CN103523208A - Method for controlling plasma flow of wing lift-rising apparatus - Google Patents
Method for controlling plasma flow of wing lift-rising apparatus Download PDFInfo
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- CN103523208A CN103523208A CN201310476784.0A CN201310476784A CN103523208A CN 103523208 A CN103523208 A CN 103523208A CN 201310476784 A CN201310476784 A CN 201310476784A CN 103523208 A CN103523208 A CN 103523208A
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Abstract
The invention relates to a method for controlling plasma flow of wing lift-rising apparatus, and the technical scheme is characterized in that: plasma exciters are laid on the suction face of the leading edge slat, the suction face of the main wing trailing edge and the suction face of the trailing edge flap. The plasma exciter is composed of electrodes in asymmetric layout and separated by an insulating material, the electrode on the upper surface of the insulating material is connected with the high-voltage terminal of a pulse plasma power supply, and the electrode on lower surface is earthed. The laying positions of the plasma exciters are at the leading edge and the trailing edge of a separation point. When the leading edge slat and the trailing edge flat are deployed during the launching and the landing of an airplane, a signal of 'on' is emitted and the plasma exciters are turned on; and when the leading edge slat and the trailing edge flat are fold by the airplane, a signal of 'off' is emitted and the plasma exciters are turned off. The method helps to effectively control the flow separation of the suction faces of the leading edge slat, the main wing trailing edge and the trailing edge flap when the airplane is in a launching or landing state, substantially improving wing lift-drag ratio and maximum lift coefficient, and further improve the launching and landing weight of the airplane and shorten the running distance.
Description
Technical field
The active Flow Control technology that the invention belongs to aircraft pneumatic design field, relates to wing method for controlling plasma flow, is specifically related to a kind of high-lift wing system method for controlling plasma flow.
Background technology
High-lift wing system design, is that large-scale airplane in transportation category and short field aircraft improve take-off weight, shorten landing ground run distance, strengthen the adaptive gordian technique in airport, has determined taking off, climb, go around and landing data of aircraft.The important breakthrough of high lift device designing technique, will play key effect to the aeroperformance of new generation large aircraft, is also worldwide the research topic of a very challenging property.Take Boeing 777 as example, and lift coefficient increases by 0.1, can reduce by 1 °, pitch attitude angle, shortens alighting gear height, and then loss of weight 636kg; While taking off, 1ift-drag ratio increases by 0.1%, can increase used load 1272kg.
The main design criteria of high lift device comprises: the maximum lift coefficient meeting design requirement, the stall angle meeting design requirement, the 1ift-drag ratio etc. that takes off meeting design requirement.The airplane design of a new generation, requirement is simpler, actv. high lift device provides better aeroperformance, and reaches better economic performance.Adopt comparatively complicated high lift device (as Boeing-737 adopts 3 sections of wing flaps) to compare with early stage seating plane, modern high lift device design all adopts single seam wing flap scheme as far as possible, to reduce driver train complexity, alleviates structural weight, strengthens reliability.High lift device leading edge slat and trailing edge flap Feng road parameter optimization, can effectively increase lift coefficient, but also can increase drag coefficient.The principal element that restriction is taken off, landing state lower wing 1ift-drag ratio improves is the flow separation that high lift device suction surface occurs, and how effectively to suppress flow separation, is a gordian technique in novel high-lift wing system design.
The passive flow control technique such as vortex generator, can effectively suppress flow separation, improve lift coefficient, but can bring the increase of resistance simultaneously, and be difficult to adapt to different mode of operations, therefore, adopting active Flow Control technology is an important development tendency.Active Flow Control can be improved by local microvariations macroscopical aerodynamic characteristic, realizes aircraft drag reduction lift-rising, driving engine pushing expansion surely, using as a new degree of freedom, incorporates the pneumatic design of following aircraft, driving engine.By active Flow Control, can under the prerequisite that does not change drag coefficient, increase lift coefficient, promote high lift device performance.The active Flow Control method of studying at present, mainly contain boundary-layer and blow/inhale, the mode such as synthesizing jet-flow, can effectively suppress flow separation, improve lift coefficient, but speed of response and the limited bandwidth of existing energisation mode are difficult to meet the needs that aircraft wing high lift device of new generation designs.
