CN112031879A - Turbine rear support plate blade and aero-engine thereof - Google Patents
Turbine rear support plate blade and aero-engine thereof Download PDFInfo
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- CN112031879A CN112031879A CN202010988348.1A CN202010988348A CN112031879A CN 112031879 A CN112031879 A CN 112031879A CN 202010988348 A CN202010988348 A CN 202010988348A CN 112031879 A CN112031879 A CN 112031879A
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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
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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Abstract
The invention discloses a turbine rear support plate blade, belongs to the technical field of gas turbine blades, and solves the technical problems that in the prior art, an engine turbine rear support plate is heavy in weight and easily causes airflow blockage. The blade is arranged on a rear force bearing casing, first blades are arranged at intervals along the circumferential direction of the rear force bearing casing, second blades are arranged between adjacent first blades at intervals, the axial length of each second blade is smaller than that of each first blade, the radial heights of the first blades and the second blades are the same, and the turbine rear support plate blade is arranged in the aero-engine. The invention is used for perfecting the use function of the turbine rear support plate blade and meeting the requirements of people on small weight and difficulty in causing airflow blockage of the turbine rear support plate blade of the engine.
Description
Technical Field
The invention belongs to the technical field of gas turbine blades, and particularly relates to a turbine rear support plate blade and an aero-engine thereof.
Background
In an aircraft engine, an inlet of a turbine rear support plate is connected with an outlet of a turbine stage, and an outlet of the turbine rear support plate is connected with an afterburner or a tail nozzle, so that the turbine rear support plate plays a role in lifting weight in the aspects of turbine outlet rectification and engine support. The rear support plate of the turbine of the high-Mach-number engine has high through-flow capacity and rectification function and meets the requirement of no flow separation in a wide attack angle range.
The turbine rear support plate in the prior art can satisfy the effect of oil feed, oil return, ventilation, support and rectification, realizes through increasing rear support plate blade number quantity and extension plate maximum thickness, but, the traditional turbine rear support plate of great maximum thickness and more blade number because of the air current circulation passageway narrows down, when the incoming flow mach number is higher, causes the airflow channel jam between the blade easily, can't satisfy the high through-flow requirement of high mach number engine to the turbine rear support plate. In the prior art, the turbine rear strut design of an aircraft engine, as shown in figures 1 and 1a,
in fig. 1, a row of blades are designed and have the same structure, so that although the requirements of oil inlet and return, ventilation, support and rectification can be met, the vane structure cannot be used in an engine with a high mach number, when the incoming flow mach number is high, the flow speed is easily over-sonic at the position of the minimum channel area between the blades, so that the channel is blocked, and the weight is large;
in fig. 1a, the problem that the mach number of the channel is supersonic is solved by adopting a design of two rows of blades, and two rows of blades are arranged along the axial direction, but the flow channel is longer and the total number of blades is large, so that the length and the weight of the engine are increased, and the thrust-weight ratio of the engine is reduced.
In view of the above, the present invention is particularly proposed.
Disclosure of Invention
The invention aims to provide a turbine rear support plate blade, which solves the technical problems that in the prior art, a turbine rear support plate of an engine is heavy and airflow is easily blocked. The technical scheme of the scheme has a plurality of technical beneficial effects, which are described as follows:
the present case provides a support plate blade behind turbine on the one hand, and it installs on back force bearing machine casket, sets up first blade along back force bearing machine casket circumference interval, sets up the second blade at the interval between adjacent first blade, and the axial length of second blade is less than first blade axial length, and first blade is the same with second blade radial height.
In a preferred or alternative embodiment, the trailing edge lines of any adjacent second vane and/or first vane coincide with each other after rotation through a preset angle to ensure the outlet flow field periodicity.
In a preferred or alternative embodiment, the first blade and the second blade have the same radial stacking rule, the preset angle is 360 °/n, and n is the total number of the support plate blades.
In a preferred or alternative embodiment, the position of the leading edge of the second blade in the axial direction of the rear bearing casing is set after the maximum thickness of the first blade, avoiding the second blade from forming a first-contraction and then-expansion channel with the first blade, and the position of the maximum thickness has the smallest flow area.
