CN103867489A - Gas compressor blade, gas compressor and aircraft engine - Google Patents
Gas compressor blade, gas compressor and aircraft engine Download PDFInfo
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- CN103867489A CN103867489A CN201210544238.1A CN201210544238A CN103867489A CN 103867489 A CN103867489 A CN 103867489A CN 201210544238 A CN201210544238 A CN 201210544238A CN 103867489 A CN103867489 A CN 103867489A
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Abstract
The invention relates to a gas compressor blade, a corresponding gas compressor and an aircraft engine. The gas compressor blade aims to solve the problem of air flow separation of an existing gas compressor blade. By designing the gas compressor blade so that the gas compressor blade is provided with a local thickness reducing part, namely a thickness curve is provided with at least three inflection points, the technical effect of reducing blade air flow separation is achieved successfully.
Description
Technical field
The present invention relates to compressor blade and there is gas compressor and the corresponding aeroengine of this blade.
Background technique
Modern civil aviation motor is towards low oil consumption, low cost, low emission target in development, and this just requires gas compressor will have high pressure ratio, high efficiency feature.Those skilled in the art will be understood that term " compressor efficiency " refers to that in diffuser grid, airspeed declines, and its kinetic energy reducing has the scaling factor that how much is converted into adiabatic compression merit.Realizing high pressure ratio, high efficiency just requires to need to have good pneumatic design performance and structural type in gas compressor design.After gas compressor one dimension, S2 is through-flow design, can obtain flow parameter on meridian plane.No matter be the flow performance of realizing on meridian plane, or realization expands to three-dimensional from meridian plane, or realize and all can realize by blade shape construction compared with the high efficiency flow of small loss.High-pressure compressor blade is divided into stator blade and rotor blade, rotor blade is to air work, make gas in moving blades, obtain kinetic energy and pressure energy flowing through, the kinetic energy that static cascade obtains air-flow from moving blades is further converted to pressure energy, increases, air-flow is entered in rear one-level simultaneously with suitable angle with the pressure that improves whole level.Blade shape construction is that the aerodynamic parameter of one dimension, S2 design is carried out to flow organization to realize pressure ratio, flow, efficiency important step up to standard.Therefore blade shape construction has important effect in whole high-pressure compressor design, and the performance of gas compressor is somewhat dependent upon the design level of blade.Fig. 1 (a) and (b) show respectively the schematic perspective view of compressor blade and dissect the schematic diagram of the primitive blade profile of acquisition along the plane A-A of the streamline in Fig. 1 (a).Arrow C represents schematic air current flow direction.Those skilled in the art will be understood that term " primitive blade profile " refers to the two-dimensional geometry leaf grating on streamline.
Past is to construct Compressor airfoil with some specific mean camber line types and thickness distribution take laboratory data as basic traditional design method, and this is suitable in the situation that low speed and load are not high.Under high-speed case, adopt in the past two circular arcs (DCA), many circular arcs (MCA) blade profile, they have also obtained certain effect.Many circular arc profiles are for traditional design method, break through to a certain extent the relative fixed mode of specific mean camber line type and thickness distribution, more can adapt to neatly the requirement of different designs condition, part has guaranteed to realize under high-speed case flowing of lower loss.Be no matter traditional blade profile or two circular arc, many circular arc profiles, its citation form as shown in Figure 2.As shown in Figure 2, conventional blade profile 10 ' relates to following parameter: leading edge l; Trailing edge t; Suction surface s; Pressure side p; The mean camber line m extending between leading edge l and trailing edge t; Thickness h, its be along perpendicular to the mean camber line m(string of a musical instrument also can) direction tolerance, self-suction face s is to the distance between pressure side p; The angle of attack+i/-i, it refers to the difference of blade profile metal angle and flow angle.
For modern age high-performance aviation axial flow compressor and fan, due to the further raising of speed of incoming flow and load, with the good flow condition of the very difficult acquisition of blade profile of conventional method design, reach the object that reduces loss.
In order to realize high pressure ratio and the designing requirement of high nargin gas compressor, blade profile camber will certainly increase, if adopt conventional blade profile, the separation of air-flow easily occurs afterbody suction surface.Term " airflow breakaway " refers to that air-flow can not be attached to blade profile surface and produce a series of vortex structure.Schematic diagram when Fig. 3 schematically shows zero-incidence (i=0).As shown in the figure, there is the vortex structure w of airflow breakaway in afterbody suction surface.
