CN117231364A - Aeroengine combustion chamber component and connection structure between aeroengine combustion chamber component and air compressor - Google Patents
Aeroengine combustion chamber component and connection structure between aeroengine combustion chamber component and air compressor Download PDFInfo
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- CN117231364A CN117231364A CN202311285702.4A CN202311285702A CN117231364A CN 117231364 A CN117231364 A CN 117231364A CN 202311285702 A CN202311285702 A CN 202311285702A CN 117231364 A CN117231364 A CN 117231364A
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- 238000002485 combustion reaction Methods 0.000 title claims abstract description 164
- 239000000463 material Substances 0.000 claims abstract description 6
- 238000007789 sealing Methods 0.000 claims description 50
- 239000000446 fuel Substances 0.000 claims description 6
- 238000002347 injection Methods 0.000 claims description 2
- 239000007924 injection Substances 0.000 claims description 2
- 239000000295 fuel oil Substances 0.000 description 3
- 239000000243 solution Substances 0.000 description 3
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000005540 biological transmission Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 229910000765 intermetallic Inorganic materials 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 239000003921 oil Substances 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
- 230000035945 sensitivity Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
- 229910006281 γ-TiAl Inorganic materials 0.000 description 1
Abstract
The application belongs to the technical field of design of a connection structure between an aeroengine combustion chamber component and a gas compressor, and particularly relates to the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the gas compressor, which comprises the following components: an outer wall of the combustion chamber; the combustion chamber inner wall is arranged in the combustion chamber outer wall; the final-stage stator component of the gas compressor is characterized in that the rear end of the outer wall of the final-stage stator component is connected with the front end of the outer wall of the combustion chamber through an annular connecting edge, and the rear end of the inner wall of the final-stage stator component is connected with the front end of the inner wall of the combustion chamber through a connecting edge; the diffuser is made of light high-temperature-resistant high-specific-strength low-plastic materials, is arranged in the outer wall of the combustion chamber, is positioned in an inlet of the outer wall of the combustion chamber, is connected with the outer wall of the combustion chamber or the inner wall of the combustion chamber through bolts by a connecting edge for supporting, the front end of the inner wall of the diffuser is connected with the front end of the inner wall of the final-stage stator assembly of the air compressor in a matched manner, and the front end of the outer wall of the diffuser is connected with the front end of the outer wall of the final-stage stator assembly of the air compressor in a matched manner.
Description
Technical Field
The application belongs to the technical field of design of connection structures between an aeroengine combustion chamber component and a gas compressor, and particularly relates to an aeroengine combustion chamber component and a connection structure between the aeroengine combustion chamber component and the gas compressor.
Background
Aiming at the requirements of high temperature resistance and light weight, at present, a diffuser in a combustion chamber of an aeroengine is mostly made of low-plasticity materials with light weight, high temperature resistance and high specific strength, such as gamma-TiAl intermetallic compounds and the like, and has poor weldability and strong notch sensitivity.
The present application has been made in view of the above-described 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 application, which is not necessarily prior art to the present application, and should not be used for evaluating the novelty and the inventive idea of the present application in the case where no clear evidence indicates that the above-mentioned content is already disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aircraft engine combustion chamber component and a connection structure between the component and a compressor, which overcomes or alleviates the technical drawbacks of at least one aspect of the known art.
The technical scheme of the application is as follows:
an aeroengine combustion chamber component and a connection structure between the aeroengine combustion chamber component and a compressor, comprising:
an outer wall of the combustion chamber;
the combustion chamber inner wall is arranged in the combustion chamber outer wall;
the final-stage stator component of the gas compressor is characterized in that the rear end of the outer wall of the final-stage stator component is connected with the front end of the outer wall of the combustion chamber through an annular connecting edge, and the rear end of the inner wall of the final-stage stator component is connected with the front end of the inner wall of the combustion chamber through a connecting edge;
the diffuser is made of light high-temperature-resistant high-specific-strength low-plastic materials, is arranged in the outer wall of the combustion chamber, is positioned in an inlet of the outer wall of the combustion chamber, is connected with the outer wall of the combustion chamber or the inner wall of the combustion chamber through bolts by a connecting edge for supporting, the front end of the inner wall of the diffuser is connected with the front end of the inner wall of the final-stage stator assembly of the air compressor in a matched manner, and the front end of the outer wall of the diffuser is connected with the front end of the outer wall of the final-stage stator assembly of the air compressor in a matched manner.
