Gas turbine disc cavity sealing cooling device

Abstract

The present invention relates to the technical field of gas turbines, and in particular to a gas turbine turbine disc cavity sealing cooling device, comprising a first guide assembly disposed between a first-stage turbine impeller and a second-stage baffle, and a second guide assembly disposed between a second-stage turbine impeller and the second-stage baffle, wherein the second-stage baffle is a symmetrical hollow structure; and further comprising a sealing assembly disposed on the inner side of the second-stage baffle. The present invention has the beneficial effect that the guide assembly can produce a suction effect on the cooling air entering the inner cavity of the impeller under the action of the centrifugal force of the rotating impeller, and a contraction flow channel is formed between the guide vanes, thereby increasing the cooling air flow rate, enhancing the heat exchange effect between the cooling air and the impeller, and improving the cooling efficiency. Furthermore, the guide vanes and the impeller are connected by bolts, and by adjusting the angle of the guide vanes and changing the minimum cross-sectional area of ​​the contraction flow channel formed by the guide vanes, the cooling air flow rate entering the inner cavity of each impeller can be adjusted.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbines, in particular to a gas turbine disc cavity sealing cooling device. Background Art

[0002] As one of the core components of a gas turbine, the turbine plays a vital role in the entire gas turbine system. The disc seal structure is a crucial and complex part of the gas turbine system, and its stability and reliability are directly related to the performance and efficiency of the gas turbine.

[0003] In the prior art, only the turbine seal or the turbine disc cavity structure is often considered separately, without comprehensively considering the cooling effect of the turbine disc cavity structure and the sealing effect of the seal structure. Summary of the Invention

[0004] In view of the above problems or problems existing in the prior art, the present invention is proposed.

[0005] To solve the above technical problems, the present invention provides the following technical solutions: a gas turbine turbine disc cavity sealing and cooling device, comprising a first guide assembly disposed between a first-stage impeller disc and a second-stage baffle of the turbine, and a second guide assembly disposed between a second-stage impeller disc and a second-stage baffle of the turbine, wherein the second-stage baffle is a symmetrical hollow structure;

[0006] Also included is a sealing assembly disposed inside the second-stage partition.

[0007] As a preferred embodiment of the gas turbine disk cavity sealing and cooling device of the present invention, the first guide assembly includes a first guide vane group arranged in a circumferential array along the first-stage turbine impeller disk, the first guide vane group being contracted in the radial direction, wherein the first guide vane group is composed of a plurality of first guide vanes having the same shape and structure;

[0008] The air inlet of each turbine first-stage moving blade corresponds to two first guide vanes.

[0009] As a preferred solution of the gas turbine turbine disc cavity sealing cooling device of the present invention, the first guide vane includes a first air inlet end and a first air outlet end, and the thickness of the first air inlet end is greater than the thickness of the first air outlet end, wherein the first air inlet end is the end away from the air inlet of the first-stage moving blade of the turbine.

[0010] As a preferred embodiment of the gas turbine disk cavity sealing and cooling device of the present invention, the second guide assembly includes a second guide vane group arranged in a circumferential array along the second-stage impeller disk, the second guide vane group being contracted in the radial direction, wherein the second guide vane group is composed of a plurality of second guide vanes having the same shape and structure;

[0011] The air inlet of each turbine second-stage moving blade corresponds to two second guide vanes.

[0012] As a preferred solution of the gas turbine turbine disc cavity sealing cooling device of the present invention, the second guide vane includes a second air inlet end and a second air outlet end, and the thickness of the second air inlet end is greater than the thickness of the second air outlet end, wherein the second air inlet end is the end away from the air inlet of the second stage moving blade of the turbine.

[0013] As a preferred solution of the gas turbine turbine disc cavity sealing cooling device of the present invention, the second-stage partition includes a base, a group of protruding wings symmetrically arranged on the base, and a first grate tooth arranged on the outer surface of the upper end of the protruding wing.

[0014] As a preferred solution of the gas turbine turbine disc cavity sealing cooling device of the present invention, a group of mounting grooves are formed at the connection between a group of protruding wings and the base, and the two mounting grooves are respectively adapted to the bosses of the corresponding turbine first-stage impeller disc or the turbine second-stage impeller disc.

[0015] As a preferred embodiment of the gas turbine turbine disc cavity sealing and cooling device of the present invention, the sealing assembly includes a second-stage sealing ring and a second-stage sealing ring cooling air hole;

[0016] The second-stage sealing ring includes a second groove arranged opposite to the top end of the extended wing, and a second grate tooth arranged at the second groove, wherein the first grate tooth and the second grate tooth are arranged opposite to each other.