As shown in Figure 4, air-flow with large positive incidence (+i) impact when blade, air-flow can accelerate in blade profile leading edge suction face portion, profile loss increases, and even produces shock wave, loss strengthens, and is more very just to start to separate in this region.
As shown in Figure 5, with large negative angle of attack, (while i) impacting blade, similarly, air-flow can accelerate in blade profile leading edge pressure side part air-flow, and profile loss increases, and even produces shock wave, and loss strengthens, and is more very just to start to separate in this region.
Therefore, there is the demand for Novel pressure mechanism of qi blade design.
Summary of the invention
The present invention proposes a kind of novel compressor blade design proposal, it solves the airflow breakaway problem of existing compressor blade well.
According to an aspect of the present invention, a kind of compressor blade is provided, the primitive blade profile of this blade has: suction surface, pressure side, leading edge, trailing edge and mean camber line, and there is thickness that measure on perpendicular to described mean camber line direction, from described suction surface to described pressure side, wherein, the continuous variation of the thickness from described leading edge to described trailing edge and thickness curve have at least three flex points.
In one embodiment, described primitive blade profile has corresponding to the local thickness of thickness curve flex point and reduces region, and the thickness that described local thickness reduces region reduces compared to the thickness of its near zone; Reduce region in described local thickness, described thickness due to one of in described pressure side and described suction surface or both simultaneously indents reduce.
According to an embodiment, described thickness curve has three flex points.In a possible embodiment, described thickness starts to increase to first gradually the first flex point from described leading edge, is then decreased to gradually Second Inflexion Point, and then increases gradually the 3rd flex point, and then is decreased to gradually described trailing edge.In a possible embodiment, the distance of described Second Inflexion Point and described leading edge is the 1/5-1/3 of chord length, and/or the distance of described Second Inflexion Point and described trailing edge is the 1/5-1/3 of chord length.In a possible embodiment, the thickness at described Second Inflexion Point place is the 60%-80% of the first flex point place thickness, and/or the thickness at described Second Inflexion Point place is the 60%-80% of the 3rd flex point place thickness.
According to another embodiment of the invention, described thickness curve has five flex points.In a possible embodiment, described thickness starts to increase to first gradually the first flex point from described leading edge, and then be decreased to gradually Second Inflexion Point, and then increase to gradually the 3rd flex point, and then be decreased to gradually the 4th flex point, and then increase to gradually the 5th flex point, and then be decreased to gradually described trailing edge.In a possible embodiment, the distance of the distance of described Second Inflexion Point and described leading edge and/or described the 4th flex point and described trailing edge is the 1/5-1/3 of chord length.In a possible embodiment, the thickness at described Second Inflexion Point place is the 60%-80% of the first flex point place thickness, and/or the thickness at described the 4th flex point place is the 60%-80% of the 5th flex point place thickness.
According to another aspect of the present invention, also provide a kind of gas compressor, it comprises compressor blade as above.
According to a further aspect of the invention, also provide a kind of aeroengine, it comprises gas compressor as above.
According to the present invention, by the primitive blade design of blade being become the continuous variation of thickness and thickness curve from described leading edge to described trailing edge there are at least three flex points, the local thickness that is particularly designed to have corresponding to thickness curve flex point reduces region, has solved well the problem of airflow breakaway.
Accompanying drawing explanation
Accompanying drawing shows nonrestrictive exemplary embodiment of the present invention, and itself and explanatory note one are used from principle of the present invention is made an explanation, and in institute's drawings attached, same or similar reference character represents same or similar feature; Wherein:
Fig. 1 (a) and (b) show respectively the schematic perspective view of existing compressor blade and dissect the schematic diagram of the primitive blade profile of acquisition along the plane A-A of streamline.
Fig. 2 schematically shows conventional blade profile, shows its some parameters.
Fig. 3-5 schematically show conventional blade profile, in zero-incidence, large positive incidence, large negative angle of attack downstream, the situation separating occur.
Fig. 6 schematically shows existing vane thickness distribution curve and the contrast of vane thickness distribution curve according to an embodiment of the invention.