According to at least one embodiment of the present application, in the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the air compressor, the outer side of the rear end of the outer wall of the final stator assembly of the air compressor is welded with the inner side of the front end of the outer wall of the combustion chamber through an annular connection edge;
the outer side of the rear end of the inner wall of the final stator assembly of the air compressor is welded with the front end of the inner wall of the combustion chamber through a connecting edge;
the inner wall of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser is in small clearance fit with the spigot part of the rear end of the inner wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not more than 0.05mm for centering;
the front end of the outer wall of the diffuser is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser and the back end spigot part of the outer wall of the final-stage stator assembly of the compressor are internally provided with W-shaped elastic sealing rings, the W-shaped elastic sealing rings are precompressed along the axial direction so as to ensure the sealing of the front end of the outer wall of the diffuser and the back end spigot part of the outer wall of the final-stage stator assembly of the compressor, the precompressed amount of the W-shaped elastic sealing rings is not more than 15 percent of the working stroke, and gaps are formed between the front end of the inner wall of the diffuser and the outer wall of the diffuser and the back end of the inner wall and the outer wall of the final-stage stator assembly of the compressor in the axial direction, and the gaps are not more than 50 percent of the working stroke of the W-shaped elastic sealing rings;
the outer side of the rear end of the outer wall of the diffuser is connected with the inner side of the front end of the outer wall of the combustion chamber through a connecting edge by bolts, wherein the nut is a supporting plate nut which is connected with the connecting edge of the inner side of the front end of the outer wall of the combustion chamber, the corresponding bolt hole on the connecting edge of the outer side of the rear end of the outer wall of the diffuser is a runway-shaped hole, the long edge of the runway-shaped hole is parallel to the radius direction of the aeroengine, and the length of the long edge is at least 2mm greater than the pitch diameter of the corresponding bolt.
According to at least one embodiment of the present application, in the above-mentioned aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the compressor, the method further comprises:
the flame tube is arranged between the outer wall of the combustion chamber and the inner wall of the combustion chamber, the rear end of the flame tube is axially connected with the front end of the turbine guide in a floating manner, and the inner wall of the turbine guide is supported on the inner side of the rear end of the inner wall of the combustion chamber through a connecting edge by a bearing;
the front cover of the outer wall of the flame tube is connected to the front end of the outer wall of the flame tube;
the front cover of the inner wall of the flame tube is connected to the front end of the inner wall of the flame tube and is bent towards each other with the front cover of the outer wall of the flame tube;
the fuel oil swirl nozzles are arranged at the head part of the flame tube along the circumferential direction, and the fuel oil spray rods of the fuel oil swirl nozzles penetrate through the front cover of the outer wall of the flame tube and the outer wall of the combustion chamber and are connected to the outer wall of the combustion chamber by using the mounting seats;
the flame tube is provided with a plurality of ignition nozzles, a flame tube is provided with a plurality of flame tubes, a supporting rod of each flame tube penetrates through the outer wall of the combustion chamber, and the flame tubes are connected to the outer wall of the combustion chamber through mounting seats.
According to at least one embodiment of the present application, in the above-mentioned aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the compressor, there is an overlapping portion in a radial direction between the front end of the outer wall front cover of the flame tube and the front end of the outer wall of the flame tube;
an overlapping part exists between the rear end of the front cover of the inner wall of the flame tube and the front end of the inner wall of the flame tube in the radial direction;
the aero-engine combustion chamber component and the connecting structure between the aero-engine combustion chamber component and the air compressor further comprise:
the flame tube outer wall positioning rods penetrate through the combustion chamber outer wall, are arranged at the radial overlapping positions between the flame tube outer wall front cover rear end and the flame tube outer wall front end, are distributed along the circumferential direction, and are provided with positioning bushings at the radial overlapping positions between the flame tube outer wall front cover rear end and the flame tube outer wall front end;
the flame tube inner wall positioning rods penetrate through the combustion chamber inner wall, are arranged at the radial overlapping positions between the rear end of the flame tube inner wall front cover and the front end of the flame tube inner wall, are distributed along the circumferential direction, and are provided with positioning bushings at the radial overlapping positions between the flame tube inner wall front cover rear end and the flame tube inner wall front end.
According to at least one embodiment of the present application, in the above-mentioned aero-engine combustion chamber component and the connection structure between the combustion chamber component and the air compressor, a supporting seat surrounding each flame tube outer wall positioning rod is formed on the inner side of the combustion chamber outer wall, and the contour of the supporting seat along the axial backward direction of the aero-engine is a subsonic airfoil;
the inner side of the inner wall of the combustion chamber is provided with a supporting seat which surrounds the positioning rods of the outer wall of each flame tube, and the contour of the supporting seat along the axial backward direction of the aero-engine is subsonic wing type.
According to at least one embodiment of the present application, in the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the compressor, the diffuser is a multi-channel diffuser, a splitter ring is disposed between an outer wall and an inner wall of the diffuser, and the diffuser is supported by a plurality of support plates along a circumferential direction, wherein a rear edge of each support plate is provided with a notch, and the notch is clamped at a front edge of the splitter ring.