[0017] As a preferred solution of the gas turbine turbine disc cavity sealing and cooling device of the present invention, the top end of the second-stage sealing ring is connected to the second-stage turbine stator blade through a hanging ear and a clamping groove.

[0018] As a preferred solution of the gas turbine turbine disc cavity sealing and cooling device of the present invention, the inner cavity cross-section of the second-stage sealing ring is in the shape of an "inverted soil".

[0019] Beneficial effects of the present invention: The present invention provides a guide assembly, which can produce a suction effect on the cooling air entering the inner cavity of the moving blade under the action of the centrifugal force of the rotating wheel, and form a contraction flow channel between the guide vanes, thereby increasing the cooling air flow rate, strengthening the heat exchange effect between the cooling air and the moving blades, and improving the cooling efficiency. Furthermore, the guide vanes and the wheel are connected by bolts. By adjusting the angle of the guide vanes and changing the minimum cross-sectional area of ​​the contraction flow channel formed by the guide vanes, the cooling air flow entering the inner cavity of each moving blade can be adjusted. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] To more clearly illustrate the technical solutions of the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. Those skilled in the art can also derive other drawings based on these drawings without inventive effort. Among them:

[0021] Figure 1 This is a schematic cross-sectional view of a gas turbine disc cavity seal cooling device.

[0022] Figure 2 Schematic diagram of the cooling air flow of the gas turbine turbine disc cavity seal cooling device.

[0023] Figure 3 Schematic diagram of the sealing assembly of the gas turbine turbine disc cavity sealing cooling device.

[0024] Figure 4 Schematic diagram of the three-dimensional structure of the roulette wheel.

[0025] Figure 5 Schematic diagram of cooling air flow for part of the wheel.

[0026] Figure 6 Schematic diagrams of the guide vane in two dimensions and three dimensions.

[0027] Figure 7 Schematic diagram of the connection between the guide vane and the wheel.

[0028] Figure 8 Schematic diagram of the countersunk hole structure opened on the wheel guide plate.

[0029] Figure 9 It is a structural schematic diagram of the second-stage partition and the second-stage sealing ring.

[0030] Figure 10 Schematic diagram of the connection between the wheel and the second-stage partition. DETAILED DESCRIPTION

[0031] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the specific embodiments of the present invention are described in detail below with reference to the accompanying drawings.

[0032] In the following description, many specific details are set forth to facilitate a full understanding of the present invention. However, the present invention may also be implemented in other ways different from those described herein. Those skilled in the art may make similar generalizations without violating the connotation of the present invention. Therefore, the present invention is not limited to the specific embodiments disclosed below.

[0033] Secondly, the term "one embodiment" or "embodiment" herein refers to a specific feature, structure, or characteristic that may be included in at least one implementation of the present invention. The phrase "in one embodiment" appearing in various places throughout this specification does not necessarily refer to the same embodiment, nor does it refer to a separate or selective embodiment that is mutually exclusive of other embodiments.

[0034] Example

[0035] Reference Figures 1 to 10 , is an embodiment of the present invention, which provides a gas turbine turbine disc cavity sealing cooling device, comprising a first guide assembly 200 disposed between a first-stage turbine impeller disc 101 and a second-stage baffle 102, and a second guide assembly 300 disposed between a second-stage turbine impeller disc 103 and the second-stage baffle 102, wherein the second-stage baffle 102 is a symmetrical hollow structure;

[0036] Also included is a sealing assembly 400 disposed inside the second-stage diaphragm 102 .

[0037] It should be noted that the turbine first-stage impeller disc 101 and the turbine second-stage impeller disc 103 form a cooling air contraction flow channel under the arrangement of the first guide component 200, the second guide component 300 and the second-stage partition 102, which can accelerate the flow rate of the cooling air when entering the inner cavity of each impeller, thereby enhancing the convection and impact cooling effect inside the impeller.

[0038] Preferably, in order to reduce the weight of the wheel, the second-stage partition 102 is designed as a hollow structure to ensure that the outer rotor cooling air and the inner stationary blade inner cavity outlet cooling air fully cool it to prevent excessive temperature from causing ablation and high-temperature thermal stress.

[0039] Furthermore, the first guide assembly 200 includes a first guide vane group 201 circumferentially arranged around the first-stage turbine impeller disk 101. The first guide vane group 201 is contracted in the radial direction. The first guide vane group 201 is composed of a plurality of first guide vanes 201a having the same shape and structure. The first guide vanes 201a function as fins, thereby enhancing the heat exchange effect between the rotor cooling air and the impeller disk.

