CN107869482B - The sharpening leading edge structure and design method of a kind of transonic fan stage leaf top primitive blade profile - Google Patents
The sharpening leading edge structure and design method of a kind of transonic fan stage leaf top primitive blade profile Download PDFInfo
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- CN107869482B CN107869482B CN201711003800.9A CN201711003800A CN107869482B CN 107869482 B CN107869482 B CN 107869482B CN 201711003800 A CN201711003800 A CN 201711003800A CN 107869482 B CN107869482 B CN 107869482B
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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
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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
- F04D29/666—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
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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
- 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
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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
- 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 methods of transonic fan stage leaf top primitive blade profile, first on the suction surface of the original primitive blade profile in leaf top, obtain the datum mark apart from original leading edge certain length；It crosses datum mark and is one and original suction surface straightway in a certain angle；It is attached between straightway and reset pressure face with the lesser arc section of radius, carries out rounding off with rounding between the straightway other end and former suction surface, and the axial position of new and old blade inlet edge is kept to remain unchanged.Since the leading edge to original primitive blade profile has carried out sharpening processing, leading edge effective radius is reduced, shock wave can be postponed, lift-off occurs, conducive to the stall margin for promoting fan.Simultaneously as the introducing of sharpening structure, so that the dilatational wave at blade inlet edge is divided into two parts, a part is the dilatational wave that leading-edge radius causes, another part is the dilatational wave that straightway and suction surface transition position cause, conducive 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 industries, are 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 technique
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
Revolving speed has to select higher revolving speed, causes 2 gases flowing at the top of fan blade to be in Supersonic state, on the top of fan blade
Portion generates 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 is sharply increased, when wavefront Mach number is greater than
When 1.5, shock wave should be 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 stretches to the upper left side of plane cascade, referred to as overhanging shock wave.Before air-flow after detached shock wave 4 flows through blade
It is divided into two when edge fillet, flows to blade back and leaf basin respectively, then just forms preceding stationary point 9.Along the air-flow of blade back flowing, flowing through
Accelerate to be supersonic speed again when leading edge and blade back curved surface, and a series of dilatational waves 7 are issued by blade back surface.A portion expansion
Wave 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 are also most strong.Experiment and theory
It calculates result of study to show to be influenced overhanging shock-wave attenuation quickly by dilatational wave 7, reaches unlimited distance, overhanging shock strength
Decrease is zero.And conduit shock wave is close to one of normal shock wave, and its blade back flow field that will affect adjacent blades, so that blade back surface
Boundary-layer separates, and the loss of bleed leaf grating increases.It can be seen that the position that conduit shock wave is located at primitive blade profile suction surface is straight
It connects and affects blade suction surface separation initial point position, to influence the flow condition of entire leaf grating.
In transonic fan stage internal rotor, conduit shock wave is in the position on suction surface and the close phase of fan operating condition
It closes.As shown in figure 3, conduit shock wave 10 is in runner compared with downstream position, with fan when fan propeller work is in blocked state
The increase of back pressure, normal shock wave elapse forward, and in high efficiency dotted state, conduit shock wave 11 is in compared with front position, work as fan work
When near nearly stall point, conduit shock wave 12 is pushed out blade path, forms detached shock wave.It is examined from the angle of fan stability
Consider, if the Forward of conduit shock wave can be postponed, is conducive to the steady operation nargin for improving fan.
Fan tip primitive blade profile leading edge thickness is bigger, and shock wave is more easy to happen lift-off, and 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 conducive to control shock loss, simultaneously
Conducive to the comprehensive stability nargin for improving fan.
In order to control shock wave, fan leaf top primitive blade inlet edge is designed more and more thinner, but works as primitive blade inlet edge wedge
When shape angle remains unchanged, 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 blade is easy to happen flutter, 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.
Summary of the invention
For in existing advanced aviation turbofan engine, inlet fan is transonic speed, and there are intense shock waves for vane tip, is
Control shock wave, blade tip thickness are answered smaller and smaller, and after leaf top thickness degree reduces to a certain extent, will affect the rigid of blade
Degree is easy so that local flutter occurs for fan blade, and in order to guarantee 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 guarantees that the structure of fan blade is complete
Whole property, the present invention provides the sharpening leading edge structures and design method of a kind 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 guarantee the structural integrity of fan blade
Property.
