CN105090098A - Transonic fan rotor blade - Google Patents
Transonic fan rotor blade Download PDFInfo
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- CN105090098A CN105090098A CN201410205351.6A CN201410205351A CN105090098A CN 105090098 A CN105090098 A CN 105090098A CN 201410205351 A CN201410205351 A CN 201410205351A CN 105090098 A CN105090098 A CN 105090098A
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Abstract
The invention discloses a transonic fan rotor blade which is a wide-chord three-dimensional composite skewed-swept blade. The transonic fan rotor blade comprises a blade root section, a blade tip section, a front blade edge, a tail blade edge and a blade chord length part. The blade chord length part is lengthened, the consistence is increased, an inlet blade tip of the front blade edge is sweptback, the root of the tail blade edge is bent, and the bending changes of all the sections in the height direction of the blade are strictly controlled through a controllable diffusion blade profile. Based on a verification conclusion, under the condition that the pressure ratio is increased by 7 percent, the flow is increased by 8 percent, the maximum relative mach number of the 95-percent blade height section of the fan rotor blade is reduced from 1.57 to 1.43, the maximum relative mach number of the 50-percent blade height section of the fan rotor blade is reduced from 1.46 to 1.41, the wave shocking intensity is obviously lowered, a boundary layer obtained after wave shocking is thinned, flow losses are greatly reduced, the efficiency of a fan rotor is improved by 5 percent, and the steady operation margin of a fan of a design point is improved by 1 percent.
Description
Technical field
The present invention relates to a kind of fan rotor blade; Refer more particularly to a kind of high load blade tip sweepback transonic fan rotor blade, belong to impeller machinery technical field.
Background technique
Fan is as one of the core component of aviation gas turbine Duct-Burning Turbofan key, and its performance impact the quality of aviation gas turbine Duct-Burning Turbofan.Along with aeropropulsion system is to the future development of high thrust weight ratio, more and more higher requirement is proposed to the performance of fan.In order to improve pressure ratio, current aviation gas turbine Duct-Burning Turbofan many employings transonic fan.So-called transonic fan refers to fan propeller root incoming flow relative Mach number subsonics, tip incoming flow relative Mach number Supersonic.Along with the raising of pressure ratio, in the middle part of fan rotor blade and cross section, tip, the numerical value of relative Mach number is also more and more higher, causes the shock strength in passage also to increase gradually.The peak value relative Mach number in fan propeller blade tip cross section had brought up to 1.5 by less than 1.4 of the past, even higher.On the one hand, fan propeller can utilize shock wave to realize high pressure ratio, and on the other hand, due to shock-boundary interaction can make shock wave after boundary layer thicken rapidly, even too high shock strength can cause the flowing after shock wave to occur being separated, and loss increases.Add the increase of suction surface and pressure side pressure reduction, Tip Clearance also can increase, and causes flow losses to increase, and fan efficiency declines.
Therefore, must design the blade profile of applicable high load transonic fan by sophistication, the reduction flow losses of trying one's best while improving pressure ratio, to improve the efficiency of fan.
A kind of stator upper end curved blade is proposed in the patent CN102927050A " a kind of end bent blades for improving gas compressor working stability " of Harbin Turbine Factory Co., Ltd., the curved blade profile of end of stator upper end offsets 0mm ~ 7.53mm relative in the middle part of stator to suction surface side, Duan Wan district accounts for 7% of stator whole height, make the inflow of air-flow and flow out the blade profile curve of blade of more fitting, with the angle automatching in gas compressor main flow region, improve the working stability of gas compressor.The end bent blades mentioned in this patent is only applicable to the stator blade of high load gas compressor.
A kind of centrifugal blade of sweepforward is proposed in the patent CN103148016A " centrifugal compressor of blade sweepforward and turbosupercharger " of Tsing-Hua University, by blade is carried out sweepforward process, reduce the load at leading edge blade tip place, thus improve the performance under centrifugal compressor small flow.Sweepforward blade can reduce the load at leading edge blade tip place, but the structure stability of rotor blade can be caused to decline, and easily causes the vibration problem of blade.
Summary of the invention
The object of the invention is to adopt low aspect ratio, controlled diffusion for the problems referred to above, plunder shape and hold a kind of transonic fan rotor blade of curved Technology design.
(1) technical problem that will solve
The present invention needs to solve the problem that existing aviation gas turbine Duct-Burning Turbofan thrust can not meet aircraft requirement, little for fan rotor blade flow, pressure ratio is low, inefficient problem, by improve fan rotor blade design, make flow improve 7%, pressure ratio improves 7%, and efficiency improves 2%.
(2) technological scheme
For achieving the above object, the present invention by the following technical solutions: a kind of transonic fan rotor blade of the present invention is a kind of three-dimensional compound flexural tensile elastic modulus of wide string, comprises blade root cross section, blade tip cross section, blade inlet edge, blade trailing edge and blade chord length.Wherein blade chord length lengthens, denseness increases, the import blade tip sweepback of blade inlet edge, blade trailing edge root end curved, and adopt controlled diffusion airfoil strictly to control to change along the camber in each cross section, leaf height direction.
