CN108180171A - Aero-engine high-pressure compressor shroud chamber bleed structure - Google Patents
Aero-engine high-pressure compressor shroud chamber bleed structure Download PDFInfo
- Publication number
- CN108180171A CN108180171A CN201611121820.1A CN201611121820A CN108180171A CN 108180171 A CN108180171 A CN 108180171A CN 201611121820 A CN201611121820 A CN 201611121820A CN 108180171 A CN108180171 A CN 108180171A
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- CN
- China
- Prior art keywords
- bleed
- pressure compressor
- circumferential
- compressor shroud
- shroud chamber
- 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.)
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Classifications
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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
- F04D29/542—Bladed diffusers
Abstract
The purpose of the present invention is to provide a kind of aero-engine high-pressure compressor shroud chamber bleed structures, and simple in structure, bleed pitot loss is low.Aero-engine high-pressure compressor shroud chamber bleed structure according to the present invention, it includes the circumferential bleed slot that both high-pressure compressor import goose neck segment trailer had been set on front side of high-pressure compressor inlet guide vane and the fixation deflector set in the circumferential bleed slot, the fixed deflector makes the circulation area in the circumferential bleed slot from sprue side to compressor shroud chamber side gradually amplify, and the circumferential direction bleed slot is used for and high-pressure compressor shroud chamber bleed air system.
Description
Technical field
The present invention relates to aero-engine high-pressure compressor shroud chamber bleed structures.
Background technology
With the proposition of " green aviation " concept, the flight economy of modern Commercial aero-engine, noise, oil consumption rate and
Life Cycle Cost is increasingly paid attention to, and the efficiency index of aero-engine high-pressure compressor is continuously improved, and requires simultaneously
Complex degree of structure is low, weight is low, maintenance cost is low.As shown in Figure 1, in general aero-engine type, before high-pressure compressor
Disk chamber bleed be from bleed after first order stator blade, successively by air entraining pipe, Middle casing hollow support plate, air collecting chamber and its with
The connecting tube that high-pressure compressor shroud chamber communicates finally just reaches high-pressure compressor shroud chamber, causes bleed flow path long in this way,
Bleed is caused to be lost, and higher, complete machine structure is complicated, weight is larger, and maintenance cost is higher.
Invention content
The purpose of the present invention is to provide a kind of aero-engine high-pressure compressor shroud chamber bleed structure, structure letters
Single, bleed pitot loss is low.
Aero-engine high-pressure compressor shroud chamber bleed structure according to the present invention, leads including high-pressure compressor import
It flows and is set in the circumferential bleed slot and the circumferential bleed slot that both high-pressure compressor import goose neck segment trailer had been set on front side of blade
Fixation deflector, the fixed deflector makes the circulation from sprue side to compressor shroud chamber side in the circumferential bleed slot
Area gradually amplifies, and the circumferential direction bleed slot is used for and high-pressure compressor shroud chamber bleed air system.
In one embodiment, the fixed deflector has the first peripheral side and the second peripheral side, first peripheral side
The state radially expanded outwardly between second peripheral side so that the sprue side of the fixed deflector is in blunt nosed knot
Structure, and the compressor shroud chamber side of the fixed deflector is in prong structure.
In one embodiment, the fixed deflector is fixed by circumferential bolt.
The beneficial effects of the invention are as follows：
One circumferential bleed is set before aero-engine high-pressure compressor inlet guide vane (IGV) at suitable position
Slot replaces traditional complicated bleed structure, and it is long unexpected with bleed circulation area to avoid bleed flow path by fixed deflector
Change the loss brought, be connected with high-pressure compressor shroud chamber bleed air system, reach cooler pan chamber and bearing pivot obturages requirement
While, before effectively reducing high-pressure compressor, after Middle casing goose neck section friction layer influence；Institute is had using scheme
Easy geometry does not influence fan root structure, greatly reduces structural complexity, and relatively traditional bleed mode can be apparent
Reducing unit body weight and cost, weight can reduce, and have benefited from its easy structure, and processing and maintenance cost have notable drop
It is low.
Description of the drawings
The above and other features of the present invention, property and advantage will pass through retouching with reference to the accompanying drawings and examples
It states and becomes readily apparent from, wherein：
Fig. 1 is the schematic diagram of existing aero-engine high-pressure compressor shroud chamber bleed structure.
