CN107869482A - A kind of the sharpening leading edge structure and design method of transonic fan stage leaf top primitive blade profile - Google Patents
A kind of the sharpening leading edge structure and design method of transonic fan stage leaf top primitive blade profile Download PDFInfo
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- CN107869482A CN107869482A CN201711003800.9A CN201711003800A CN107869482A CN 107869482 A CN107869482 A CN 107869482A CN 201711003800 A CN201711003800 A CN 201711003800A CN 107869482 A CN107869482 A CN 107869482A
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- blade profile
- leading edge
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- primitive
- primitive blade
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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
- F04D29/26—Rotors specially for elastic fluids
-
- 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
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
- F04D29/666—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps by means of rotor construction or layout, e.g. unequal distribution of blades or vanes
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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
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
- F04D29/667—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps by influencing the flow pattern, e.g. suppression of turbulence
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/302—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor characteristics related to shock waves, transonic or supersonic flow
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
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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 relates to a kind of leading edge design method of transonic fan stage leaf top primitive blade profile, first on the suction surface of the original primitive blade profile in leaf top, obtains the datum mark apart from original leading edge certain length；Cross datum mark and be one and original suction surface straightway in a certain angle；It is attached between straightway and reset pressure face with the less arc section of radius, rounding off is carried out with rounding between the straightway other end and former suction surface, and keeps the axial location of new and old blade inlet edge to keep constant.Due to having carried out sharpening processing to the leading edge of original primitive blade profile, leading edge effective radius is reduced, shock wave can be postponed lift-off occurs, beneficial to the stall margin of lifting fan.Simultaneously as the introducing of sharpening structure so that the dilatational wave at blade inlet edge is divided into two parts, the dilatational wave that a part triggers for leading-edge radius, another part is the dilatational wave that straightway triggers with suction surface transition position, beneficial to the Mach number for reducing conduit shock wave front, the efficiency of raising primitive blade profile.
Description
Technical field
The invention belongs to transonic fan stage blade industry, it is related to the blade construction and design method of a kind of transonic fan stage,
More particularly to a kind of transonic fan stage leaf top primitive blade profile sharpening leading edge structure and design method
Background technology
In the present age advanced aviation turbofan engine, as shown in figure 1, the fan 1 of transonic fan stage rotor is towards high pass
The direction of stream ability, high efficiency and high stability is developed, due to needing to take into account the acting ability of coaxial low-pressure turbine, low-pressure shaft
Rotating speed has to select higher rotating speed, causes 2 gases flowing at the top of fan blade to be in Supersonic state, on the top of fan blade
Portion produces shock wave.Shock wave has certain pressurization first, can effectively improve the pressurization on leaf top, still, shock wave meeting
Certain flow losses are introduced, with the increase of shock front Mach number, shock loss sharply increases, when wavefront Mach number is more than
When 1.5, reply shock wave is controlled.
Shock wave structure in transonic cascade is as shown in Fig. 2 the edge of primitive blade 5 has a detached shock wave 4.This road
The lower half of shock wave 4 stretches to the blade back of adjacent blades, and is generally adjacent to the shape of normal shock wave, referred to as conduit shock wave；Lift-off
The upper semisection of shock wave 4, then the upper left side of plane cascade is stretched to, referred to as overhanging shock wave.Before air-flow after detached shock wave 4 flows through blade
It is divided into two during edge fillet, flows to blade back and leaf basin respectively, then just forms preceding stationary point 9.The air-flow flowed along blade back, is flowing through
Accelerate again when leading edge and blade back curved surface as supersonic speed, and a series of dilatational waves 7 are sent by blade back surface.A portion expands
Ripple 7 and overhanging shock interaction simultaneously make its decrease, a part of dilatational wave and conduit shock interaction, and change the intensity of conduit shock wave.
