CN113513372B - Double-wall turbine guide blade with small air guiding amount - Google Patents
Double-wall turbine guide blade with small air guiding amount Download PDFInfo
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- CN113513372B CN113513372B CN202110858804.5A CN202110858804A CN113513372B CN 113513372 B CN113513372 B CN 113513372B CN 202110858804 A CN202110858804 A CN 202110858804A CN 113513372 B CN113513372 B CN 113513372B
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- 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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- 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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a double-wall turbine guide blade with small air guiding quantity, which belongs to the technical field of aero-engines and comprises a blade front edge, a middle chord blade basin, a blade back middle chord and a blade tail edge, wherein the blade front edge is cooled by a double-layer partition plate impact air-entrainment film, the middle chord blade basin and the blade back middle chord are cooled by double-wall impact air-entrainment films, and the blade tail edge is cooled by a radial turbulence column. The technology can be applied to the design of the guider of the gas turbine with high circulation parameters and small cooling air quantity, the cooling system has the advantages of small heat exchange area and high cooling effect, is very suitable for aero-engines and gas turbines with high circulation parameters, and is favorable for the thermal management and the cooling design optimization of the guider of the turbine. The small air entraining amount can be realized, the performance of the engine is improved, the flowing and the impact of a small channel are adopted, the higher heat exchange coefficient of the inner cavity can be realized, the position and the angle of the partition plate hole are adjusted, the temperatures of different positions of the blade are adjusted, and the metal temperature gradient of the blade is reduced.
Description
Technical Field
The invention belongs to the technical field of aero-engines, and particularly relates to a double-wall turbine guide vane with small air guiding quantity.
Background
In order to improve the performance and power of aero-engines and gas turbine engines and continuously improve the cycle parameters of the engines, the temperature resistance of metal materials of turbine blades cannot meet the requirement of improving the cycle parameters of the engines due to the increase of the temperature resistance of the metal materials of the turbine blades, and the thermal protection requirements of the turbine blades are increasing day by day. The most common thermal protection mode of the turbine blade is to cool the blade, the inner part of the blade is usually hollowed, and different cooling structures are arranged in the inner cavity of the blade according to different position requirements, so that the requirements of reducing the metal temperature of the blade and prolonging the service life of the turbine blade are met.
The existing turbine guide vane cooling mainly comprises cooling forms such as impact, snakelike convection, air film, turbulence columns and the like, and a plurality of cooling structure composite cooling designs are adopted according to the structure and the environmental influence of the turbine guide vane. For a high-cycle parameter turbine guider, the front edge of a blade is cooled by adopting an impact and an air film, a middle chord region of the blade adopts an impact and boss cooling structure, and a tail edge region adopts a spoiler column cooling mode. At present, the technology is mature and has been widely applied to large, medium and small aircraft engines. However, with the further increase of the turbine front temperature of the aircraft engine, the cooling mode has the limit of reducing the metal temperature of the blade by optimizing the inner cavity cooling structure to strengthen heat exchange, and only can passively adopt full-air film cooling to reduce the blade temperature, and simultaneously, the cold air bleed ratio is also increased, so that the performance of the engine is reduced.
The prior art has at least the following disadvantages:
1. the impact cavity structure is adjusted to improve the heat exchange capacity of the inner cavity of the blade, so that the potential is low, a large amount of cold air is needed to improve the heat exchange coefficient, the metal temperature of the blade body is reduced, and the performance of an engine is sacrificed;
2. the flow partition regulation difficulty is high, and the metal temperature of the blade cannot be finely controlled;
3. the existing double-wall cooling design can realize high-efficiency heat exchange, but the cold air quantity of different areas of the blade cannot be accurately controlled, and the utilization rate of the cold air is lower.
Disclosure of Invention
The invention aims to provide a double-wall turbine guide blade with small air guiding quantity aiming at the defects of the prior art, and solves the technical problem that a gas turbine guide device is not cooled sufficiently under the conditions of high circulation parameters and small air guiding quantity.
According to one aspect of the invention, the double-wall turbine guide vane with small air entrainment amount comprises a vane front edge, a middle chord vane basin, a vane back middle chord and a vane tail edge, wherein the vane front edge is used for cooling a double-layer partition plate impact air film, the middle chord vane basin and the vane back middle chord are used for cooling the double-layer wall impact air film, and the vane tail edge is used for cooling a radial turbulence column.