In a preferred or alternative embodiment, the second vane axial length is less than 1/2 the first vane length, and the second vane trailing edge thickness and its axial position coincide with the first vane trailing edge to ensure periodicity of the outlet flow field.
In a preferred or alternative embodiment, the outlet configuration angles of the first and second vanes are within 5 degrees to ensure that the vane outlet airflow angle meets the requirements of the downstream component airflow angle; and/or the presence of a gas in the gas,
the forward wedge angle of the first blade is greater than 80 ° to accommodate a wide operating range of angles of attack.
In a preferred or alternative embodiment, the second blades are arranged in a compressor cascade configuration, with the second blades preceded by a diverging flow path to accommodate high mach number oncoming flow.
In a preferred or alternative embodiment, the first and/or second vanes have an inlet configuration angle that is the same as the inlet airflow angle, avoiding flow losses from flow separation.
In a preferred or alternative embodiment, the pitch of the first and second vanes is the same, meeting the requirement of the periodicity of the vane outlet flow field.
On the other hand, the aeroengine comprises a rear bearing casing, and the rear bearing casing is provided with the turbine rear support plate blades which are partially or completely arranged.
Compared with the prior art, the technical scheme provided by the invention has the following beneficial effects:
the technical scheme of this case ingenious utilizes hydrodynamics's theory of operation, under the prerequisite that satisfies engine operating requirement, sets up a plurality of axial length and is less than its second blade between first blade, satisfies under the high flux condition of air current and satisfy the requirement to engine mach number, can reduce the whole weight of engine, improves the compactness of the thrust-weight ratio and the structure of engine. Compared with the traditional blade, the weight of the installed blade is greatly reduced, and the economic cost is saved.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is an expanded view of a prior art blade;
FIG. 1a is an expanded view of a prior art two row blade;
FIG. 2 is a schematic view of the installation of the turbine rear stay blade of the present invention;
FIG. 3 is an expanded view of the first and second blades of the aft bearing cartridge of the present invention cut radially equal radii;
FIG. 4 is a schematic view of the control of the radial stacking law of the trailing edges of the first blade and the second blade of the turbine rear support plate blade according to the present invention;
FIG. 5 is a schematic view of a first vane angle setting of the turbine rear stay vane of the present invention;
wherein:
1. mounting the casing on a rear bearing machine casing; 2. a rear bearing case lower case; 11. a first blade; 12. A second blade; 3. an inlet airflow angle; 4. an inlet configuration angle; 5. an outlet construction angle; 6. mounting angles, 9 and a first blade tail edge line; 10. a second blade trailing edge line; a. a circumferential stack angle.
Detailed Description
The embodiments of the present invention are described below with reference to specific embodiments, and other advantages and effects of the present invention will be easily understood by those skilled in the art from the disclosure of the present specification. It is to be understood that the described embodiments are merely exemplary of the invention, and not restrictive of the full scope of the invention. The invention is capable of other and different embodiments and of being practiced or of being carried out in various ways, and its several details are capable of modification in various respects, all without departing from the spirit and scope of the present invention. It is to be noted that the features in the following embodiments and examples may be combined with each other without conflict. 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.
It is noted that various aspects of the embodiments are described below within the scope of the appended claims. It should be apparent that the aspects described herein may be embodied in a wide variety of forms and that any specific structure and/or function described herein is merely illustrative. Based on the disclosure, one skilled in the art should appreciate that one aspect described herein may be implemented independently of any other aspects and that two or more of these aspects may be combined in various ways. For example, an apparatus may be implemented and/or a method practiced using any number of the aspects set forth herein. Additionally, such an apparatus may be implemented and/or such a method may be practiced using other structure and/or functionality in addition to one or more of the aspects set forth herein.
It should be noted that the drawings provided in the following embodiments are only for illustrating the basic idea of the present invention, and the drawings only show the components related to the present invention rather than the number, shape and size of the components in practical implementation, and the type, quantity and proportion of the components in practical implementation can be changed freely, and the layout of the components can be more complicated.
In addition, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that aspects may be practiced without these specific details. In order that those skilled in the art will better understand the disclosure, the invention will be described in further detail with reference to the accompanying drawings and specific embodiments. The terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless otherwise specified.
