CN116085316A - Interstage bleed air structure of air compressor in aero-engine and assembly method thereof - Google Patents
Interstage bleed air structure of air compressor in aero-engine and assembly method thereof Download PDFInfo
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- CN116085316A CN116085316A CN202310223874.2A CN202310223874A CN116085316A CN 116085316 A CN116085316 A CN 116085316A CN 202310223874 A CN202310223874 A CN 202310223874A CN 116085316 A CN116085316 A CN 116085316A
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/60—Mounting; Assembling; Disassembling
- F04D29/64—Mounting; Assembling; Disassembling of axial pumps
- F04D29/644—Mounting; Assembling; Disassembling of axial pumps especially adapted for elastic fluid pumps
- F04D29/646—Mounting or removal of fans
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- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The utility model relates to a compressor interstage bleed structure and assembly method in aeroengine, wherein, the structure includes N-1 level stator casing, N level rotor casing, N level stator casing N+1 level rotor casing, wherein, only be located N-1 level stator casing that the forefront is in the low pressure position is the structure of folio, the three box that the rest is located the rear end high pressure position is the full-circle structure, low in processing cost, be difficult for causing the air current leakage, can guarantee interstage bleed efficiency, and full-circle structural design, can guarantee its internal rotation, stator blade radial gap' S uniformity, can reduce the risk of bumping between rotor, stator blade, in addition, the design is in N level stator casing and the N level stator back annular installation limit that the bending extends backward forms annular clamping groove cooperation with N level rotor casing rear end and blocks N level stator blade S N The lugs at the two ends of the upper edge can be used for conveniently realizing the assembly of the structure on the basis of ensuring the whole ring structure.
Description
Technical Field
The application belongs to the technical field of air bleed structure design between air compressor stages in an aeroengine, and particularly relates to an air bleed structure between air compressor stages in an aeroengine and an assembly method thereof.
Background
The compressor is one of the core components of the aeroengine, not only is to provide high-temperature and high-pressure compressed air for the combustion chamber, but also is designed to provide heated or cooled air for other components of the aeroengine, such as turbine cooling air, heating air of an anti-icing system, etc., and a typical post-bleed air structure of an nth stage stator blade is shown in fig. 1, and comprises an nth-1 stage stator outer casing, an nth stage rotor inner casing and an nth stage rotor casing, wherein:
the front end of the outer casing of the N-1-level stator is internally provided with an annular clamping groove of the N-1-level stator for clamping the S-level stator blade S of the N-1-level stator N-1 Lugs at two ends of the upper edge, and the outer wall of the lugs is provided with air holes;
the inner-layer casing of the N-stage rotor is positioned in the rear end of the outer-layer casing of the N-1-stage stator, and the front end is coated with the N-stage rotor blade R N The rear end is internally provided with an N-stage stator annular clamping groove for clamping the N-stage stator blade S N The lugs at the two ends of the upper edge are provided with the front annular connecting edge of the N-th level rotor outside the front end and are connected in the outer casing of the N-1-th level stator through bolts to form a cantilever structure;
the N-stage rotor casing is coated with N-stage rotor blades, an annular air guide channel is formed at the front end and the rear end of the N-stage rotor inner-layer casing, an N-stage rotor rear annular connecting edge is arranged outside the rear end, and the N-stage rotor rear annular connecting edge is connected to the rear end of the N-1-stage stator outer-layer casing through bolts;
an annular air introducing cavity Q is formed among the N-1 level stator outer casing, the N level rotor inner casing and the N level rotor casing.
When the aeroengine works, air can be led from the main flow passage of the air compressor to the annular air-leading cavity Q through the annular air-leading passage, and then led out through the air-leading hole by utilizing a pipeline.
In the typical N-stage stator blade rear bleed air structure, in order to facilitate assembly and design of an N-1 stage stator outer-layer casing and an N-stage rotor inner-layer casing to be of a split structure, the following defects exist:
1) The N-1-level stator outer-layer casing and the N-level rotor inner-layer casing are of split structures, so that the processing cost is high, the assembly difficulty is high, the segmented stator blades are required to be cut in along a circumferential clamping groove, air flow leakage is easy to occur, and the interstage air-entraining efficiency is reduced;
2) The inner-layer casing of the Nth-stage rotor stator is of a cantilever structure, which is easy to cause the rotor blade R of the Nth-stage rotor N The radial clearance is different, which is not beneficial to controlling the radial clearance, and the N-th stage rotor blade R is added N Nth stage stator blade S N Risk of rubbing between them.
