CN107762973A - Steady blade and its trailing edge groove forming method are expanded in a kind of compressor angular region - Google Patents
Steady blade and its trailing edge groove forming method are expanded in a kind of compressor angular region Download PDFInfo
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- CN107762973A CN107762973A CN201710980936.9A CN201710980936A CN107762973A CN 107762973 A CN107762973 A CN 107762973A CN 201710980936 A CN201710980936 A CN 201710980936A CN 107762973 A CN107762973 A CN 107762973A
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- trailing edge
- angular region
- compressor
- blade
- expanded
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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
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
- F04D29/384—Blades characterised by form
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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
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
- F04D29/384—Blades characterised by form
- F04D29/386—Skewed blades
-
- 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
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
- F04D29/388—Blades characterised by construction
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The present invention is to provide a kind of compressor angular region to expand steady blade and its trailing edge groove forming method.Trailing edge groove is provided with compressor blade root angular region, the trailing edge groove originates in blade trailing edge, width is 10%~30% times of chord length, is highly 0.5%~1.5% times of chord length.The present invention is to form local jet effect in root of blade angular region using compressor blade trailing edge pressure face and suction surface side pressure difference, suppresses the radial migration process of angular region low energy fluid, so as to obtain the compressor blade with wider wide adaptability and stable operation range.
Description
Technical field
The present invention relates to a kind of gas compressor blade chip architecture, specifically a kind of calming the anger with angular region bathtub construction
Machine blade construction.The present invention also relates to be a kind of compressor blade trailing edge groove forming method.
Background technology
Flowing can be efficiently distributed in 60%~80% leaf exhibition position among blade by modern high-performance compressor blade
Place is put, thus the low energy fluid near end wall and angular region turns into the main source that flowing is blocked and lost, and influence compressor
An important factor for stability.Corner separation is a kind of a kind of complicated three-dimensional for being widely present in gas compressor moving blade and stator blade angular region
Flow separation phenomenon, the reduction of raising and aspect ratio in particular with modern multistage compressor load, the angular region stream of moving-stator blade
It is dynamic more complicated.
According to whether extra increase energy, corner separation control method is broadly divided into active control and passive control, commonly used
Active Control Method include boundary layer suction technology, plasma controls etc., its be mainly characterized by by using injection or
The method of other external sources suppresses boundary-layer separation, and its control effect is obvious, strong adaptability, but needs to increase extra means
And control system.Passive control methods need not add any external energy, pass through the geometry of changing section blade path
Improve the flow field structure inside original compressor blade row and expand surely to realize.The typical passive control structure in compressor blade row inside is such as
Vortex generator, wing fence, gap blade etc., passive control methods have the advantages of need not adding extra means, but need
Design passive control structure.The technological difficulties passively controlled essentially consist in：(1) technique that additional accessory structure adds compressor
Complexity, structural reliability reduce, and accessory structure, which comes off, increases the possibility of compressor blade damage；(2) on the other hand, quilt
Dynamic control effect and the change of specific compressor flow field structure are closely related, and same structure is difficult that compressor majority operating mode is all produced
Raw control effect, the heightened awareness dependent on designer to compressor flow field structure.Compressor gap blade technology is near
A kind of compressor flowing passive control methods that year rises, it is some by being set between compressor blade suction surface and pressure face
Individual gap forms jet to blow down suction surface boundary-layer, improves compressor load level and stability.The control effect of this method
Fruit depends on the blade surface pressure of gap both ends present position, is typically chosen the larger blade both sides of pressure difference, therefore is difficult real
Now compressor varying load process is well adapted to.Penetrated in addition, gap blade needs to crack in compressor blade position with realizing
Effect is flowed, all there is high requirement to flow angle, seam shape, roughness, processing technology requires higher, and is possible to drop
The margin of safety of low blades.
The radial migration effect of compressor angular region low energy fluid is the build-in attribute of compressor flow field, is due to low energy stream
Caused by body is accumulated to a certain extent in angular region.Off-design behaviour state is in particularly with high load capacity compressor blade row, radially
Migration effect is especially notable.Large scale concentrates the radial migration of vortex system to cause the low energy stream between end wall and suction surface boundary-layer
Body transports, and forms angular region stall event, and it is the crucial evolutionary process that corner region flow deteriorates.Compressor angular region trailing edge is nearby deposited
It is the major incentive to form stall zone towards suction surface backflow phenomenon in pressure.
