflow speed at the outlet of a compressor is high, generally between 130m/s and 170m/s, and the number of the compressor can reach about 220m/s, at the moment, stable ignition and combustion are difficult to organize in a combustion chamber, and high-speed airflow can bring about large pressure loss when flowing in the combustion chamber. During design, before the combustion of the combustor structure, the air flow speed is generally required to be reduced to 1/5 of the air flow at the outlet of the air compressor, so that the stable combustion of the combustor structure is facilitated, and the diffuser is used for reducing the air flow speed at the outlet of the air compressor, converting the kinetic energy of the air flow into static pressure as much as possible, forming a stable outlet flow field, reducing the total pressure loss of the combustor and further reducing the oil consumption rate of the engine.
Generally, increasing the length of the diffuser is beneficial to reduce the velocity of the outlet gas flow and thus the total pressure loss in the combustion chamber, but in order to reduce the overall length and weight of the engine, the diffuser needs to be as short and compact as possible, so a diffuser structure with excellent performance must compromise the contradiction between pressure loss and length.
In order to take account of the contradiction between the pressure loss and the length of the diffuser structure, the applicant previously filed a patent with the patent number of CN114263933A and the patent name of the combined multi-channel diffuser of the gas turbine and the diffuser air inlet structure thereof.
In the application of the technology, the applicant finds that with the development of the advanced gas turbine, the mach number Ma of the air flow at the outlet of the diffuser is further increased, the total pressure loss of the diffuser is increased sharply, and meanwhile, the flow separation is easy to occur, such as: the flow field stability at the air inlet of the combustion chamber of the gas turbine is also influenced to a certain extent by the structures such as the splitter and the support plate supporting the splitter. For this reason, it is necessary to optimize these structures so as to stabilize the air flow field at the intake port of the combustor.
Disclosure of Invention
In order to stabilize the air flow field at the air inlet of the combustion chamber of the multi-channel diffusion structure, the scheme provides a hollow support plate of the air inlet of the combustion chamber of the gas turbine.
The technical scheme adopted by the invention is as follows:
a hollow support plate for an air inlet of a combustion chamber of a gas turbine comprises a support plate main body;
a front diffusion channel is arranged at the air inlet of the combustion chamber of the gas turbine; an annular flow divider is arranged in the front diffusion channel; the flow divider divides the preposed diffusion channel into an inner diversion channel and an outer diversion channel;
the support plate main body is arranged in the inner diversion channel, and is connected with and supports the flow divider; the support plate main body is in a wing shape, and a first air entraining channel and a second air entraining channel are arranged in the support plate main body; the first bleed air channel is used for guiding air on the front side of the support plate main body to the outside of the front diffusion channel; the second air-entraining channel can guide the air at the position, close to the wall surface of the flow divider, of the outer flow-dividing channel to the rear side of the support plate main body and lead the air out towards the rear side of the inner flow-dividing channel.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: a plurality of first bleed air inlets are formed in the wing head of the support plate main body and communicated with the first bleed air channel.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the first bleed air channel is also communicated with a first bleed air outlet, the first bleed air outlet is positioned outside the front diffusion channel, and air led out from the first bleed air outlet can be used for cooling parts of the gas turbine.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the gas turbine combustor comprises a combustor casing, and the combustor casing comprises a diffuser outer ring and a diffuser inner ring; the preposed diffusion channel is formed between the outer ring and the inner ring of the diffuser; the first bleed air outlet is arranged on the inner side of the inner ring of the diffuser.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the plurality of first bleed air inlets are arranged in a straight line along the width direction of the inner diversion channel.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the flow direction of the air led out from the second bleed air channel is a horizontal flow direction, or the flow direction of the air led out from the second bleed air channel has the same trend with the air flow direction in the inner branch flow channel.
As an alternative structure or a supplementary design of the hollow support plate of the gas turbine combustion chamber air inlet: the supporting plate main body is internally provided with a plurality of second air-entraining channels, a plurality of second air-entraining outlets are arranged at the head and the tail of the wing, and the second air-entraining outlets are correspondingly communicated with the second air-entraining channels.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: each second air-entraining passage is also correspondingly communicated with a second air-entraining inlet which is arranged on the wall surface of the outer ring side of the flow divider.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the plurality of second bleed air outlets are arranged in a straight line along the width direction of the inner diversion channel; the second bleed air inlets are arranged along the airflow direction in the outer flow dividing channel.