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 a compressor interstage bleed air structure in an aircraft engine and a method of assembling the same that overcomes or mitigates at least one of the known technical drawbacks.
The technical scheme of the application is as follows:
in one aspect, there is provided an interstage bleed air structure for a compressor in an aircraft engine, comprising:
the N-1-level stator casing is of a split structure and is internally provided with an N-1-level stator ring shapeA clamping groove for clamping the N-1 th stator blade S N-1 Lugs at two ends of the upper edge, and a rear annular connecting edge of the N-1-th stage stator casing is arranged outside the rear end;
the Nth-stage rotor casing is of a full-ring structure and is coated with the Nth-stage rotor blades R N The front end is provided with an N-th-stage rotor casing front annular connecting edge, and the rear end is provided with an N-th-stage stator front annular clamping groove;
the N-stage stator casing is of a full-ring structure, the outer wall of the N-stage stator casing is provided with a gas-guiding hole, the front end of the N-stage stator casing is provided with a front annular connecting edge, the rear end of the N-stage stator casing is provided with a rear annular connecting edge, and the front end of the N-stage stator casing is internally provided with a rear annular mounting edge; the front annular connecting edge of the N-stage stator casing, the front annular connecting edge of the N-stage rotor casing and the rear annular connecting edge of the N-1-stage stator casing are connected through bolts; the rear annular mounting edge of the N-th stator is bent and extends backwards, part of the mounting edge is abutted against the outer side of the N-th rotor casing, a rear annular clamping groove of the N-th stator is formed in the rear end of the mounting edge and is matched with the front annular clamping groove of the N-th rotor to clamp the S-th stator blades N Lugs at two ends of the upper edge;
the N+1st-stage rotor casing is of a full-ring structure and is coated with N+1st-stage rotor blades R N+1 An annular air guide channel is formed between the front end and the rear end of the annular installation edge of the Nth stage stator, and an annular connecting edge behind the (n+1) th stage rotor casing is arranged outside the rear end; the rear annular connecting edge of the N+1th-stage rotor casing is connected with the rear annular connecting edge of the N-stage stator casing through bolts;
the N-stage stator casing and the N-stage stator rear annular mounting edge, and the N+1-stage rotor casing rear annular connecting edge form an annular air-guiding cavity Q, and the annular air-guiding cavity Q is communicated with an air-guiding hole and an annular air-guiding channel.
According to at least one embodiment of the present application, in the above-mentioned interstage bleed air structure of the air compressor in the aeroengine, the rear annular connecting edge of the N-1-th stage stator casing and the front annular connecting edge of the N-th stage rotor casing are positioned at the seam allowance.
According to at least one embodiment of the application, in the interstage bleed air structure of the air compressor in the aeroengine, the rear annular connecting edge of the N-1-stage stator casing and the front annular connecting edge of the N-stage stator casing are positioned at the seam allowance.
According to at least one embodiment of the application, in the interstage bleed air structure of the air compressor in the aeroengine, the outer edge of the rear annular connecting edge of the N-1 stage stator is provided with an annular folded edge, and the annular folded edge is clamped at the peripheries of the front annular connecting edge of the N stage rotor and the front annular connecting edge of the N stage stator.
According to at least one embodiment of the present application, in the above-mentioned air bleed structure between compressor stages in an aeroengine, the annular connecting edge behind the n+1st stage rotor casing and the annular connecting edge inter-spigot behind the N stage stator casing are positioned.
According to at least one embodiment of the present application, in the above-mentioned air compressor interstage bleed air structure in an aeroengine, further comprises:
the inlet end of the air guide pipe is communicated with the air guide hole, and the outlet end of the air guide pipe is communicated with the part of the turbine and the pivot of the aircraft or the aeroengine, which need cooling or sealing air;
the middle part of the air bleed pipe is provided with a regulating valve, and the regulating valve controls the air bleed amount according to actual requirements by an external control system;
the air-inducing pipe and the corresponding air-inducing holes are provided with a plurality of positions along the circumferential direction.
