Gas turbine capable of reducing turbine stator blade cooling air pressure loss

Abstract

The invention discloses a gas turbine capable of reducing turbine stator blade cooling air pressure loss, which comprises a shell assembly, the interior of the shell assembly is rotatably connected with a gas turbine main shaft, and the outer side of the shell assembly is fixedly provided with a plurality of gas turbine combustion chambers; the shell assembly is sequentially composed of an air inlet cylinder, a pressure middle cylinder, a pressure exhaust cylinder, a turbine cylinder and an exhaust cylinder. A plurality of groups of turbine stator blades are fixed in the turbine cylinder; a turbine stationary blade cooling structure is arranged in the turbine stationary blade, the turbine stationary blade cooling structure comprises a stationary blade cooling main channel which is arranged in the turbine stationary blade and extends from the root of the turbine stationary blade to a blade tip, and a plurality of branch cooling channels communicated with the stationary blade cooling main channel are arranged in the turbine stationary blade. In the turbine stator blade of the gas turbine, after cooling air enters the stator blade cooling main channel from the root of the turbine stator blade, part of the cooling air can be shunted to enter the spiral branch cooling channels, and under the action of centrifugal force, the cooling air can be more evenly distributed to all areas of the turbine stator blade.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbines, and in particular to a gas turbine capable of reducing the pressure loss of cooling air for turbine stator blades. Background Art

[0002] In a gas turbine, turbine stator cooling air pressure loss refers to the pressure drop caused by flow resistance when the cooling air flows through the cooling channel of the turbine stator (guide vane). This loss is one of the parameters that need to be focused on in the design and operation of the gas turbine, and directly affects the efficiency, reliability and life of the unit. Cooling air is usually taken from the compressor outlet. Pressure loss will cause the available energy of this part of the air to decrease, and the compressor may need to consume more work to maintain the air supply pressure, thereby indirectly reducing the cycle efficiency of the unit. Uneven pressure loss may cause unstable cooling air flow, causing vibration or fatigue damage inside the blades. Excessive pressure loss may cause cooling system components (such as ducts and nozzles) to be subjected to greater stress, increasing the risk of leakage or failure. Summary of the Invention

[0003] The object of the present invention is to provide a gas turbine capable of reducing the pressure loss of cooling air for turbine stator blades. By optimizing the cooling structure thereof, the pressure loss of cooling air for turbine stator blades can be significantly reduced and the cooling effect can be improved.

[0004] To achieve the above object, the present invention provides the following technical solutions:

[0005] A gas turbine capable of reducing the pressure loss of cooling air for turbine stator blades comprises a casing assembly, wherein a gas turbine main shaft is rotatably connected inside the casing assembly, and a plurality of gas turbine combustion chambers are fixed outside the casing assembly;

[0006] The housing assembly is composed of the intake cylinder, the compression cylinder, the compression and exhaust cylinder, the turbine cylinder and the exhaust cylinder in sequence;

[0007] The output end of the gas turbine combustion chamber is connected to the turbine cylinder, and the input end of the gas turbine combustion chamber is connected to the compression and exhaust cylinder;

[0008] There are multiple sets of tail section compressor and static pressure blades fixed in the exhaust cylinder, and multiple sets of turbine stator blades fixed in the turbine cylinder;

[0009] The turbine stator blade has a turbine stator blade cooling structure inside, which includes a main stator blade cooling channel arranged inside the turbine stator blade and extending from the root to the blade tip, and a plurality of branch cooling channels connected to the main stator blade cooling channel inside the turbine stator blade;

[0010] The surface of the turbine stator blade is provided with a variable aperture air film hole structure, which includes an air film cooling hole arranged on the surface of the turbine stator blade and connected to the branch cooling channel, and a memory variable aperture mechanism is provided in the air film cooling hole.

[0011] Preferably, the cross-sectional area of ​​the branch cooling channel is 5-10% of the cross-sectional area of ​​the main cooling channel of the stator blade, and the sum of the cross-sectional areas of the branch cooling channels in each turbine stator blade is 50-60% of the cross-sectional area of ​​the main cooling channel of the stator blade.