The content of the invention
It is an object of the invention to provide a kind of compressor angular region expansion with wider wide adaptability and stable operation range
Steady blade.Expand the trailing edge groove forming method of steady blade the present invention also aims to provide a kind of compressor angular region.
Steady blade is expanded in the compressor angular region of the present invention：Trailing edge groove, the trailing edge are provided with compressor blade root angular region
Groove originates in blade trailing edge, width is 10%~30% times of chord length, is highly 0.5%~1.5% times of chord length.
Steady blade is expanded in the compressor angular region of the present invention to be included：
1st, trailing edge groove is rectangle, ladder type or wave-shaped groove.
2nd, the chamfer dimesion of trailing edge groove is in 0.5mm~1.5mm.
The trailing edge groove forming method that steady blade is expanded in the compressor angular region of the present invention is：
Step 1, turned back according to the diffusion factor of compressor blade row, geometry angle, cascade solidity, free stream Mach number, pass through
Angular region stall criterion determines compressor angular region stall pattern；
Step 2, the width and height for determining according to the locus of compressor angular region stall, yardstick angular region trailing edge bathtub construction
Degree；
Step 3, according to processing technology and intensity requirement, determine bathtub construction chamfer dimesion；
Step 4, numerical simulation is carried out, the modification of trailing edge slot structure parameter is carried out according to numerical simulation result.
The invention provides a kind of gas compressor blade chip architecture with angular region bathtub construction, the side pressure of compressor blade two is utilized
Difference is formed about nature jet in blade trailing edge root to destroy angular region stall condition, so as to improve compressor steady operation side
Boundary.
The present invention is to form part in root of blade angular region using compressor blade trailing edge pressure face and suction surface side pressure difference
Jet effect, suppress the radial migration process of angular region low energy fluid, so as to obtain with wider wide adaptability and steady operation model
The compressor blade enclosed.
The present invention principle be：
Angular region stall event is the major reason for influenceing compressor steady operation, because blade petiolarea transverse-pressure gradient is made
With, end wall and suction surface boundary-layer, in compressor blade row angular region, constantly accumulation inevitably produces stall event, especially when
Stall condition is even more serious when compressor is operated in off-design behaviour.It is outside by introducing to flow the general principle passively controlled
Structure produces extra vortex structure and interfered with prophyll piece petiolarea spiral structure, plays a part of suppressing stall.Therefore, passive control
Method processed must have clear and definite control object (certain class vortex in petiolarea), due to the vortex position, intensity, chi of compressor petiolarea
The parameters such as degree exist with compressor carrying out practically operating mode to be contacted, therefore the position of external structure, parameter etc. are required for basis
Control object carries out Fine design, and general not wide for the scope of application of compressor operating mode.Current flowing passively controls
Method is based primarily upon diagonal regions Passage Vortex and is intervened (vortex generator), and suction surface boundary-layer is blown down near separation initial point
(gap blade) etc..The present invention be directed to originate in the low energy fluid radial migration process of trailing edge suction surface during the stall of angular region
Implement to intervene, destroy angular region stall formation condition from jet effect using angular region trailing edge bathtub construction, reach and expand steady purpose.
The present invention compared with prior art the advantages of：
1) because angular region trailing edge low energy fluid radial migration is the flow phenomenon of angular region stall forming process generally existing, production
Raw position and form are relatively stable, and compared to existing passive control methods, the present invention passes through diagonal regions low energy fluid Radial Flow
Implement to intervene, there is more preferable operating mode applicability.
2) compared to existing flow control method, the present invention need not introduce extra energy device, simple in construction.
3) it is small to the structural modification of former compressor blade compared to existing flow control method, trailing edge bathtub construction, very
Extremely can be directly simple using the tenon structure of original root of blade, technological requirement.
Brief description of the drawings
Fig. 1 a- Fig. 1 b are compressor angular region stall schematic diagrames, and wherein Fig. 1 a are compressor blade row stall group；Fig. 1 b are to calm the anger
Internal structure is rolled into a ball in machine stall.
Fig. 2 a- Fig. 2 b are trailing edge bathtub construction flow field control effect diagrams, and wherein Fig. 2 a are prototype stator blade limiting streamline；
Fig. 2 b are with the stator blade limiting streamline for expanding steady structure.