The technical solution that the present invention is taken by realization its technical purpose are as follows:
The leading edge design method of a kind of transonic fan stage leaf top primitive blade profile, which is characterized in that the method 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 position of blade profile is identical；
SS4. at the reference point location, rounding is carried out between the straightway and the suction surface of original primitive blade profile
Transition processing makes rounding off between the two.
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 diameter.It is weaker to the control action of shock wave if distance between the two is too short, if this apart from too long, will affect wind
The structural intergrity of fan leaf, therefore the distance should control 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, it will lead to that new leading-edge radius is smaller, the control effect of shock wave be more obvious, but will lead to the straightway and former suction surface
Junction Curvature varying it is excessive, influence the efficiency of blade.The angle is smaller, and it is bigger to will lead to new leading-edge radius, to shock wave
Control effect it is 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 influencing shock wave.New leading-edge radius reduces, and is conducive to control
Shock strength processed is conducive to control profile loss so that shock-wave spot moves back.
Preferably, by adjusting the distance between leading edge point of the datum mark and original primitive blade profile and straightway
Angle between original primitive blade profile suction surface, is adjusted the radius of the arc section.
Preferably, in step SS3 when setting the arc section, it should ensure that the deflection angle at new and old up-front metal geometry angle
No more than 5 °.Due to having carried out sharpening processing to original primitive blade profile leading edge, up-front metal geometry angle can be made to deflect,
The presence of the deflection angle will increase leading edge import geometry angle, and in the case where identical incoming flow, effective angle of attack compares protophyll at leaf top
Type has certain reduction, the stall margin of Yi Tisheng 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, which is characterized in that the leading edge design side of the sharpening leading edge above-mentioned transonic fan stage leaf top primitive blade profile according to the present invention
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
With above-mentioned sharpening leading edge structure of the 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
Lift-off occurs, conducive to the stall margin for promoting fan.Simultaneously as the introducing of " sharpening " structure, so that swollen at blade inlet edge
Swollen wavelength-division is cut into two parts, and a part is the dilatational wave that leading-edge radius causes, and another part is straightway and suction surface transition position
The dilatational wave of initiation, this process is also conducive to reduce the Mach number of conduit shock wave front, conducive to the efficiency for improving primitive blade profile.
Detailed description of the invention
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 condition inferior lobe top grid；
Fig. 4 is the sharpening leading edge structure schematic diagram of transonic fan stage leaf top primitive blade profile of the invention；
The influence schematic diagram of sharpening blade Fig. 5 of the 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
Diameter 18, straight line 19, detached shock wave 20, attached shock 21, dilatational wave 22, dilatational wave 23, dilatational wave 24
Specific embodiment
The present invention will be further described in detail below with reference to the embodiments, following embodiment be explanation of the invention and
The invention is not limited to following embodiments.
As shown in figure 4, the sharpening leading edge structure of transonic fan stage leaf top primitive blade profile of the invention, as follows
It arrives: first on the suction surface of the original primitive blade profile in leaf top, obtaining the datum mark apart from original leading edge certain length 17；Later,
Cross the straightway 19 that this datum mark is one Yu original suction surface in a certain angle 13；Then, straightway 19 and reset pressure face it
Between be attached with the arc section that Radius 14 is less than original leading-edge radius, 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 axis of new and old blade inlet edge
It is remained unchanged to position.
Wherein, the length 17 between datum mark and the leading edge point of 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 will affect fan blade apart from too long
Structural intergrity, therefore length 17 should control between 4-6 times of old leading-edge radius 15.
The angle 13 that straightway 19 and former suction surface are formed, angle 13 is bigger, and new leading-edge radius 14 is smaller, the control to shock wave
Effect processed is more obvious, but makes straightway 19 and the junction Curvature varying of former suction surface big, influences the efficiency of blade.Angle 13
Smaller, new leading-edge radius 14 is bigger, unobvious to the control effect of shock wave.It is proposed that the value range of angle 13 7 °~
Between 12 °.