[0012] Description: After the cooling air enters the main cooling channel of the turbine stator blade from the root of the turbine stator blade, as it flows inside the main cooling channel, part of the cooling air will be diverted into the spiral branch cooling channel. Due to the spiral structure of the branch cooling channel, the cooling air will generate a certain centrifugal force during the flow process, so that the cooling air can be more evenly distributed to various areas of the turbine stator blade.

[0013] Preferably, the helix angle of the branch cooling channel is 80-85°, and the helix radius of the branch cooling channel is 3-5 times its own diameter;

[0014] The angle between the spiral axis around which the spiral extension of the branch cooling channel is arranged and the main cooling channel for the stationary blade is 35 to 45 degrees.

[0015] Note: The spiral flow path prolongs the contact time between the cooling air and the turbine stator blades, improving the cooling efficiency.

[0016] Preferably, the branch cooling channel has a variable cross-section structure, the end of the branch cooling channel connected to the stationary blade cooling main channel is the end with a larger cross-section, and the cross-sectional area of ​​the end with a smaller cross-section of the branch cooling channel is 90% of the end with a larger cross-section.

[0017] Description: The variable cross-section design of the branch cooling channel gradually increases the flow velocity of the cooling air in the branch cooling channel, reducing the turbulence and eddy current caused by excessive flow velocity changes, thereby reducing flow resistance.

[0018] Preferably, the inner side walls of the main cooling channel and the branch cooling channel of the stator blade are coated with a thermal barrier coating, which is a yttria-stabilized zirconia coating.

[0019] Description: Thermal barrier coatings can significantly reduce the air flow resistance of the main cooling channels and branch cooling channels inside turbine stator blades.

[0020] Preferably, a plurality of turbine stator blade connecting rings coaxially arranged therewith are fixed to the inner side wall of the turbine cylinder, and each turbine stator blade in each group is fixedly connected to the inner side of the turbine stator blade connecting ring;

[0021] The interior of the turbine stator blade connecting ring is a hollow structure, and the interior of the turbine stator blade connecting ring is connected to the stator blade cooling main channel;

[0022] The interior of the turbine stator blade connecting ring sleeve is connected to the interior of the compression and exhaust cylinder through the turbine stator blade cooling channel.

[0023] Description: The cavity structure inside the turbine stator blade connecting ring can evenly distribute the cooling air to the main stator cooling channel of each turbine stator blade.

[0024] Preferably, the memory variable aperture mechanism comprises a memory variable aperture support ring fixed in the film cooling hole, wherein the memory variable aperture support ring has a plurality of aperture constraint memory springs extending along its axial direction fixed therein;

[0025] The aperture-constrained memory spring consists of a spring fixed end and a spring deformation end. The spring fixed end is fixedly connected to the inner side wall of the memory variable aperture support ring, and the spring deformation end can deform and warp along the radial direction of the memory variable aperture support ring.

[0026] Description: When the temperature of a certain area on the surface of the turbine stator blade rises or the cooling air pressure changes, the aperture constraint memory spring of the memory alloy material can produce corresponding deformation according to the temperature change, thereby adjusting the aperture size of the film cooling hole;

[0027] When the surface of the turbine stator blade is in a condition with a high thermal load, such as the leading edge and trailing edge of the turbine stator blade, the warping deformation of the aperture-constrained memory spring will cause the deformed ends of the spring to move away from each other, thereby increasing the aperture of the film cooling hole, increasing the outflow of cooling air, forming a thicker air film, and improving the cooling effect; when the surface of the turbine stator blade is in a condition with a low thermal load, the warping deformation of the aperture-constrained memory spring will cause the deformed ends of the spring to rebound close to each other, thereby reducing the aperture of the film cooling hole, avoiding waste of cooling air.

[0028] Preferably, one end of the film cooling hole on the surface of the turbine stator blade is a trumpet-shaped expanding section, and the length of the expanding section is consistent with the diameter of the film cooling hole.

[0029] Description: The gradually expanding hole structure of the film cooling hole enables the cooling air to transition smoothly when it flows out, reduces the resistance of the cooling air when it flows out, and reduces the pressure loss caused by sudden expansion.

[0030] Preferably, a plurality of groups of first-stage compression and static pressure blades are fixed in the compression cylinder;

[0031] The parts of the gas turbine main shaft in the compression cylinder, the compression exhaust cylinder and the turbine cylinder are fixed with multiple groups of first-stage compression aerodynamic blades, tail-stage compression aerodynamic blades and turbine blades in sequence.