The turbine rear support plate blades shown in fig. 2 to 5 are mounted on a rear bearing casing, the rear bearing casing comprises a rear bearing casing upper casing 1 and a rear bearing casing lower casing 2, and the rear bearing casing is arranged adjacent to the rotor.
The rear bearing casing is shaped, for example, as a circular ring, is arranged adjacent to the turbine rotor, and the first blades 11 and the second blades 12 are installed toward the center of the rear bearing casing. According to the technical scheme, on the premise of meeting the working requirements of the engine, by skillfully utilizing the working principle of hydromechanics, as shown in fig. 2 and 3, a plurality of second blades 12 with axial lengths smaller than that of the first blades 11 are arranged between the first blades 11, so that the requirements on the Mach number of the engine are met under the condition of meeting the high-flow condition of airflow, the overall weight of the engine can be reduced, and the thrust-weight ratio and the compactness of the structure of the engine are improved. Compared with the driven blade, the weight of the installed blade is greatly reduced, and the economic cost is saved.
As a specific embodiment provided in the present disclosure, trailing edge lines of any adjacent second blade 12 and/or first blade 11 coincide with each other after rotating by a preset angle, so as to ensure periodicity of the outlet flow field. For example, the preset angle of rotation is 360 °/n, and n is the total number of the supported plate blades, by adopting the method that the radial stacking rules of the first blade 11 and the second blade 12 are the same. Generally, the preset angle is a rotation angle of one pitch, as shown in fig. 4, after the second blade 12 rotates by one pitch, the trailing edge lines of the first blade 11 and the second blade 12 can be overlapped with each other, and the whole meets the requirement of the periodicity of the outlet flow field.
As an embodiment provided by the present disclosure, as shown in fig. 2 and 3, the front edge position of the second blade 12 in the axial direction of the rear bearing casing is set after the maximum thickness of the first blade 11, so as to avoid the second blade 12 and the first blade 11 from forming a first-contraction and then-expansion channel, and the flow area at the maximum thickness position is the minimum, so as to avoid the second blade 12 and the first blade 11 from forming a first-contraction and then-expansion channel. When the mach number of the incoming flow is high, the air flow enters the channel of the support plate to accelerate all the time, and the mach number of the channel at the position with the maximum thickness is easy to overspeed, so that the channel is blocked.
In the technical field of aeroengines, high fluidity means that the Mach number of a cascade incoming flow is not less than 0.5, and the flow of gas flowing through a unit area is large. When the Mach number of the channel is more than 1, the channel is blocked, the air flow is not smooth, the air flow velocity is increased, the pressure is reduced, and the matching of the engine is influenced. Through the structure setting of present case, improve the high circulation nature of air current, avoided the air current to appear flowing easily when flowing through first blade 11 and second blade 12 and blockked up the phenomenon. Thereby solve traditional design for satisfying the requirement of oil feed, oil return, ventilation and supporting role, can't satisfy the requirement of air current high flow property, realize first blade 11 and second blade 12 to the common rectification effect of air current.
It should be noted that, because the airflow channel formed between the first blades 11 is gradually expanded after the maximum thickness, and the axial length of the second blade 12 is smaller than the length of the first blade 11, the continuous expansion rule of the airflow channel is not changed, the channel mach number between the second blade 12 and the second blade 12 is gradually reduced along the flow direction after the maximum thickness position, and the ultrasonic phenomenon does not occur between the second blade 12 and the second blade 12.
As an embodiment provided by the present disclosure, the axial length of the second vane 12 is less than 1/2 of the length of the first vane 11, and the thickness of the trailing edge of the second vane 12 and the axial position thereof are consistent with the trailing edge of the first vane 11, so as to ensure the periodicity of the outlet flow field. On the premise of meeting the air flow circulation, the weight of the blades can be reduced to the maximum extent, and the weight of the engine is reduced as a whole.
As a specific embodiment provided by the present disclosure, the outlet structure angle 5 of the first blade 11 and the second blade 12 is within 5 °, and the installation angle 6 may satisfy the angle required by the blade design specification, so as to ensure that the angle of the outlet structure angle 5 of the blade satisfies the requirement of the airflow angle of the downstream component; and/or the presence of a gas in the gas,
the front wedge angle of the first blade 11 is larger than 80 degrees, namely the included angle of the upper tangent line and the lower tangent line of the front circle of the first blade or the front circle of the second blade, the universality of the blade is improved, and the flow field is not separated in a larger attack angle range so as to meet the requirement of an afterburner.