[0040] The air inlet of each turbine first-stage rotor blade 101a corresponds to two first guide vanes 201a.

[0041] It should be noted that each first guide vane 201a has two adjacent first guide vanes 201a. The two first guide vanes 201a acting on the air inlet of the same turbine first-stage moving blade 101a with one adjacent first guide vane 201a are called a grouped guide vane, and together with another adjacent first guide vane 201a are called an adjacent guide vane group.

[0042] Furthermore, the first guide vane 201a includes a first air inlet end 201a-1 and a first air outlet end 201a-2, and the thickness of the first air inlet end 201a-1 is greater than the thickness of the first air outlet end 201a-2, wherein the first air inlet end 201a-1 is the end away from the air inlet of the first-stage turbine blade 101a, and the air inlet ends of adjacent guide vane groups are in contact, so that the cooling air can only enter the internal cavity of each blade through the contraction channel formed by the first guide vane, thereby cooling the blade.

[0043] It should be noted that each first air inlet end 201a-1 and the first-stage impeller disk 101 of the turbine are provided with corresponding countersunk holes, and the two are connected by bolts. The thickness of the first air inlet end 201a-1 is greater than the thickness of the first air outlet end 201a-2, so that the bolts will not be exposed, and further the outer side of the second-stage partition 102 can match the shape of the first guide vane 201a.

[0044] Preferably, the thickness of the first air inlet end 201a-1 is greater than the thickness of the first air outlet end 201a-2. The countersunk hole is set at the first air inlet end 201a-1 to ensure the connection strength of the bolt. At the same time, the position of the bolt connection is set at the first air inlet end 201a-1, so that when the angle of the first guide vane 201a is adjusted, the first air inlet end 201a-1 only needs to rotate a very small angle to achieve a large change in the area of ​​the contraction flow channel at the air outlet end of the grouped guide vanes. Furthermore, adjusting the angle can change the minimum cross-sectional area of ​​the contraction flow channel, thereby adjusting the cooling air flow of each blade.

[0045] Furthermore, the second guide assembly 300 includes a second guide vane group 301 arranged in a circumferential array along the second-stage impeller disk. The second guide vane group 301 is contracted in the radial direction. The second guide vane group 301 is composed of a plurality of second guide vanes 301a having the same shape and structure. The second guide vanes 301a function as fins, thereby enhancing the heat exchange effect between the rotor cooling air and the disk.

[0046] The air inlet of each turbine second-stage rotor blade 103a corresponds to two second guide vanes 301a.

[0047] It should be noted that each second guide vane 301a has two adjacent second guide vanes 301a. The two second guide vanes 301a acting on the air inlet of the same turbine second-stage moving blade 103a with one adjacent second guide vane 301a are called a grouped guide vane, and together with another adjacent second guide vane 301a are called an adjacent guide vane group.

[0048] Furthermore, the second guide vane 301a includes a second air inlet end 301a-1 and a second air outlet end 301a-3, and the thickness of the second air inlet end 301a-1 is greater than the thickness of the second air outlet end 301a-3, wherein the second air inlet end 301a-1 is the end away from the air inlet of the second-stage turbine blade 103a, and the air inlet ends of adjacent guide vane groups are in contact, so that the cooling air can only enter the internal cavity of each blade through the contraction channel formed by the second guide vane, thereby cooling the blade.

[0049] It should be noted that each second air inlet end 301a-1 and the second-stage impeller disk 103 of the turbine are provided with corresponding countersunk holes, and the two are connected by bolts. The thickness of the second air inlet end 301a-1 is greater than the thickness of the second air outlet end 301a-3, so that the bolts will not be exposed, and further the outer side of the second-stage partition 102 can match the shape of the second guide vane 301a.

[0050] Preferably, the thickness of the second air inlet end 301a-1 is greater than the thickness of the second air outlet end 301a-3. The countersunk hole is set at the second air inlet end 301a-1 to ensure the connection strength of the bolt. At the same time, the position of the bolt connection is set at the second air inlet end 301a-1, so that when the angle of the second guide vane 301a is adjusted, the second air inlet end 301a-1 only needs to rotate a very small angle to achieve a large change in the area of ​​the contraction flow channel at the air outlet end of the grouped guide vanes. Furthermore, adjusting the angle can change the minimum cross-sectional area of ​​the contraction flow channel, thereby adjusting the cooling air flow of each blade.