According to a preferred embodiment of the invention, the front edge of the blade is provided with a first cooling air cavity, a second cooling air cavity, a first impact cavity and a second impact cavity along the length direction of the blade, an impact-hole-provided partition plate is arranged between the first cooling air cavity and the first impact cavity, an impact-hole-provided partition plate is arranged between the first impact cavity and the second cooling air cavity, an impact-hole-free partition plate is arranged between the second cooling air cavity and the second impact cavity, and the first cooling air cavity, the second cooling air cavity, the first impact cavity and the second impact cavity are all provided with air film holes uniformly distributed along the length direction of the blade.
According to a preferred embodiment of the invention, the axial extensions of the impingement holes of the two impingement hole partitions in the leading edge of the blade form a leading edge impingement compared to the outer wall surface of the first impingement cavity.
According to the preferred embodiment of the invention, the middle chord blade basin is sequentially provided with a third impact cavity, a third cooling air cavity and a fourth impact cavity, an impact hole-free partition plate is arranged between the third impact cavity and the third cooling air cavity, an impact hole-equipped partition plate is arranged between the third cooling air cavity and the fourth impact cavity, impact hole-free partition plates are arranged between the fourth impact cavity and the first cooling air cavity, and air film holes uniformly distributed along the length direction of the blade are formed in the third impact cavity, the third cooling air cavity and the fourth impact cavity.
According to the preferred embodiment of the invention, a fourth cooling air cavity and a fifth impact cavity are arranged on the blade back middle chord, a partition plate with an impact hole is arranged between the fourth cooling air cavity and the fifth impact cavity, and a partition plate with an impact hole is arranged between the fourth cooling air cavity and the second impact cavity.
According to the preferred embodiment of the invention, the blade tail edge is provided with a fifth cooling air cavity, a turbulent flow column cavity and a tail edge middle split seam, the fifth cooling air cavity is U-shaped, a partition plate with an impact hole is arranged between one end of the fifth cooling air cavity and the third impact cavity, a partition plate with an impact hole is arranged between the other end of the fifth cooling air cavity and the fifth impact cavity, the fifth cooling air cavity is communicated with the turbulent flow column cavity through an air film hole, and the turbulent flow column cavity is communicated with the tail edge middle split seam.
According to the preferred embodiment of the invention, the top end of the guide blade is respectively provided with a cold air inlet corresponding to the first cooling air cavity, the second cooling air cavity, the third cooling air cavity, the fourth cooling air cavity, the fifth cooling air cavity and the turbulent flow column cavity.
According to the preferred embodiment of the invention, the impact holes of the baffle plate with the impact holes are uniformly distributed along the length direction of the baffle plate with the impact holes, and the included angle between the axial direction of the impact holes and the plane of the baffle plate with the impact holes is 30-90 degrees.
Furthermore, the extension direction of the impact holes is intersected with the outer wall surface of the double-layer blade, so that part of cold air entering from the cooling air cavity impacts the outer wall surface of the double-layer blade through the impact holes with the impact hole partition plates, and the effect of impact cooling of the cold air is achieved.
Compared with the prior art, the invention has the beneficial effects that:
1. the invention is mainly applied to aeroengines and gas turbine guide blades, and mainly comprises blade front edge double-layer partition board impact air-entraining film cooling, blade body cheek area to middle part double-layer wall impact air-entraining film cooling, and blade tail edge radial turbulence column cooling.
2. The small air entraining amount can be realized, the performance of the engine is improved, the flowing and the impact of a small channel are adopted, the higher heat exchange coefficient of the inner cavity can be realized, the position and the angle of the partition plate hole are adjusted, the temperatures of different positions of the blade are adjusted, and the metal temperature gradient of the blade is reduced.
Drawings
In order to facilitate understanding for those skilled in the art, the present invention will be further described with reference to the accompanying drawings.
FIG. 1 is a schematic cross-sectional view of a guide vane of the present invention;
FIG. 2 is a schematic view of a blade tip cold air inlet according to the present invention;
fig. 3 is a schematic view of the cross-sectional structure a-a of fig. 2.
