CN116291756A - Guide vane of turbine guide of aeroengine - Google Patents
Guide vane of turbine guide of aeroengine Download PDFInfo
- Publication number
- CN116291756A CN116291756A CN202211737708.6A CN202211737708A CN116291756A CN 116291756 A CN116291756 A CN 116291756A CN 202211737708 A CN202211737708 A CN 202211737708A CN 116291756 A CN116291756 A CN 116291756A
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- CN
- China
- Prior art keywords
- blade body
- blade
- guide vane
- aeroengine
- mounting head
- Prior art date
- 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.)
- Pending
Links
- 238000009434 installation Methods 0.000 claims abstract description 11
- 238000003466 welding Methods 0.000 claims abstract description 11
- 238000001816 cooling Methods 0.000 claims description 15
- 238000013461 design Methods 0.000 abstract description 12
- 238000005266 casting Methods 0.000 description 25
- 239000013078 crystal Substances 0.000 description 13
- 238000000034 method Methods 0.000 description 8
- 238000012797 qualification Methods 0.000 description 7
- 238000004364 calculation method Methods 0.000 description 3
- 238000004519 manufacturing process Methods 0.000 description 3
- 238000012360 testing method Methods 0.000 description 3
- 238000011161 development Methods 0.000 description 2
- 238000012545 processing Methods 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 235000003283 Pachira macrocarpa Nutrition 0.000 description 1
- 241001083492 Trapa Species 0.000 description 1
- 235000014364 Trapa natans Nutrition 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000010485 coping Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000008030 elimination Effects 0.000 description 1
- 238000003379 elimination reaction Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
- 235000009165 saligot Nutrition 0.000 description 1
- 238000007493 shaping process Methods 0.000 description 1
Images
Classifications
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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
-
- 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
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E10/00—Energy generation through renewable energy sources
- Y02E10/20—Hydro energy
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the technical field of aeroengine turbine director guide vane design, and in particular relates to an aeroengine turbine director guide vane, which comprises: two edge plates, two installation heads and blade body, wherein: the blade body is provided with three groups of large, medium and small areas; the two mounting heads can accommodate three groups of blade bodies with large, medium and small areas; the blade root and the blade tip of the blade body can be respectively inserted into one mounting head and are connected with the mounting heads in a welding way; the two edge plates are provided with holes; each mounting head is correspondingly and integrally formed in one forming hole.
Description
Technical Field
The application belongs to the technical field of design of turbine guide vanes of aeroengines, and particularly relates to a turbine guide vane of an aeroengine.
Background
In the development process of the aeroengine, certain uncertainty of performance calculation, such as calculation model simplification, calculation errors and the like, and errors in the machining and assembly process, inevitably cause a gap between the area first assembly requirement of the turbine guide and the area required in the actual test run process, and the area of the guide is required to be designed to be adjustable so as to meet the requirement of overall performance.
The common method for adjusting the area of the turbine guide is to design and manufacture three guide vanes with large, medium and small areas in the development stage, wherein the guide vanes with medium area meet the design point requirement, and the guide vanes with larger and smaller areas are used for meeting the group requirements of slip and performance in the actual test run process, referring to the design position in fig. 1.
Because the guide vane of the turbine guide vane is not very sensitive to the change of the inlet attack angle, the area of the guide vane can be changed by adopting a method for changing the mounting angle of the blade profile, the guide vanes with three areas share the same rough material, and the angle relation between the blade body and the edge plate is ensured by the tool fixture.
Currently, the change in the turbine guide vane group is mainly achieved in two ways:
one is that the blade body of the guide vane, edge plate are integrally molded and cast, the guide vane with large, medium and small different areas is processed by the mould respectively;
another kind of guide vane components of a whole that can function independently casting, blade body, marginal board welding combination, as shown in fig. 2, design can contain the installation head of big, little group's guide vane turned angle on the blade body, as shown in fig. 3, the installation head contour line is group's loose piece line of containing, its mould loose piece design, according to engine performance demand, only need rotate the installation head mould angle, can make the blade of corresponding group, and need not to throw the mould newly, reducible casting mould number shortens the process cycle, this kind of technical scheme has following defect in practical application:
1) The profile of the mounting head and the hole patterns of the flange plate have tolerance deviations, if interference or welding gaps are insufficient during assembly, the assembly is often carried out after partial grinding, step differences exist between the blade body and the flange plate at welding positions during welding assembly, and a recess in the R region of the near end wall of the blade body is formed, as shown in fig. 4, the recess can strengthen the horseshoe vortex strength of the end wall of the front edge, and the damage of the guide blade influenced by the horseshoe vortex is aggravated;
2) For the low-consistency and large-chord-length flange, even if the deformation of the flange profile is well controlled, the profile tolerance is unavoidable, so that the flange profile deviates from the theoretical profile, as shown by the flange profile I and the flange profile II in fig. 5, the profile of the region R of the blade body near the end wall deviates from the flange profile, the profile curve is not smooth, and the pneumatic loss of the flange near the wall is increased.