[0032] Preferably, a multi-stage diversion mechanism is provided inside the side wall of the housing assembly located at the pressure-discharging cylinder, and the multi-stage diversion mechanism includes a diversion accommodating ring shell fixed inside the inner side wall of the pressure-discharging cylinder and coaxially arranged therewith, the interior of the diversion accommodating ring shell is a hollow structure, a diversion control flow pipe is fixed inside the diversion accommodating ring shell, one end of the diversion control flow pipe is connected to the interior of the pressure-discharging cylinder through the diversion control input pipe, and a diversion control discharge pipe is fixed to the other end of the diversion control flow pipe;

[0033] The flow splitting control discharge pipe is connected to the interior of the turbine stator blade connecting ring sleeve;

[0034] The diversion control flow pipe is provided with a diversion control valve;

[0035] There are multiple flow-dividing accommodating annular casings arranged along the axis of the compression and discharge cylinder. The flow-dividing control input pipes in each flow-dividing accommodating annular casing are connected to the compressor and static pressure blades of each level of the tail section inside the compression and discharge cylinder in a one-to-one correspondence.

[0036] Description: The multi-stage diversion mechanism can accurately perform flow control and pressure regulation, adjust the supply of cooling air, and allow the cooling air to enter the stationary blade cooling channel at a more appropriate pressure and flow rate, avoiding additional pressure loss caused by excessively high or low pressure.

[0037] Compared with the prior art, the beneficial effects of the present invention are embodied in the following aspects:

[0038] 1. The present invention has a reasonable structural design. Inside the gas turbine stator blade, cooling air enters the main cooling channel of the stator blade from the root of the turbine stator blade. As it flows inside the main cooling channel, part of the cooling air is diverted into the spiral branch cooling channel. Due to the spiral structure of the branch cooling channel, the cooling air generates a certain centrifugal force during the flow process, so that the cooling air can be more evenly distributed to various areas of the turbine stator blade.

[0039] 2. In the technical solution of the present invention, the variable cross-section design of the branch cooling channel gradually increases the flow velocity of the cooling air in the branch cooling channel, reducing the turbulence and eddy currents caused by excessive flow velocity changes, thereby reducing flow resistance. The spiral flow path prolongs the contact time between the cooling air and the turbine stator blades, improving cooling efficiency. This structure can guide the cooling air to flow in an orderly manner, avoiding the flow dead zones and pressure unevenness that are prone to occur in traditional straight channels or simple curved channels, and effectively reducing the pressure loss of the cooling air inside the stator blades.

[0040] 3. In the technical solution of the present invention, when the temperature of a certain area on the surface of the turbine stator blade increases or the cooling air pressure changes, the aperture-constrained memory spring of the memory alloy material can produce corresponding deformation according to the temperature change, thereby adjusting the aperture size of the film cooling hole;

[0041] 4. In the technical solution of the present invention, when the turbine stator blade is in a high heat load condition, such as the leading and trailing edges of the turbine stator blade, the warping deformation of the aperture-constrained memory spring will increase the aperture of the film cooling hole, increase the outflow of cooling air, form a thicker air film on the surface of the turbine stator blade, and improve the cooling effect; while in a low heat load condition, due to temperature changes, the aperture of the film cooling hole will be appropriately reduced to avoid waste of cooling air. Through this adaptive variable aperture control, the cooling air can be accurately distributed according to the actual needs of various parts of the turbine stator blade, which not only improves the cooling efficiency but also reduces the overall pressure loss;

[0042] 5. In the technical solution of the present invention, the gradually expanding hole structure of the film cooling hole enables the cooling air to transition smoothly when it flows out, reduces the resistance when the cooling air flows out, and reduces the pressure loss caused by sudden expansion;

[0043] 6. In the technical solution of the present invention, the outside air first enters the compression cylinder and the exhaust cylinder through the intake cylinder to undergo multi-stage compression. During each stage of compression, the pressure and temperature of the air gradually increase. The multi-stage diversion mechanism can divert a portion of the compressed air as cooling air according to the operating conditions of the gas turbine and the actual requirements of the stator blade cooling. The remaining air continues to enter the next stage of compression as the mainstream air, and finally enters the gas turbine combustion chamber to participate in combustion.

