CN102817873B - 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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- CN102817873B CN102817873B CN201210285334.9A CN201210285334A CN102817873B CN 102817873 B CN102817873 B CN 102817873B CN 201210285334 A CN201210285334 A CN 201210285334A CN 102817873 B CN102817873 B CN 102817873B
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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 Investigation of Stepped Tip Gap for aerial engine fan and gas compressor, it can realize the stability-enhancement synergistic of fan and gas compressor, is specially adapted to high performance turbine gas turbine engine.
Background technique
In turbomachine, blade tip clearance is introduced for avoiding touching between rotation blade and casing mill, 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 forms leakage flow through blade tip clearance, and simultaneously owing to being subject to the impact of main flow, this leakage flow is present in compressor rotor Ye Ding region with the form in tip leakage whirlpool usually.For fan/axial flow compressor, the negative effect major embodiment of tip leakage stream is for producing leakage loss and blocking, and the former can reduce compressor efficiency, and the latter can reduce voltage rise ability and the stable operation range of gas compressor.The motor of medium thrust, medium boost ratio, blade height is comparatively large, and the loss caused by tip clearance is also not bery serious.Along with the increase of pressure ratio, blade height significantly shortens, after high-pressure compressor what Ye Gaoyou shorten to 20-30mm, the loss that such tip clearance causes becomes more remarkable.According to actual measurement, tip clearance relative value (i.e. gap/blade height) increases by 1%, and efficiency about reduces by 1%; And efficiency reduces by 1%, oil consumption rate about increases by 2%.In addition, increasing result of study shows, the unstability of modern high performance gas compressor is that the stall precursor flowing produced by clearance leakage of blade tip triggers mostly.To the military requirement of high thrust weight ratio, modern advanced aero engine requires that axial flow compressor overall pressure tatio improves constantly and progression (or number of blade) constantly reduces.This just causes the stage load of axial flow compressor more and more higher, and tip leakage is more serious, and the receptance of gas compressor to tip clearance progressively strengthens, and the ratio shared by the loss that blade tip clearance causes is more and more higher.
Can find out, blade tip clearance size and layout to annulus wall boundary layer and and the interaction of blade boundary layer play a very important role.If gap control must be got well, rotor voltage rise, efficiency and stall margin all can obtain improvement in various degree; Otherwise if excesssive gap, or layout is unreasonable, peaked area will be again a serious aerodynamic loss source and the stall zone that takes the lead in.The pneumatic design of modern aeroengine advanced person and test method have made compressor efficiency up to more than 88%.If want to improve engine performance further, need to reduce flow leakage as far as possible, reduce the end wall loss in runner.Along with becoming increasingly abundant to tip leakage flowing understanding, people start to consider to take control measure to degenerate and degradation problem to the stable operation nargin slowing down leakage loss and bring, as the treated casing of being used widely in the actual model of many motors, utilize the change of tip clearance to improve rotor performance.
The patent documentation of notification number CN102162472A discloses a kind of many arc slot casing treatments being applied in axial flow compressor rotor blade tip petiolarea.The treatment trough of described many arc slot casing treatments adopts bicircular arcs type in radial direction and adopts the Combination Design of circular arc type in circumference.By the geometrical construction form of rational design treatment groove, namely adopt bicircular arcs type in R direction (radial direction) and adopt the Combination Design of circular arc type at θ (circumference).
Notification number CN101691869 patent document 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 footpath flow air compressor three parts coaxially connect successively; The casing wall of described axial and radial flowing compressor shaft flow rotor is processed with circumferential equally distributed axial chute, and axial chute turns to the inclination in 30 °-60 ° at the clockwise son of radial direction.
Traditional peripheral groove processor box as shown in Figure 1, open several straight troughs along the circumference of gas compressor on casing, practical application effect shows, no matter incoming flow is equal uniform flow or inlet distortions occurs, gas compressor stability margin is all improved, because peripheral groove can realize processing easily, therefore for the performance improving motor, there is certain meaning.But the shortcoming of this kind of processor box is that the improvement of stability margin is to lose the efficiency of gas compressor for cost.
Therefore, need the rational deployment seeking a kind of rotor blade tip clearance badly, reach the dual purpose expanding stable operation range and raise the efficiency.
Summary of the invention
Not only technical problem to be solved by this invention is to provide a kind of reasonable in design, achieves the Investigation of Stepped Tip Gap that stable operation nargin promotes but also do not sacrifice compressor efficiency, simple and practical aero-engine compressor.