As a specific embodiment provided by the present disclosure, the second blade 12 is disposed in a compressor cascade structure, and an expanded flow channel is formed in front of the second blade 12 to adapt to a high mach number incoming flow.
As an embodiment provided in the present disclosure, the angle between the inlet structure angle 4 of the first blade 11 and/or the second blade 12 and the inlet airflow angle 3 is the same, so as to avoid the flow loss caused by the airflow separation.
As the specific implementation mode provided by the scheme, the grid pitches of the first blade 11 and the second blade 12 are the same, and the requirement of the periodicity of the flow field at the outlet of the blades is met.
On the other hand, the aeroengine comprises a rear bearing casing, and the rear bearing casing is provided with the turbine rear support plate blades which are partially or completely arranged.
The turbine rear support plate blade provided by the invention is described in detail above. The principles and embodiments of the present invention are explained herein using specific examples, which are presented only to assist in understanding the core concepts of the present invention. It should be noted that, for those skilled in the art, it is possible to make various improvements and modifications to the invention without departing from the inventive concept, and those improvements and modifications also fall within the scope of the claims of the invention.
Claims (10)
1. The rear support plate blade of the turbine is mounted on a rear force bearing casing and is characterized in that first blades (11) are arranged at intervals along the circumferential direction of the rear force bearing casing, second blades (12) are arranged between the adjacent first blades (11) at intervals, the axial length of each second blade (12) is smaller than that of each first blade (11), and the radial heights of the first blades (11) and the second blades (12) are the same.
2. The turbine rear strut blade as claimed in claim 1, wherein trailing edge lines of any adjacent second blade (12) and/or first blade (11) coincide with each other after rotation by a preset angle to ensure outlet flow field periodicity.
3. The turbine rear strut blade according to claim 2, characterized in that the first blade (11) and the second blade (12) are radially stacked in the same manner, with a predetermined angle of 360 °/n, n being the total number of strut blades.
4. The turbine rear strut blade according to claim 1 or 2, characterized in that the position of the front edge of the second blade (12) in the axial direction of the rear bearing casing is arranged behind the maximum thickness of the first blade (11), so that the second blade (12) and the first blade (11) are prevented from forming a first-contraction and then-expansion channel, and the flow area is minimum at the position of the maximum thickness.
5. The turbine rear strut blade as claimed in claim 4, wherein said second blade (12) axial length is less than 1/2 of said first blade (11) length, the trailing edge thickness of said second blade (12) and its axial position being coincident with the trailing edge of said first blade (11) to ensure the periodicity of the outlet flow field.
6. The turbine rear strut blade as claimed in claim 1, wherein the outlet structure angle (5) of the first and second blades (11, 12) is within 5 degrees to ensure that the blade outlet flow angle meets the downstream component flow angle requirements; and/or the presence of a gas in the gas,
the forward wedge angle of the first blade (11) is greater than 80 degrees to accommodate a wide operating range of angles of attack.
7. The turbine aft strut blade as in claim 6, wherein said second blades (12) are arranged in a compressor cascade configuration, said second blades (12) defining diverging flow passages therebetween for accommodating high mach number oncoming flow.
8. The turbine rear strut blade according to claim 7, characterized in that the inlet structure angle (4) of the first blade (11) and/or the second blade (12) is the same angle as the inlet flow angle (3), avoiding flow losses due to flow separation.
9. The turbine rear strut blade as claimed in claim 1, wherein the first blade (11) and the second blade (12) have the same pitch, and meet the requirement of the periodicity of the blade outlet flow field.