[0044] 7. In the technical solution of the present invention, the multi-stage diversion mechanism can accurately perform flow control and pressure regulation, adjust the supply of cooling air, and allow the cooling air to enter the stator cooling channel at a more appropriate pressure and flow rate, avoiding additional pressure loss caused by excessively high or low pressure. The multi-stage diversion design enables the cooling air and mainstream air to be optimized separately, thereby improving the efficiency of the air intake system and providing a stable and reliable source of cooling air for the stator cooling system. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 It is a front view of the present invention;

[0046] Figure 2 It is a structural schematic diagram of the turbine stator blade cooling structure of the present invention;

[0047] Figure 3 yes Figure 2 A top view of

[0048] Figure 4 2. It is a top view of the variable aperture air film hole structure of the present invention;

[0049] Figure 5 This is a schematic structural diagram of the memory variable aperture support ring of the present invention;

[0050] Figure 6 It is a structural schematic diagram of the multi-stage diversion mechanism of the present invention.

[0051] In the figure, 10-housing assembly, 11-intake cylinder, 12-compression cylinder, 121-first stage compression and static pressure blades, 13-compression and exhaust cylinder, 131-tail stage compression and static pressure blades, 14-turbine cylinder, 141-turbine static blades, 142-turbine static blade connecting ring sleeve, 15-exhaust cylinder, 16-gas turbine main shaft, 161-first stage compression and dynamic pressure blades, 162-tail stage compression and dynamic pressure blades, 163-turbine dynamic blades, 17-gas turbine combustion chamber, 20-turbine static blade cooling structure, 21 - Main channel for stationary blade cooling, 22 - branch cooling channel, 30 - variable aperture air film hole structure, 31 - air film cooling hole, 32 - memory variable aperture mechanism, 321 - memory variable aperture support ring, 322 - aperture constraint memory spring, 3221 - spring fixed end, 3222 - spring deformation end, 40 - multi-stage diversion mechanism, 41 - diversion containing ring shell, 421 - diversion control flow pipe, 422 - diversion control input pipe, 423 - diversion control discharge pipe, 4210 - diversion control valve. DETAILED DESCRIPTION

[0052] The following combination Figures 1 to 6 The present invention is described in detail. For the convenience of description, the directions mentioned below are defined as follows: the up, down, left, right, front and back directions mentioned below are consistent with the up, down, left, right, front and back directions of the projection relationship of each main view or structural schematic diagram itself.

[0053] Example 1:

[0054] A gas turbine capable of reducing the pressure loss of cooling air for turbine stator blades, such as Figure 1 As shown,

[0055] It includes a housing assembly 10, a gas turbine main shaft 16 is rotatably connected to the inside of the housing assembly 10, and a plurality of gas turbine combustion chambers 17 are fixed to the outside of the housing assembly 10;

[0056] The housing assembly 10 is composed of an intake cylinder 11, a compression cylinder 12, a compression and exhaust cylinder 13, a turbine cylinder 14 and an exhaust cylinder 15 in sequence;

[0057] The output end of the gas turbine combustion chamber 17 is connected to the turbine cylinder 14, and the input end of the gas turbine combustion chamber 17 is connected to the compression and exhaust cylinder 13;

[0058] A plurality of groups of first-stage compression and static pressure blades 121 are fixed in the compression cylinder 12, a plurality of groups of tail-stage compression and static pressure blades 131 are fixed in the compression cylinder 13, and a plurality of groups of turbine stator blades 141 are fixed in the turbine cylinder 14;

[0059] The first-stage compressor and static pressure blades 121, the tail-stage compressor and static pressure blades 131, and the turbine stator blades 141 are all made by metal 3D printing technology;

[0060] The plurality of first-stage compressor and static pressure blades 121 arranged circumferentially around the compression center cylinder 12 form a group, and the plurality of groups of first-stage compressor and static pressure blades 121 are arranged along the axis direction of the compression center cylinder 12;

[0061] The plurality of tail section compressor and static pressure blades 131 arranged circumferentially around the compression and discharge cylinder 13 form a group, and the plurality of groups of tail section compressor and static pressure blades 131 are arranged along the axis direction of the compression and discharge cylinder 13;

[0062] The plurality of turbine stator blades 141 arranged circumferentially around the turbine cylinder 14 form a group, and the plurality of groups of turbine stator blades 141 are arranged along the axis of the turbine cylinder 14 ;

[0063] The gas turbine main shaft 16 is located in the compression cylinder 12, the exhaust cylinder 13 and the turbine cylinder 14. A plurality of groups of first-stage compression aerodynamic blades 161, tail-stage compression aerodynamic blades 162 and turbine rotor blades 163 are fixed in sequence.