Plasma flow control is a kind of novel active flow control method based on plasma excitation, utilize aerial discharge to produce the controlled disturbance that in the process of plasma, stream field applies, change speed and the vorticity boundary condition in flow field, and then realize flowing and control, its main advantage is: do not change wing aerodynamic configuration, when not working, wing aeroperformance is not had to harmful effect; Exciter response speed is fast, in microsecond magnitude, can Rapid Implementation flows and controls; Excitation frequency band is wide, between 10Hz-100kHz, can meet the needs that different conditions current downflow is controlled.On high-lift wing system, apply plasma flow control technology, will significantly improve maximum lift coefficient and 1ift-drag ratio.
Summary of the invention
The technical matters solving
For fear of the deficiencies in the prior art part, the present invention proposes a kind of high-lift wing system method for controlling plasma flow.
Technical scheme
A method for controlling plasma flow, is characterized in that step is as follows:
Step 1: lay respectively Plasma Actuator at leading edge slat suction surface, main wing trailing edge suction surface and trailing edge flap suction surface; Described Plasma Actuator forms by being insulated the asymmetric layout electrode that material separates, and the upper surface electrode of insulating material is connected with pulse plasma power high-pressure side, the lower surface electrode ground connection of insulating material; Described Plasma Actuator lay position at leading edge and the trailing edge of separation point;
Step 2: when taking off, land while opening leading edge slat, trailing edge flap, send " opening " signal, open plasma exiter, the upper surface electrode of Plasma Actuator is connected with ground terminal with the high-pressure side of pulse plasma power respectively with lower surface electrode, the output voltage waveforms of power supply is nanosecond pulse, voltage is between 1～30kV, and discharge frequency is between 1～5kHz;
When aircraft is packed up leading edge slat, trailing edge flap, send " pass " signal, close Plasma Actuator.
The upper surface electrode width d of described insulating material
1be 1～10mm, the lower surface electrode width d of insulating material
2be 1～10mm, the separation delta d of upper surface electrode and lower surface electrode is 0～2mm, the thickness h of upper surface electrode and lower surface electrode
ebe 0.001～0.035mm, insulation thickness h
dbe 0.2～2mm.
Described upper surface electrode and lower surface electrode all adopt copper electrode.
Described interelectrode insulating material adopts polyimide.
Beneficial effect
The high-lift wing system method for controlling plasma flow that the present invention proposes, the separated flow that can effectively suppress leading edge of a wing slat suction surface, main wing trailing edge suction surface and trailing edge flap suction surface, and the exciter response time is short, bandwidth, for improving wing maximum lift coefficient and 1ift-drag ratio, there is vital function.Numerical simulation and wind tunnel experiment show, when speed of incoming flow is 0.2Ma, in wing main wing leading edge and trailing edge flap suction surface, apply nanosecond pulse plasma excitation, and when the angle of attack is 16 °, airfoil lift coefficient increases 22.6%, and drag coefficient reduces 19.3%.
Major advantage is that effect is rapid, excitation frequency band is wide, there is no movable parts, simple in structure, energy consumption is lower, can solve that other flow control meanses can not solve or the insoluble problems such as excitation of controlling quickly, for realizing mobile control of real-time adaptive, provide good basic condition, can effectively suppress high lift device flow separation, improve wing maximum lift coefficient and 1ift-drag ratio, and then improve the taking off of aircraft, landing weight, shorten ground run distance.