According to at least one embodiment of the present application, in the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the air compressor, the outer side of the rear end of the outer wall of the final stator assembly of the air compressor is welded with the inner side of the front end of the outer wall of the combustion chamber through an annular connection edge;
the outer side of the rear end of the inner wall of the final stator assembly of the air compressor is welded with the front end of the inner wall of the combustion chamber through a connecting edge;
the inner wall of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
w-shaped elastic sealing rings are arranged in the front ends of the inner wall and the outer wall of the diffuser and the front ends and the back ends of the inner wall and the outer wall of the final-stage stator assembly of the compressor, so that the sealing of the front ends of the inner wall and the outer wall of the diffuser and the front ends of the back ends of the inner wall and the outer wall of the final-stage stator assembly of the compressor is ensured, the precompression amount of the W-shaped elastic sealing rings is not more than 15% of the working stroke, and gaps are formed between the front ends of the inner wall and the outer wall of the diffuser and the back ends of the inner wall and the outer wall of the final-stage stator assembly of the compressor in the axial direction, and the gaps are not more than 50% of the working stroke of the W-shaped elastic sealing rings;
the diffuser outer wall rear end outside is connected with the combustion chamber outer wall front end inboard through the connection limit with the bolt, and the tang cooperation between this department connection limit, the fit clearance is not greater than 0.05mm in radial, and is centering, in addition, wherein, this department nut is the layer board nut, connect on the inboard connection edge of combustion chamber outer wall front end, and the corresponding bolt hole on diffuser outer wall rear end outside connection edge is runway type hole, the long limit in runway type hole is on a parallel with aeroengine radial direction, and long limit length is greater than corresponding bolt pitch diameter at least 2mm.
According to at least one embodiment of the present application, in the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the air compressor, the outer side of the rear end of the outer wall of the final stator assembly of the air compressor is welded with the inner side of the rear end of the outer wall of the combustion chamber through an annular connection edge;
the rear end of the inner wall of the final-stage stator assembly of the air compressor, the front end of the inner wall of the combustion chamber and the rear end of the inner wall of the diffuser are connected through bolts, the rear end of the outer annular connecting edge of the inner wall of the diffuser is matched with a spigot between the rear end of the inner wall of the final-stage stator assembly of the air compressor and the front end of the inner wall of the combustion chamber, the maximum interference of the fit clearance in the radial direction is not more than 0.1mm, the maximum clearance is not more than 0.025mm, and centering is carried out;
the front end of the inner wall of the diffuser is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final stator component of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front ends of the inner wall and the outer wall of the diffuser and the front and rear ends of the inner wall and the outer wall of the final-stage stator assembly of the compressor are respectively internally provided with a W-shaped elastic sealing ring, and the W-shaped elastic sealing rings are precompressed along the axial direction so as to ensure the sealing of the front ends of the inner wall and the outer wall of the diffuser and the front ends of the inner wall and the rear ends of the outer wall of the final-stage stator assembly of the compressor, wherein the precompressed amount of the W-shaped elastic sealing rings is not more than 15% of the working stroke, and gaps are formed between the front ends of the inner wall and the outer wall of the diffuser and the rear ends of the inner wall of the final-stage stator assembly of the compressor in the axial direction, and the gaps are not more than 50% of the working stroke of the W-shaped elastic sealing rings.
Drawings
FIGS. 1-2 are schematic diagrams of an aero-engine combustor component and an example first connection structure between the aero-engine combustor component and a compressor according to an embodiment of the present application;
FIG. 3 is a schematic diagram of an aero-engine combustor component and an example two of a connection structure between the aero-engine combustor component and a compressor provided by an embodiment of the application;
FIG. 4 is a schematic view of an aero-engine combustor component and an example three of a connection structure between the aero-engine combustor component and a compressor provided by an embodiment of the application;
wherein:
1-the outer wall of the combustion chamber; 2-the inner wall of the combustion chamber; 3-a compressor final stage stator assembly; 4-diffuser; 5-W-type elastic sealing ring; 6-a flame tube; 7-turbine guide; 8-a front cover of the outer wall of the flame tube; 9-a front cover of the inner wall of the flame tube; 10-a fuel swirl nozzle; 11-ignition electric nozzle; 12-a flame tube outer wall positioning rod; 13-a flame tube inner wall positioning rod.
For the purpose of better illustrating the embodiments, certain elements of the drawings are omitted, enlarged or reduced in size 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 application.
Detailed Description
In order to make the technical solution of the present application and its advantages more clear, the technical solution of the present application will be further and completely described in detail 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 and not for limitation of the present application. It should be noted that, for convenience of description, only the part related to the present application is shown in the drawings, and other related parts may refer to the general design, and the embodiments of the present application and the technical features of the embodiments may be combined with each other to obtain new embodiments without conflict.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of the application should be given the ordinary meaning as understood by one of ordinary skill in the art to which the application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in the description of the present application are merely used for indicating relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and that the relative positional relationships may be changed when the absolute position of the object to be described is changed, thus not being construed as limiting the application. The terms "first," "second," "third," and the like, as used in the description of the present application, 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 application are not to be construed as limiting the amount absolutely, but rather as existence of at least one. As used in this description of the application, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term as such, but does not exclude other elements or articles from the list of elements or articles that appear after the term.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description of the present application are used in a broad sense, and for example, the connection may be a fixed connection, a removable connection, or an integral connection; 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 application is described in further detail below with reference to fig. 1 to 4.