Conduit shock wave and suction surface point of intersection 8, the local Mach number highest of conduit shock wave front, therefore shock wave is also most strong.Experiment and theory
Calculate result of study to show, influenceed overhanging shock-wave attenuation quickly by dilatational wave 7, reach unlimited distance, overhanging shock strength
Decrease is zero.And conduit shock wave is close to one of normal shock wave, and it can influence the blade back flow field of adjacent blades so that blade back surface
Boundary-layer separates, the loss increase of bleed leaf grating.As can be seen here conduit shock wave be located at primitive blade profile suction surface position it is straight
Connect and affect blade suction surface separation initial point position, so as to influence the flow condition of whole leaf grating.
In transonic fan stage internal rotor, the position that conduit shock wave is on suction surface and the close phase of fan operating condition
Close.As shown in figure 3, when fan propeller is operated in blocked state, conduit shock wave 10 is in runner compared with downstream position, with fan
The increase of back pressure, normal shock wave are elapsed forward, and in high efficiency dotted state, conduit shock wave 11 is in compared with front position, works as fan work
When near nearly stall point, conduit shock wave 12 is pushed out blade path, forms detached shock wave.Examined from the angle of fan stability
Consider, if the reach of conduit shock wave can be postponed, be advantageous to improve the steady operation nargin of fan.
Edge thickness is bigger before the primitive blade profile of fan tip, and the easier generation lift-off of shock wave, shock loss is bigger, and easier
Cause fan unstability.Fan tip primitive blade profile leading edge is smaller, and shock-wave spot more rearward, is advantageous to control shock loss, simultaneously
Beneficial to the comprehensive stability nargin for improving fan.
In order to control shock wave, the more and more thinner of fan leaf top primitive blade inlet edge design, but work as primitive blade inlet edge wedge
When shape angle keeps constant, the blade thickness of primitive blade is also constantly thinned, and after reaching to a certain degree, can reduce the rigidity of blade,
So that flutter easily occurs for blade, therefore, blade inlet edge thickness is thinned and is restricted, and in order to solve this contradiction, invents herein
A kind of blade inlet edge sharpening structure.
The content of the invention
For in existing advanced aviation turbofan engine, inlet fan is transonic speed, and vane tip has intense shock wave, is
After control shock wave, blade tip thickness should be less and less, and leaf top thickness degree reduces to a certain extent, the firm of blade can be influenceed
Degree, easily cause fan blade that local flutter occurs, in order to ensure the structural intergrity of compressor blade, blade tip thickness is not
What can be subtracted is too small, for this technical problem, in order to control the shock wave structure of vane tip, and ensures that the structure of fan blade is complete
Whole property, the invention provides a kind of the sharpening leading edge structure and design method of transonic fan stage leaf top primitive blade profile, in front of the blade
Edge has carried out sharpening processing, it is therefore an objective to can control the shock wave structure of vane tip, and can ensures the structural integrity of fan blade
Property.
The present invention is realizes that the technical scheme that its technical purpose is taken is：
A kind of leading edge design method of transonic fan stage leaf top primitive blade profile, it is characterised in that methods described includes following
Step：
SS1. a datum mark, the datum mark and original are chosen on the suction surface of the original primitive blade profile in transonic fan stage leaf top
The distance between leading edge point of primordium member blade profile is the several times of original primitive blade profile leading-edge radius；
SS2. using the datum mark as starting point, a straightway in an acute angle with the suction surface of original primitive blade profile is done,
The straightway intersects with original primitive blade profile leading edge；
SS3. before being less than original primitive blade profile using Radius between the straightway and the pressure face of original primitive blade profile
The arc section of edge radius smoothly transits, and the arc section is formed as the new leading edge of primitive blade profile, the new leading edge and original primitive
The leading edge axial location of blade profile is identical；
SS4. at the reference point location, to carrying out rounding between the straightway and the suction surface of original primitive blade profile
Transition processing, make rounding off therebetween.
Preferably, the distance between leading edge point of the datum mark and original primitive blade profile is original primitive blade profile leading edge half
4-6 times of footpath.It is weaker to the control action of shock wave if distance therebetween is too short, if this apart from long, can influence wind
The structural intergrity of fan leaf, therefore the distance should be controlled between 4-6 times of original primitive blade profile leading-edge radius.