3) Along with the rise of the temperature resistant technology of the guide vane, the cooling channel structure in the blade body is developed to millimeter or even micrometer level, and single crystal casting is used, and the mounting head is formed on the blade body, so that great difficulty is brought to the single crystal growth and the single crystal integrity casting of the blade body due to the magnitude difference of casting dimensions, and the qualification rate of the single crystal blade body is lower.
The present application has been made in view of the existence of the above-mentioned technical drawbacks.
It should be noted that the above disclosure of the background art is only for aiding in understanding the inventive concept and technical solution of the present invention, which is not necessarily prior art to the present application, and should not be used for evaluating the novelty and the creativity of the present application in the case where no clear evidence indicates that the above content has been disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aeroengine turbine guide vane that overcomes or mitigates at least one technical disadvantage of known existing aspects.
The technical scheme of the application is as follows:
an aeroengine turbine guide vane comprising: two edge plates, two installation heads and blade body, wherein:
the blade body is provided with three groups of large, medium and small areas;
the two mounting heads can accommodate three groups of blade bodies with large, medium and small areas;
the blade root and the blade tip of the blade body can be respectively inserted into one mounting head and are connected with the mounting heads in a welding way;
the two edge plates are provided with holes;
each mounting head is correspondingly and integrally formed in one forming hole.
According to at least one embodiment of the present application, in the turbine guide vane of an aeroengine, the two mounting heads and the corresponding flange plates are integrally cast and formed.
According to at least one embodiment of the present application, in the turbine guide vane of the aeroengine, the blade body is internally provided with a cooling channel.
The application has at least the following beneficial technical effects:
the utility model provides an aeroengine turbine director stator, its design installation head corresponds the cast shaping in the type hole of limit plate, can effectively reduce change group processing mould, reduction in production cost, and do not have the installation head on the blade body, with the limit plate welded position not receive the coping influence, the blade body is misplaced in radial direction no matter during the welding equipment, all there is not the recess in the blade body near-end wall R district, as shown in fig. 7, can effectively reduce the guide vane damage risk that the water chestnut vortex influences, in addition, even to low consistence, big chord length limit plate, even there is profile tolerance, also do not have the problem that profile curve is not smooth, guarantee the pneumatic properties of guide vane, as shown in fig. 8.
For the aeroengine turbine guide vane disclosed by the embodiment, the person skilled in the art can also understand that the mounting head is designed to correspond to the hole middle flange plate of the casting forming flange plate to be close to the flat plate structure, the cooling structure is few in type, low in casting difficulty and high in qualification rate, the flange plate is large in structural size and is equivalent to the mounting head in size, the casting size cannot have magnitude difference, the casting structure size of the cooling channel in the blade body is the same magnitude, and the cooling channel is designed to be separated from the mounting head, so that the casting difficulty of the micro-size cooling blade body of the single crystal can be reduced, the growth of the single crystal of the blade body is facilitated, and the qualification rate of the integrity of the single crystal casting is improved.
According to the guide vane of the turbine guide vane of the aeroengine, the mounting head is designed to correspond to the hole middle edge plate of the casting forming edge plate to be close to the flat plate structure, the cooling structure is few in type, the casting difficulty is low, the qualification rate is high, the structural size of the edge plate is larger, the size of the edge plate is equal to that of the mounting head, the casting size cannot have magnitude difference, the casting structure size of the cooling channel in the blade body is the same magnitude, the cooling channel is designed to be separated from the mounting head, the casting difficulty of the micro-size cooling blade body of the single crystal can be reduced, the growth of the single crystal of the blade body is facilitated, and the qualification rate of the integrity of the single crystal casting is improved.
Drawings
FIG. 1 is a schematic illustration of three different area vanes, large, medium, and small, in turbine pilot area adjustment;
FIG. 2 is a schematic illustration of a turbine guide vane split casting, blade airfoil, and platform weld assembly;
FIG. 3 is a schematic illustration of a mounting head on a blade body designed to contain large and small sets of vane angles of rotation;
FIG. 4 is a schematic view of a blade body and flange having a step at the weld to form a recess in the blade body near end wall R region;
FIG. 5 is a schematic view of a blade edge profile deviating from a theoretical profile, resulting in deviation of the blade edge profile in the region R from the blade edge profile, resulting in a non-smooth profile curve;
FIG. 6 is a schematic view of an aircraft engine turbine guide vane provided by an embodiment of the present application;
FIG. 7 is a schematic illustration of an aircraft engine turbine guide vane eliminating blade body near end wall R region recession provided by an embodiment of the present application;
fig. 8 is a schematic diagram of an aircraft engine turbine guide vane elimination flange plate profile deviating from a theoretical profile, wherein no non-smooth profile curve is caused by the profile deviation.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solution of the present application and the advantages thereof more apparent, the technical solution of the present application will be more fully described in detail below with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application, which are for explanation of the present application, not for limitation of the present application. It should be noted that, for convenience of description, only the portion relevant to the present application is shown in the drawings, and other relevant portions may refer to a general design, and without conflict, the embodiments and technical features in the embodiments may be combined with each other to obtain new embodiments.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of this application should be given the ordinary meaning as understood by one of ordinary skill in the art to which this application belongs. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in this description are merely used to indicate relative directions or positional relationships, and do not imply that a device or element must have a particular orientation, be configured and operated in a particular orientation, and that the relative positional relationships may be changed when the absolute position of the object being described is changed, and thus should not be construed as limiting the present application. The terms "first," "second," "third," and the like, as used in the description herein, are used for descriptive purposes only and are not to be construed as indicating or implying any particular importance to the various components. The use of the terms "a," "an," or "the" and similar referents in the description of the invention are not to be construed as limited in number to the precise location of at least one. As used in this description, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term and that is listed after the term and its equivalents, without excluding other elements or articles.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description herein are to be construed broadly and refer to either a fixed connection, a removable connection, or an integral connection, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can also be communicated with the inside of two elements, and the specific meaning of the two elements can be understood by a person skilled in the art according to specific situations.