Fig. 7 schematically shows blade profile according to an embodiment of the invention.
Fig. 8-Figure 10 schematically shows the airflow state under zero-incidence, large positive incidence, large negative angle of attack according to the blade design of the present invention of Fig. 7.
Figure 11 (a) shows the simulation result in the flow field of traditional blade profile.
Figure 11 (b) shows the simulation result in the flow field of blade profile according to an embodiment of the invention.
Embodiment
Those skilled in the art are known according to the conventional mass conservation law in aerodynamics field: G=ρ VA, and wherein, G is mass flow rate, ρ is gas density, V is airspeed, flows (for example shown in Figure 9) for primitive blade profile, and the airflow area between A two adjacent blades.DA/A is area change gradient, according to above-mentioned formula, and in the situation that density p is certain ,-dA/A=dV/V.In gas compressor, kinetic energy transfers pressure energy to.It is the variation that the variation of speed transfers pressure to.There is following mode,
accordingly, be understood that, in the time of mobile between two adjacent blades of air-flow (reference example primitive blade profile as shown in Figure 9 flows), adverse pressure gradient and the basic positive correlation of area change gradient.In blade grid passage flows, in the time that area variable gradient is large, adverse pressure gradient is large, easily produces airflow breakaway.Exactly because why conventional blade profile easily separates and just starts to separate and to start to separate air-flow adverse pressure gradient from leading edge pressure side region when large negative angle of attack very large caused from leading edge suction surface when large positive incidence in blade profile afterbody suction surface.From blade profile geometrically, be because blade profile arrives throat region in leading edge portion, and trailing edge regional gas stream circulation " area change rate " is too fast caused.
Conventional blade profile thickness distributes as shown in the curve M in Fig. 6, this thickness distribution of only having a flex point M1 that is similar to parabolic, for the large or large positive incidence in this turning and large negative angle of attack, it is the appearance that easily causes segregation phenomenon, wherein transverse axis represents mean camber line or string of a musical instrument length B, longitudinal axis representative thickness h, " thickness " of the present invention h is along perpendicular to mean camber line m(or the string of a musical instrument) direction tolerance, self-suction face s is to the distance between pressure side p.
Thinking of the present invention is just from the correlation angle of air-flow and blade profile area change, invented a kind of blade profile slowing down at leading edge and/or trailing edge region area variance ratio.Namely blade profile thickness is not the thickness distribution form that tradition only exists a flex point M1, but from blade profile leading edge to trailing edge, thickness curve has at least three flex points, for example, from blade profile leading edge to trailing edge, vane thickness can first reduce to increase again, and then reducing again increases again.Curve N in Fig. 6 shows a kind of exemplary according to thickness distribution curve of the present invention, five flex points of its total N1-N5.Corresponding blade profile for example can be as shown in the reference character 10 in Fig. 7.The thickness distribution of this form has slowed down air-flow flow area variance ratio in leading edge portion and trailing edge part, can effectively reduce airflow breakaway.
After adopting blade profile of the present invention, be conducive to reduce the pressure gradient on suction surface, pressure side, air-flow can effectively be attached to blade profile surface, the generation of separation delay; Can reduce the airspeed of blade profile leading edge portion, improve shock wave structure, reduce loss, improve compressor efficiency simultaneously.Adopt blade profile of the present invention flow schematic diagram afterwards as shown in Fig. 8-10.Figure 11 (a) shows the simulation result in the flow field of traditional blade profile.Figure 11 (b) shows according to the simulation result in the flow field of blade profile of the present invention.Described simulation result is to have calculated by the Numeca of CFD simulation software.By comparison diagram 11(a) and Figure 11 (b), can obviously see effect of the present invention, in other words, can effectively suppress airflow breakaway according to the blade profile of the present invention's design, and then can promote operating range, efficiency, the nargin of gas compressor.
Here it should be noted that, " area of reduced thickness " of the present invention is actually " part " area of reduced thickness corresponding to thickness curve flex point, the region that thickness reduces compared with himself near zone, does not comprise for example, this region of " thickness reduces " on the whole from thickness maximum region (flex point M1) to trailing edge in common blade for example.
