CN114135340A - Binary channels refrigerated turbine bladed disk - Google Patents
Binary channels refrigerated turbine bladed disk Download PDFInfo
- Publication number
- CN114135340A CN114135340A CN202111470846.8A CN202111470846A CN114135340A CN 114135340 A CN114135340 A CN 114135340A CN 202111470846 A CN202111470846 A CN 202111470846A CN 114135340 A CN114135340 A CN 114135340A
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- Prior art keywords
- turbine
- blade
- disc
- low
- pressure
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- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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Links
- 238000001816 cooling Methods 0.000 claims abstract description 44
- 230000009977 dual effect Effects 0.000 claims description 7
- 238000010146 3D printing Methods 0.000 claims description 3
- 235000017166 Bambusa arundinacea Nutrition 0.000 claims description 3
- 235000017491 Bambusa tulda Nutrition 0.000 claims description 3
- 241001330002 Bambuseae Species 0.000 claims description 3
- 235000015334 Phyllostachys viridis Nutrition 0.000 claims description 3
- 239000011425 bamboo Substances 0.000 claims description 3
- 229910003460 diamond Inorganic materials 0.000 claims 1
- 239000010432 diamond Substances 0.000 claims 1
- 239000000112 cooling gas Substances 0.000 description 15
- 239000007789 gas Substances 0.000 description 13
- 230000000694 effects Effects 0.000 description 4
- 239000007787 solid Substances 0.000 description 3
- QNRATNLHPGXHMA-XZHTYLCXSA-N (r)-(6-ethoxyquinolin-4-yl)-[(2s,4s,5r)-5-ethyl-1-azabicyclo[2.2.2]octan-2-yl]methanol;hydrochloride Chemical compound Cl.C([C@H]([C@H](C1)CC)C2)CN1[C@@H]2[C@H](O)C1=CC=NC2=CC=C(OCC)C=C21 QNRATNLHPGXHMA-XZHTYLCXSA-N 0.000 description 2
- 125000004122 cyclic group Chemical group 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 238000013459 approach Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 239000002699 waste material Substances 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a dual-channel cooled turbine blade disc, which comprises an annular turbine disc and two radial plates arranged on the side surfaces of the turbine disc, wherein cross ribs are arranged in a disc cavity formed by the turbine disc and the two radial plates, a plurality of blades are arranged on the periphery of the turbine disc, and each blade is provided with a low-pressure cooling channel and a high-pressure cooling channel; a low-pressure air inlet of the low-pressure cooling channel is communicated with the disc cavity, a plurality of low-pressure outlets of the low-pressure cooling channel are arranged on the side surface of each blade, and a turbulence column is arranged in each of the plurality of low-pressure outlets of each blade to form a turbulence column group; the high-pressure air inlet of the high-pressure cooling channel is arranged on the radial plate, the high-pressure air outlet is arranged on the outer edge of the blade, and the high-pressure cooling channel is formed in the blade in an inverted N shape.
Description
Technical Field
The invention relates to the technical field of gas turbines, in particular to a dual-channel cooled turbine blade disc.
Background
With the population growth and global economic development in the next 20 years, the world aviation demand is increasing, and energy safety and carbon emission reduction are important issues for advanced aviation gas turbine designs. In modern gas turbines, up to 20% of the core gas flow is used for cooling and sealing the turbine rotor, which results in 5% of available energy waste and fuel loss, and the reduction of cooling gas has great significance for saving energy of advanced aeroengines. With the continuously improved aviation performance, the design temperature of a turbine gas inlet continuously rises, the contradiction between temperature stress reduction and cold air reduction is more and more serious under the condition that the inlet temperature of a fifth generation of fighter exceeds 2000K, and the structural design of combining a traditional solid turbine disc with split type cooling blades in a turbine rotor approaches the heat exchange design limit.
The blisk structure combines the blades and the rotating disc into a complete component by using an integrated forming technology, the design of tenons and mortises is reduced, the problem of failure caused by fracture of the tenons and the mortises is avoided, the structure is more compact, but the structure cannot bear a high-temperature environment and is only used in aero-engine components with relatively low temperature of fans and gas compressors.
Disclosure of Invention
Aiming at the defects in the prior art, the invention provides a double-channel cooled turbine blade disc, and aims to solve the problem that the existing blisk structure cannot bear a high-temperature environment.