New leading-edge radius 14 is the important geometric parameter for influencing shock wave.New leading-edge radius 14 reduces, and is conducive to control shock wave
Intensity is conducive to control profile loss so that 21 position of shock wave moves back.It is proposed that new leading-edge radius 14 and old leading-edge radius 15
Ratio value range between 0.4-0.6.It can be to new leading-edge radius 14 by adjusting length 17 and the value of angle 13
Size be adjusted.
Sharpening processing is carried out to leading edge, up-front metal geometry angle can be made to deflect, deflection angle 16 should not exceed 5 °.
Due to the presence of deflection angle 15, import geometry angle can be made to increase, effective angle of attack phase in the case where identical incoming flow, at leaf top
There are certain reduction, the stall margin of Yi Tisheng primitive blade profile than prophyll type.
As shown in figure 5, " sharpening " blade reduces leading edge effective radius, shock wave 21 can be postponed, lift-off occurs, be conducive to be promoted
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, a part is the dilatational wave 24 that leading-edge radius causes, and a part is the dilatational wave 22 that straight line 19 and suction surface transition position cause,
This process is also conducive to reduce the Mach number of conduit shock wave front, conducive to the efficiency for improving primitive blade profile.
In addition, it should be noted that, the specific embodiments described in this specification, the shape of parts and components are named
Title etc. can be different.It is all that the equivalent or simple change that the structure, feature and principle is done is conceived according to the invention patent, include
In in the scope of protection of the patent of the present invention.Those skilled in the art can be to described specific embodiment
It does various modifications or additions or is substituted in a similar manner, without departing from structure of the invention or surmount this power
Range defined in sharp claim, is within the scope of protection of the invention.
Claims (8)
1. a kind of leading edge design method of transonic fan stage leaf top primitive blade profile, which is characterized in that the method includes following steps
It is rapid:
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 of diameter 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 position it is identical；
SS4. at the reference point location, to progress rounding transition between the straightway and the suction surface of original primitive blade profile
Processing, makes rounding off between the two.
2. leading edge design method according to claim 1, which is characterized 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. leading edge design method according to claim 1, which is characterized in that the suction of the straightway and original primitive blade profile
Angle between power face is 7 °~12 °.
4. leading edge design method according to claim 1, which is characterized in that the arc section radius is original primitive blade profile
0.4-0.6 times of leading-edge radius.
5. leading edge design method according to claim 1, which is characterized in that by adjusting the datum mark and original primitive
Angle between the distance between leading edge point of blade profile and the straightway and the suction surface of original primitive blade profile, to described
The radius of arc section is adjusted.
6. leading edge design method according to claim 1, which is characterized in that in step SS3 when setting the arc section,
It should ensure that the deflection angle at new leading edge and old up-front metal geometry angle is not more than 5 °.
7. a kind of sharpening leading edge structure of transonic fan stage leaf top primitive blade profile, which is characterized in that the sharpening leading edge is according to power
Benefit requires the leading edge design method of 1 to 6 described in any item transonic fan stage leaves top primitive blade profile to obtain.
8. a kind of transonic fan stage blade, which is characterized in that the fan blade has fan leaf top primitive blade profile, the fan
Leaf top primitive blade profile has sharpening leading edge structure as claimed in claim 7.
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| Application Number | Priority Date | Filing Date | Title |
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| 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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| 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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| CN107869482B true CN107869482B (en) | 2019-03-19 |
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| US11506059B2 (en) | 2020-08-07 | 2022-11-22 | Honeywell International Inc. | Compressor impeller with partially swept leading edge surface |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| 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 |
| JP6030853B2 (en) * | 2011-06-29 | 2016-11-24 | 三菱日立パワーシステムズ株式会社 | Turbine blade and axial turbine |
| CN103790639B (en) * | 2013-12-26 | 2016-11-16 | 北京理工大学 | A kind of turbine petiolarea blade inlet edge edge strip repaiies type method |
| CN105332952B (en) * | 2015-11-02 | 2017-10-31 | 南京航空航天大学 | A kind of adjustable stator design method of small camber |
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