Preferably, the angle between the straightway and the suction surface of original primitive blade profile is 7 °~12 °.The angle is got over
Greatly, new leading-edge radius can be caused smaller, it is more obvious to the control effect of shock wave, but the straightway and former suction surface can be caused
Junction Curvature varying it is excessive, influence the efficiency of blade.The angle is smaller, new leading-edge radius can be caused bigger, to shock wave
Control effect unobvious.
Preferably, the arc section radius is 0.4-0.6 times of original primitive blade profile leading-edge radius.The arc section is formed
For the new leading edge of primitive blade profile, new leading-edge radius is the important geometric parameter for influenceing shock wave.New leading-edge radius reduces, and is advantageous to control
Shock strength processed so that moved after shock-wave spot, beneficial to control profile loss.
Preferably, the distance between leading edge point by adjusting the datum mark and original primitive blade profile, and straightway
Angle between original primitive blade profile suction surface, is adjusted to the radius of the arc section.
Preferably, in step SS3 when setting the arc section, it should ensure that the deflection angle at the metal geometry angle of new and old leading edge
No more than 5 °.Due to having carried out sharpening processing to original primitive blade profile leading edge, the metal geometry angle of leading edge can be caused to deflect,
The presence of the deflection angle, leading edge import geometry angle can be increased, in the case of identical incoming, effective angle of attack compares protophyll at leaf top
Type has certain reduction, easily lifts the stall margin of primitive blade profile.
According to another aspect of the present invention, a kind of sharpening leading edge knot of transonic fan stage leaf top primitive blade profile is additionally provided
Structure, it is characterised in that the sharpening leading edge according to the present invention above-mentioned transonic fan stage leaf top primitive blade profile leading edge design side
Method obtains.
According to another aspect of the present invention, a kind of transonic fan stage leaf top primitive blade profile, the primitive blade profile are additionally provided
Above-mentioned sharpening leading edge structure with the present invention.
According to another aspect of the present invention, a kind of transonic fan stage blade is additionally provided, the fan blade has this hair
Bright said fans leaf top primitive blade profile.
Compared with the existing technology, the leading edge design method of transonic fan stage leaf top primitive blade profile of the invention, due to original
The leading edge of primordium member blade profile has carried out " sharpening " processing, and the blade after " sharpening " reduces leading edge effective radius, can postpone shock wave
Generation lift-off, beneficial to the stall margin of lifting fan.Simultaneously as the introducing of " sharpening " structure so that swollen at blade inlet edge
Swollen wavelength-division is cut into two parts, and the dilatational wave that a part triggers for leading-edge radius, another part is straightway and suction surface transition position
The dilatational wave of initiation, this process is also beneficial to the Mach number for reducing conduit shock wave front, beneficial to the efficiency for improving primitive blade profile.
Brief description of the drawings
Fig. 1 is existing transonic fan stage rotor structure schematic diagram；
Fig. 2 is the shock wave structure schematic diagram in existing supersonic speed fan leaf top grid；
Fig. 3 is the shock wave form schematic diagram in existing transonic fan stage difference operating mode inferior lobe top grid；
Fig. 4 is the sharpening leading edge structure schematic diagram of the transonic fan stage leaf top primitive blade profile of the present invention；
Influence schematic diagram of the sharpening blade of Fig. 5 present invention to shock wave structure.
Label declaration：Fan blade 1, top primitive blade profile 2, rotary shaft 3, detached shock wave 4, primitive blade 5, stagnation streamline
6, dilatational wave 7, conduit shock wave and suction surface intersection point 8, preceding stationary point 9, blocked state conduit shock wave 10, best efficiency point conduit shock wave
11, nearly stall point conduit shock wave 12, angle 13, new leading-edge radius 14, old leading-edge radius 15, deflection angle 16, length 17, switching half
Footpath 18, straight line 19, detached shock wave 20, attached shock 21, dilatational wave 22, dilatational wave 23, dilatational wave 24
Embodiment
With reference to embodiment, the present invention is described in further detail, following examples be explanation of the invention and
The invention is not limited in following examples.
