CN202914391U - Step-shaped interstitial structure of gas compressor of aircraft engine - Google Patents
Step-shaped interstitial structure of gas compressor of aircraft engine Download PDFInfo
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- CN202914391U CN202914391U CN 201220397519 CN201220397519U CN202914391U CN 202914391 U CN202914391 U CN 202914391U CN 201220397519 CN201220397519 CN 201220397519 CN 201220397519 U CN201220397519 U CN 201220397519U CN 202914391 U CN202914391 U CN 202914391U
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Abstract
The utility model provides a step-shaped interstitial structure of a gas compressor of an aircraft engine and relates to the extending stability technology of a high load fan or blower gas compressor of the aircraft engine. The gas compressor of the aircraft engine comprises a rotor blade and a casing shell, wherein a circumferential step groove with a certain depth and width is formed at the corresponding position of the internal lateral wall of the casing shell through processing. Different step-shaped interstitial structural configuration is obtained by optimizing and matching a gap size and a locating position of the circumferential step groove and complex flow interaction in the step groove and a top area of the gas compressor is utilized to effectively control the size and position of a blocking mass in the top area of the gas compressor, so that a flow area of main flow of a channel is enlarged, the stable operating margin of the gas compressor is improved, and the performance of the gas compressor is improved simultaneously.
Description
Technical field
The utility model relates to a kind of Investigation of Stepped Tip Gap for aerial engine fan and gas compressor, and 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 be subject to the impact of main flow, this leakage flow is present in zone, compressor rotor leaf top with the form in tip leakage whirlpool usually simultaneously.For fan/axial flow compressor, the negative effect major embodiment of tip leakage stream is for producing leakage loss and obstruction, 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 pressure ratio, blade height is larger, and is also not bery 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 Efficiency Decreasing 1%, 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 produces mostly.Modern advanced aero engine 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 more and more higher, 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 more and more higher.
Can find out that blade tip clearance size and layout play a very important role 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.Modern aeroengine advanced person's pneumatic design and test method have made compressor efficiency up to more than 88%.If want further to improve engine performance, need to 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 the degeneration of stable operation nargin and hydraulic performance decline problem that leakage loss is brought, process such as the casing of in the actual model of many motors, being used widely, 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 described many arc slot casing treatments is radially adopting two circular arc types and in the Combination Design that circumferentially adopts circular arc type.By the geometrical construction form of rational design treatment trough, namely adopt two circular arc types and in the Combination Design of θ (circumferentially) employing circular arc type in R direction (radially).
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 successively coaxial connection of footpath three parts of flow air compressor; Casing wall at described axial and radial flowing compressor shaft flow rotor is processed with circumferential equally distributed axial chute, and axial chute turns at radially clockwise son and is 30 ° ~ 60 ° inclination.
Traditional peripheral groove processor box as shown in Figure 1, on casing, circumferentially to open several straight troughs along gas compressor, practical application effect shows, no matter incoming flow is uniform-flow or the import distortion occurs, the gas compressor stability margin all is improved, because peripheral groove can be realized processing easily, therefore has certain meaning for the performance of improving motor.But the shortcoming of this class processor box be stability margin improvement take the loss gas compressor efficient as cost.
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.
The model utility content
Technical problem to be solved in the utility model provides a kind of reasonable in design, has namely realized that stable operation nargin promotes the Investigation of Stepped Tip Gap of not sacrificing again compressor efficiency, simple and practical aero-engine compressor.
The utility model 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: described casing shell madial wall is processed into the ladder peripheral groove with certain depth and width on the corresponding position, 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 described in the utility model equates with gas compressor blade top gap length.
Be preferably, peripheral groove fluting described in the utility model is positioned at the leaf apical axis to 60%～108% scope of chord length.
Be preferably, Investigation of Stepped Tip Gap described in the utility model 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 is made as the leaf apical axis to 0.3% of chord length, and make the peripheral groove fluting be positioned at the leaf apical axis to 60%～108% scope of chord length; In order to obtain higher stable operation range, be preferably blade tip clearance is made as the leaf apical axis to 0.6% of chord length, and make the peripheral groove fluting be positioned at the leaf apical axis to 90%～108% scope of chord length.
The utility model is compared with prior art and had the following advantages and effect: novel Investigation of Stepped Tip Gap layout of the present utility model is simpler, realizes more easily processing, and can improve compressor efficiency when improving gas compressor stable operation nargin; In addition, by the optimum organization of blade tip clearance size with the ladder slotting position, can realize the rationalization to the compressor rotor endwall flow, reach the target that improves gas compressor stable operation nargin or Capability of Compressor.Utilization interacts at the Complex Flows of step trough and gas compressor top area, effectively control gas compressor top area is blocked 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 Investigation of Stepped Tip Gap of the described aero-engine compressor of the utility model one embodiment.