In another aspect, a method for assembling an interstage bleed air structure of a compressor in an aircraft engine is provided, comprising:
the N-1 stage stator casing and the N-1 stage stator blade S N-1 Assembling;
the Nth-stage rotor case is dropped and assembled with the Nth-stage stator blade S N Assembling;
the Nth stage stator case is dropped and assembled with the Nth stage stator blade S N Assembling;
and (5) the n+1st-stage rotor case is dropped.
Drawings
FIG. 1 is a schematic illustration of a prior art aircraft engine compressor interstage bleed air configuration;
FIG. 2 is a schematic illustration of a compressor interstage bleed air configuration in an aircraft engine provided in an embodiment of the present application;
FIG. 3 is an N-1 th stator blade S of a compressor interstage bleed structure in an aircraft engine provided in an embodiment of the present application N-1 Assembly process schematicA figure;
FIG. 4 is a schematic illustration of an assembly process for a compressor interstage bleed air structure in an aircraft engine provided in an embodiment of the present application;
wherein:
1-N-1-th stage stator casing; 2-nth stage rotor case; 3-Nth stage stator case; 4 n+1st stage rotor case.
For better illustration of the present embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the actual product size, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present application.
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. 1-2.
In one aspect, there is provided an interstage bleed air structure for a compressor in an aircraft engine, comprising:
the N-1-level stator casing 1 is of a split structure and is internally provided with an N-1-level stator annular clamping groove for clamping the N-1-level stator blades S N-1 Lugs at two ends of the upper edge, and a rear annular connecting edge of the N-1-th stage stator casing is arranged outside the rear end;
the Nth-stage rotor casing 2 is of a full-ring structure and is coated with the Nth-stage rotor blades R N The front end is provided with an N-th-stage rotor casing front annular connecting edge, and the rear end is provided with an N-th-stage stator front annular clamping groove;
the N-stage stator casing 3 is of a full-ring structure, the outer wall of the N-stage stator casing is provided with a gas-guiding hole, the front end of the N-stage stator casing is provided with a front annular connecting edge, the rear end of the N-stage stator casing is provided with a rear annular connecting edge, and the front end of the N-stage stator casing is internally provided with a rear annular mounting edge; n-stage stator casing front annular connecting edge, N-stage rotor casing front annular connecting edge and N-1-stage stator casing rear annular connecting edge are communicatedConnecting through bolts; the rear annular mounting edge of the N-th stator is bent and extends backwards, part of the mounting edge is abutted against the outer side of the N-th rotor casing 2, a rear annular clamping groove of the N-th stator is formed in the rear end of the mounting edge and is matched with the front annular clamping groove of the N-th rotor to clamp the N-th stator blades S N Lugs at two ends of the upper edge;
the N+1st-stage rotor casing 4 is of a full-ring structure and is coated with the N+1st-stage rotor blade R N+1 An annular air guide channel is formed between the front end of the inner side and the rear end of the annular installation edge of the N-th stage stator, and the rear end is provided with an annular connecting edge behind the n+1-th stage rotor casing; the rear annular connecting edge of the N+1th-stage rotor casing is connected with the rear annular connecting edge of the N-stage stator casing through bolts;
the N-stage stator casing 3 and the N-stage stator rear annular mounting edge, and the N+1st stage rotor casing 4 and the N+1st stage rotor rear annular connecting edge form an annular air-guiding cavity Q, and the annular air-guiding cavity Q is communicated with an air-guiding hole and an annular air-guiding channel.
According to the interstage bleed air structure of the air compressor in the aeroengine disclosed by the embodiment, when the aeroengine works, the annular bleed air channel can bleed air from the main flow channel of the air compressor to the annular bleed air cavity Q, and then the bleed air is led out through the bleed air holes by utilizing pipelines, and the number and the distribution positions of the bleed air holes and the bleed air pipes of the bleed air holes can be designed according to actual requirements.