As an alternative structure or a supplementary design of the hollow support plate of the gas inlet of the gas turbine combustion chamber: the second air-entraining channel comprises a first section channel and a tail section channel which are communicated with each other; the first section of channel is inclined and not vertical to the direction of the airflow in the outer shunting channel; the tail-section channel is horizontal, or the air flow direction at the outlet of the tail-section channel has the same trend with the air flow direction in the inner diversion channel.
The invention has the beneficial effects that:
1. according to the scheme, through reasonable technical measures, the structure of the first air-entraining channel and the second air-entraining channel is arranged in the hollow support plate, and air is entrained through the air-entraining channels, so that the support plate main body is smoother for an air flow field entering a combustion chamber, the caused blockage is reduced, the stability of the flow field in the multi-channel diffuser is improved, meanwhile, the flow field distortion effect in the circumferential direction is also hindered, and the performance of the diffuser is improved;
2. the airflow entering the inner diversion passage can generate disturbance wake at the wing tail of the inner diversion passage, and then vortex masses are generated between the wing tail of the support plate main body and the sudden-expansion diffuser, and the vortex masses can influence the airflow stability at the outlet of the front diffuser.
3. Due to the air viscosity factor, a boundary flow field with a low air flow speed and a certain thickness is formed on the wall surface of the airflow channel, and the hollow support plate air-entraining channel structure in the scheme of the invention has the advantages that the thickness of the boundary flow field at the front edge of the diffuser support plate, the inner wall of the outer channel of the diffuser and the tail edge of the diffuser support plate is restrained to a certain extent by the suction effect of the air-entraining channel on the boundary flow field near the wall surface of the outer shunting channel, so that the stability of the front diffuser is improved;
4. the first bleed air channel in the scheme of the invention can lead the airflow to other positions, such as turbine blades of a gas turbine, electronic systems, fuel pipelines and other high-temperature components, and can be used for cooling the high-temperature components without leading air from an outer annular channel of the combustion chamber and causing the performance of the combustion chamber to be reduced;
5. according to the scheme of the invention, the hollow support plate guides air, the channels finely control the structural parameters of the air-guiding channels at each position, the air-guiding airflow flow at each position of the channels is effectively controlled, the fine regulation and control effect on the flow field in the whole diffuser channel is achieved, and the adaptability and fault tolerance of the preposed multi-channel diffuser are further improved;
6. the hollow support plate in the scheme of the invention further reduces the flow speed in the channel through the drainage function of the air-entraining channel, simultaneously reduces the flow loss of the preposed diffuser and improves the overall performance of the combustion chamber;
7. the inner diversion channel and the outer diversion channel in the scheme of the invention both adopt an equal pressure gradient design, and the flow proportion of the inner channel and the outer channel can be finely controlled by controlling the characteristic parameters of the inner channel and the outer channel, so that the pressure loss in the channels can be kept to be optimal under the condition of no flow separation.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below.
FIG. 1 is a block diagram of the gas turbine combustor at the inlet port in this scenario;
FIG. 2 is an overall structural view of the gas turbine combustor in the present embodiment;
FIG. 3 is an enlarged structural view of a portion A of FIG. 1;
FIG. 4 is a comparison of flow fields for solid and hollow plates.
In the figure: 1-a pre-diffusion channel; 11-an internal flow splitting channel; 12-an outer diversion channel; 2-sudden expansion of the diffusion channel; 21-outer ring channel; 22-inner ring channel; 23-a pressure expansion zone; 3-a combustion chamber casing; 31-the diffuser outer ring; 32-diffuser inner ring; 4-a flame tube; 5-a flow divider; 6-a fairing; 7-a plate body; 71-a first bleed air channel; 711-first bleed air inlet; 712-a first bleed air outlet; 72-a second bleed air passage; 721-a second bleed air inlet; 722-a second bleed air outlet; 81-stagnant zone; 82-a turbulent flow zone; 9-boundary flow field.
Detailed Description
The technical solutions in the embodiments will be described clearly and completely with reference to the accompanying drawings, and the described embodiments are only a part of the embodiments, but not all embodiments, and all other embodiments obtained by those skilled in the art without creative efforts will belong to the protection scope of the present solution based on the embodiments in the present solution.