In order to achieve the purpose, the invention adopts the following technical scheme:
the turbine blade disc comprises an annular turbine disc and two radial plates arranged on the side faces of the turbine disc, cross ribs are arranged in a disc cavity formed by the turbine disc and the two radial plates, a plurality of blades are arranged on the periphery of the turbine disc, and each blade is provided with a low-pressure cooling channel and a high-pressure cooling channel;
a low-pressure air inlet of the low-pressure cooling channel is communicated with the disc cavity, a plurality of low-pressure outlets of the low-pressure cooling channel are arranged on the side surface of each blade, and a turbulence column is arranged in each of the plurality of low-pressure outlets of each blade to form a turbulence column group;
the high-pressure air inlet of the high-pressure cooling channel is arranged on the radial plate, the high-pressure air outlet is arranged on the outer edge of the blade, and the high-pressure cooling channel is formed in the blade in an inverted N shape.
Furthermore, the turbulence column is a columnar body, and the cross section of the turbulence column is in any one of a circular shape, a round bamboo joint shape, an oval shape, a water drop shape, a rectangular shape and a rhombic shape.
Furthermore, the crossed ribs comprise a plurality of first rib plates and second rib plates, one ends of the adjacent first rib plates and second rib plates are crossed and fixed on the inner periphery of the turbine disc, and the other ends of the adjacent first rib plates and second rib plates are respectively fixed on the disc centers of the two radial plates.
Furthermore, the adjacent first rib plates and the second rib plates are arranged in a crossed manner.
Further, the blades are arranged on the outer periphery of the turbine disc in a circular array in the center of the turbine disc.
Further, turbine disc and a plurality of blade are through 3D printing and are integrative setting.
The invention has the beneficial effects that: in this scheme, because traditional turbine bladed disk is solid, can't bear the high temperature environment, the inefficacy problem that thermal deformation and high temperature creep lead to makes it unable use in turbine rotor, and this scheme has solved the problem that can't bear the high temperature through the disk chamber that turbine disk and two radials formed to it subtracts heavy effect obvious, is fit for high-speed high maneuvering's small-size aeroengine. Adopt high-pressure cooling gas to cool off the blade front end, low pressure cooling gas cools off disk chamber and blade trailing edge, through setting up low pressure cooling channel and high pressure cooling channel, can give turbine dish cooling fast. Through the high-pressure cooling channel who falls "N" shape, reached cyclic utilization high-pressure cooling gas's purpose, avoided high-pressure cooling gas's pressure to reduce fast, guarantee that exhaust high-pressure cooling gas's pressure is invariably greater than the pressure of the high temperature high pressure gas of blade outer edge to prevent that the gas from flowing into high-pressure cooling channel in, cause the damage to the blade. The cross ribs enhance the internal heat exchange of the turbine disc and also increase the structural strength of the turbine disc.
In addition to the technical problems addressed by the present invention, the technical features constituting the technical solutions, and the advantageous effects brought by the technical features of the technical solutions described above, other technical problems that the present invention can solve, other technical features included in the technical solutions, and advantageous effects brought by the technical features will be described in further detail in the detailed description.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention and not to limit the invention. In the drawings:
FIG. 1 is a schematic illustration of a dual path cooled turbine disk according to the present invention.
FIG. 2 is a schematic illustration of a turbine disk configuration with the webs removed.
FIG. 3 is a partial left sectional view of a dual path cooled turbine disk in accordance with the present invention.
Wherein: 1. a blade; 2. a turbine disk; 3. a first rib plate; 4. a second rib plate; 5. a disc core; 6. a high pressure gas outlet; 7. a high pressure cooling channel; 8. a high pressure air inlet; 9. a web; 10. a disc cavity; 11. a low pressure air inlet; 12. a low pressure cooling channel; 13. and a low-pressure air outlet.
Detailed Description
In order to make the aforementioned objects, features and advantages of the present invention comprehensible, embodiments accompanied with figures are described in detail below. It is to be understood that the described embodiments are merely a few embodiments of the invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1-3, the present invention provides a dual-channel cooled turbine blade disc, which includes an annular turbine disc 2 and two radial plates 9 disposed on the side of the turbine disc 2, cross ribs are disposed in a disc cavity 10 formed by the turbine disc 2 and the two radial plates 9, a plurality of blades 1 are disposed on the outer circumference of the turbine disc 2, and each blade 1 is provided with a low-pressure cooling channel 12 and a high-pressure cooling channel 7.