Wherein in order to improve load, fan rotor blade chord length is lengthened 30%; Blade tip denseness is the ratio of blade chord length and pitch, increases to 1.2 by 1.0.
Wherein in order to realize the control to flowing, adopt controlled diffusion airfoil, maximum camber position is moved to blade tip gradually afterwards by blade root, root maximum camber chordwise location is 50%, be 65% (reaching extreme value) in 90% leaf eminence maximum camber chordwise location, 90% to 100% leaf high maximum camber chordwise location moves forward gradually, and blade tip maximum camber chordwise location is 60%.
Wherein in order to reduce the maximum relative Mach number in blade tip cross section, reducing shock loss, improving the structure stability of blade, the blade tip sweepback of blade inlet edge simultaneously.
Wherein in order to reduce tail, reduce loss, the root of blade exit trailing edge does the curved design of end.
Compared with prior art, the invention has the beneficial effects as follows:
Curve guide impeller draws through checking: when pressure ratio improves 7%, flow improves 8%, the maximum relative Mach number in fan rotor blade 95% leaf height cross section have decreased to 1.43 by 1.57, the 50% maximum relative Mach number in leaf height cross section have decreased to 1.41 by 1.46, shock strength significantly reduces, and the boundary layer after shock wave is thinning, and flow losses reduce greatly, the efficiency of fan propeller improves 5%, and the stable operation nargin of design point fan improves 1%.
Accompanying drawing explanation
Fig. 1 is blade profile schematic diagram.
2-blade tip cross section, 1-blade root cross section 3-blade inlet edge 4-blade trailing edge 5-blade tip sweepback 14-blade chord length
Fig. 2 is the right elevation of Fig. 1.
6-root end is curved
Fig. 3 embodiment relative Mach number isopleth cloud atlas
12-shock wave 13-blade rotary direction, 9-Mach number isopleth 10-Mach number numerical value 11-boundary layer, 8-95% leaf height cross section, 7-50% leaf height cross section 15-pitch 16-tail
Embodiment
The present invention is further described by specific embodiment below in conjunction with Fig. 1 ~ Fig. 3.Below implement just descriptive, be not determinate, protection scope of the present invention can not be limited with this.
From Fig. 1 ~ Fig. 3, a kind of transonic fan rotor blade of the present invention is a kind of three-dimensional compound flexural tensile elastic modulus of wide string, comprises blade root cross section 1, blade tip cross section 2, blade inlet edge 3, blade trailing edge 4 and blade chord length 14.Described blade chord length 14 lengthens, denseness increase, the import blade tip sweepback 5 of blade inlet edge 3, the root end curved 6 of blade trailing edge 4, and adopt controlled diffusion airfoil strictly to control to change along the camber in each cross section, leaf height direction.
Wherein in order to improve load, fan rotor blade chord length 14 is lengthened 30%, blade tip denseness is blade chord length 14 and the ratio of pitch 15, increases to 1.2 by 1.0.
Because the Mach number of incoming flow is higher, in the design process, in order to reduce the shock strength in blade tip cross section 2, edge 3 adopts blade tip sweepback 5 to design in front of the blade, blade tip sweepback 5 design can reduce the Mach-number component of incoming flow edge 3 normal direction in front of the blade, effectively reduce shock wave 12 intensity, reduce shock loss, and blade tip sweepback 5 design can improve the structure stability of blade.
Wherein in order to realize the control to flowing, adopt controlled diffusion airfoil, maximum camber position is moved to blade tip gradually afterwards by blade root, blade root cross section 1 maximum camber chordwise location is 50%, be 65% (reaching extreme value) in 90% leaf eminence maximum camber chordwise location, 90% to 100% leaf high maximum camber chordwise location moves forward gradually, and blade tip cross section 2 maximum camber chordwise location is 60%.
In conjunction with the result of Flow Field Calculation, wherein in order to reduce tail 16, reduce loss, blade trailing edge 4 does curved 6 designs of root end.
Concrete implementation methods is: based on streamline curvature method through-flow reconstruct, continue with any moulding of three dimendional blade, then three-dimensional flow field numerical calculation is carried out to the result of moulding, again result of calculation is fed back to the iterative process forming closed loop in through-flow reconstruct and any moulding of blade, until obtain satisfied aeroperformance, and be equipped with Structural Strength Design, vibration, surge margin checking computations.
Design result obtains checking, and when pressure ratio improves 7%, flow improves 8%, and efficiency improves 5%, and the stable operation nargin of design point fan improves 1%.Can be seen by the change of Mach number numerical value 10 in the cloud atlas of Mach number isopleth 9, the maximum relative Mach number in fan rotor blade 95% leaf height cross section 8 have decreased to 1.43 by 1.57, the maximum relative Mach number in 50% leaf height cross section 7 have decreased to 1.41 by 1.46, shock wave 12 intensity significantly reduces, boundary layer 11 after shock wave is thinning, and flow losses reduce greatly.