[0064] The plurality of first-stage aerodynamic pressure blades 161 arranged circumferentially around the gas turbine main shaft 16 constitute a group, and the plurality of groups of first-stage aerodynamic pressure blades 161 are arranged along the axis of the gas turbine main shaft 16;

[0065] The plurality of tail section aerodynamic pressure blades 162 arranged circumferentially around the gas turbine main shaft 16 form a group, and the plurality of groups of tail section aerodynamic pressure blades 162 are arranged along the axis direction of the gas turbine main shaft 16;

[0066] The plurality of turbine blades 163 arranged circumferentially around the gas turbine main shaft 16 constitute a group, and the plurality of groups of turbine blades 163 are arranged along the axial direction of the gas turbine main shaft 16;

[0067] A plurality of groups of first-stage compressor static pressure blades 121 and a plurality of groups of first-stage compressor dynamic pressure blades 161 are arranged alternately along the axis of the gas turbine main shaft 16;

[0068] A plurality of groups of tail section compressor static pressure blades 131 and a plurality of groups of tail section compressor dynamic pressure blades 162 are arranged alternately along the axis direction of the gas turbine main shaft 16;

[0069] The plurality of groups of turbine stationary blades 141 and the plurality of groups of turbine moving blades 163 are arranged alternately along the axis of the gas turbine main shaft 16;

[0070] An adjacent set of first-stage compression dynamic pressure blades 161 and a set of first-stage compression static pressure blades 121 form a stage;

[0071] An adjacent set of tail section aerodynamic pressure blades 162 and a set of tail section static pressure blades 131 form a stage;

[0072] An adjacent set of turbine stationary blades 141 and a set of turbine moving blades 163 form a stage;

[0073] The gas turbine is a gas turbine having a seventeen-stage compressor;

[0074] like Figure 2 As shown, the turbine stator blade 141 has a turbine stator blade cooling structure 20 inside. The turbine stator blade cooling structure 20 includes a stator blade cooling main channel 21 provided inside the turbine stator blade 141 and extending from the root to the blade tip. The turbine stator blade 141 has a plurality of branch cooling channels 22 connected to the stator blade cooling main channel 21.

[0075] A variable aperture film hole structure 30 is provided on the surface of the turbine stator blade 131. The variable aperture film hole structure 30 includes an air film cooling hole 31 provided on the surface of the turbine stator blade 141 and connected to the branch cooling channel 22. A memory variable aperture mechanism 32 is provided in the air film cooling hole 31.

[0076] like Figure 2 As shown, a plurality of turbine stator blade connecting rings 142 arranged coaxially therewith are fixed to the inner wall of the turbine cylinder 14, and each turbine stator blade 141 in each group is fixedly connected to the inner side of the turbine stator blade connecting ring 142;

[0077] The interior of the turbine stator blade connecting ring sleeve 142 is a hollow structure, and the interior of the turbine stator blade connecting ring sleeve 142 is connected to the stator blade cooling main channel 21;

[0078] The interior of the turbine stator blade connecting ring sleeve 142 is connected to the interior of the compression and exhaust cylinder 13 through the turbine stator blade cooling channel.

[0079] Example 2:

[0080] Based on Example 1, the cross-sectional area of ​​the branch cooling channel 22 is 10% of the cross-sectional area of ​​the main stationary blade cooling channel 21, and the total cross-sectional area of ​​each branch cooling channel 22 in each turbine stationary blade 141 is 60% of the cross-sectional area of ​​the main stationary blade cooling channel 21.

[0081] The helix angle of the branch cooling channel 22 is 85°, and the helix radius of the branch cooling channel 22 is 5 times its own diameter;

[0082] The included angle between the spiral axis around which the branch cooling channel 22 spirally extends and the main vane cooling channel 21 is 45°.