An aeroengine combustion chamber component and a connection structure between the aeroengine combustion chamber component and a compressor, comprising:
a combustion chamber outer wall 1;
a combustion chamber inner wall 2 provided in the combustion chamber outer wall 1;
the rear end of the outer wall of the final stator component 3 of the air compressor is connected with the front end of the outer wall 1 of the combustion chamber through an annular connecting edge, and the rear end of the inner wall of the final stator component is connected with the front end of the inner wall 2 of the combustion chamber through a connecting edge;
the diffuser 4 is made of a low-plastic material with light weight, high temperature resistance and high specific strength, is arranged in the outer wall 1 of the combustion chamber, is positioned in the inlet of the outer wall 1 of the combustion chamber, is supported by being connected with the outer wall 1 of the combustion chamber or the inner wall 2 of the combustion chamber through a connecting edge through bolts, is in matched connection with a spigot between the front end of the inner wall and the rear end of the inner wall of the final stator assembly 3 of the compressor, and is in matched connection with a spigot between the front end of the outer wall and the rear end of the outer wall of the final stator assembly 3 of the compressor.
For the aeroengine combustion chamber component and the connection structure between the aeroengine combustion chamber component and the air compressor disclosed by the embodiment, those skilled in the art can understand that the design is that under the condition that the diffuser 4 is made of a light high-temperature resistant high-strength low-plasticity material, the connection between the combustion chamber outer wall 1 and the combustion chamber inner wall 2 is realized through the stator of the air compressor final-stage stator assembly 3, the transmission between the axial forces is carried out, the diffuser 4 is designed to be supported by the connection edge through the bolt connection of the combustion chamber outer wall 1 or the combustion chamber inner wall 2, the front end of the inner wall of the diffuser 4 is designed to be connected with the spigot between the rear end of the inner wall of the air compressor final-stage stator assembly 3 in a matched manner, and the front end of the outer wall is connected with the spigot between the rear end of the outer wall of the air compressor final-stage stator assembly 3 in a matched manner, so that the positioning is realized, and the direct welding connection between the diffuser 4 and the combustion chamber outer wall 1 and the combustion chamber inner wall 2 can be avoided, and the damage at the stress concentration part and the surface defect part can be avoided.
In the aeroengine combustion chamber component and the connecting structure between the aeroengine combustion chamber component and the air compressor disclosed in the embodiment, in the first example, as shown in fig. 1, the outer side of the rear end of the outer wall of the air compressor final-stage stator component 3 is welded and connected with the inner side of the front end of the outer wall 1 of the combustion chamber through an annular connecting edge;
the outer side of the rear end of the inner wall of the final stator assembly 3 of the air compressor is welded with the front end of the inner wall 2 of the combustion chamber through a connecting edge;
the inner wall 2 of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser 4 is in small clearance fit with the seam allowance part of the rear end of the inner wall of the final stator assembly 3 of the compressor in the radial direction, and the maximum fit clearance of the cold state is not more than 0.05mm, so that centering is performed;
the front end of the outer wall of the diffuser 4 is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final-stage stator assembly 3 of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm, so that the radial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor can be compensated;
the front end of the outer wall of the diffuser 4 and the seam allowance part of the rear end of the outer wall of the final-stage stator assembly 3 of the compressor are internally provided with W-shaped elastic sealing rings 5,W type elastic sealing rings 5 which are precompressed along the axial direction so as to ensure that the precompressed amount of the W-shaped elastic sealing rings 5 is not more than 15% of the working stroke, and gaps are formed between the front end of the inner wall of the diffuser 4 and the rear end of the outer wall of the final-stage stator assembly 3 of the compressor and the rear end of the inner wall of the outer wall of the final-stage stator assembly 3 of the compressor in the axial direction so as to compensate the axial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor, and the gaps at the positions are not more than 50% of the working stroke of the W-shaped elastic sealing rings 5 can be designed so as to ensure that the compressed amount of the W-shaped elastic sealing rings 5 is not more than 50% of the working stroke and avoid damaging the W-shaped elastic sealing rings 5;
the outer side of the rear end of the outer wall of the diffuser 4 is connected with the inner side of the front end of the outer wall 1 of the combustion chamber through a connecting edge by bolts, wherein the nuts are supporting plate nuts which are connected to the connecting edge of the inner side of the front end of the outer wall 1 of the combustion chamber, the corresponding bolt holes on the connecting edge of the outer side of the rear end of the outer wall of the diffuser 4 are runway-shaped holes, the long edges of the runway-shaped holes are parallel to the radial direction of an aeroengine, so that the stress concentration of the connecting edge of the outer side of the rear end of the outer wall of the diffuser 4 at the bolt holes can be reduced, the length of the long edges is larger than the diameter of the corresponding bolts by at least 2mm, and a margin can be reserved for radial deformation of the diffuser 4.