[0051] Furthermore, the second-stage partition 102 includes a base 102a, a group of protruding wings 102b symmetrically arranged on the base 102a, and first comb teeth 102c arranged on the outer surface of the upper end of the protruding wings 102b.

[0052] Furthermore, a group of mounting grooves 102d are formed at the connection between a group of protruding wings 102b and the base 102a, and the two mounting grooves 102d are respectively adapted to the bosses of the corresponding first-stage turbine impeller 101 or the second-stage turbine impeller 103, wherein the first-stage impeller and the second-stage impeller are connected to the second-stage partition 102 by bolts.

[0053] Furthermore, the sealing assembly 400 includes a second-stage sealing ring 401 and a second-stage sealing ring cooling air hole 402;

[0054] The second-stage sealing ring 401 includes a second groove 401a arranged opposite to the top of the extended wing 102b, and a second grate 401b arranged at the second groove 401a, wherein the first grate 102c and the second grate 401b are arranged opposite to each other.

[0055] It should be noted that the second-stage sealing ring 401 is a hollow structure, which reduces weight while ensuring that the cooling air at the stator cavity outlet fully cools the second-stage sealing ring 401. The arrangement of the first grate teeth 102c and the second grate teeth 401b forms a grate seal structure to reduce cold air leakage.

[0056] It should be further explained that the grate sealing structure is only provided on the second-stage baffle 102 and the second-stage sealing ring 401. No sealing structure is provided in the internal cooling channel formed by the second-stage baffle 102 and the second-stage sealing ring 401, ensuring that the cooling air fully cools the baffle and the sealing ring. In addition, due to the gradual decrease in the temperature and pressure of the mainstream gas along the flow direction, the grate sealing density of the rear of the turbine first-stage impeller disk 101 is less than the grate sealing density of the front of the turbine second-stage impeller disk 103. This ensures that the outflow rate of cold air from the rear disc cavity of the first-stage impeller is greater than that from the front disc cavity of the second-stage impeller. This effectively prevents the intrusion of mainstream gas while minimizing the amount of cold air leakage from the disc cavity seal.

[0057] Preferably, the density of the comb teeth sealing structure varies along the flow direction, so that the sealing cold air outflow rate between different stages is different, the sealing cold air outflow rate of the front stage is large, and the sealing cold air outflow rate of the rear stage is small, which not only ensures the sealing effect, but also minimizes the mixing intensity of the sealing cold air and the mainstream gas, reduces flow losses, and improves turbine efficiency.

[0058] Furthermore, the top end of the second-stage sealing ring 401 is connected to the second-stage turbine vane 104 through a hanging ear and a slot.

[0059] Furthermore, the inner cross-section of the second-stage sealing ring 401 is in the shape of an inverted earth.

[0060] It should be noted that the cooling air at the outlet of the inner cavity of the stator blade cools the second-stage sealing ring 401 and is discharged into the cooling channel formed between the second-stage sealing ring 401 and the second-stage partition plate 102 through the second-stage sealing ring cooling air hole 402, fully cooling the second-stage partition plate 102 and the second-stage sealing ring 401, and is discharged after being sealed through the second-stage partition plate 102 and the grate teeth on the outside of the second-stage sealing ring 401, thereby preventing the intrusion of external mainstream fuel gas and cooling the second-stage stator blade edge plate and the third-stage moving blade edge plate.

[0061] During use, high-pressure cooling air from the compressor enters the rotor impeller, and the first cooling air enters the chamber between the first-stage turbine impeller 101 and the outer side of the second-stage partition 102 through the cooling holes on the first-stage turbine impeller 101. Under the action of the centrifugal force of the high-speed rotation of the rotor impeller, the high-pressure cooling air passes through the contraction flow channel composed of the first guide vane 201a and the second-stage partition 102. The high-pressure cooling air enters the blade cavity through the impeller groove and the cooling holes at the root of the first-stage moving blade to cool the first-stage moving blade. It should be noted that the principle of cooling the second-stage moving blade is the same as that of cooling the first-stage moving blade, which will not be repeated here.

[0062] After cooling the second-stage vane cooling air, it passes through the inner cavity of the second-stage vane and enters the chamber formed by the second-stage vane 104 and the second-stage seal ring 401. After cooling the second-stage seal ring 401, it passes through the second-stage seal ring cooling holes 402 and enters the cooling chamber between the second-stage seal ring 401 and the second-stage baffle 102. The cooling air undergoes sufficient convection heat exchange in the cooling channel formed by the second-stage baffle 102 and the second-stage seal ring 401 before entering the mainstream fuel flow through the grate seal structure on the outer end surfaces of the second-stage seal ring 401 and the second-stage baffle 102. This prevents the intrusion of mainstream fuel gas and cools the vane lip.