[0083] Example 3:

[0084] On the basis of Example 2, Figure 2As shown, the branch cooling channel 22 has a variable cross-section structure. The end of the branch cooling channel 22 connected to the stationary blade cooling main channel 21 is the end with a larger cross-section, and the cross-sectional area of ​​the end with a smaller cross-section of the branch cooling channel 22 is 90% of the end with a larger cross-section.

[0085] Example 4:

[0086] Based on Example 3, the inner walls of the main cooling channel 21 and the branch cooling channel 22 of the stator blade are coated with a thermal barrier coating. The thermal barrier coating used in this application is the existing technology, which is the existing technology yttria-stabilized zirconia coating doped with 0.2% Nd2O3, 0.1% Gd2O3, and 0.1% Sc2O3 by mass.

[0087] Example 5:

[0088] On the basis of Example 4, Figure 4 As shown, the memory variable aperture mechanism 32 includes a memory variable aperture support ring 321 fixed in the air film cooling hole 31, and the memory variable aperture support ring 321 has multiple aperture constraint memory springs 322 extending along its axial direction fixed therein.

[0089] like Figure 4 As shown, the aperture constraint memory spring 322 consists of a spring fixed end 3221 and a spring deforming end 3222. The spring fixed end 3221 is fixedly connected to the inner side wall of the memory variable aperture support ring 321, and the spring deforming end 3222 can deform and warp along the radial direction of the memory variable aperture support ring 321.

[0090] At room temperature, the deformed ends 3222 of each spring clip are close to each other, causing the channel formed by multiple aperture-constrained memory spring clips 322 to narrow. When the temperature rises, the deformed ends 3222 of each aperture-constrained memory spring clip 322 will deform and warp, and the deformed ends 3222 of each spring clip will move away from each other.

[0091] like Figure 4 As shown, the film cooling hole 31 is located on the surface of the turbine stator blade 141 and has a trumpet-shaped expanding section at one end, and the length of the expanding section is consistent with the diameter of the film cooling hole 31 .

[0092] Example 6:

[0093] On the basis of Example 5, Figure 1 As shown, a multi-stage flow dividing mechanism 40 is provided inside the side wall of the housing assembly 10 at the pressure discharge cylinder 13. Figure 6As shown, the multi-stage diversion mechanism 40 includes a diversion accommodating annular casing 41 fixed to the inner side wall of the compression and exhaust cylinder 13 and arranged coaxially therewith. The interior of the diversion accommodating annular casing 41 is a hollow structure. A diversion control flow pipe 421 is fixed in the diversion accommodating annular casing 41. One end of the diversion control flow pipe 421 is connected to the interior of the compression and exhaust cylinder 13 through a diversion control input pipe 422. A diversion control discharge pipe 423 is fixed to the other end of the diversion control flow pipe 421. The diversion control discharge pipe 423 is connected to the interior of the turbine stator blade connecting ring sleeve 142. A diversion control valve 4210 is provided on the diversion control flow pipe 421.

[0094] There are multiple diversion accommodating annular casings 41 arranged along the axial direction of the compression and discharge cylinder 13, and the diversion control input pipes 422 in each diversion accommodating annular casing 41 are connected to the tail stage compressor and static pressure blades 131 of each level inside the compression and discharge cylinder 13 respectively.

[0095] Example 7:

[0096] The difference from Example 6 is that the cross-sectional area of ​​the branch cooling channel 22 is 7% of the cross-sectional area of ​​the main vane cooling channel 21, and the total cross-sectional area of ​​each branch cooling channel 22 in each turbine vane 141 is 55% of the cross-sectional area of ​​the main vane cooling channel 21.

[0097] The helix angle of the branch cooling channel 22 is 82°, and the helix radius of the branch cooling channel 22 is 4 times its own diameter;

[0098] The included angle between the spiral axis around which the branch cooling channel 22 spirally extends and the main vane cooling channel 21 is 40°.

[0099] Example 8:

[0100] The difference from Example 6 is that the cross-sectional area of ​​the branch cooling channel 22 is 5% of the cross-sectional area of ​​the main vane cooling channel 21, and the sum of the cross-sectional areas of the branch cooling channels 22 in each turbine vane 141 is 50% of the cross-sectional area of ​​the main vane cooling channel 21.

[0101] The helix angle of the branch cooling channel 22 is 80°, and the helix radius of the branch cooling channel 22 is 3 times its own diameter;

[0102] The included angle between the spiral axis around which the branch cooling channel 22 spirally extends and the main vane cooling channel 21 is 35°.