Example one disclosed aircraft engine combustor component and its connection structure with a compressor, as shown in fig. 1, further includes:
the flame tube 6 is arranged between the outer wall 1 of the combustion chamber and the inner wall 2 of the combustion chamber, the rear end of the flame tube is axially connected with the front end of the turbine guide 7 in a floating manner, and the inner wall of the turbine guide 7 is supported on the inner side of the rear end of the inner wall 2 of the combustion chamber through a connecting edge by a pre-rotation nozzle;
the front cover 8 of the outer wall of the flame tube is connected to the front end of the outer wall of the flame tube 6;
the front cover 9 of the inner wall of the flame tube is connected to the front end of the inner wall of the flame tube 6, and is bent towards each other with the front cover 8 of the outer wall of the flame tube;
the fuel swirl nozzles 10 are circumferentially arranged at the head of the flame tube 6, and oil injection rods of the fuel swirl nozzles penetrate through the front cover 8 of the outer wall of the flame tube and the outer wall 1 of the combustion chamber and are connected to the outer wall 1 of the combustion chamber by using mounting seats;
the plurality of ignition nozzles 11 are circumferentially arranged on the flame tube 6, and their struts are arranged so as to penetrate the combustion chamber outer wall 1, and are connected to the combustion chamber outer wall 1 by means of mounting seats.
In the first disclosed aero-engine combustion chamber component and the connection structure between the component and the air compressor, as shown in fig. 2, an overlapping part exists between the rear end of the front cover 8 of the outer wall of the flame tube and the front end of the outer wall of the flame tube 6 in the radial direction;
an overlapping part exists between the rear end of the front cover 9 of the inner wall of the flame tube and the front end of the inner wall of the flame tube 6 in the radial direction;
the aero-engine combustion chamber component and the connecting structure between the aero-engine combustion chamber component and the air compressor further comprise:
the flame tube outer wall positioning rods 12 penetrate through the combustion chamber outer wall 1, are arranged at the radial overlapping positions between the rear end of the flame tube outer wall front cover 8 and the front end of the flame tube 6 outer wall, are distributed along the circumferential direction, and are provided with positioning bushings at the radial overlapping positions between the rear end of the flame tube outer wall front cover 8 and the front end of the flame tube 6 outer wall;
the flame tube inner wall positioning rods 13 penetrate through the combustion chamber inner wall 2, are arranged at the radial overlapping positions between the rear end of the flame tube inner wall front cover 9 and the front end of the flame tube inner wall 6, are distributed along the circumferential direction, are arranged at the radial overlapping positions between the rear end of the flame tube inner wall front cover 9 and the front end of the flame tube inner wall 6, so that the flame tube 6 can be positioned at multiple points on the combustion chamber outer wall 1 and the combustion chamber inner wall 2 by utilizing the flame tube outer wall positioning rods 12 and the flame tube inner wall positioning rods 13, and certain radial deformation is allowed.
In an exemplary disclosed combustion chamber component of an aero-engine and a connection structure between the combustion chamber component and a gas compressor, as shown in fig. 2, a supporting seat surrounding each flame tube outer wall positioning rod 12 is formed on the inner side of the combustion chamber outer wall 1, and the supporting seat is provided with subsonic airfoils along the axial backward outline of the aero-engine so as to reduce the pressure loss of passing gas flow;
the inner side of the inner wall 2 of the combustion chamber is provided with a supporting seat which surrounds each flame tube outer wall positioning rod 13, and the supporting seat is provided with subsonic wing profiles along the axial backward outline of the aero-engine so as to reduce the pressure loss of the passing air flow.
In the second example, as shown in fig. 3, the outer side of the rear end of the outer wall of the final stator assembly 3 of the compressor is welded and connected with the inner side of the front end of the outer wall 1 of the combustion chamber through an annular connecting edge;
the outer side of the rear end of the inner wall of the final stator assembly 3 of the air compressor is welded with the front end of the inner wall 2 of the combustion chamber through a connecting edge;
the inner wall 2 of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser 4 is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final-stage stator assembly 3 of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm, so that the radial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor can be compensated;
the front end of the outer wall of the diffuser 4 is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final-stage stator assembly 3 of the compressor in the radial direction, and the cold state maximum fit clearance is not less than 1mm so as to compensate the radial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor;
the front ends of the inner wall and the outer wall of the diffuser 4 and the front ends of the inner wall and the rear ends of the outer wall of the final-stage stator assembly 3 of the compressor are internally provided with W-shaped elastic sealing rings 5,W-shaped elastic sealing rings 5 which are precompressed along the axial direction so as to ensure the sealing of the front ends of the inner wall and the outer wall of the diffuser 4 and the front ends of the inner wall and the rear ends of the outer wall of the final-stage stator assembly 3 of the compressor, the precompressed amount of the W-shaped elastic sealing rings 5 is not more than 15% of the working stroke, and gaps exist between the front ends of the inner wall and the outer wall of the diffuser 4 and the rear ends of the inner wall and the outer wall of the final-stage stator assembly 3 of the compressor in the axial direction so as to compensate the axial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor, and the gaps at the positions can be designed to be not more than 50% of the working stroke of the W-shaped elastic sealing rings 5 so as to ensure the compression amount of the W-shaped elastic sealing rings 5 not to exceed 50% of the working stroke and avoid damage to the W-shaped elastic sealing rings 5;
the outer side of the rear end of the outer wall of the diffuser 4 is connected with the inner side of the front end of the outer wall 1 of the combustion chamber through connecting edges by bolts, the seam allowance is matched between the connecting edges, the fit clearance is not more than 0.05mm in radial direction, centering is carried out, in addition, the nut is a supporting plate nut which is connected with the connecting edges of the inner side of the front end of the outer wall 1 of the combustion chamber, the corresponding bolt holes on the connecting edges of the outer side of the rear end of the outer wall of the diffuser 4 are racetrack-shaped holes, the long edges of the racetrack-shaped holes are parallel to the radial direction of the aeroengine, so that stress concentration of the connecting edges of the outer side of the rear end of the outer wall of the diffuser 4 at the bolt holes can be reduced, and the length of the long edges is at least 2mm larger than the middles of the corresponding bolts, so that allowance can be reserved for radial deformation of the diffuser 4. .