As shown in fig. 1 to 4, the gas turbine combustor includes a flame tube 4, a combustor casing 3, a fairing 6, and the like, where the flame tube 4 is located inside the combustor casing 3, and a cavity inside the flame tube 4 is a combustion chamber, the fairing 6 is arranged outside the combustor casing 3, and a gap between the fairing 6 and an inner wall of the combustor casing 3 forms a sudden expansion diffusion channel 2, and the sudden expansion diffusion channel 2 includes components such as an outer ring channel 21, an inner ring channel 22, and a diffusion area 23. In addition, the air inlet of the gas turbine combustion chamber is positioned at the inner ring side of the combustion chamber casing 3, and a front diffusion channel 1 is arranged at the air inlet of the gas turbine combustion chamber; an annular flow divider 5 is arranged in the front diffusion channel 1; the flow divider 5 divides the preposed diffusion channel 1 into an inner flow dividing channel 11 and an outer flow dividing channel 12; the inner diversion channel 11 is used to divert air into the inner ring channel 22 and the outer diversion channel 12 is used to divert air into the outer ring channel 21. And the diffuser zone 23 is located at the end of the forward diffuser channel 1 and effects a split of the gas flow.
Because the flow splitter 5 arranged in the front diffusion channel 1 is supported by the support plate main body 7 arranged in the inner flow splitting channel 11, and about 36 support plate main bodies 7 are arranged in the circumferential direction of the flow splitter 5, when air entering an air inlet of a combustion chamber of a gas turbine flows into the flow splitter 5 and the support plate main bodies 7, due to factors such as air viscosity, a boundary flow field 9 with a certain thickness and a slow air flow speed is formed on the wall surface of the air flow channel, the boundary flow field 9 is distributed on the wall surfaces of the inner ring side and the outer ring side of the flow splitter 5, the boundary flow field 9 is also distributed on the outer wall surface of the support plate main body 7, if the support plate main body 7 adopts a solid wing shape, air flow is stopped at the wing head part impacting the support plate main body 7, and a stagnation region 81 is formed, as shown in (a) in fig. 4, the stagnation region 81 will cause the wing head part of the support plate main body 7 to form the boundary flow field 9 with a higher thickness, and the boundary 9 at the wing body surface of the wing head part also increases the thickness of the flow field 9. In addition, the air flows in the inner flow path 11 divided by the strut body 7 will rejoin at the wing tails of the strut body 7 and generate vortex masses, which in turn form turbulent flow regions 82, which turbulent flow regions 82 will cause turbulence in the bleed air flow and turbulent flow dispersion problems at the diffuser region 23. The formation of these boundary flow fields 9, stagnation areas 81 and turbulent flow areas 82 all contribute to a certain extent to the stability of the air flow in the entire combustion chamber.
Example 1
In order to reduce the thickness of the boundary flow field 9 formed on the wall surfaces of the flow splitter 5 and the plate body 7, and to eliminate or reduce the stagnation region 81 and the turbulent flow region 82, and to increase the stability of the air flow in the combustion chamber, the hollow plate in the present embodiment is designed, as shown in fig. 1 to 4.
The hollow support plate of the gas turbine combustion chamber air inlet of the embodiment comprises a support plate main body 7, and a first bleed air channel 71 and a second bleed air channel 72 are arranged inside the support plate main body 7.
The first bleed air channel 71 is used for guiding air at the front side of the support plate main body 7 to the outside of the front diffusion channel 1; a plurality of first bleed air inlets 711 are arranged at the wing heads of the support plate main body 7, and each first bleed air inlet 711 is communicated with the first bleed air channel 71. The first bleed air channel 71 is further communicated with a first bleed air outlet 712, the first bleed air outlet 712 is located outside the front diffuser channel 1, and air led out from the first bleed air outlet 712 can be used for cooling parts of the gas turbine. The preposed diffusion channel 1 is formed between a diffuser outer ring 31 and a diffuser inner ring 32; the first bleed air outlet 712 is disposed inboard of the diffuser inner ring 32. The plurality of first bleed air inlets 711 are arranged in a line in the width direction of the inner diversion passage 11.
In this embodiment, a plurality of first bleed air inlets 711 are provided at the wing head of the support plate main body 7, so that the air flow at the wing head of the support plate main body 7 can be directly led out through the first bleed air channel 71, thereby eliminating or reducing the stagnation area 81 at the wing head of the support plate main body 7, and reducing the influence of the excessive thickness of the boundary flow field 9 at this position on the stability of the air flow field of the whole combustion chamber, and the first bleed air channel 71 can guide the air flow to other positions, such as to the turbine blades of the gas turbine, the electronic system, the fuel pipeline and other high temperature components, and can be used for cooling these high temperature components without leading air from the outer ring channel 21 of the combustion chamber and causing the performance reduction of the combustion chamber.