[0063] In summary, the beneficial effect of the gas turbine turbine disc cavity sealing cooling device of the present invention is that the guide component is set up, which can produce a suction effect on the cooling air entering the inner cavity of the moving blade under the action of the centrifugal force of the rotating wheel, and a contraction flow channel is formed between the guide vanes, which increases the cooling air flow rate, strengthens the heat exchange effect between the cooling air and the moving blades, and improves the cooling efficiency. Furthermore, the guide vanes and the wheel are connected by bolts. By adjusting the angle of the guide vanes and changing the minimum cross-sectional area of ​​the contraction flow channel formed by the guide vanes, the cooling air flow entering the inner cavity of each moving blade can be adjusted.

[0064] It should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit the present invention. Although the present invention has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical solutions of the present invention may be modified or replaced by equivalents without departing from the spirit and scope of the technical solutions of the present invention, which should all be included in the scope of the claims of the present invention.

Claims

1. A gas turbine disc cavity seal cooling device, characterized in that: include, A first flow guide assembly (200) is provided between a first-stage impeller disc (101) of a turbine and a second-stage partition plate (102), and a second flow guide assembly (300) is provided between a second-stage impeller disc (103) of a turbine and a second-stage partition plate (102), wherein the second-stage partition plate (102) is a symmetrical hollow structure; Also included is a sealing assembly (400) disposed inside the second-stage partition (102); The first guide assembly (200) comprises a first guide vane group (201) arranged in a circumferential array along the first-stage impeller disk (101) of the turbine, the first guide vane group (201) being contracted in a radial direction, wherein the first guide vane group (201) is composed of a plurality of first guide vanes (201a) having the same shape and structure; The air inlet of each turbine first-stage rotor blade (101a) corresponds to two of the first guide vanes (201a); The first guide vane (201a) comprises a first air inlet end (201a-1) and a first air outlet end (201a-2), and the thickness of the first air inlet end (201a-1) is greater than the thickness of the first air outlet end (201a-2), wherein the first air inlet end (201a-1) is an end away from the air inlet of the first-stage turbine blade (101a); The second-stage partition (102) comprises a base (102a), a group of protruding wings (102b) symmetrically arranged on the base (102a), and first comb teeth (102c) arranged on the outer surface of the upper end of the protruding wings (102b); A group of mounting grooves (102d) are formed at the connection between a group of the extended wings (102b) and the base (102a), and two of the mounting grooves (102d) are respectively adapted to the corresponding bosses of the turbine first-stage impeller disc (101) or the turbine second-stage impeller disc (103); The sealing assembly (400) comprises a second-stage sealing ring (401) and a second-stage sealing ring cooling air hole (402); The second-stage sealing ring (401) comprises a second groove (401a) arranged opposite to the top of the extended wing (102b), and second grate teeth (401b) arranged at the second groove (401a), wherein the first grate teeth (102c) and the second grate teeth (401b) are arranged opposite to each other.

2. The gas turbine turbine disc cavity seal cooling device according to claim 1, wherein: The second guide assembly (300) comprises a second guide vane group (301) arranged in a circumferential array along the second-stage impeller disc, the second guide vane group (301) being contracted in a radial direction, wherein the second guide vane group (301) is composed of a plurality of second guide vanes (301a) having the same shape and structure; The air inlet of each turbine second-stage moving blade (103a) corresponds to two of the second guide vanes (301a).

3. The gas turbine turbine disc cavity seal cooling device according to claim 2, wherein: The second guide vane (301a) comprises a second air inlet end (301a-1) and a second air outlet end (301a-3), and the thickness of the second air inlet end (301a-1) is greater than the thickness of the second air outlet end (301a-3), wherein the second air inlet end (301a-1) is an end away from the air inlet of the second-stage turbine blade (103a).

4. The gas turbine turbine disc cavity seal cooling device according to claim 3, wherein: The top end of the second-stage sealing ring (401) is connected to the second-stage stationary blade (104) of the turbine via a hanging ear and a slot.

5. The gas turbine turbine disc cavity seal cooling device according to claim 4, wherein: The inner cavity cross-section of the second-stage sealing ring (401) is in the shape of an inverted earth.

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