As shown in figure 4, the sharpening leading edge structure of the transonic fan stage leaf top primitive blade profile of the present invention, as follows
Arrive：First on the suction surface of the original primitive blade profile in leaf top, the datum mark apart from original leading edge certain length 17 is obtained；Afterwards,
Cross the straightway 19 that this datum mark is one and original suction surface in a certain angle 13；Then, straightway 19 and reset pressure face it
Between be less than the arc section of original leading-edge radius with Radius 14 and be attached, and between the other end of straightway 19 and former suction surface
Rounding off is carried out with the rounding of radius 18, arc section forms new blade inlet edge, but needs to keep the axle of new and old blade inlet edge
Keep constant to position.
Wherein, the length 17 between the leading edge point of datum mark and original primitive blade profile, i.e. arc section formed new leading edge with
The length of the intersection point of straightway and suction surface, 4-6 times of about original leading-edge radius 15.Datum mark and original primitive blade profile
Length 17 between leading edge point is weaker to the control action of shock wave apart from too short, and length 17 can influence fan blade apart from long
Structural intergrity, therefore length 17 should be controlled between 4-6 times of old leading-edge radius 15.
The angle 13 that straightway 19 is formed with former suction surface, angle 13 is bigger, and new leading-edge radius 14 is smaller, the control to shock wave
Effect processed is more obvious, but make it that the junction Curvature varying of straightway 19 and former suction surface is big, influences the efficiency of blade.Angle 13
Smaller, new leading-edge radius 14 is bigger, to the control effect unobvious of shock wave.It is proposed that the span of angle 13 7 °~
Between 12 °.
New leading-edge radius 14 is the important geometric parameter for influenceing shock wave.New leading-edge radius 14 reduces, and is advantageous to control shock wave
Intensity so that shock wave 21 moves behind position, beneficial to control profile loss.It is proposed that new leading-edge radius 14 and old leading-edge radius 15
Ratio span between 0.4-0.6.Can be to new leading-edge radius 14 by the value for adjusting length 17 and angle 13
Size be adjusted.
Sharpening processing is carried out to leading edge, the metal geometry angle of leading edge can be caused to deflect, deflection angle 16 not should be greater than 5 °.
Due to the presence of deflection angle 15, import geometry angle can be caused to increase, in the case of identical incoming, the effective angle of attack phase at leaf top
There is certain reduction than prophyll type, easily lift the stall margin of primitive blade profile.
As shown in figure 5, " sharpening " blade reduces leading edge effective radius, shock wave 21 can be postponed lift-off occurs, beneficial to lifting
The stall margin of fan.Simultaneously as the introducing of " sharpening " structure so that the dilatational wave 23 at blade inlet edge is divided into two
Point, the dilatational wave 24 that a part triggers for leading-edge radius a, part is the dilatational wave 22 that straight line 19 triggers with suction surface transition position,
This process is also beneficial to the Mach number for reducing conduit shock wave front, beneficial to the efficiency for improving primitive blade profile.
Furthermore, it is necessary to illustrate, the specific embodiment described in this specification, the shape of its parts and components, it is named
Title etc. can be different.All equivalent or simple changes done according to construction, feature and principle described in inventional idea of the present invention, include
In in the protection domain of patent of the present invention.Those skilled in the art can be to described specific embodiment
Make various modifications or supplement or substituted using similar mode, without departing from structure of the invention or surmount this power
Scope defined in sharp claim, protection scope of the present invention all should be belonged to.
Claims (9)
1. a kind of leading edge design method of transonic fan stage leaf top primitive blade profile, it is characterised in that methods described includes following step
Suddenly：
SS1. a datum mark, the datum mark and original base are chosen on the suction surface of the original primitive blade profile in transonic fan stage leaf top
The distance between leading edge point of first blade profile is the several times of original primitive blade profile leading-edge radius；
SS2. using the datum mark as starting point, a straightway in an acute angle with the suction surface of original primitive blade profile is done, it is described
Straightway intersects with original primitive blade profile leading edge；
SS3. it is less than original primitive blade profile leading edge half using Radius between the straightway and the pressure face of original primitive blade profile
The arc section in footpath smoothly transits, and the arc section is formed as the new leading edge of primitive blade profile, the new leading edge and original primitive blade profile
Leading edge axial location it is identical；
SS4. at the reference point location, to carrying out rounding transition between the straightway and the suction surface of original primitive blade profile
Processing, makes rounding off therebetween.