The low-pressure air inlet 11 of the low-pressure cooling channel 12 is communicated with the disc cavity 10, the side parts of the blades 1 are provided with a plurality of low-pressure air outlets 13 of the low-pressure cooling channel 12, and turbulence columns are arranged in the low-pressure air outlets 13 of each blade 1 to form a turbulence column group. A high-pressure air inlet 8 of a high-pressure cooling channel 7 is arranged on a radial plate 9, a high-pressure air outlet 6 is arranged on the outer periphery of the blade 1, and the high-pressure cooling channel 7 is formed in an inverted 'N' shape in the blade 1.
The crossed ribs comprise a plurality of first rib plates 3 and second rib plates 4, one ends of the adjacent first rib plates 3 and second rib plates 4 are crossed and fixed on the inner periphery of the turbine disc 2, and the other ends of the adjacent first rib plates 3 and second rib plates 4 are respectively fixed on the disc centers 5 of the two radial plates 9. The adjacent first rib 3 and second rib 4 are arranged crosswise.
The turbulence column is a columnar body, and the cross section of the turbulence column is in any one of a circular shape, a round bamboo joint shape, an oval shape, a drop shape, a rectangular shape and a rhombic shape. A plurality of blades 1 are arranged on the outer circumference of the turbine disc 2 in a circular array around the center of the turbine disc 2. Traditional turbine dish 2 and blade 1 adopt connections such as tenon, tongue-and-groove, compare with this scheme, blade 1 and turbine dish 2 are integrative setting through 3D printing, make the juncture more firm, and structural stress distributes more evenly.
To make more efficient use of the cooling gas for cooling, the following two cooling principles are set forth:
high-pressure cooling gas enters the high-pressure cooling channel 7 from a high-pressure gas inlet 8 at the side edge of the web plate 9, flows in the inverted N-shaped high-pressure cooling channel 7, and is discharged from a high-pressure gas outlet 6 on the outer edge of the blade 1 after exchanging heat with the blade 1. The low-pressure cooling gas enters the disc cavity 10 through the disc center 5 opening of the radial plate 9, and flows into the low-pressure cooling channel 12 of each blade 1 after the pressure of the low-pressure cooling gas is increased by centrifugal force generated by rotation and then flows out from the low-pressure cooling outlet after the heat exchange with the turbulent flow column group is enhanced.
In this scheme, because traditional turbine bladed disk is solid, can't bear the high temperature environment, the inefficacy problem that thermal deformation and high temperature creep lead to makes it unable use in turbine rotor, and this scheme has solved the problem that can't bear the high temperature through the disk chamber 10 that turbine disk 2 and two radials 9 formed to it subtracts heavy effect obvious, is fit for high-speed high maneuvering's small-size aeroengine. Adopt high-pressure cooling gas to cool off blade 1 front end, low pressure cooling gas cools off disk chamber 10 and blade 1 trailing edge, through setting up low pressure cooling channel 12 and high pressure cooling channel 7, can give turbine disc 2 cooling fast. Through the high-pressure cooling channel 7 of "N" shape of falling, reached cyclic utilization high-pressure cooling gas's purpose, avoided high-pressure cooling gas's pressure to reduce fast, guarantee that exhaust high-pressure cooling gas's pressure is invariably greater than the pressure of the high temperature high pressure gas of blade 1 outer edge to prevent that the gas from flowing back into high-pressure cooling channel 7 in, cause the damage to blade 1. The cross ribs enhance the internal heat exchange of the turbine disc 2 and also increase the structural strength of the turbine disc 2.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit the same; while the invention has been described in detail and with reference to the foregoing embodiments, it will be understood by those skilled in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present invention.
Claims (6)
1. The dual-channel cooled turbine blade disc is characterized by comprising an annular turbine disc (2) and two radial plates (9) arranged on the side faces of the turbine disc (2), wherein cross ribs are arranged in a disc cavity (10) formed by the turbine disc (2) and the two radial plates (9), a plurality of blades (1) are arranged on the periphery of the turbine disc (2), and each blade (1) is provided with a low-pressure cooling channel (12) and a high-pressure cooling channel (7);
a low-pressure air inlet (11) of the low-pressure cooling channel (12) is communicated with the disc cavity (10), a plurality of low-pressure air outlets (13) of the low-pressure cooling channel (12) are arranged on the side surface of each blade (1), and turbulence columns are arranged in the plurality of low-pressure air outlets (13) of each blade (1) to form a turbulence column group;
high-pressure air inlet (8) of high pressure cooling channel (7) set up on radials (9), and high-pressure gas outlet (6) set up on the outward flange of blade (1), high pressure cooling channel (7) form in blade (1) and are "N" shape of falling.