The present invention solves the problems of the technologies described above the Investigation of Stepped Tip Gap into a kind of aero-engine compressor, described aero-engine compressor comprises rotor blade and casing shell, its structural feature is: the ladder peripheral groove with certain depth and width is processed in the described corresponding position of casing housing interior side-wall, obtains different Investigation of Stepped Tip Gap layouts by Optimized Matching gap length and ladder peripheral groove slotting position.
Be preferably, the degree of depth of peripheral groove of the present invention and gas compressor blade top clearance t 1 equal and opposite in direction.
Be preferably, peripheral groove fluting of the present invention is positioned at 60% ~ 108% scope of leaf apical axis to chord length.
Be preferably, Investigation of Stepped Tip Gap of the present invention is relevant with the target that gas compressor design is pursued, in order to obtain higher pressure ratio and efficiency, preferably blade tip clearance t1 is set to leaf apical axis to 0.3% of chord length, and makes peripheral groove fluting be positioned at 60% ~ 108% scope of leaf apical axis to chord length; In order to obtain higher stable operation range, preferably blade tip clearance t1 is set to leaf apical axis to 0.6% of chord length, and makes peripheral groove fluting be positioned at 90% ~ 108% scope of leaf apical axis to chord length.
The present invention compared with the existing technology has the following advantages and effect: novel Investigation of Stepped Tip Gap layout of the present invention is more simple, realizes processing more easily, and can improve compressor efficiency while raising gas compressor stable operation nargin; In addition, by the optimum organization of blade tip clearance size and ladder slotting position, the rationalization to compressor rotor endwall flow can be realized, reach the target improving gas compressor stable operation nargin or Capability of Compressor.The present invention utilizes and interacts at the Complex Flows of step trough and gas compressor top area, effective control gas compressor top area blocks size and the position of group, improve the circulation area of passage main flow, while raising gas compressor stable operation nargin, improve the performance of gas compressor.
Accompanying drawing explanation
Fig. 1 is the schematic diagram of traditional peripheral groove processor casing structure.
Fig. 2 is the schematic diagram of the Investigation of Stepped Tip Gap of aero-engine compressor described in one embodiment of the invention.
Label declaration: 1-rotor blade, 2-casing shell, 3-step trough.
Embodiment
Below, the present invention is described in further detail in conjunction with the embodiments, and following examples are explanation of the invention and the present invention is not limited to following examples.
Embodiment 1: as shown in Figure 2, and the aero-engine compressor described in the 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, realize the tissue to compressor rotor leaf top zone Complex Flows and regulation and control, compressor casing shell 2 has been offered the degree of depth and the sizable step trough 3 of blade tip clearance, for the difference pursued a goal in gas compressor design to determine blade tip clearance size and step trough slotting position, design different rotor leaf top Investigation of Stepped Tip Gaps.If pursue higher pressure ratio and efficiency in design, then rotor blade tip clearance t1 is set to rotor leaf apical axis to 0.3% of chord length, in 60% ~ 108% scope of leaf apical axis to chord length, introduces the degree of depth step trough suitable with blade tip clearance simultaneously; If pursue higher stable operation range, then rotor blade tip clearance t1 is set to rotor leaf apical axis to 0.6% of chord length, in 90% ~ 108% scope of leaf apical axis to chord length, introduces the degree of depth step trough suitable with blade tip clearance simultaneously.
During work, by the pressure gradient of rotor blade 1 leaf top suction surface and pressure side both sides, the low energy be positioned near rotor blade trailing edge that whirlpool, top clearance and shock wave interaction produce can be blocked group and be brought adjacent rotor blades top clearance into by step trough, the height loss entering adjacent blades top clearance block group with casing viscous boundary layer interaction process in dissipate and be known as the obstruction group of moderate losses, thus the blocking effectively eliminated near rotor blade pressure side, a kind of pneumostop effect is produced to rotor Ye Ding region.Therefore, Investigation of Stepped Tip Gap can regulate and control size and the position that group is blocked in rotor Ye Ding region, improves the circulation area of passage main flow, thus improves the pressure ratio of compressor rotor, efficiency and range of flow.
In sum, the present invention can be directly used in aviation gas turbine and start fan/machine gas compressor, improves the efficiency of gas compressor while the stable operation nargin improving fan/machine gas compressor.
Thinking of the present invention is from rationalization's gas compressor blade top zone Complex Flows, explore a kind of Novel leaf top Investigation of Stepped Tip Gap, design a kind of step trough processor box, break " treated casing can expand gas compressor stable operation nargin; but to lower efficiency " traditional concept, this also becomes the aim of the present invention's design.