For the interstage bleed air structure of the air compressor in the aeroengine disclosed by the embodiment, as can be understood by those skilled in the art, the interstage bleed air structure comprises an N-1-th stator casing 1, an N-th rotor casing 2, an N-th stator casing 3 and an N+1-th rotor casing 4, wherein the N-1-th stator casing 1 which is only positioned at the foremost end and is positioned at the low pressure position is of a split structure, the three casings which are positioned at the rear end and are positioned at the high pressure position are of a whole ring structure, the processing cost is low, air leakage is not easy to be caused, the interstage bleed air efficiency can be ensured, the consistency of radial gaps of inner rotor blades and stator blades can be ensured by the whole ring structure design, the risk of collision between rotor blades and stator blades can be reduced, and in addition, the N-th stator rear annular mounting edge which is designed in the N-th stator casing 3 and extends backwards is matched with the annular clamping groove formed at the rear end of the N-th rotor casing 2 to clamp the N-th stator blades S N Lugs at both ends of the upper edge for securingOn the basis of the integral ring structure, the structure is convenient to assemble, the specific assembling process can refer to the assembling method of the interstage bleed air structure of the air compressor in the aero-engine disclosed by the application, and the annular mounting edge part of the nth stage stator is attached to the outer side of the nth stage rotor case 2, so that the overall stability of the structure can be enhanced.
In some alternative embodiments, in the above-mentioned interstage bleed air structure of the air compressor in the aeroengine, the rear annular connecting edge of the N-1-th stage stator casing and the front annular connecting edge of the N-th stage rotor casing are positioned at the seam allowance.
In some alternative embodiments, in the above-mentioned interstage bleed air structure of the air compressor in the aeroengine, the rear annular connecting edge of the N-1-th stage stator casing and the front annular connecting edge of the N-th stage stator casing are positioned at the seam allowance.
In some optional embodiments, in the above-mentioned interstage bleed air structure of the air compressor in the aeroengine, the outer edge of the N-1-stage stator rear annular connecting edge is provided with an annular folded edge, and the annular folded edge is clamped on the outer periphery of the N-stage rotor front annular connecting edge and the N-stage stator front annular connecting edge so as to be capable of positioning.
In some alternative embodiments, in the above-mentioned interstage bleed air structure of the air compressor in the aeroengine, the rear annular connecting edge of the n+1st stage rotor and the rear annular connecting edge inter-spigot of the N stage stator casing 3 are positioned.
In some alternative embodiments, the interstage bleed air structure of the compressor in the aeroengine further comprises:
the inlet end of the air guide pipe is communicated with the air guide hole, and the outlet end of the air guide pipe is communicated with the part of the turbine and the pivot of the aircraft or the aeroengine, which need cooling or sealing air;
the middle part of the air bleed pipe is provided with a regulating valve, and the regulating valve controls the air bleed amount according to actual requirements by an external control system;
the air-inducing pipe and the corresponding air-inducing holes are provided with a plurality of positions along the circumferential direction.
In another aspect, a method for assembling an interstage bleed air structure of a compressor in an aircraft engine is provided, comprising:
the N-1 stage stator case 1 and the N-1 stage stator blade S N-1 Assembly, see fig. 3;
the Nth stage rotor case 2 is dropped and assembled with the Nth stage stator blade S N Assembly, see fig. 4;
the Nth stage stator case 3 is dropped and assembled with the Nth stage stator blade S N Assembly, see fig. 4;
the n+1st stage rotor casing 4 is dropped, see fig. 4.
For the method for assembling the inter-stage bleed air structure of the air engine in the air engine disclosed in the above embodiment, which is used for realizing the corresponding assembly of the inter-stage bleed air structure of the air engine in the air engine disclosed in the above embodiment, the description is simpler, the specific relevant places can be referred to the relevant description of the inter-stage bleed air structure part of the air engine in the air engine, the technical effects of the relevant parts of the inter-stage bleed air structure of the air engine can also be referred to, and the description is omitted here.