2. The dual channel cooled turbine blisk according to claim 1, wherein the turbulator is a cylinder, and the cross-sectional shape of the turbulator is any one of circular, round bamboo joint, oval, drop, rectangular, and diamond.
3. The dual channel cooled turbine disk according to claim 1, wherein the cross ribs comprise a plurality of first ribs (3) and second ribs (4), one end of each of the first ribs (3) and the second ribs (4) adjacent to each other is fixed to the inner periphery of the turbine disk (2) in a crossing manner, and the other end is fixed to the disk center (5) of each of the two radial plates (9).
4. The dual channel cooled turbine disk according to claim 3, wherein adjacent first ribs (3) and second ribs (4) are interdigitated.
5. The dual channel cooled turbine blisk according to claim 1, characterised in that several of said blades (1) are arranged on the outer circumference of the turbine disk (2) in a circular array with the centre of the turbine disk (2).
6. The dual channel cooled turbine blisk according to claim 1, characterised in that the turbine disk (2) and several blades (1) are integrated by 3D printing.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111470846.8A CN114135340A (en) | 2021-12-03 | 2021-12-03 | Binary channels refrigerated turbine bladed disk |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111470846.8A CN114135340A (en) | 2021-12-03 | 2021-12-03 | Binary channels refrigerated turbine bladed disk |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114135340A true CN114135340A (en) | 2022-03-04 |
Family
ID=80387684
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111470846.8A Pending CN114135340A (en) | 2021-12-03 | 2021-12-03 | Binary channels refrigerated turbine bladed disk |
Country Status (1)
| Country | Link |
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| CN (1) | CN114135340A (en) |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104929779A (en) * | 2015-04-30 | 2015-09-23 | 中国科学院工程热物理研究所 | Turbine disk connecting structure and gas turbine engine with same |
| CN105275499A (en) * | 2015-06-26 | 2016-01-27 | 中航空天发动机研究院有限公司 | Disc center air inlet structure of double-radial-plate turbine disc with centrifugal pressurization effect and sealing effect |
| CN106065785A (en) * | 2016-07-21 | 2016-11-02 | 中国航空动力机械研究所 | Cooling blades of turbine rotor |
| CN107060889A (en) * | 2017-04-19 | 2017-08-18 | 西北工业大学 | A kind of double disc turbine disks with disk chamber turbulence columns |
| CN108374692A (en) * | 2018-01-25 | 2018-08-07 | 南方科技大学 | Turbine wheel disc and turbine engine |
| CN113550795A (en) * | 2021-08-25 | 2021-10-26 | 中国航发湖南动力机械研究所 | Gas turbine suitable for all territories |
| CN113550794A (en) * | 2021-09-10 | 2021-10-26 | 中国航发湖南动力机械研究所 | Multi-cavity efficient cooling structure and cooling method for turbine rotor blade |
-
2021
- 2021-12-03 CN CN202111470846.8A patent/CN114135340A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104929779A (en) * | 2015-04-30 | 2015-09-23 | 中国科学院工程热物理研究所 | Turbine disk connecting structure and gas turbine engine with same |
| CN105275499A (en) * | 2015-06-26 | 2016-01-27 | 中航空天发动机研究院有限公司 | Disc center air inlet structure of double-radial-plate turbine disc with centrifugal pressurization effect and sealing effect |
| CN106065785A (en) * | 2016-07-21 | 2016-11-02 | 中国航空动力机械研究所 | Cooling blades of turbine rotor |
| CN107060889A (en) * | 2017-04-19 | 2017-08-18 | 西北工业大学 | A kind of double disc turbine disks with disk chamber turbulence columns |
| CN108374692A (en) * | 2018-01-25 | 2018-08-07 | 南方科技大学 | Turbine wheel disc and turbine engine |
| CN113550795A (en) * | 2021-08-25 | 2021-10-26 | 中国航发湖南动力机械研究所 | Gas turbine suitable for all territories |
| CN113550794A (en) * | 2021-09-10 | 2021-10-26 | 中国航发湖南动力机械研究所 | Multi-cavity efficient cooling structure and cooling method for turbine rotor blade |
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