CN102817873A - Ladder-shaped gap structure for gas compressor of aircraft engine - Google Patents
Ladder-shaped gap structure for gas compressor of aircraft engine Download PDFInfo
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- CN102817873A CN102817873A CN2012102853349A CN201210285334A CN102817873A CN 102817873 A CN102817873 A CN 102817873A CN 2012102853349 A CN2012102853349 A CN 2012102853349A CN 201210285334 A CN201210285334 A CN 201210285334A CN 102817873 A CN102817873 A CN 102817873A
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Abstract
The invention provides a ladder-shaped gap structure for a gas compressor of an aircraft engine, and relates to the stability expanding technology of a high-load fan or a machine pressure gas machine of the aircraft engine. The gas machine of the aircraft engine comprises a blade and a cartridge receiver housing, wherein the position which corresponds to the inner side wall of the cartridge receiver housing is provided with a ladder-shaped peripheral groove with a certain depth and width in a machining way, the different ladder-shaped gap structure distributions can be obtained by optimizing the size of a matching gas and the position of the ladder-shaped peripheral groove, and the size and the position of a blockage group at the top region of the gas compressor can be effectively controlled due to the interaction of complex flow between the ladder-shaped groove and the top region of the gas compressor, so that the flowing area of the channel main flow can be improved, the stability work margin of the gas compressor can be improved, and the performance of the gas compressor can be improved.
Description
Technical field
The present invention relates to a kind of scalariform gap structure that is used for aerial engine fan and gas compressor, the stability-enhancement synergistic that it can realize fan and gas compressor is specially adapted to the high-performance aero gas turbine engine.
Background technique
Blade tip clearance is introduced for avoiding bumping mill between rotation blade and the casing in the turbomachine, and its size is about 0.3%～1% compressor rotor leaf apical axis to chord length.Under the differential pressure action of blade tip clearance both sides, segment fluid flow passes blade tip clearance and forms leakage flow, owing to receive the influence of main flow, this leakage flow is present in zone, compressor rotor leaf top with the form in tip leakage whirlpool usually simultaneously.As far as fan/axial flow compressor, the negative effect of tip leakage stream mainly is presented as generation leakage loss and obstruction, and the former can reduce compressor efficiency, and the latter can reduce the voltage rise ability and the stable operation range of gas compressor.The motor of medium thrust, medium pressure ratio, blade height is bigger, and is also not really serious by the loss that tip clearance causes.Along with the increase of pressure ratio, blade height significantly shortens, behind the high-pressure compressor what Ye Gaoyou shorten to 20 ~ 30mm, the loss that causes of tip clearance becomes more remarkable like this.According to actual measurement, tip clearance relative value (being gap/blade height) increases by 1%, and efficient reduces by 1% approximately; And efficient reduces by 1%, and oil consumption rate increases by 2% approximately.In addition, increasing result of study shows that the unstability of modern high performance gas compressor is to be triggered by the stall tendency that the blade tip clearance leakage flow is produced mostly.Modern advanced aeroengine requires the axial flow compressor overall pressure tatio to improve constantly and constantly minimizing of progression (or number of blade) to the military requirement of high thrust weight ratio.This just causes the stage load of axial flow compressor increasingly high, and tip leakage is more serious, and gas compressor progressively strengthens the receptance of tip clearance, and the shared ratio of the loss that blade tip clearance causes is increasingly high.
Can find out that blade tip clearance size and layout play crucial effect to annulus wall boundary layer and with the interaction of blade boundary layer.If gap control must be got well, rotor voltage rise, efficient and stall margin all can obtain improvement in various degree; Otherwise, if excesssive gap, or layout is unreasonable, and peaked area will be again a serious aerodynamic loss source and the stall zone that takes the lead in.Advanced pneumatic design and the test method of modern aeroengine made compressor efficiency up to more than 88%.If want further to improve engine performance, need reduce flow leakage as far as possible, reduce the end wall loss in the runner.Along with becoming increasingly abundant of understanding that tip leakage is flowed; People begin to consider to take control measure to slow down degeneration of stable operation nargin and decreased performance problem that leakage loss is brought; As the casing of in the actual model of many motors, being used widely processing, utilize the variation of tip clearance to improve rotor performance.