Fig. 3 is that the structural representation of steady blade is expanded in the compressor angular region of the present invention.
Fig. 4 is Fig. 3 partial sectional view.
Fig. 5 is that the combined effect figure of steady blade is expanded in the compressor angular region of the present invention.
Embodiment
Illustrate below and the present invention is described in more detail.
With reference to Fig. 1 a- Fig. 1 b, Fig. 2 a- Fig. 2 b compressor angular region stall structural representation, angular region stall footpath is substantially determined
To the position of flowing, included according to the design parameter of compressor blade row：Free stream Mach number, diffusion factor, fluid deflection, blade
The parameters such as hub ratio tentatively judge angular region stall form and end wall backflow zone position, using numerical simulation means, it is determined that horizontal two
Secondary stream scope and end wall circulating flow strength, determine respectively on this basis movable vane or stator blade angular region trailing edge bathtub construction width,
The basic physical dimensions such as height.In Fig. 1 a- Fig. 1 b and Fig. 2 a- Fig. 2 b, A represents stall zone, B represents radial direction ring whirlpool, C represents tail
Edge backflow, D represent wall chorista, E represents suction surface separation whirlpool, the slow-witted material trailing edge slot structure jets of F.
The trailing edge groove of the present invention determines as follows：
(1), turned back the parameters such as angle, cascade solidity, free stream Mach number according to the diffusion factor of compressor blade row, geometry,
By angular region stall criterion, compressor angular region stall pattern is determined.
(2) width and height of angular region trailing edge bathtub construction, are determined according to the locus of compressor angular region stall, yardstick
Degree.
(3), according to processing technology and intensity requirement, bathtub construction chamfer dimesion is determined.
(4) slot structure parameter modification, is carried out according to numerical simulation result.
More than 0.5, working speed designs in 60%-90% for step (1), (2), (3), the compressor diffusion factor of (4)
Rotating speed.
With reference to Fig. 3 and Fig. 4, the present invention is using compressor blade trailing edge pressure face and suction surface side pressure difference at root of blade angle
Area forms local jet effect, suppresses compressor angular region stall.In figs. 3 and 4,1 represent wheel rim, 2 represent blade inlet edge, 3
Blade trailing edge is represented, 4 trailing edge groove is represented, 5 represents wheel hub, 7 represents blade profile.The present invention is in compressor stator piece root angular region
Trailing edge groove is provided with, the trailing edge groove originates in blade trailing edge, width is 10%~30% times of chord length, is highly 0.5%~1.5%
Times chord length.Trailing edge groove is rectangle, ladder type or wave-shaped groove.The chamfer dimesion of trailing edge groove is in 0.5mm~1.5mm.
Two kinds of specific axial slot forms (rising as high as the banks and curve slot structure) can be selected.Finally required according to Strength co-mputation,
Determine slot structure chamfering form.
Combined effect of the steady blade in compressor is expanded in the compressor angular region that Fig. 5 gives the present invention, wherein 10 representatives are opened
There are the stator blade of trailing edge groove, 11 to represent movable vane piece.
Claims (4)
1. steady blade is expanded in a kind of compressor angular region, it is characterized in that：Trailing edge groove, the trailing edge are provided with compressor blade root angular region
Groove originates in blade trailing edge, width is 10%~30% times of chord length, is highly 0.5%~1.5% times of chord length.
2. steady blade is expanded in compressor angular region according to claim 1, it is characterized in that：The trailing edge groove be rectangle, ladder type or
Wave-shaped groove.
3. steady blade is expanded in compressor angular region according to claim 1 or 2, it is characterized in that：The chamfer dimesion of trailing edge groove exists
0.5mm~1.5mm.