Claims (5)
1. a transonic fan rotor blade is a kind of three-dimensional compound flexural tensile elastic modulus of wide string, comprise blade root cross section (1), blade tip cross section (2), blade inlet edge (3), blade trailing edge (4) and blade chord length (14), it is characterized in that: blade chord length (14) lengthening, denseness increase, the import blade tip sweepback (5) of blade inlet edge (3), the root end curved (6) of blade trailing edge (4), and adopt controlled diffusion airfoil strictly to control to change along the camber in each cross section, leaf height direction.
2. a kind of transonic fan rotor blade according to claim 1, is characterized in that: fan rotor blade chord length (14) lengthens 30%.
3. a kind of transonic fan rotor blade according to claim 1, is characterized in that: blade tip denseness increases to 1.2 by 1.0.
4. a kind of transonic fan rotor blade according to claim 1, is characterized in that: edge (3) adopts blade tip sweepback (5) design in front of the blade.
5. a kind of transonic fan rotor blade according to claim 1, it is characterized in that: adopt controlled diffusion airfoil, maximum camber position is moved to blade tip gradually afterwards by blade root, blade root cross section (1) maximum camber chordwise location is 50%, be 65% in 90% leaf eminence maximum camber chordwise location, 90% to 100% leaf high maximum camber chordwise location moves forward gradually, and blade tip cross section (2) maximum camber chordwise location is 60%.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410205351.6A CN105090098A (en) | 2014-05-09 | 2014-05-09 | Transonic fan rotor blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410205351.6A CN105090098A (en) | 2014-05-09 | 2014-05-09 | Transonic fan rotor blade |
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| Publication Number | Publication Date |
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| CN105090098A true CN105090098A (en) | 2015-11-25 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN201410205351.6A Pending CN105090098A (en) | 2014-05-09 | 2014-05-09 | Transonic fan rotor blade |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113153816A (en) * | 2021-03-29 | 2021-07-23 | 北京航空航天大学 | Controllable deformation fan and design method thereof |
| CN114673686A (en) * | 2022-04-15 | 2022-06-28 | 苏州浪潮智能科技有限公司 | Design method of fan and corresponding fan |
| WO2024002212A1 (en) * | 2022-06-30 | 2024-01-04 | 中国航发商用航空发动机有限责任公司 | Anti-spin blade and manufacturing method therefor, aviation engine, and aircraft |
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| CN101173680A (en) * | 2007-11-29 | 2008-05-07 | 北京航空航天大学 | Forward-sweeping big and small blade transonic impeller and axial flow air compressor and inclined flow air compressor |
| EP2080909A1 (en) * | 2006-11-02 | 2009-07-22 | Mitsubishi Heavy Industries, Ltd. | Transonic airfoil and axial flow rotary machine |
| CN103423193A (en) * | 2013-04-18 | 2013-12-04 | 哈尔滨汽轮机厂有限责任公司 | Secondary first-stage blade for transonic compressor on high-speed gas turbine |
| KR101346085B1 (en) * | 2011-12-23 | 2013-12-31 | 한국항공우주연구원 | Transonic Compressor Rotor Design Method, And Transonic Compressor Rotor |
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2014
- 2014-05-09 CN CN201410205351.6A patent/CN105090098A/en active Pending
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP2080909A1 (en) * | 2006-11-02 | 2009-07-22 | Mitsubishi Heavy Industries, Ltd. | Transonic airfoil and axial flow rotary machine |
| CN101173680A (en) * | 2007-11-29 | 2008-05-07 | 北京航空航天大学 | Forward-sweeping big and small blade transonic impeller and axial flow air compressor and inclined flow air compressor |
| KR101346085B1 (en) * | 2011-12-23 | 2013-12-31 | 한국항공우주연구원 | Transonic Compressor Rotor Design Method, And Transonic Compressor Rotor |
| CN103423193A (en) * | 2013-04-18 | 2013-12-04 | 哈尔滨汽轮机厂有限责任公司 | Secondary first-stage blade for transonic compressor on high-speed gas turbine |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113153816A (en) * | 2021-03-29 | 2021-07-23 | 北京航空航天大学 | Controllable deformation fan and design method thereof |
| CN113153816B (en) * | 2021-03-29 | 2022-12-06 | 北京航空航天大学 | Controllable deformation fan and design method thereof |
| CN114673686A (en) * | 2022-04-15 | 2022-06-28 | 苏州浪潮智能科技有限公司 | Design method of fan and corresponding fan |
| CN114673686B (en) * | 2022-04-15 | 2024-01-26 | 苏州浪潮智能科技有限公司 | Fan design method and corresponding fan |
| WO2024002212A1 (en) * | 2022-06-30 | 2024-01-04 | 中国航发商用航空发动机有限责任公司 | Anti-spin blade and manufacturing method therefor, aviation engine, and aircraft |
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Application publication date: 20151125 |
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