The patent documentation of notification number CN102162472A discloses a kind of many arc slot casing treatments that are applied in axial flow compressor rotor blade tip petiolarea.The treatment trough of said many arc slot casing treatments is radially adopting two circular arc types and in the modular design that circumferentially adopts circular arc type.Through the geometrical construction form of rational Design Treatment groove, promptly adopt the modular design of circular arc type at the two circular arc types of R direction (radially) employing and at θ (circumferentially).
The patent documentation of notification number CN101691869 discloses a kind of axial and radial flowing compressor with axial chute processor casing structure; This axial and radial flowing compressor comprises shaft flow rotor; Axial flow stator and footpath flow air compressor; And shaft flow rotor, axial flow stator and the coaxial successively connection of footpath three parts of flow air compressor; On the casing wall of described axial and radial flowing compressor shaft flow rotor, be processed with circumferential equally distributed axial chute, axial chute turns at radially clockwise son and is 30 ° ~ 60 ° inclination.
Traditional peripheral groove processor box is as shown in Figure 1; Be in the casing upper edge gas compressor circumferentially open several straight troughs; Practical application effect shows no matter incoming flow is equal uniform flow or the import distortion takes place, and the gas compressor stability margin all has improvement; Because peripheral groove can be realized processing easily, therefore has certain meaning for the performance of improving motor.But the shortcoming of this type processor box is that the improvement of stability margin is a cost with the efficient of losing gas compressor.
Therefore, need the rational deployment of seeking a kind of rotor blade tip clearance badly, reach the dual purpose that enlarges stable operation range and raise the efficiency.
Summary of the invention
Technical problem to be solved by this invention provides a kind of reasonable in design, has promptly realized that stable operation nargin promotes the scalariform gap structure of not sacrificing compressor efficiency, simple and practical aero-engine compressor again.
The present invention solves the problems of the technologies described above the scalariform gap structure into a kind of aero-engine compressor; Said aero-engine compressor comprises rotor blade and casing shell; Its structural feature is: described casing shell madial wall is processed into the ladder peripheral groove with certain depth and width on the corresponding position, obtains different scalariform gap structure layouts through optimization of matching gap length and ladder peripheral groove slotting position.
Be preferably, the degree of depth of peripheral groove of the present invention equates with gas compressor blade top gap length t1.
Be preferably, peripheral groove fluting of the present invention is positioned at 60%～108% scope of leaf apical axis to chord length.
Be preferably; Scalariform gap structure according to the invention is relevant with the target that the gas compressor designing institute is pursued; In order to obtain higher pressure ratio and efficient, be preferably blade tip clearance t1 is made as the leaf apical axis to 0.3% of chord length, and make the peripheral groove fluting be positioned at 60%～108% scope of leaf apical axis to chord length; In order to obtain higher stable operation range, be preferably the leaf top clearance is made as the leaf apical axis to 0.6% of chord length, and make the peripheral groove fluting be positioned at 90%～108% scope of leaf apical axis to chord length.
The present invention compares with existing technology and has the following advantages and effect: novel scalariform gap structure layout of the present invention is simpler, realizes processing more easily, and can when improving gas compressor stable operation nargin, improve compressor efficiency; In addition, through of the optimum organization of blade tip clearance size, can realize the mobile rationalization in compressor rotor end is reached the target that improves gas compressor stable operation nargin or gas compressor performance with the ladder slotting position.The complicated flow that the present invention is utilized in step trough and gas compressor top area interacts; Effectively control gas compressor top area is blocked the size and the position of group; Improve the circulation area of passage main flow, when improving gas compressor stable operation nargin, improved the performance of gas compressor.
Description of drawings
Fig. 1 is the schematic representation of traditional peripheral groove processor casing structure.
Fig. 2 is the schematic representation of the scalariform gap structure of the said aero-engine compressor of one embodiment of the invention.
Label declaration: 1-rotor blade, 2-casing shell, 3-step trough.
Embodiment
Below, in conjunction with embodiment the present invention is done further detailed description, following examples are to explanation of the present invention and the present invention is not limited to following examples.
Embodiment 1: as shown in Figure 2, the described aero-engine compressor of present embodiment comprises rotor blade 1 and casing shell 2, and the madial wall of casing shell 2 is provided with the less step trough of the degree of depth 3.