4. the trailing edge groove forming method of steady blade is expanded in a kind of compressor angular region, it is characterized in that：
Step 1, turned back according to the diffusion factor of compressor blade row, geometry angle, cascade solidity, free stream Mach number, pass through angular region
Stall criterion determines compressor angular region stall pattern；
Step 2, the width and height for determining according to the locus of compressor angular region stall, yardstick angular region trailing edge bathtub construction；
Step 3, according to processing technology and intensity requirement, determine bathtub construction chamfer dimesion；
Step 4, numerical simulation is carried out, the modification of trailing edge slot structure parameter is carried out according to numerical simulation result.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710980936.9A CN107762973B (en) | 2017-10-20 | 2017-10-20 | Compressor corner region stability-expanding blade and trailing edge groove forming method thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710980936.9A CN107762973B (en) | 2017-10-20 | 2017-10-20 | Compressor corner region stability-expanding blade and trailing edge groove forming method thereof |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN107762973A true CN107762973A (en) | 2018-03-06 |
| CN107762973B CN107762973B (en) | 2020-06-16 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN201710980936.9A Active CN107762973B (en) | 2017-10-20 | 2017-10-20 | Compressor corner region stability-expanding blade and trailing edge groove forming method thereof |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110821851A (en) * | 2019-11-22 | 2020-02-21 | 南京航空航天大学 | Multistage axial compressor expands steady structure based on sawtooth trailing edge blade |
| CN111692128A (en) * | 2020-05-21 | 2020-09-22 | 西安交通大学 | Structure is handled to transonic compressor combination machine casket |
| FR3104636A1 (en) * | 2019-12-17 | 2021-06-18 | Safran Aircraft Engines | IMPROVED DESIGN STATOR SECTOR FOR AIRCRAFT TURBOMACHINE |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4720239A (en) * | 1982-10-22 | 1988-01-19 | Owczarek Jerzy A | Stator blades of turbomachines |
| SU1763679A1 (en) * | 1990-12-29 | 1992-09-23 | Московский энергетический институт | Vane cascade of turbine |
| EP2267274A2 (en) * | 2009-06-22 | 2010-12-29 | Rolls-Royce plc | A compressor blade |
| GB2483059A (en) * | 2010-08-23 | 2012-02-29 | Rolls Royce Plc | An aerofoil blade with a set-back portion |
| CN102588337A (en) * | 2012-02-29 | 2012-07-18 | Tcl空调器(中山)有限公司 | Axial wind wheel structure and axial fan |
| US8585350B1 (en) * | 2011-01-13 | 2013-11-19 | George Liang | Turbine vane with trailing edge extension |
| US20150240648A1 (en) * | 2014-02-27 | 2015-08-27 | Rolls-Royce Deutschland Ltd & Co Kg | Group of blade rows |
| CN105240322A (en) * | 2015-11-04 | 2016-01-13 | 北京航空航天大学 | Method for forming S-shaped channels on blade root to control corner separation of compressor stator |
-
2017
- 2017-10-20 CN CN201710980936.9A patent/CN107762973B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4720239A (en) * | 1982-10-22 | 1988-01-19 | Owczarek Jerzy A | Stator blades of turbomachines |
| SU1763679A1 (en) * | 1990-12-29 | 1992-09-23 | Московский энергетический институт | Vane cascade of turbine |
| EP2267274A2 (en) * | 2009-06-22 | 2010-12-29 | Rolls-Royce plc | A compressor blade |
| GB2483059A (en) * | 2010-08-23 | 2012-02-29 | Rolls Royce Plc | An aerofoil blade with a set-back portion |
| US8585350B1 (en) * | 2011-01-13 | 2013-11-19 | George Liang | Turbine vane with trailing edge extension |
| CN102588337A (en) * | 2012-02-29 | 2012-07-18 | Tcl空调器(中山)有限公司 | Axial wind wheel structure and axial fan |
| US20150240648A1 (en) * | 2014-02-27 | 2015-08-27 | Rolls-Royce Deutschland Ltd & Co Kg | Group of blade rows |
| CN105240322A (en) * | 2015-11-04 | 2016-01-13 | 北京航空航天大学 | Method for forming S-shaped channels on blade root to control corner separation of compressor stator |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110821851A (en) * | 2019-11-22 | 2020-02-21 | 南京航空航天大学 | Multistage axial compressor expands steady structure based on sawtooth trailing edge blade |
| FR3104636A1 (en) * | 2019-12-17 | 2021-06-18 | Safran Aircraft Engines | IMPROVED DESIGN STATOR SECTOR FOR AIRCRAFT TURBOMACHINE |
| CN111692128A (en) * | 2020-05-21 | 2020-09-22 | 西安交通大学 | Structure is handled to transonic compressor combination machine casket |
| CN111692128B (en) * | 2020-05-21 | 2021-12-10 | 西安交通大学 | Structure is handled to transonic compressor combination machine casket |
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| Publication number | Publication date |
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| CN107762973B (en) | 2020-06-16 |
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