According to the blade profile embodiment in Fig. 7, the thickness that the local thickness of this blade profile 10 reduces region reduce due to described pressure side and described suction surface simultaneously indent produce.In other mode of execution, described local thickness reduces region also can be by indent and producing separately one of in described pressure side and suction surface.
In addition, the thickness of the blade profile 10 in Fig. 7 has five flex points, being described thickness starts to increase to first gradually for example thickness biggest place nearest apart from leading edge of the first flex point N1(from described leading edge), and then be decreased to gradually Second Inflexion Point N2(for example apart from the minimum place of the nearest thickness of leading edge), and then increase to gradually the 3rd flex point N3(for example apart from leading edge time near thickness biggest place, the maximum ga(u)ge place of whole blade profile in other words), and then be decreased to gradually the 4th flex point N4(for example apart from the minimum place of the nearest thickness of trailing edge), and then increase to gradually for example thickness biggest place nearest apart from trailing edge of the 5th flex point N5(), and then be decreased to gradually described trailing edge, curve N in Fig. 6 is shown.That is to say, blade profile reduces region in leading edge and trailing edge Dou You local thickness.
In an optional embodiment, Second Inflexion Point N2 is the chord length 1/5-1/3 of (being connected the straightway of leading edge and trailing edge) with distance and/or described the 4th flex point of described leading edge with the distance of described trailing edge.In another optional embodiment, the thickness at described Second Inflexion Point N2 place is the 60%-80% of the first flex point N1 place thickness, and/or the thickness at described the 4th flex point N4 place is the 60%-80% of the 5th flex point N5 place thickness.
As the replacement scheme of above-mentioned five flex points, in a further embodiment, blade profile thickness curve can only have three flex points, for example, only have local thickness one of in leading edge and trailing edge and reduce region.Now, described thickness for example starts to increase to first gradually the first flex point N1 from described leading edge, is then decreased to gradually Second Inflexion Point N2, and then increases gradually the 3rd flex point N3, and then is decreased to gradually described trailing edge.Certainly, understandable, there is local thickness and reduce the situation in region for leading edge only, for example the 3rd flex point N3 can represent the position of maximum ga(u)ge.The situation that has local thickness and reduce region for trailing edge only, for example the first flex point N1 can represent the position of maximum ga(u)ge.In an optional embodiment, for example there is local thickness and reduce the situation in region for leading edge only, the distance of described Second Inflexion Point N2 and described leading edge is the 1/5-1/3 of chord length.The situation that has local thickness and reduce region for trailing edge only, the distance of described Second Inflexion Point N2 and described trailing edge can be also the 1/5-1/3 of chord length.In another optional embodiment, the thickness at described Second Inflexion Point N2 place is the 60%-80% of the first flex point N1 place thickness, and/or the thickness at described Second Inflexion Point N2 place is the 60%-80% of the 3rd flex point N3 place thickness.
Those skilled in the art should be understood that, according to the present invention, on the basis of aforementioned instruction, flex point quantity, corner position, local thickness reduce the parameters such as the thickness size in region and can for example, design and choose according to actual conditions (Mach number, corner) etc.
The present invention also provides the gas compressor that comprises foregoing compressor blade and the aeroengine that comprises this gas compressor.
Claims (12)
1. a compressor blade, the primitive blade profile of this blade has: suction surface, pressure side, leading edge, trailing edge and mean camber line, and there is thickness that measure on perpendicular to described mean camber line direction, from described suction surface to described pressure side, it is characterized in that, the thickness from described leading edge to described trailing edge changes continuously and thickness curve has at least three flex points.
2. compressor blade as claimed in claim 1, is characterized in that, described primitive blade profile has corresponding to the local thickness of thickness curve flex point and reduces region, and the thickness that described local thickness reduces region reduces compared to the thickness of its near zone; Reduce region in described local thickness, described thickness due to one of in described pressure side and described suction surface or both simultaneously indents reduce.
3. compressor blade as claimed in claim 1, is characterized in that, described thickness curve has three flex points.
4. compressor blade as claimed in claim 3, is characterized in that, described thickness starts to increase to first gradually the first flex point from described leading edge, is then decreased to gradually Second Inflexion Point, and then increases gradually the 3rd flex point, and then is decreased to gradually described trailing edge.