Fig. 2 is the schematic diagram according to the aero-engine high-pressure compressor shroud chamber bleed structure of the present invention.
Fig. 3 is the bleed flow path schematic diagram of the aero-engine high-pressure compressor shroud chamber bleed structure.
Fig. 4 is the schematic diagram of deflector set in circumferential bleed slot.
Specific embodiment
With reference to specific embodiments and the drawings, the invention will be further described, elaborates in the following description more
Details to facilitate a thorough understanding of the present invention, still the present invention obviously can be come with a variety of other manners different from this description it is real
It applies, those skilled in the art can make similar popularization according to practical situations without violating the connotation of the present invention, drill
It unravels silk, therefore should not be limited the scope of the invention with the content of this specific embodiment.
As shown in Figures 2 to 4, in one embodiment of this invention, aero-engine high-pressure compressor shroud chamber bleed knot
Structure includes the both aft mounted circumferential bleeds of high-pressure compressor import goose neck section 12 of 11 front side of high-pressure compressor inlet guide vane
The fixation deflector 15 set in slot 13 and circumferential bleed slot 13, fixed deflector 15 make in circumferential bleed slot 13 from sprue
The circulation area of side to compressor shroud chamber side is gradually amplified, and circumferential bleed slot 13 is used for and high-pressure compressor shroud chamber bleed system
System, the air-flow of introducing are obturaged for cooler pan chamber and bearing pivot.
It is described to obturage the overheated gas that bearing or gear is mainly prevented to generate, prevent main flow area gas from leaking and provide cold
But rotor and blade, balancing axial thrust.
As shown in figure 4, fixed deflector 15 has the first peripheral side 151 and the second peripheral side 152, the first peripheral side 151
And second state radially expanded outwardly between peripheral side 152, i.e., more extend outwardly, the interval of the two is bigger, and more inside
The interval of the two is smaller so that it is in blunt nosed structure that the sprue side 153 of fixed deflector 15, which is radial outside, and fixes deflector
15 compressor shroud chamber side 154 is in prong structure, and so-called tip refers to compressor shroud chamber side 154 relative to sprue side
153 is more thin, and the spacing distance of the first peripheral side 151 and the second peripheral side 152 is more small, is an opposite term, and
It is non-to refer to that compressor shroud chamber side 154 is pointed structures.Due to the structure of fixed deflector 15, from sprue to compressor shroud
Chamber, the circulation area of circumferential bleed slot 13 are gradually amplified, and make flow slowing down diffusion, to reduce the pressure loss of air-flow to the greatest extent.
As shown in Fig. 2, fixed deflector 15 passes through its hole 151 to be fixed on casing by circumferential bolt 14.
Fig. 3 shows bleed flow path schematic diagram, and IGV leading air-flow roads and not bleed flow path is shown, it can be seen that
Its effectively reduce high-pressure compressor before, after Middle casing goose neck section friction layer influence.
Using preceding solution, reduced by the bleed before high-pressure compressor inlet guide vane, and using deflector
The pitot loss caused by circulation area sudden enlargement, very low pitot loss ensure that bleed air system has following effect in this programme
Fruit：
1. because bleed position is before high-pressure compressor inlet guide vane, can be provided enough for high-pressure compressor shroud chamber
The gas of pressure, to meet the obturaging with gas demand of shroud chamber cooling requirement and bearing pivot；
2. bleed position in high-pressure compressor import goose neck segment trailer, can effectively reduce the shadow of inlet air flow boundary-layer
It rings；
3. using scheme that there is easy geometry, fan root structure is not influenced, greatly reduces structural complexity；
4. unit body weight and cost can be substantially reduced by using the relatively traditional bleed mode of bleed air system, weight can be with
Nearly 20kg is reduced, has benefited from its easy structure, processing and maintenance cost have and significantly reduce.
Although the present invention is disclosed as above with preferred embodiment, it is not for limiting the present invention, any this field skill
Art personnel without departing from the spirit and scope of the present invention, can make possible variation and modification.Therefore, it is every without departing from
The content of technical solution of the present invention, any modification that technical spirit according to the present invention makees above example, equivalent variations
And modification, it each falls within the protection domain that the claims in the present invention are defined.