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 (6)
1. An interstage bleed air structure of a compressor in an aircraft engine, comprising:
the N-1-level stator casing (1) is of a split structure and is internally provided with an N-1-level stator annular clamping groove for clamping the N-1-level stator blade S N-1 Lugs at two ends of the upper edge, and a rear annular connecting edge of the N-1-th stage stator casing is arranged outside the rear end;
the Nth-stage rotor casing (2) is of a full-ring structure and is coated with the Nth-stage rotor blade R N Front end is provided with a front annular connection of an N-th level rotor casingThe rear end of the edge is provided with an N-th-stage stator front annular clamping groove;
the N-stage stator casing (3) is of a full-ring structure, the outer wall of the N-stage stator casing is provided with a gas-guiding hole, the front end of the N-stage stator casing is provided with a front annular connecting edge, the rear end of the N-stage stator casing is provided with a rear annular connecting edge, and the front end of the N-stage stator casing is internally provided with a rear annular mounting edge; the front annular connecting edge of the N-stage stator casing, the front annular connecting edge of the N-stage rotor casing and the rear annular connecting edge of the N-1-stage stator casing are connected through bolts; the rear annular mounting edge of the N-th stator is bent and extends backwards, part of the mounting edge is attached to the outer side of the N-th rotor casing (2), a rear annular clamping groove of the N-th stator is formed in the rear end of the mounting edge and is matched with the front annular clamping groove of the N-th rotor casing to clamp the N-th stator blades S N Lugs at two ends of the upper edge;
the N+1st-stage rotor casing (4) is of a full-ring structure and is coated with the N+1st-stage rotor blade R N+1 An annular air guide channel is formed between the front end and the rear end of the annular installation edge of the Nth stage stator, and an annular connecting edge behind the (n+1) th stage rotor casing is arranged outside the rear end; the rear annular connecting edge of the N+1th-stage rotor casing is connected with the rear annular connecting edge of the N-stage stator casing through bolts;
the N-stage stator casing (3) and the N-stage stator rear annular mounting edge, and the N+1th stage rotor casing (4) and the N+1th stage rotor casing rear annular connecting edge form an annular air-guiding cavity Q, and the annular air-guiding cavity Q is communicated with an air-guiding hole and an annular air-guiding channel.
2. An aircraft engine mid-compressor interstage bleed air structure as claimed in claim 1, wherein,
the front end spigot of the N-1 level stator case (1) and the front end of the N level rotor case (2) are positioned.
3. An aircraft engine mid-compressor interstage bleed air structure as claimed in claim 1, wherein,
the outer edge of the rear annular connecting edge of the N-1 stage stator is provided with an annular folded edge which is clamped at the periphery of the front annular connecting edge of the N stage rotor and the front annular connecting edge of the N stage stator.
4. An aircraft engine mid-compressor interstage bleed air structure as claimed in claim 1, wherein,
the rear annular connecting edge of the N+1th-stage rotor casing and the seam allowance of the rear annular connecting edge of the N-stage stator casing are positioned.
5. An aircraft engine mid-compressor interstage bleed air structure as claimed in claim 1, wherein,
further comprises:
the inlet end of the air guide pipe is communicated with the air guide hole, and the outlet end of the air guide pipe is communicated with the part of the turbine and the pivot of the aircraft or the aeroengine, which need cooling or sealing air;
the middle part of the air bleed pipe is provided with a regulating valve, and the regulating valve controls the air bleed amount according to actual requirements by an external control system;
the air-inducing pipe and the corresponding air-inducing holes are provided with a plurality of positions along the circumferential direction.
6. A method of assembling an interstage bleed air structure of a compressor in an aircraft engine, comprising:
the N-1-th stage stator case (1) and the N-1-th stator blade S N-1 Assembling;
dropping the Nth stage rotor casing (2) and the Nth stage stator blade S N Assembling;
the Nth stage stator case (3) is dropped and assembled with the Nth stage stator blade S N Assembling;
and (5) the (4) th-level rotor case (1) is arranged in a falling mode.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310223874.2A CN116085316A (en) | 2023-03-09 | 2023-03-09 | Interstage bleed air structure of air compressor in aero-engine and assembly method thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310223874.2A CN116085316A (en) | 2023-03-09 | 2023-03-09 | Interstage bleed air structure of air compressor in aero-engine and assembly method thereof |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116085316A true CN116085316A (en) | 2023-05-09 |
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ID=86204655
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310223874.2A Pending CN116085316A (en) | 2023-03-09 | 2023-03-09 | Interstage bleed air structure of air compressor in aero-engine and assembly method thereof |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116085316A (en) |
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2023
- 2023-03-09 CN CN202310223874.2A patent/CN116085316A/en active Pending
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