[0103] In actual application, the present invention firstly passes the air intake cylinder into the compression cylinder and the exhaust cylinder to carry out multi-stage compression of the air. During each stage of compression, the pressure and temperature of the air gradually increase. The multi-stage diversion mechanism can divert a portion of the compressed air as cooling air according to the operating conditions of the gas turbine and the actual requirements of the stator blade cooling. The remaining air continues to enter the next stage of compression as the mainstream air, and finally enters the gas turbine combustion chamber to participate in combustion.

[0104] The cooling air of the compression and discharge cylinder section 13 enters the shunt control flow pipe 421 through the shunt control input pipe 422, and the flow rate through the shunt control flow pipe 421 can be controlled by adjusting the shunt control valve 4210. The cooling air in the shunt control flow pipe 421 is then discharged through the shunt control discharge pipe 423, and the cooling air discharged through the shunt control discharge pipe 423 then enters the interior of the turbine stator blade connecting ring sleeve 142. The cooling air inside the turbine stator blade connecting ring sleeve 142 then enters the stator blade cooling main channel 21 from the root of each turbine stator blade 141. As the cooling air flows inside the stator blade cooling main channel 21, part of the cooling air will be diverted into the spiral branch cooling channel 22. Due to the spiral structure of the branch cooling channel 22, the cooling air will generate a certain centrifugal force during the flow process, so that the cooling air can be more evenly distributed to various areas of the turbine stator blade 141, so that the cooling air and the turbine stator blade 141 can fully exchange heat, thereby realizing the cooling of the turbine stator blade 141.

[0105] During the flow of the branch cooling channels 22, the cooling air is discharged from each film cooling hole 31. The film cooling hole 31 can adjust its aperture size in real time according to temperature changes. When the surface of the turbine stator blade 141 is under a high heat load, such as the leading and trailing edges of the turbine stator blade, the warping deformation of the aperture constraint memory spring 322 causes the deformed ends 3222 of the spring to move away from each other, thereby increasing the aperture of the film cooling hole 31, increasing the outflow of cooling air, forming a thicker air film, and improving the cooling effect.

[0106] When the surface of the turbine stator blade 141 is in a low heat load condition, the warping deformation of the aperture-constrained memory spring 322 causes the deformed ends 3222 of the spring to move closer to each other and rebound, thereby reducing the aperture of the film cooling hole 31 and avoiding waste of cooling air.

[0107] Through this adaptive variable aperture control of the air film cooling hole 31, the cooling air can be accurately distributed according to the actual needs of each part of the turbine stator blade 141, which not only improves the cooling efficiency but also reduces the overall pressure loss.

Claims

1. A gas turbine capable of reducing turbine vane cooling air pressure loss, comprising a casing assembly (10), a gas turbine main shaft (16) being rotatably connected to the inside of the casing assembly (10), and a plurality of gas turbine combustion chambers (17) being fixed to the outside of the casing assembly (10); The housing assembly (10) is composed of an intake cylinder (11), a compression cylinder (12), a compression and exhaust cylinder (13), a turbine cylinder (14) and an exhaust cylinder (15) in sequence; The output end of the gas turbine combustion chamber (17) is connected to the turbine cylinder (14), and the input end of the gas turbine combustion chamber (17) is connected to the compression and exhaust cylinder (13); A plurality of groups of tail section compressor and static pressure blades (131) are fixed in the compression and exhaust cylinder (13), and a plurality of groups of turbine stationary blades (141) are fixed in the turbine cylinder (14); It is characterized by: The turbine stator blade (141) has a turbine stator blade cooling structure (20) therein, the turbine stator blade cooling structure (20) comprising a stator blade cooling main channel (21) arranged inside the turbine stator blade (141) and extending from the root to the blade tip, and the turbine stator blade (141) has a plurality of branch cooling channels (22) therein that are in communication with the stator blade cooling main channel (21); The surface of the turbine stator blade (131) is provided with a variable aperture air film hole structure (30), and the variable aperture air film hole structure (30) includes an air film cooling hole (31) provided on the surface of the turbine stator blade (141) and connected to the branch cooling channel (22), and a memory variable aperture mechanism (32) is provided in the air film cooling hole (31).

2. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The cross-sectional area of ​​the branch cooling channel (22) is 5 to 10% of the cross-sectional area of ​​the main stationary blade cooling channel (21), and the sum of the cross-sectional areas of the branch cooling channels (22) in each turbine stationary blade (141) is 50 to 60% of the cross-sectional area of ​​the main stationary blade cooling channel (21).

3. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The helical angle of the branch cooling channel (22) is 80 to 85 degrees, and the helical radius of the branch cooling channel (22) is 3 to 5 times its own diameter; The angle between the spiral axis around which the branch cooling channel (22) spirally extends and the main stationary blade cooling channel (21) is 35 to 45 degrees.

4. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The branch cooling channel (22) is a variable cross-section structure, the end of the branch cooling channel (22) connected to the stationary blade cooling main channel (21) is the end with a larger cross-section, and the cross-sectional area of ​​the end with a smaller cross-section of the branch cooling channel (22) is 90% of the cross-sectional area of ​​the end with a larger cross-section.

5. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The inner side walls of the main cooling channel (21) of the stationary blade and the branch cooling channel (22) are coated with a thermal barrier coating, which is a yttria-stabilized zirconia coating.

6. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: A plurality of turbine stator blade connecting rings (142) coaxially arranged therewith are fixed to the inner side wall of the turbine cylinder (14), and each of the turbine stator blades (141) in each group is fixedly connected to the inner side of the turbine stator blade connecting ring (142); The interior of the turbine stator blade connecting ring sleeve (142) is a hollow structure, and the interior of the turbine stator blade connecting ring sleeve (142) is connected to the stator blade cooling main channel (21); The interior of the turbine stator blade connecting ring sleeve (142) is connected to the interior of the compression and discharge cylinder (13) through a turbine stator blade cooling channel.

7. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The memory variable aperture mechanism (32) comprises a memory variable aperture support ring (321) fixed in the air film cooling hole (31), wherein a plurality of aperture constraint memory springs (322) extending along the axial direction of the memory variable aperture support ring (321) are fixed in the memory variable aperture support ring (321); The aperture-constraining memory spring clip (322) consists of a spring clip fixed end (3221) and a spring clip deforming end (3222); the spring clip fixed end (3221) is fixedly connected to the inner side wall of the memory variable aperture support ring (321); and the spring clip deforming end (3222) is capable of deforming and warping along the radial direction of the memory variable aperture support ring (321).

8. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 7, characterized in that: One end of the film cooling hole (31) on the surface of the turbine stator blade (141) is a trumpet-shaped expanding section, and the length of the expanding section is consistent with the diameter of the film cooling hole (31).

9. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: A plurality of groups of first-stage air compression and static pressure blades (121) are fixed in the compression cylinder (12); The gas turbine main shaft (16) is located in the compression cylinder (12), the compression exhaust cylinder (13) and the turbine cylinder (14), and a plurality of groups of first-stage compression aerodynamic blades (161), tail-stage compression aerodynamic blades (162) and turbine blades (163) are fixed in sequence.

10. A gas turbine capable of reducing turbine vane cooling air pressure loss according to claim 1, characterized in that: The housing assembly (10) is provided with a multi-stage diversion mechanism (40) inside the side wall of the pressure-discharging cylinder (13). The multi-stage diversion mechanism (40) includes a diversion accommodating ring shell (41) fixed inside the inner side wall of the pressure-discharging cylinder (13) and arranged coaxially therewith. The interior of the diversion accommodating ring shell (41) is a hollow structure. A diversion control flow pipe (421) is fixed inside the diversion accommodating ring shell (41). One end of the diversion control flow pipe (421) is connected to the interior of the pressure-discharging cylinder (13) through a diversion control input pipe (422). The other end of the diversion control flow pipe (421) is fixed with a diversion control discharge pipe (423). The flow splitting control discharge pipe (423) is in communication with the interior of the turbine stator blade connecting ring sleeve (142); The diversion control flow pipe (421) is provided with a diversion control valve (4210); A plurality of the flow-dividing accommodating annular casings (41) are arranged along the axial direction of the compression-discharging cylinder (13), and the flow-dividing control input pipes (422) in each of the flow-dividing accommodating annular casings (41) are connected to the respective stages of the tail section compression and static pressure blades (131) inside the compression-discharging cylinder (13) in a one-to-one correspondence.

Images