Label declaration: 1-rotor blade, 2-casing shell, 3-step trough.
Embodiment
Below, in conjunction with the embodiments the utility model being described in further detail, following examples are to explanation of the present utility model and the utility model 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 tissue and the regulation and control of compressor rotor leaf top zone Complex Flows, the degree of depth and the sizable step trough 3 of blade tip clearance have been offered at compressor casing shell 2, determine blade tip clearance size and step trough slotting position for the difference that pursues a goal in the gas compressor design, design different rotor leaf top Investigation of Stepped Tip Gaps.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 with blade tip clearance at the leaf apical axis simultaneously; If pursue higher stable operation range, then the rotor blade tip 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 with blade tip clearance at the leaf apical axis simultaneously.
During work, pressure gradient by rotor blade 1 leaf top suction surface and pressure side both sides, step trough can produce whirlpool, top clearance and shock wave interaction is positioned near the rotor blade trailing edge low energy and blocks group and bring the adjacent rotor blades top clearance into, the height loss that enters the adjacent blades top clearance block group with casing viscous boundary layer interaction process in dissipate and be known as the obstruction group of medium loss, thereby effectively eliminated near the obstruction the rotor blade pressure side, zone, rotor leaf top has been produced a kind of pneumostop effect.Therefore, Investigation of Stepped Tip Gap can be regulated and control zone, rotor leaf top and block 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 utility model can be directly used in aviation gas turbine and start fan/machine gas compressor, improves the efficient of gas compressor in the stable operation nargin that improves fan/machine gas compressor.
Thinking of the present utility model is from rationalization's gas compressor blade top zone Complex Flows, explored a kind of Novel leaf top Investigation of Stepped Tip Gap, design a kind of step trough processor box, broken " casing is processed and can be enlarged gas compressor stable operation nargin; but to lower efficiency " traditional concept, this also becomes the aim of the utility model 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, feature and principle are done according to the utility model patent, be included in the protection domain of the utility model patent.The utility model person of ordinary skill in the field can make various modifications or replenishes or adopt similar mode to substitute described specific embodiment; only otherwise depart from structure of the present utility model or surmount this scope as defined in the claims, all should belong to protection domain of the present utility model.
Claims (5)
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, offers the ladder peripheral groove with predetermined depth and width by processing on the madial wall of described casing shell.
2. the Investigation of Stepped Tip Gap 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.
3. the Investigation of Stepped Tip Gap of aero-engine compressor according to claim 1 and 2 is characterized in that, described peripheral groove fluting is positioned at the leaf apical axis to 60%～108% scope of chord length.
4. the Investigation of Stepped Tip Gap of aero-engine compressor according to claim 3, it is characterized in that, blade tip clearance is made as the leaf apical axis to 0.3% of chord length, in 60%～108% scope of chord length, introduce the degree of depth step trough suitable with blade tip clearance at the leaf apical axis simultaneously, can obtain higher pressure ratio and efficient.
5. the Investigation of Stepped Tip Gap of aero-engine compressor according to claim 3, it is characterized in that, blade tip clearance is made as the leaf apical axis to 0.6% of chord length, in 90%～108% scope of chord length, introduce the degree of depth step trough suitable with blade tip clearance at the leaf apical axis simultaneously, can obtain higher stable operation range.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201220397519 CN202914391U (en) | 2012-08-10 | 2012-08-10 | Step-shaped interstitial structure of gas compressor of aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201220397519 CN202914391U (en) | 2012-08-10 | 2012-08-10 | Step-shaped interstitial structure of gas compressor of aircraft engine |
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| Publication Number | Publication Date |
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| CN202914391U true CN202914391U (en) | 2013-05-01 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 201220397519 Withdrawn - After Issue CN202914391U (en) | 2012-08-10 | 2012-08-10 | Step-shaped interstitial structure of gas compressor of aircraft engine |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102817873A (en) * | 2012-08-10 | 2012-12-12 | 势加透博（北京）科技有限公司 | Ladder-shaped gap structure for gas compressor of aircraft engine |
-
2012
- 2012-08-10 CN CN 201220397519 patent/CN202914391U/en not_active Withdrawn - After Issue
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102817873A (en) * | 2012-08-10 | 2012-12-12 | 势加透博（北京）科技有限公司 | Ladder-shaped gap structure for gas compressor of aircraft engine |
| CN102817873B (en) * | 2012-08-10 | 2015-07-15 | 势加透博（北京）科技有限公司 | Ladder-shaped gap structure for gas compressor of aircraft engine |
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| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| AV01 | Patent right actively abandoned |
Granted publication date: 20130501 Effective date of abandoning: 20150715 |
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| RGAV | Abandon patent right to avoid regrant |