In the figure: 1. a blade leading edge; 101. a first cooling air cavity; 102. a second cooling air cavity; 103. a first impingement cavity; 104. leading edge impact; 105. a second impingement cavity; 2. a medium chord leaf basin; 201. a third cooling air cavity; 202. a third impingement cavity; 203. a fourth impingement cavity; 3. the blade back middle chord; 301. a fourth cooling air cavity; 302. a fifth impingement cavity; 4. a trailing edge of the blade; 401. a fifth cooling air cavity; 402. a turbulator column cavity; 403. splitting in the tail edge; 5. a baffle plate with an impact hole; 6. a non-impact hole partition plate; 7. a gas film hole; 8. a central cavity; 9. and a blade top end mounting plate.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the technical solutions of the present invention are further described in detail below by way of examples with reference to the accompanying drawings. In the specification, the same or similar reference numerals denote the same or similar components. The following description of the embodiments of the present invention with reference to the accompanying drawings is intended to explain the general inventive concept of the present invention and should not be construed as limiting the invention.
Furthermore, in the following detailed description, for purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of the embodiments of the disclosure. It may be evident, however, that one or more embodiments may be practiced without these specific details. In other instances, well-known structures and devices are shown in schematic form in order to simplify the drawing.
Referring to fig. 1 to 3, according to a general technical concept of the present invention, a preferred embodiment of the present invention discloses a specific technical solution of a double-walled turbine guide vane with a small air bleed amount as follows:
a double-wall turbine guide blade with small air guiding quantity comprises a blade front edge 1, a middle chord blade basin 2, a blade back middle chord 3 and a blade tail edge 4, a narrow space is formed between double walls, a double-wall turbine guide blade flow path in the limited space is formed, cold air enters the blade from cold air inlets A-F of the upper end wall of the blade, as shown in figure 3, the cold air enters the blade along a cooling channel, the cooling channel is a narrow space and has a high heat exchange coefficient when cooling flows, and in addition, ribs, turbulence columns and the like can be arranged on the cooling channel to strengthen heat exchange. The air conditioning that gets into cooling channel passes through the baffle and gets into the impact chamber, and the baffle is equipped with the percussion hole, and the baffle percussion hole can change the impact position, strengthens regional heat transfer ability, and guide blade is 8 formula structures of central cavity, is favorable to alleviateing whole weight, and a plurality of blades pass through blade top mounting panel 9 interconnect and form the loop configuration.
The blade front edge 1 is provided with a first cooling air cavity 101, a second cooling air cavity 102, a first impact cavity 103 and a second impact cavity 105 along the length direction of the blade, the first cooling air cavity 101 and the first impact cavity 103 are separated by a partition plate 5 with impact holes, the first impact cavity 103 and the second cooling air cavity 102 are separated by a partition plate 5 with impact holes, the second cooling air cavity 102 and the second impact cavity 105 are separated by a partition plate 6 without impact holes, the first cooling air cavity 101, the second cooling air cavity 102, the first impact cavity 103 and the second impact cavity 105 are all provided with air film holes 7 uniformly distributed along the length direction of the blade, cold air respectively enters the first cooling air cavity 101 and the second cooling air cavity 102 from cold air inlets A and E of the upper end wall of the blade, a part of the cold air of the first cooling air cavity 101 and the second cooling air cavity 102 enters the first impact cavity 103 through the impact holes, the intersection of the outer walls of the blades in the first impact cavity 103 by two convection is the front edge impact position 104, because the temperature of the blade leading edge 1 is higher, the design properly increases the cold air flow at the blade leading edge 1 and effectively reduces the temperature at the position.
The middle chord blade basin 2 is sequentially provided with a third impact cavity 202, a third cooling air cavity 201 and a fourth impact cavity 203, the third impact cavity 202 and the third cooling air cavity 201 are separated by a non-impact hole partition plate 6, the third cooling air cavity 201 and the fourth impact cavity 203 are separated by a partition plate 5 with impact holes, the fourth impact cavity 203 and the first cooling air cavity 101 are separated by the non-impact hole partition plate 6, and air film holes 7 uniformly distributed along the length direction of the blade are formed in the third impact cavity 202, the third cooling air cavity 201 and the fourth impact cavity 203.