10. An aircraft engine comprising a rear stressed casing on which is mounted a turbine rear strut blade as claimed in any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010988348.1A CN112031879A (en) | 2020-09-18 | 2020-09-18 | Turbine rear support plate blade and aero-engine thereof |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010988348.1A CN112031879A (en) | 2020-09-18 | 2020-09-18 | Turbine rear support plate blade and aero-engine thereof |
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| CN112031879A true CN112031879A (en) | 2020-12-04 |
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| CN202010988348.1A Pending CN112031879A (en) | 2020-09-18 | 2020-09-18 | Turbine rear support plate blade and aero-engine thereof |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113982707A (en) * | 2021-11-04 | 2022-01-28 | 中国航发沈阳黎明航空发动机有限责任公司 | Aeroengine unloading cavity exhaust steering support plate |
| CN114542216A (en) * | 2022-02-25 | 2022-05-27 | 中国航发沈阳发动机研究所 | Design method of turbine support plate blade with supporting and flow guiding functions and blade |
| CN117874929A (en) * | 2024-03-12 | 2024-04-12 | 中国航发四川燃气涡轮研究院 | Design method for profile of swirler vane with flow stability |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2844001A (en) * | 1953-01-06 | 1958-07-22 | Gen Electric | Flow straightening vanes for diffuser passages |
| US4989406A (en) * | 1988-12-29 | 1991-02-05 | General Electric Company | Turbine engine assembly with aft mounted outlet guide vanes |
| US20030059291A1 (en) * | 2001-09-27 | 2003-03-27 | Koshoffer John Michael | Method and apparatus for reducing distortion losses induced to gas turbine engine airflow |
| US20050262847A1 (en) * | 2004-05-28 | 2005-12-01 | Koshoffer John M | Method and apparatus for gas turbine engines |
| US20140338357A1 (en) * | 2012-09-06 | 2014-11-20 | United Technologies Corporation | Cavity swirl fuel injector for an augmentor section of a gas turbine engine |
| CN105179028A (en) * | 2015-04-22 | 2015-12-23 | 北京航空航天大学 | Turbine back-bearing-force casing and gate-leaf integrated structure |
| CN207715195U (en) * | 2017-12-12 | 2018-08-10 | 中国航发沈阳发动机研究所 | A kind of rectification support plate and the aero-turbine rear housing with it |
| CN110761855A (en) * | 2019-10-11 | 2020-02-07 | 中国航发沈阳发动机研究所 | Gas turbine engine rear casing |
-
2020
- 2020-09-18 CN CN202010988348.1A patent/CN112031879A/en active Pending
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2844001A (en) * | 1953-01-06 | 1958-07-22 | Gen Electric | Flow straightening vanes for diffuser passages |
| US4989406A (en) * | 1988-12-29 | 1991-02-05 | General Electric Company | Turbine engine assembly with aft mounted outlet guide vanes |
| US20030059291A1 (en) * | 2001-09-27 | 2003-03-27 | Koshoffer John Michael | Method and apparatus for reducing distortion losses induced to gas turbine engine airflow |
| US20050262847A1 (en) * | 2004-05-28 | 2005-12-01 | Koshoffer John M | Method and apparatus for gas turbine engines |
| US20140338357A1 (en) * | 2012-09-06 | 2014-11-20 | United Technologies Corporation | Cavity swirl fuel injector for an augmentor section of a gas turbine engine |
| CN105179028A (en) * | 2015-04-22 | 2015-12-23 | 北京航空航天大学 | Turbine back-bearing-force casing and gate-leaf integrated structure |
| CN207715195U (en) * | 2017-12-12 | 2018-08-10 | 中国航发沈阳发动机研究所 | A kind of rectification support plate and the aero-turbine rear housing with it |
| CN110761855A (en) * | 2019-10-11 | 2020-02-07 | 中国航发沈阳发动机研究所 | Gas turbine engine rear casing |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113982707A (en) * | 2021-11-04 | 2022-01-28 | 中国航发沈阳黎明航空发动机有限责任公司 | Aeroengine unloading cavity exhaust steering support plate |
| CN114542216A (en) * | 2022-02-25 | 2022-05-27 | 中国航发沈阳发动机研究所 | Design method of turbine support plate blade with supporting and flow guiding functions and blade |
| CN117874929A (en) * | 2024-03-12 | 2024-04-12 | 中国航发四川燃气涡轮研究院 | Design method for profile of swirler vane with flow stability |
| CN117874929B (en) * | 2024-03-12 | 2024-06-04 | 中国航发四川燃气涡轮研究院 | Design method for profile of swirler vane with flow stability |
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Application publication date: 20201204 |