In addition, the specific embodiment described in this specification, the shape, institute's title of being named etc. of its component can be different.All equivalences of doing according to structure, feature and the principle described in inventional idea of the present invention or simple change, be included in the protection domain of patent of the present invention.Those skilled in the art can make various amendment or supplement or adopt similar mode to substitute to described specific embodiment; only otherwise depart from structure of the present invention or surmount this scope as defined in the claims, protection scope of the present invention all should be belonged to.
Claims (3)
1. the Investigation of Stepped Tip Gap of an aero-engine compressor, described aero-engine compressor comprises rotor blade and casing shell, it is characterized in that, the ladder peripheral groove with certain depth and width is offered by processing in the described corresponding position of casing housing interior side-wall, obtains different Investigation of Stepped Tip Gaps by Optimized Matching gap length and step trough slotting position; The peripheral groove degree of depth of this Investigation of Stepped Tip Gap and gas compressor blade top clearance t 1 equal and opposite in direction; Described peripheral groove fluting is positioned at 60% ~ 108% scope of leaf apical axis to chord length.
2. the Investigation of Stepped Tip Gap of aero-engine compressor according to claim 1, it is characterized in that, blade tip clearance t1 is set to leaf apical axis to 0.3% of chord length, in 60% ~ 108% scope of leaf apical axis to chord length, introduce the degree of depth step trough suitable with blade tip clearance simultaneously, higher pressure ratio and efficiency can be obtained.
3. the Investigation of Stepped Tip Gap of aero-engine compressor according to claim 1, it is characterized in that, blade tip clearance t1 is set to leaf apical axis to 0.6% of chord length, in 90% ~ 108% scope of leaf apical axis to chord length, introduce the degree of depth step trough suitable with blade tip clearance simultaneously, higher stable operation range can be obtained.
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| Application Number | Priority Date | Filing Date | Title |
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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 CN102817873A (en) | 2012-12-12 |
| CN102817873B true CN102817873B (en) | 2015-07-15 |
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Families Citing this family (9)
| 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 |
| 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 |
| CN106289784B (en) * | 2016-08-02 | 2019-01-18 | 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 |
| 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 |
| FR3065994B1 (en) * | 2017-05-02 | 2019-04-19 | Safran Aircraft Engines | BLOWER ROTOR TURBOMACHINE AND REDUCER DRIVING A LOW PRESSURE COMPRESSOR SHAFT |
| CN110630565B (en) * | 2018-06-25 | 2025-08-22 | 约克广州空调冷冻设备有限公司 | Axial flow fan air guide ring and axial flow fan |
| DE102018132978A1 (en) * | 2018-12-19 | 2020-06-25 | Ebm-Papst Mulfingen Gmbh & Co. Kg | Turbo compressor with adapted meridian contour of the blades and compressor wall |
| CN112283167B (en) * | 2020-11-20 | 2022-04-01 | 西安热工研究院有限公司 | A design method for the treatment of a circumferential grooved casing for an axial flow compressor |
| 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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| US5137419A (en) * | 1984-06-19 | 1992-08-11 | Rolls-Royce Plc | Axial flow compressor surge margin improvement |
| CN1096347A (en) * | 1993-03-04 | 1994-12-14 | Abb管理有限公司 | Has the centrifugal compressor that can make the stable casing that flows |
| US5762470A (en) * | 1993-03-11 | 1998-06-09 | Central Institute Of Aviation Motors (Ciam) | Anti-stall tip treatment means |
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| DE102008010283A1 (en) * | 2008-02-21 | 2009-08-27 | Mtu Aero Engines Gmbh | Circulation structure for a turbocompressor |
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| US5137419A (en) * | 1984-06-19 | 1992-08-11 | Rolls-Royce Plc | Axial flow compressor surge margin improvement |
| US4781530A (en) * | 1986-07-28 | 1988-11-01 | Cummins Engine Company, Inc. | Compressor range improvement means |
| CN1096347A (en) * | 1993-03-04 | 1994-12-14 | Abb管理有限公司 | Has the centrifugal compressor that can make the stable casing that flows |
| US5762470A (en) * | 1993-03-11 | 1998-06-09 | Central Institute Of Aviation Motors (Ciam) | Anti-stall tip treatment means |
| CN1281953A (en) * | 1999-07-15 | 2001-01-31 | 株式会社日立制作所 | Turbomachine |
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| CN102817873A (en) | 2012-12-12 |
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