The blade back middle chord 3 is provided with a fourth cooling air cavity 301 and a fifth impact cavity 302, the fourth cooling air cavity 301 and the fifth impact cavity 302 are separated by a partition plate with impact holes 5, the fourth cooling air cavity 301 and the second impact cavity 105 are separated by the partition plate with impact holes 5, and the fourth cooling air cavity 301 and the fifth impact cavity 302 are not provided with air film holes 7.
The blade trailing edge 4 is equipped with fifth cooling air cavity 401, spoiler column chamber 402 and trailing edge median slit 403, and fifth cooling air cavity 401 is the U-shaped, separates with taking impact hole baffle 5 between fifth cooling air cavity 401 one end and the third impact chamber 202, separates with taking impact hole baffle 5 between fifth cooling air cavity 401 other end and the fifth impact chamber 302, communicates with spoiler column chamber 402 with diaphragm hole 7 between fifth cooling air cavity 401 and the spoiler column chamber 402, spoiler column chamber 402 and trailing edge median slit 403 communicate.
The cold flows in the cross section of the blade body as shown in fig. 1 below.
The cold air of the front edge 1 of the blade enters a first cooling air cavity 101 and a second cooling air cavity 102 inside the blade from cold air inlets A and E, enters a first impact cavity 103 of the blade through a partition plate 5 with impact holes, and can pass through the angle of the impact holes to adjust the position of the cold air to impact the inner cavity of the front edge 1 of the blade, so that the effect of locally strengthening heat exchange is achieved, and the cold air flows out through a front edge air film hole 7 after impacting the front edge 1 of the blade.
In the blade middle chord blade basin 2, cold air enters a third cooling air cavity 201 through a cold air inlet B, impacts a fourth impact cavity 203 of the middle chord blade basin 2 through a partition plate 5 with an impact hole, and after impact heat exchange is finished, the cold air flows out from an air film hole 7.
Because the blade back middle chord 3 can not be provided with the air film hole 7, after cold air enters the fourth cooling air cavity 301 from the cold air inlet D, the cold air is divided into two paths, and a part of the cold air enters the second impact cavity 105 in the direction of the blade front edge 1 to play a role in strengthening heat exchange and flows out from the air film hole 7; the other part of the cold air enters the fifth impact cavity 302 along the impact partition plate, is mixed with the cold air of the fifth cooling air cavity 401, enters the spoiler cavity 402 and is discharged through the tail cleft.
The above-mentioned embodiments are intended to illustrate the objects, technical solutions and advantages of the present invention in further detail, and it should be understood that the above-mentioned embodiments are only exemplary embodiments of the present invention, and are not intended to limit the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (7)
1. A double-wall turbine guide blade with small air guiding quantity comprises a blade front edge (1), a middle chord blade basin (2), a blade back middle chord (3) and a blade tail edge (4), and is characterized in that the blade front edge (1) is cooled by a double-layer partition impact air-entraining film, the middle chord blade basin (2) and the blade back middle chord (3) are cooled by a double-wall impact air-entraining film, and the blade tail edge (4) is cooled by a radial air-entraining column;
the blade front edge (1) is provided with a first cooling air cavity (101), a second cooling air cavity (102), a first impact cavity (103) and a second impact cavity (105) along the length direction of the blade, a partition plate (5) with an impact hole is arranged between the first cooling air cavity (101) and the first impact cavity (103), the partition plate (5) with the impact hole is arranged between the first impact cavity (103) and the second cooling air cavity (102), a non-impact-hole partition plate (6) is arranged between the second cooling air cavity (102) and the second impact cavity (105), and air film holes (7) uniformly distributed along the length direction of the blade are formed in the first cooling air cavity (101), the second cooling air cavity (102), the first impact cavity (103) and the second impact cavity (105).
2. The double-walled turbine guide vane with small air entrainment amount according to claim 1, characterized in that the axial extensions of the impingement holes of the two impingement hole partition plates (5) in the leading edge (1) of the vane form leading edge impingement (104) compared to the outer wall surface of the first impingement cavity (103).