Claims (3)
1. aero-engine high-pressure compressor shroud chamber bleed structure, which is characterized in that including high-pressure compressor import water conservancy diversion leaf
What is set in the circumferential bleed slot and the circumferential bleed slot that both high-pressure compressor import goose neck segment trailer had been set on front side of piece consolidates
Determine deflector, the fixed deflector makes the circulation area from sprue side to compressor shroud chamber side in the circumferential bleed slot
Gradually amplify, the circumferential direction bleed slot is used for and high-pressure compressor shroud chamber bleed air system.
2. aero-engine high-pressure compressor shroud chamber bleed structure as described in claim 1, which is characterized in that the fixation is led
Flowing plate has the first peripheral side and the second peripheral side, radially to extending out between first peripheral side and second peripheral side
The state opened so that the sprue side of the fixed deflector is in blunt nosed structure, and the compressor shroud of the fixed deflector
Chamber side is in prong structure.
3. aero-engine high-pressure compressor shroud chamber bleed structure as described in claim 1, which is characterized in that the fixation is led
Flowing plate is fixed by circumferential bolt.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201611121820.1A CN108180171B (en) | 2016-12-08 | 2016-12-08 | Aero-engine high-pressure compressor shroud chamber bleed structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201611121820.1A CN108180171B (en) | 2016-12-08 | 2016-12-08 | Aero-engine high-pressure compressor shroud chamber bleed structure |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN108180171A true CN108180171A (en) | 2018-06-19 |
| CN108180171B CN108180171B (en) | 2019-09-17 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201611121820.1A Active CN108180171B (en) | 2016-12-08 | 2016-12-08 | Aero-engine high-pressure compressor shroud chamber bleed structure |
Country Status (1)
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| CN (1) | CN108180171B (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110966261A (en) * | 2018-09-30 | 2020-04-07 | 中国航发商用航空发动机有限责任公司 | Gas-entraining structure and method for casing of gas compressor and aircraft engine |
| CN111396196A (en) * | 2019-01-02 | 2020-07-10 | 中国航发商用航空发动机有限责任公司 | S-shaped switching section of gas compressor and turbofan engine |
| CN113217120A (en) * | 2020-01-21 | 2021-08-06 | 中国航发商用航空发动机有限责任公司 | High-pressure turbine cooling air supply system and aircraft engine |
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| US20160245230A1 (en) * | 2015-02-23 | 2016-08-25 | Rolls-Royce Deutschland Ltd & Co Kg | Engine cowling of a gas turbine with thrust-reversing device and cross-sectionally adjustable outlet nozzle |
| CN106194846A (en) * | 2016-07-12 | 2016-12-07 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of double-layered case structure compressor and there is its aero-engine |
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| FR2887924A1 (en) * | 2005-06-30 | 2007-01-05 | Snecma | Guide for air flow between compressor and combustion chamber of aircraft turbine engine has independent rectifier supported by diffuser |
| CN102889133A (en) * | 2012-10-24 | 2013-01-23 | 哈尔滨东安发动机（集团）有限公司 | Gas-leading sealing structure of gas turbine |
| CN203685310U (en) * | 2013-12-11 | 2014-07-02 | 中航商用航空发动机有限责任公司 | Air entraining structure of eddy reduction device |
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| US20160245230A1 (en) * | 2015-02-23 | 2016-08-25 | Rolls-Royce Deutschland Ltd & Co Kg | Engine cowling of a gas turbine with thrust-reversing device and cross-sectionally adjustable outlet nozzle |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110966261A (en) * | 2018-09-30 | 2020-04-07 | 中国航发商用航空发动机有限责任公司 | Gas-entraining structure and method for casing of gas compressor and aircraft engine |
| CN111396196A (en) * | 2019-01-02 | 2020-07-10 | 中国航发商用航空发动机有限责任公司 | S-shaped switching section of gas compressor and turbofan engine |
| CN113217120A (en) * | 2020-01-21 | 2021-08-06 | 中国航发商用航空发动机有限责任公司 | High-pressure turbine cooling air supply system and aircraft engine |
| CN113217120B (en) * | 2020-01-21 | 2023-08-08 | 中国航发商用航空发动机有限责任公司 | High-pressure turbine cooling air supply system and aeroengine |
Also Published As
| Publication number | Publication date |
|---|---|
| CN108180171B (en) | 2019-09-17 |
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