In the third embodiment, as shown in fig. 4, the outer side of the rear end of the outer wall of the final stator assembly 3 of the compressor is welded and connected with the inner side of the rear end of the outer wall 1 of the combustion chamber through an annular connecting edge;
the rear end of the inner wall of the final-stage stator assembly 3 of the air compressor, the front end of the inner wall 2 of the combustion chamber and the rear end of the inner wall of the diffuser 4 are connected through an outward annular connecting edge by bolts, the outward annular connecting edge of the rear end of the inner wall of the diffuser 4 is matched with a spigot between the rear end of the inner wall of the final-stage stator assembly 3 of the air compressor and the outward annular connecting edge of the front end of the inner wall 2 of the combustion chamber, the maximum interference of the fit clearance in the radial direction is not more than 0.1mm, the maximum clearance is not more than 0.025mm, and centering is carried out;
the front end of the inner wall of the diffuser 4 is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final-stage stator assembly 3 of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm, so that the radial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor can be compensated;
the front end of the outer wall of the diffuser 4 is in large clearance fit with the seam allowance part of the rear end of the outer wall of the final-stage stator assembly 3 of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm, so that the radial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor can be compensated;
the front ends of the inner wall and the outer wall of the diffuser 4 and the front ends of the inner wall and the rear ends of the inner wall and the outer wall of the final-stage stator assembly 3 of the compressor are respectively provided with a W-shaped elastic sealing ring 5,W type elastic sealing ring 5 which is precompressed along the axial direction so as to ensure that the sealing of the front ends of the inner wall and the outer wall of the diffuser 4 and the front ends of the rear ends of the inner wall and the outer wall of the final-stage stator assembly 3 of the compressor is ensured, the precompressed amount of the W-shaped elastic sealing ring 5 is not more than 15% of the working stroke, and gaps exist between the front ends of the inner wall and the outer wall of the diffuser 4 and the rear ends of the inner wall and the outer wall of the final-stage stator assembly 3 of the compressor in the axial direction so as to compensate the axial deformation between the diffuser 4 and the final-stage stator assembly 3 of the compressor, and the gaps of the W-shaped elastic sealing ring 5 can be designed to be not more than 50% of the working stroke so as to ensure that the compression amount of the W-shaped elastic sealing ring 5 is not more than 50% of the working stroke, and damage to the W-shaped elastic sealing ring 5 is avoided.
In some alternative embodiments, in the above-mentioned aeroengine combustion chamber component and the connection structure between the component and the compressor, the diffuser 4 is a multi-channel diffuser, a splitter ring is disposed between an outer wall and an inner wall of the diffuser, and a plurality of support plates are supported along the circumferential direction, where a notch is formed on a trailing edge of each support plate and is clamped on a leading edge of the splitter ring.
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 shown in the 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 related technical features without departing from the principle of the present application, and those changes or substitutions will fall within the scope of the present application.
Claims (8)
1. An aeroengine combustion chamber component and a connection structure between the aeroengine combustion chamber component and a compressor, which is characterized by comprising:
a combustion chamber outer wall (1);
a combustion chamber inner wall (2) arranged in the combustion chamber outer wall (1);
the rear end of the outer wall of the final stator component (3) of the air compressor is connected with the front end of the outer wall (1) of the combustion chamber through an annular connecting edge, and the rear end of the inner wall of the final stator component is connected with the front end of the inner wall (2) of the combustion chamber through a connecting edge;
the diffuser (4) is made of light high-temperature-resistant high-strength low-plasticity materials, is arranged in the outer wall (1) of the combustion chamber, is positioned in the inlet of the outer wall (1) of the combustion chamber, is connected with the outer wall (1) of the combustion chamber or the inner wall (2) of the combustion chamber through bolts by a connecting edge to support, is matched and connected with a spigot between the front end of the inner wall and the rear end of the inner wall of the final stator assembly (3) of the compressor, and is matched and connected with a spigot between the front end of the outer wall and the rear end of the outer wall of the final stator assembly (3) of the compressor.
2. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 1,
the outer side of the rear end of the outer wall of the final stator component (3) of the air compressor is welded with the inner side of the front end of the outer wall (1) of the combustion chamber through an annular connecting edge;
the outer side of the rear end of the inner wall of the final stator component (3) of the air compressor is welded with the front end of the inner wall (2) of the combustion chamber through a connecting edge;
the inner wall (2) of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser (4) is in small clearance fit with the seam allowance part of the rear end of the inner wall of the final stator assembly (3) of the compressor in the radial direction, and the cold state maximum fit clearance is not more than 0.05mm for centering;
the front end of the outer wall of the diffuser (4) is in large clearance fit with the spigot part of the rear end of the outer wall of the final stator assembly (3) of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser (4) and the back end spigot part of the outer wall of the final-stage stator assembly (3) of the compressor are internally provided with W-shaped elastic sealing rings (5), the W-shaped elastic sealing rings (5) are precompressed along the axial direction to ensure the sealing of the front end of the outer wall of the diffuser (4) and the back end spigot part of the outer wall of the final-stage stator assembly (3) of the compressor, the precompressed amount of the W-shaped elastic sealing rings (5) is not more than 15% of the working stroke, and gaps exist between the front end of the inner wall and the outer wall of the diffuser (4) and the back ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor in the axial direction, and the gaps are not more than 50% of the working stroke of the W-shaped elastic sealing rings (5);
the outer side of the rear end of the outer wall of the diffuser (4) is connected with the inner side of the front end of the outer wall (1) of the combustion chamber through a connecting edge by bolts, wherein the nuts are supporting plate nuts which are connected to the connecting edge of the inner side of the front end of the outer wall (1) of the combustion chamber, the corresponding bolt holes on the connecting edge of the outer side of the rear end of the outer wall of the diffuser (4) are runway-shaped holes, the long sides of the runway-shaped holes are parallel to the radius direction of the aeroengine, and the length of the long sides is at least 2mm greater than the middles of the corresponding bolts.
3. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 2,
further comprises:
the flame tube (6) is arranged between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber, the rear end of the flame tube is axially connected with the front end of the turbine guide (7) in a floating mode, and the inner wall of the turbine guide (7) is supported on the inner side of the rear end of the inner wall (2) of the combustion chamber through a connecting edge by a bearing;
a front cover (8) of the outer wall of the flame tube is connected to the front end of the outer wall of the flame tube (6);
a front cover (9) of the inner wall of the flame tube is connected to the front end of the inner wall of the flame tube (6) and is bent in opposite directions with the front cover (8) of the outer wall of the flame tube;
the fuel swirl nozzles (10) are arranged at the head part of the flame tube (6) along the circumferential direction, and the fuel injection rods of the fuel swirl nozzles penetrate through the front cover (8) of the outer wall of the flame tube and the outer wall (1) of the combustion chamber and are connected to the outer wall (1) of the combustion chamber by using the mounting seats;
the plurality of ignition nozzles (11) are circumferentially arranged on the flame tube (6), and the supporting rods of the ignition nozzles penetrate through the outer wall (1) of the combustion chamber and are connected to the outer wall (1) of the combustion chamber by using the mounting seats.
4. An aircraft engine combustor component and connection with a compressor thereof according to claim 3,
an overlapping part exists between the rear end of the front cover (8) of the outer wall of the flame tube and the front end of the outer wall of the flame tube (6) in the radial direction;
an overlapping part exists between the rear end of the front cover (9) of the inner wall of the flame tube and the front end of the inner wall of the flame tube (6) in the radial direction;
the aero-engine combustion chamber component and the connecting structure between the aero-engine combustion chamber component and the air compressor further comprise:
a plurality of flame tube outer wall positioning rods (12) penetrate through the combustion chamber outer wall (1), are arranged at radial overlapping positions between the rear end of the flame tube outer wall front cover (8) and the flame tube (6) outer wall front end, are distributed along the circumferential direction, and are provided with positioning bushings at radial overlapping positions between the flame tube outer wall front cover (8) rear end and the flame tube (6) outer wall front end;
the flame tube inner wall positioning rods (13) penetrate through the combustion chamber inner wall (2), are arranged at radial overlapping positions between the rear end of the flame tube inner wall front cover (9) and the front end of the flame tube inner wall (6), are distributed along the circumferential direction, and are provided with positioning bushings at radial overlapping positions between the flame tube inner wall front cover (9) rear end and the flame tube inner wall front end.
5. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 4,
a supporting seat surrounding each flame tube outer wall positioning rod (12) is formed on the inner side of the combustion chamber outer wall (1), and the contour of the supporting seat along the axial backward direction of the aero-engine is a subsonic airfoil;
the inner side of the inner wall (2) of the combustion chamber is provided with a supporting seat which surrounds each flame tube outer wall positioning rod (13), and the contour of the supporting seat along the axial backward direction of the aero-engine is subsonic wing type.
6. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 1,
the diffuser (4) is a multichannel diffuser, a flow distribution ring is arranged between the outer wall and the inner wall of the multichannel diffuser, and a plurality of support plates are supported along the circumferential direction, wherein the rear edge of each support plate is provided with a notch, and the notch is clamped at the front edge of the flow distribution ring.
7. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 1,
the outer side of the rear end of the outer wall of the final stator component (3) of the air compressor is welded with the inner side of the front end of the outer wall (1) of the combustion chamber through an annular connecting edge;
the outer side of the rear end of the inner wall of the final stator component (3) of the air compressor is welded with the front end of the inner wall (2) of the combustion chamber through a connecting edge;
the inner wall (2) of the combustion chamber is divided into a front section and a rear section, and the two sections are connected through an outward annular connecting edge by bolts;
the front end of the inner wall of the diffuser (4) is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final stator assembly (3) of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser (4) is in large clearance fit with the spigot part of the rear end of the outer wall of the final stator assembly (3) of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front ends of the inner wall and the outer wall of the diffuser (4) and the front ends and the rear ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor are internally provided with W-shaped elastic sealing rings (5), the W-shaped elastic sealing rings (5) are precompressed along the axial direction so as to ensure the sealing of the front ends of the inner wall and the outer wall of the diffuser (4) and the front ends and the rear ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor, the precompressed amount of the W-shaped elastic sealing rings (5) is not more than 15% of the working stroke, and gaps are formed between the front ends of the inner wall and the outer wall of the diffuser (4) and the rear ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor in the axial direction and are not more than 50% of the working stroke of the W-shaped elastic sealing rings (5);
the diffuser (4) outer wall rear end outside and the combustion chamber outer wall (1) front end inboard between, connect with the bolt through the limit of connecting, the tang cooperation between this department is connected the limit, the fit clearance is not greater than 0.05mm in radial, centering is carried out, in addition, wherein, this department nut is the layer board nut, connect on the inboard connection edge of combustion chamber outer wall (1) front end, and the corresponding bolt hole on diffuser (4) outer wall rear end outside connection edge is runway type hole, the long limit in runway type hole is on a parallel with aeroengine radial direction, and long limit length is greater than corresponding bolt pitch diameter at least 2mm.
8. The aircraft engine combustor component and connection structure with a compressor thereof according to claim 1,
the outer side of the rear end of the outer wall of the final stator assembly (3) of the air compressor is welded with the inner side of the rear end of the outer wall (1) of the combustion chamber through an annular connecting edge;
the rear end of the inner wall of the final-stage stator assembly (3) of the air compressor, the front end of the inner wall (2) of the combustion chamber and the rear end of the inner wall of the diffuser (4) are connected through bolts through outward annular connecting edges, the outward annular connecting edges of the rear end of the inner wall of the diffuser (4) are matched with the rabbets between the rear end of the inner wall of the final-stage stator assembly (3) of the air compressor and the outward annular connecting edges of the front end of the inner wall (2) of the combustion chamber, the maximum interference of the fit clearance in the radial direction is not more than 0.1mm, and the maximum clearance is not more than 0.025mm, so that centering is carried out;
the front end of the inner wall of the diffuser (4) is in large clearance fit with the seam allowance part of the rear end of the inner wall of the final stator assembly (3) of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front end of the outer wall of the diffuser (4) is in large clearance fit with the spigot part of the rear end of the outer wall of the final stator assembly (3) of the compressor in the radial direction, and the maximum fit clearance of the cold state is not less than 1mm;
the front ends of the inner wall and the outer wall of the diffuser (4) and the front ends and the rear ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor are internally provided with W-shaped elastic sealing rings (5), the W-shaped elastic sealing rings (5) are precompressed along the axial direction so as to ensure the sealing of the front ends and the rear ends of the inner wall and the outer wall of the diffuser (4) and the front ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor, the precompressed amount of the W-shaped elastic sealing rings (5) is not more than 15% of the working stroke, and gaps are reserved between the front ends of the inner wall and the outer wall of the diffuser (4) and the rear ends of the inner wall and the outer wall of the final-stage stator assembly (3) of the compressor in the axial direction, and are not more than 50% of the working stroke of the W-shaped elastic sealing rings (5).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311285702.4A CN117231364A (en) | 2023-10-07 | 2023-10-07 | Aeroengine combustion chamber component and connection structure between aeroengine combustion chamber component and air compressor |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311285702.4A CN117231364A (en) | 2023-10-07 | 2023-10-07 | Aeroengine combustion chamber component and connection structure between aeroengine combustion chamber component and air compressor |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117231364A true CN117231364A (en) | 2023-12-15 |
Family
ID=89084168
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311285702.4A Pending CN117231364A (en) | 2023-10-07 | 2023-10-07 | Aeroengine combustion chamber component and connection structure between aeroengine combustion chamber component and air compressor |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117231364A (en) |
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2023
- 2023-10-07 CN CN202311285702.4A patent/CN117231364A/en active Pending
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