In order to improve the performance of gas compressor; Realization is to the tissue and the regulation and control of compressor rotor leaf top zone complicated flow; On compressor casing shell 2, offered the sizable step trough 3 of the degree of depth and blade tip clearance; Confirm blade tip clearance size and step trough slotting position to the difference that is pursued one's goal in the gas compressor design, design different rotor leaf top scalariform gap structures.If pursue higher pressure ratio and efficient in the design, then rotor blade tip clearance t1 is made as rotor leaf apical axis to 0.3% of chord length, in 60%～108% scope of chord length, introduce the degree of depth step trough suitable at the leaf apical axis simultaneously with blade tip clearance; If pursue higher stable operation range, then rotor leaf top clearance is made as rotor leaf apical axis to 0.6% of chord length, in 90%～108% scope of chord length, introduce the degree of depth step trough suitable at the leaf apical axis simultaneously with blade tip clearance.
During work; Pressure gradient by rotor blade 1 leaf top suction surface and pressure side both sides; Step trough can block near the low energy rotor blade trailing edge of being positioned at that whirlpool, top clearance and shock wave interaction produce group and brings the adjacent rotor blades top clearance into; The height loss that gets into the adjacent blades top clearance block group with casing viscous boundary layer interaction process in dissipate and become the obstruction group of medium loss; Thereby eliminated near the obstruction the rotor blade pressure side effectively, zone, rotor leaf top has been produced a kind of pneumostop effect.Therefore, the scalariform gap structure can be regulated and control zone, rotor leaf top and block the size and the position of group, improves the circulation area of passage main flow, thereby has improved pressure ratio, efficient and the range of flow of compressor rotor.
In sum, the present invention can directly be used for aviation gas turbine and start fan/machine gas compressor, in the stable operation nargin that improves fan/machine gas compressor, improves the efficient of gas compressor.
Thinking of the present invention is from rationalization's gas compressor blade top zone complicated flow; Explored a kind of novel leaf top scalariform gap structure; Design a kind of step trough processor box; Broken the traditional concept of " casing is handled can enlarge gas compressor stable operation nargin, but will lower efficiency ", this also becomes the aim of the present invention's design.
In addition, the specific embodiment described in this specification, the shape of its component, institute's title of being named etc. can be different.Allly conceive equivalence or the simple change that described structure, characteristic and principle are done, include in the protection domain of patent of the present invention according to patent of the present invention.Person of ordinary skill in the field of the present invention can make various modifications or replenishes or adopt similar mode to substitute described specific embodiment; Only otherwise depart from structure of the present invention or surmount the defined scope of these claims, all should belong to protection scope of the present invention.
Claims (5)
1. the scalariform gap structure of an aero-engine compressor; Said aero-engine compressor comprises rotor blade and casing shell; It is characterized in that; On the corresponding position of described casing shell madial wall, offer ladder peripheral groove, obtain different scalariform gap structures through optimization of matching gap length and step trough slotting position with certain depth and width through processing.
2. the scalariform gap structure of aero-engine compressor according to claim 1 is characterized in that, the degree of depth of peripheral groove equates with gas compressor blade top gap length t1.
3. the scalariform gap structure of aero-engine compressor according to claim 1 and 2 is characterized in that, said peripheral groove fluting is positioned at 60%～108% scope of leaf apical axis to chord length.
4. the scalariform gap structure of aero-engine compressor according to claim 3; It is characterized in that; Blade tip clearance t1 is made as the leaf apical axis to 0.3% of chord length; Simultaneously in 60%～108% scope of chord length, introduce the degree of depth step trough suitable, can obtain higher pressure ratio and efficient with blade tip clearance at the leaf apical axis.
5. the scalariform gap structure of aero-engine compressor according to claim 3; It is characterized in that; The leaf top clearance is made as the leaf apical axis to 0.6% of chord length; Simultaneously in 90%～108% scope of chord length, introduce the degree of depth step trough suitable, can obtain higher stable operation range with blade tip clearance at the leaf apical axis.