3. The double-wall turbine guide blade with small air guiding quantity according to claim 1, wherein the middle chord blade basin (2) is sequentially provided with a third impingement cavity (202), a third cooling air cavity (201) and a fourth impingement cavity (203), a non-impingement-hole partition plate (6) is arranged between the third impingement cavity (202) and the third cooling air cavity (201), a partition plate (5) with impingement holes is arranged between the third cooling air cavity (201) and the fourth impingement cavity (203), the fourth impingement cavity (203) and the first cooling air cavity (101) are provided with non-impingement-hole partition plates (6), and the third impingement cavity (202), the third cooling air cavity (201) and the fourth impingement cavity (203) are all provided with air film holes (7) uniformly distributed along the length direction of the blade.
4. The double-walled turbine guide vane with small air induction amount according to claim 3, characterized in that a fourth cooling air cavity (301) and a fifth impingement cavity (302) are arranged on the blade back center chord (3), a baffle plate (5) with impingement holes is arranged between the fourth cooling air cavity (301) and the fifth impingement cavity (302), and the baffle plate (5) with impingement holes is arranged between the fourth cooling air cavity (301) and the second impingement cavity (105).
5. The double-walled turbine guide blade with small air entrainment volume according to claim 4, characterized in that the trailing edge (4) of the blade is provided with a fifth cooling air cavity (401), a spoiler column cavity (402) and a trailing edge median split seam (403), the fifth cooling air cavity (401) is U-shaped, a partition plate (5) with an impact hole is arranged between one end of the fifth cooling air cavity (401) and the third impact cavity (202), a partition plate (5) with an impact hole is arranged between the other end of the fifth cooling air cavity (401) and the fifth impact cavity (302), the fifth cooling air cavity (401) is communicated with the spoiler column cavity (402) through an air film hole (7), and the spoiler column cavity (402) is communicated with the spoiler trailing edge median split seam (403).
6. The double-walled turbine guide vane with small air entrainment amount according to claim 5, characterized in that the guide vane tip is provided with cold air inlets corresponding to the first cooling air chamber (101), the second cooling air chamber (102), the third cooling air chamber (201), the fourth cooling air chamber (301), the fifth cooling air chamber (401) and the spoiler column chamber (402), respectively.
7. The double-walled turbine guide vane with small air entrainment quantity according to claim 6, characterized in that the impingement holes of the baffle plate (5) with impingement holes are uniformly distributed along the length direction thereof, and the included angle between the axial direction of the impingement holes and the plane of the baffle plate (5) with impingement holes is 30-90 °.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202110858804.5A CN113513372B (en) | 2021-07-28 | 2021-07-28 | Double-wall turbine guide blade with small air guiding amount |
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| CN202110858804.5A CN113513372B (en) | 2021-07-28 | 2021-07-28 | Double-wall turbine guide blade with small air guiding amount |
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| CN113513372B true CN113513372B (en) | 2022-05-31 |
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Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114151138B (en) * | 2021-10-20 | 2023-05-05 | 中国航发四川燃气涡轮研究院 | Interlayer combined cooling structure of turbine rotor blade |
| CN114017131B (en) * | 2021-11-12 | 2023-06-02 | 中国航发沈阳发动机研究所 | Variable geometry low pressure turbine guide vane half-layer plate cooling structure |
| CN114382553B (en) * | 2021-12-26 | 2024-06-18 | 西北工业大学 | High-blockage-ratio rib laminate cooling structure and cooling method for middle chord zone of turbine blade |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2001140602A (en) * | 1999-11-12 | 2001-05-22 | Mitsubishi Heavy Ind Ltd | Gas turbine stationary blade |
| CN202417612U (en) * | 2011-12-27 | 2012-09-05 | 中航商用航空发动机有限责任公司 | Turbine guide blade |
| CN205064004U (en) * | 2015-09-07 | 2016-03-02 | 沈阳航空航天大学 | Turbine guide blade structure under heat transfer effect is assisted to heat pipe |
| US10907479B2 (en) * | 2018-05-07 | 2021-02-02 | Raytheon Technologies Corporation | Airfoil having improved leading edge cooling scheme and damage resistance |
| CN109812301A (en) * | 2019-03-06 | 2019-05-28 | 上海交通大学 | A kind of turbo blade double wall cooling structure with horizontal communication hole |
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