The second bleed air channel 72 can guide the air in the outer diversion channel 12 close to the wall surface of the diverter 5 to the rear side of the support plate body 7 and lead out towards the rear side of the inner diversion channel 11. The flow direction of the air discharged from the second bleed air passage 72 is a horizontal flow direction, or the flow direction of the air discharged from the second bleed air passage 72 is approximately the same as the flow direction of the air in the inner flow dividing passage 11. A plurality of second bleed air channels 72 are arranged in the support plate main body 7, a plurality of second bleed air outlets 722 are arranged at the head and the tail of the wing, and the second bleed air outlets 722 are correspondingly communicated with the second bleed air channels 72. Each second bleed air channel 72 is also correspondingly communicated with a second bleed air inlet 721, and the second bleed air inlet 721 is arranged on the wall surface of the outer ring side of the flow divider 5. The plurality of second bleed air outlets 722 are arranged in a line in the width direction of the inner diversion passage 11; a plurality of second bleed air inlets 721 are arranged in the direction of the airflow in the outer distribution channel 12. The second bleed air passage 72 includes a first passage and a second passage that are communicated with each other; the first section of the channel is inclined and not perpendicular to the direction of the airflow in the outer shunting channel 12; the tail section channel is horizontal, or the tail section channel is converged to the airflow direction in the inner diversion channel (11) (namely, the airflow direction at the outlet of the tail section channel and the airflow direction in the inner diversion channel (11) have the same trend).
In the embodiment, the drainage through the second drainage channel utilizes the suction effect generated by the second drainage inlet on the boundary flow field 9 on the near-wall surface of the external flow distribution channel 12, so that the thickness of the boundary flow field 9 at the two drainage inlets on the side wall surface of the outer ring of the flow divider 5 is suppressed to a certain extent, and the stability of the front diffuser is further improved.
In addition, the second bleed air outlet 722 can be horizontal, or the second bleed air outlet 722 can lead the air flow out in the direction of the air flow in the inner diversion channel 11 (i.e. the direction of the air led out by the second bleed air channel has the same trend as the direction of the air flow in the inner diversion channel), so that the turbulent flow region 82 at the wing tail of the support plate body 7 is eliminated or eliminated, and the air flow direction at the wing tail of the support plate body 7 is more uniform and stable, as shown in fig. 4 (b).
Example 2
In order to reduce the thickness of the boundary flow field 9 formed on the wall surface of the plate body 7 and eliminate or reduce the stagnation region 81, the hollow plate in the present embodiment is designed as shown in fig. 1 to 4.
The hollow support plate of the gas turbine combustion chamber air inlet comprises a support plate main body 7; the support plate main body 7 is in a wing shape, and a first air guide channel 71 is arranged inside the support plate main body 7; the first bleed air channel 71 is used for guiding air on the front side of the main plate body 7 to the outside of the front diffusion channel 1.
A plurality of first bleed air inlets 711 are arranged at the wing heads of the support plate main body 7, and each first bleed air inlet 711 is communicated with the first bleed air channel 71. The first bleed air channel 71 is further communicated with a first bleed air outlet 712, the first bleed air outlet 712 is located outside the front diffuser channel 1, and air led out from the first bleed air outlet 712 can be used for cooling parts of the gas turbine.
The gas turbine combustion chamber comprises a combustion chamber casing 3, and the combustion chamber casing 3 comprises a diffuser outer ring 31 and a diffuser inner ring 32; the preposed diffusion channel 1 is formed between a diffuser outer ring 31 and a diffuser inner ring 32; the first bleed air outlet 712 is disposed inboard of the diffuser inner ring 32. The plurality of first bleed air inlets 711 are arranged in a line in the width direction of the inner diversion passage 11.
Example 3
In order to reduce the thickness of the boundary flow field 9 formed on the wall surface of the flow splitter 5, and to eliminate or reduce the turbulent flow region 82, and to increase the stability of the air flow in the combustion chamber, a hollow support plate in the present embodiment is designed, as shown in fig. 1 to 4.
The hollow support plate of the gas inlet of the combustion chamber of the gas turbine comprises a support plate main body 7; the support plate main body 7 is in a wing shape, and a second air guide channel 72 is arranged inside the support plate main body 7; the second bleed air channel 72 can guide the air in the outer diversion channel 12 close to the wall surface of the diverter 5 to the rear side of the support plate body 7 and lead out towards the rear side of the inner diversion channel 11.
The flow direction of the air introduced from the second bleed air passage 72 is the horizontal flow direction, or the flow direction of the air introduced from the second bleed air passage 72 is approximately the same as the flow direction in the inner diversion passage 11. A plurality of second bleed air channels 72 are arranged in the support plate main body 7, a plurality of second bleed air outlets 722 are arranged at the head and the tail of the wing, and the second bleed air outlets 722 are correspondingly communicated with the second bleed air channels 72.