2. leading edge design method according to claim 1, it is characterised in that before the datum mark and original primitive blade profile
The distance between edge point is 4-6 times of original primitive blade profile leading-edge radius.
3. the leading edge design method according to the claims, it is characterised in that the straightway and original primitive blade profile
Suction surface between angle be 7 °~12 °.
4. the leading edge design method according to the claims, it is characterised in that the arc section radius is original primitive
0.4-0.6 times of blade profile leading-edge radius.
5. the leading edge design method according to the claims, it is characterised in that by adjust the datum mark with it is original
Angle between the distance between leading edge point of primitive blade profile, and the straightway and the suction surface of original primitive blade profile, it is right
The radius of the arc section is adjusted.
6. the leading edge design method according to the claims, it is characterised in that setting the arc section in step SS3
When, it should ensure that the deflection angle at the metal geometry angle of new and old leading edge is not more than 5 °.
7. a kind of sharpening leading edge structure of transonic fan stage leaf top primitive blade profile, it is characterised in that the sharpening leading edge is according to power
Profit requires that the leading edge design method of the transonic fan stage leaf top primitive blade profile described in 1 to 6 any one obtains.
8. a kind of transonic fan stage leaf top primitive blade profile, it is characterised in that the primitive blade profile has cutting described in claim 7
Sharp leading edge structure.
9. a kind of transonic fan stage blade, it is characterised in that the fan blade has the fan leaf top base described in claim 8
First blade profile.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711003800.9A CN107869482B (en) | 2017-10-24 | 2017-10-24 | The sharpening leading edge structure and design method of a kind of transonic fan stage leaf top primitive blade profile |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711003800.9A CN107869482B (en) | 2017-10-24 | 2017-10-24 | The sharpening leading edge structure and design method of a kind of transonic fan stage leaf top primitive blade profile |
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| Publication Number | Publication Date |
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| CN107869482A true CN107869482A (en) | 2018-04-03 |
| CN107869482B CN107869482B (en) | 2019-03-19 |
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| CN201711003800.9A Active CN107869482B (en) | 2017-10-24 | 2017-10-24 | The sharpening leading edge structure and design method of a kind of transonic fan stage leaf top primitive blade profile |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US11506059B2 (en) | 2020-08-07 | 2022-11-22 | Honeywell International Inc. | Compressor impeller with partially swept leading edge surface |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3820918A (en) * | 1972-01-21 | 1974-06-28 | N A S A | Supersonic fan blading |
| US4408957A (en) * | 1972-02-22 | 1983-10-11 | General Motors Corporation | Supersonic blading |
| CN102852560A (en) * | 2011-06-29 | 2013-01-02 | 株式会社日立制作所 | Supersonic turbine moving blade and axial-flow turbine |
| CN103790639A (en) * | 2013-12-26 | 2014-05-14 | 北京理工大学 | Method for edge strip shape modifying of front edge of end area blade of turbine |
| CN105332952A (en) * | 2015-11-02 | 2016-02-17 | 南京航空航天大学 | Small-bend adjustable stator design method |
-
2017
- 2017-10-24 CN CN201711003800.9A patent/CN107869482B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3820918A (en) * | 1972-01-21 | 1974-06-28 | N A S A | Supersonic fan blading |
| US4408957A (en) * | 1972-02-22 | 1983-10-11 | General Motors Corporation | Supersonic blading |
| CN102852560A (en) * | 2011-06-29 | 2013-01-02 | 株式会社日立制作所 | Supersonic turbine moving blade and axial-flow turbine |
| CN103790639A (en) * | 2013-12-26 | 2014-05-14 | 北京理工大学 | Method for edge strip shape modifying of front edge of end area blade of turbine |
| CN105332952A (en) * | 2015-11-02 | 2016-02-17 | 南京航空航天大学 | Small-bend adjustable stator design method |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US11506059B2 (en) | 2020-08-07 | 2022-11-22 | Honeywell International Inc. | Compressor impeller with partially swept leading edge surface |
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| CN107869482B (en) | 2019-03-19 |
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