5. compressor blade as claimed in claim 4, is characterized in that, the distance of described Second Inflexion Point and described leading edge is the 1/5-1/3 of chord length, and/or the distance of described Second Inflexion Point and described trailing edge is the 1/5-1/3 of chord length.
6. compressor blade as claimed in claim 4, is characterized in that, the thickness at described Second Inflexion Point place is the 60%-80% of the first flex point place thickness, and/or the thickness at described Second Inflexion Point place is the 60%-80% of the 3rd flex point place thickness.
7. compressor blade as claimed in claim 1, is characterized in that, described thickness curve has five flex points.
8. compressor blade as claimed in claim 7, it is characterized in that, described thickness starts to increase to first gradually the first flex point from described leading edge, and then be decreased to gradually Second Inflexion Point, and then increase to gradually the 3rd flex point, and then be decreased to gradually the 4th flex point, and then increase to gradually the 5th flex point, and then be decreased to gradually described trailing edge.
9. compressor blade as claimed in claim 8, is characterized in that, the distance of the distance of described Second Inflexion Point and described leading edge and/or described the 4th flex point and described trailing edge is the 1/5-1/3 of chord length.
10. compressor blade as claimed in claim 8, is characterized in that, the thickness at described Second Inflexion Point place is the 60%-80% of the first flex point place thickness, and/or the thickness at described the 4th flex point place is the 60%-80% of the 5th flex point place thickness.
11. 1 kinds of gas compressors, is characterized in that, comprise the compressor blade as described in aforementioned arbitrary claim.
12. 1 kinds of aeroengines, is characterized in that, comprise gas compressor as claimed in claim 11.
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| CN201210544238.1A CN103867489B (en) | 2012-12-14 | 2012-12-14 | Compressor blade, compressor and aero-engine |
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| CN201210544238.1A CN103867489B (en) | 2012-12-14 | 2012-12-14 | Compressor blade, compressor and aero-engine |
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| CN103867489A true CN103867489A (en) | 2014-06-18 |
| CN103867489B CN103867489B (en) | 2017-06-16 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106089801A (en) * | 2016-08-11 | 2016-11-09 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of compressor blade formative method |
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| JPH09256997A (en) * | 1996-03-25 | 1997-09-30 | Senshin Zairyo Riyou Gas Jienereeta Kenkyusho:Kk | Moving blade for axial flow compressor |
| US6116856A (en) * | 1998-09-18 | 2000-09-12 | Patterson Technique, Inc. | Bi-directional fan having asymmetric, reversible blades |
| US20060275134A1 (en) * | 2005-06-01 | 2006-12-07 | Honda Motor Co., Ltd. | Blade of axial flow-type rotary fluid machine |
| US20070158495A1 (en) * | 2003-04-04 | 2007-07-12 | Hubbard Adrian A | High lift and high strength aerofoil |
| CN102483072A (en) * | 2009-09-04 | 2012-05-30 | 西门子公司 | Compressor blade for an axial compressor |
| CN102753835A (en) * | 2009-12-07 | 2012-10-24 | 法雷奥热系统公司 | Fan propeller, in particular for a motor vehicle |
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2012
- 2012-12-14 CN CN201210544238.1A patent/CN103867489B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH09256997A (en) * | 1996-03-25 | 1997-09-30 | Senshin Zairyo Riyou Gas Jienereeta Kenkyusho:Kk | Moving blade for axial flow compressor |
| US6116856A (en) * | 1998-09-18 | 2000-09-12 | Patterson Technique, Inc. | Bi-directional fan having asymmetric, reversible blades |
| US20070158495A1 (en) * | 2003-04-04 | 2007-07-12 | Hubbard Adrian A | High lift and high strength aerofoil |
| US20060275134A1 (en) * | 2005-06-01 | 2006-12-07 | Honda Motor Co., Ltd. | Blade of axial flow-type rotary fluid machine |
| CN102483072A (en) * | 2009-09-04 | 2012-05-30 | 西门子公司 | Compressor blade for an axial compressor |
| CN102753835A (en) * | 2009-12-07 | 2012-10-24 | 法雷奥热系统公司 | Fan propeller, in particular for a motor vehicle |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106089801A (en) * | 2016-08-11 | 2016-11-09 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of compressor blade formative method |
| CN106089801B (en) * | 2016-08-11 | 2018-08-24 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of compressor blade formative method |
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