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| CN201210285334.9A CN102817873B (en) | 2012-08-10 | 2012-08-10 | Ladder-shaped gap structure for gas compressor of aircraft engine |
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| CN201210285334.9A CN102817873B (en) | 2012-08-10 | 2012-08-10 | Ladder-shaped gap structure for gas compressor of aircraft engine |
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| CN102817873A true CN102817873A (en) | 2012-12-12 |
| CN102817873B CN102817873B (en) | 2015-07-15 |
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Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105298923A (en) * | 2014-06-17 | 2016-02-03 | 中国科学院工程热物理研究所 | Front seam and rear groove type casing treatment stability enhancement device for gas compressor |
| CN105570187A (en) * | 2015-12-11 | 2016-05-11 | 哈尔滨东安发动机（集团）有限公司 | Control method for dimensions of rotor tips of gas compressor |
| CN106289784A (en) * | 2016-08-02 | 2017-01-04 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of inlet distortion stagnation pressure rake structure |
| CN106286394A (en) * | 2016-10-14 | 2017-01-04 | 中国科学院工程热物理研究所 | A kind of compressor communication type shrinkage joint treated casing method and device |
| CN110651112A (en) * | 2017-05-02 | 2020-01-03 | 赛峰飞机发动机公司 | Turbomachine having a fan rotor and a reduction gearbox driving a shaft of a low-pressure compressor |
| CN112283167A (en) * | 2020-11-20 | 2021-01-29 | 西安热工研究院有限公司 | Circumferential groove type casing treatment design method for axial flow compressor |
| CN113167287A (en) * | 2018-12-19 | 2021-07-23 | 依必安派特穆尔芬根有限两合公司 | Turbocompressor with adjusted radial profile of the vanes and the compressor wall |
| CN117948290A (en) * | 2024-03-22 | 2024-04-30 | 河北冀力重型机械设备有限公司 | High-air-volume high-pressure axial flow fan |
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| US4238170A (en) * | 1978-06-26 | 1980-12-09 | United Technologies Corporation | Blade tip seal for an axial flow rotary machine |
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Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105298923B (en) * | 2014-06-17 | 2018-01-02 | 中国科学院工程热物理研究所 | Slot type treated casing expands stabilization device after being stitched before compressor |
| CN105298923A (en) * | 2014-06-17 | 2016-02-03 | 中国科学院工程热物理研究所 | Front seam and rear groove type casing treatment stability enhancement device for gas compressor |
| CN105570187B (en) * | 2015-12-11 | 2020-04-28 | 哈尔滨东安发动机（集团）有限公司 | Compressor rotor blade tip size control method |
| CN105570187A (en) * | 2015-12-11 | 2016-05-11 | 哈尔滨东安发动机（集团）有限公司 | Control method for dimensions of rotor tips of gas compressor |
| CN106289784A (en) * | 2016-08-02 | 2017-01-04 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of inlet distortion stagnation pressure rake structure |
| CN106286394A (en) * | 2016-10-14 | 2017-01-04 | 中国科学院工程热物理研究所 | A kind of compressor communication type shrinkage joint treated casing method and device |
| CN106286394B (en) * | 2016-10-14 | 2018-08-10 | 中国科学院工程热物理研究所 | A kind of compressor communication type shrinkage joint treated casing method and device |
| CN110651112A (en) * | 2017-05-02 | 2020-01-03 | 赛峰飞机发动机公司 | Turbomachine having a fan rotor and a reduction gearbox driving a shaft of a low-pressure compressor |
| CN113167287A (en) * | 2018-12-19 | 2021-07-23 | 依必安派特穆尔芬根有限两合公司 | Turbocompressor with adjusted radial profile of the vanes and the compressor wall |
| CN113167287B (en) * | 2018-12-19 | 2023-06-30 | 依必安派特穆尔芬根有限两合公司 | Turbo compressor with tailored meridian profile of blades and compressor wall |
| CN112283167A (en) * | 2020-11-20 | 2021-01-29 | 西安热工研究院有限公司 | Circumferential groove type casing treatment design method for axial flow compressor |
| CN117948290A (en) * | 2024-03-22 | 2024-04-30 | 河北冀力重型机械设备有限公司 | High-air-volume high-pressure axial flow fan |
| CN117948290B (en) * | 2024-03-22 | 2024-06-18 | 河北冀力重型机械设备有限公司 | Axial flow fan |
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