The present application is described in further detail below with reference to fig. 1-8.
An aeroengine turbine guide vane comprising: two edge plates, two installation heads and blade body, wherein:
the blade body is provided with three groups of large, medium and small areas, the medium area group meets the requirement of a design point, and the larger and smaller areas are used for meeting the group requirements of slip and performance in the actual test run process;
the two mounting heads can accommodate the blade bodies of three groups of large, medium and small areas, the contour lines are group movable block accommodating lines, the thickness of the contour lines is consistent with the distance delta between the group movable block accommodating lines, and the group movable block accommodating lines are designed to be capable of accommodating the minimum design of the variable group + -alpha DEG;
the blade root and the blade tip of the blade body can be respectively inserted into one mounting head and are connected with the mounting heads in a welding way;
the two edge plates are provided with holes;
each mounting head is correspondingly and integrally formed in one hole and integrally cast with the corresponding flange plate.
For the aeroengine turbine guide vane disclosed in the above embodiment, it can be understood by those skilled in the art that the design and installation head corresponds to the hole of the casting forming flange plate, so that the processing mould of the variable component can be effectively reduced, the production cost is reduced, the installation head is not arranged on the blade body, the welding position of the blade body and the flange plate is not affected by grinding, no matter how the blade body is dislocated in the radial direction during welding and assembling, no recess exists in the R region of the near end wall of the blade body, as shown in fig. 7, the damage risk of the vane influenced by horseshoe vortex can be effectively reduced, and in addition, for the low-consistency and large-chord length flange plate, even if the contour tolerance exists, the problem of unsmooth profile curve is avoided, and the aerodynamic performance of the vane is ensured, as shown in fig. 8.
For the aeroengine turbine guide vane disclosed by the embodiment, the person skilled in the art can also understand that the mounting head is designed to correspond to the hole middle flange plate of the casting forming flange plate to be close to the flat plate structure, the cooling structure is few in type, low in casting difficulty and high in qualification rate, the flange plate is large in structural size and is equivalent to the mounting head in size, the casting size cannot have magnitude difference, the casting structure size of the cooling channel in the blade body is the same magnitude, and the cooling channel is designed to be separated from the mounting head, so that the casting difficulty of the micro-size cooling blade body of the single crystal can be reduced, the growth of the single crystal of the blade body is facilitated, and the qualification rate of the integrity of the single crystal casting is improved.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred.
Having thus described the technical aspects of the present application with reference to the preferred embodiments illustrated in the accompanying drawings, it should be understood by those skilled in the art that the scope of the present application is not limited to the specific embodiments, and those skilled in the art may make equivalent changes or substitutions to the relevant technical features without departing from the principles of the present application, and those changes or substitutions will now fall within the scope of the present application.
Claims (3)
1. An aeroengine turbine guide vane, comprising: two edge plates, two installation heads and blade body, wherein:
the blade body is provided with three groups of large, medium and small areas;
the two mounting heads can accommodate three groups of blade bodies with large, medium and small areas;
the blade root and the blade tip of the blade body can be respectively inserted into one mounting head and are connected with the mounting heads in a welding way;
the two edge plates are provided with holes;
each mounting head is correspondingly and integrally formed in one forming hole.
2. The aircraft engine turbine guide vane of claim 1,
and the two mounting heads and the corresponding edge plates are integrally cast and formed.
3. The aircraft engine turbine guide vane of claim 1,
cooling channels are designed in the blade body.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211737708.6A CN116291756A (en) | 2022-12-31 | 2022-12-31 | Guide vane of turbine guide of aeroengine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211737708.6A CN116291756A (en) | 2022-12-31 | 2022-12-31 | Guide vane of turbine guide of aeroengine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116291756A true CN116291756A (en) | 2023-06-23 |
Family
ID=86789541
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211737708.6A Pending CN116291756A (en) | 2022-12-31 | 2022-12-31 | Guide vane of turbine guide of aeroengine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116291756A (en) |
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2022
- 2022-12-31 CN CN202211737708.6A patent/CN116291756A/en active Pending
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