Each second bleed air channel 72 is also correspondingly communicated with a second bleed air inlet 721, and the second bleed air inlet 721 is arranged on the wall surface of the outer ring side of the flow divider 5. The plurality of second bleed air outlets 722 are arranged in a line in the width direction of the inner diversion passage 11; a plurality of second bleed air inlets 721 are arranged in the direction of the airflow in the outer distribution channel 12. The second bleed air passage 72 includes a first passage and a second passage that are communicated with each other; the first section of the channel is inclined and not perpendicular to the direction of the airflow in the outer shunting channel 12; the tail section channel is horizontal or convergent to the direction of the gas flow in the inner flow dividing channel 11.
The above examples are merely for clearly illustrating the examples and are not intended to limit the embodiments; and are neither required nor exhaustive of all embodiments. And obvious variations or modifications of this technology may be resorted to while remaining within the scope of the technology.
Claims (10)
1. The utility model provides a hollow extension board of gas turbine combustion chamber air inlet which characterized in that: comprises a main plate body (7);
a preposed diffusion channel (1) is arranged at an air inlet of the combustion chamber of the gas turbine; an annular flow divider (5) is arranged in the front diffusion channel (1); the flow divider (5) divides the preposed diffusion channel (1) into an inner flow dividing channel (11) and an outer flow dividing channel (12);
the support plate body (7) is arranged in the inner diversion channel (11) and is connected with and supports the flow divider (5); the support plate main body (7) is wing-shaped, and a first bleed air channel (71) and a second bleed air channel (72) are formed in the support plate main body (7); the first bleed air channel (71) is used for guiding air at the front side of the support plate main body (7) to the outside of the front diffusion channel (1); the second air guide channel (72) can guide the air at the position, close to the wall surface of the flow divider (5), of the outer diversion channel (12) to the rear side of the support plate main body (7) and lead out towards the rear side of the inner diversion channel (11).
2. The gas turbine combustor air inlet hollow plate of claim 1, wherein: a plurality of first bleed air inlets (711) are formed in the wing head of the support plate main body (7), and each first bleed air inlet (711) is communicated with the first bleed air channel (71).
3. The gas turbine combustor air inlet hollow strip of claim 2, wherein: the first bleed air channel (71) is also communicated with a first bleed air outlet (712), the first bleed air outlet (712) is positioned outside the front diffusion channel (1), and air led out from the first bleed air outlet (712) can be used for cooling parts of the gas turbine.
4. The gas turbine combustor air inlet hollow plate of claim 3, wherein: the gas turbine combustion chamber comprises a combustion chamber casing (3), and the combustion chamber casing (3) comprises a diffuser outer ring (31) and a diffuser inner ring (32); the preposed diffusion channel (1) is formed between a diffuser outer ring (31) and a diffuser inner ring (32); the first bleed air outlet (712) is disposed inboard of a diffuser inner ring (32).
5. The gas turbine combustor air inlet hollow plate of claim 2, wherein: the first bleed air inlets (711) are arranged in a line in the width direction of the inner diversion passage (11).
6. The gas turbine combustor air inlet hollow plate of claim 1, wherein: the flow direction of the air led out from the second bleed air channel (72) is a horizontal flow direction, or the flow direction of the air led out from the second bleed air channel (72) and the air flow direction in the inner diversion channel (11) have the same trend.
7. The gas turbine combustor air inlet hollow plate of claim 6, wherein: the support plate main body (7) is internally provided with a plurality of second bleed air channels (72), a plurality of second bleed air outlets (722) are arranged at the head and the tail of the wing, and the second bleed air outlets (722) are correspondingly communicated with the second bleed air channels (72).
8. The gas turbine combustor air inlet hollow plate of claim 7, wherein: each second bleed air channel (72) is also correspondingly communicated with a second bleed air inlet (721), and the second bleed air inlets (721) are arranged on the wall surface of the outer ring side of the flow divider (5).
9. The gas turbine combustor air inlet hollow plate of claim 7, wherein: the second bleed air outlets (722) are arranged in a straight line along the width direction of the inner diversion channel (11); a plurality of second bleed air inlets (721) are arranged in the direction of the air flow in the outer distribution channel (12).
10. The gas turbine combustor air inlet hollow plate of claim 6, wherein: the second air-entraining channel (72) comprises a first section channel and a tail section channel which are communicated with each other; the first section of channel is inclined and not vertical to the direction of the airflow in the outer shunting channel (12); the tail section channel is horizontal, or the air flow direction at the outlet of the tail section channel has the same trend with the air flow direction in the inner diversion channel (11).
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