Accompanying drawing explanation
Fig. 1: be the conceptual scheme that Plasma Actuator lays at leading edge of a wing slat suction surface, main wing trailing edge suction surface and trailing edge flap suction surface;
S1, S2 first and second group exiter for laying at leading edge slat suction surface, M1, M2 first and second group exiter for laying at main wing trailing edge suction surface, F1, F2, F3 first, second and third group Plasma Actuator for laying at trailing edge flap suction surface;
Fig. 2: the sparking voltage-current waveform that is Plasma Actuator while connecting nanosecond pulse plasma electrical source;
Fig. 3: be to lay the photo of Plasma Actuator at the leading edge of a wing and trailing edge flap;
Fig. 4: be the lift coefficient comparison diagram that applies nanosecond pulse plasma excitation front and back;
Fig. 5: be the drag coefficient comparison diagram that applies nanosecond pulse plasma excitation front and back;
1 is upper surface electrode, and 2 is insulating material, and 3 is lower surface electrode, and 4 is plasma
The specific embodiment
Now in conjunction with the embodiments, the invention will be further described for accompanying drawing:
The control method of the present embodiment, lays Plasma Actuator at leading edge slat suction surface, main wing trailing edge suction surface and trailing edge flap suction surface.Plasma Actuator forms by being insulated the asymmetric layout electrode that material separates, and insulating material upper surface electrode is connected with pulse plasma power high-pressure side, lower surface electrode ground connection.Plasma Actuator lay position at leading edge and the trailing edge of separation point.Take off, land while opening leading edge slat, trailing edge flap, send " opening " signal, open plasma exiter, when aircraft is packed up leading edge slat, trailing edge flap, sends " pass " signal, closes Plasma Actuator.The output wave shape of pulse plasma power is sinusoidal wave pulse or nanosecond pulse high pressure.The pulsed plasma excitation producing after plasma excitation electrode two ends apply high pressure, with leading edge slat suction surface, main wing trailing edge suction surface and the coupling of trailing edge flap suction surface separated flow, flow separation be can effectively suppress, wing maximum lift coefficient and 1ift-drag ratio improved.
Specific as follows:
A, at the leading edge of a wing and trailing edge flap, lay Plasma Actuator;
Plasma Actuator lay position, as Fig. 1: laying two groups of Plasma Actuators at leading edge slat suction surface, is the first Plasma Actuator S1, the second Plasma Actuator S2; At main wing trailing edge suction surface, laying two groups of Plasma Actuators, is C grade gas ions exiter M1, the 4th Plasma Actuator M2; At trailing edge flap suction surface, laying three groups of Plasma Actuators, is the 5th Plasma Actuator F1, the 6th Plasma Actuator F2, the 7th Plasma Actuator F3;
The layout of Plasma Actuator as shown in Figure 3, upper surface electrode width d
1for 3mm(can choose between 1-10mm), lower surface electrode width d
2for 3mm(can choose between 1-10mm), upper and lower surface electrode separation delta d is that 1mm(can choose between 0-2mm), upper and lower surface electrode thickness h
ebeing 0.018mm(can choose between 0.001-0.035mm), insulation thickness h
dfor 0.5mm(can choose between 0.2-2mm).Electrode adopts copper electrode, between electrode, with insulating material, separates, and insulating material is polyimide.
C, when taking off, land while opening leading edge slat, trailing edge flap, send " opening " signal, open plasma exiter, the upper surface electrode of Plasma Actuator is connected with ground terminal with the high-pressure side of pulse plasma power respectively with lower surface electrode, the output voltage waveforms of power supply is nanosecond pulse, voltage is between 1～30kV, and discharge frequency is between 1～5kHz, and sparking voltage-electric current as shown in Figure 4.
When aircraft is packed up leading edge slat, trailing edge flap, send " pass " signal, close Plasma Actuator.
In low speed wind tunnel quantitative appraisal nanosecond pulse plasma excitation suppress the effect of high-lift wing system flow separation, as shown in Figure 5, speed of incoming flow is that 0.2Ma, the angle of attack are 16 °, when driving voltage, frequency are respectively 10kV, 3kHz, airfoil lift coefficient increases 22.6%, and drag coefficient reduces 19.3%.
Claims (4)
1. a high-lift wing system method for controlling plasma flow, is characterized in that step is as follows:
Step 1: lay respectively Plasma Actuator at leading edge slat suction surface, main wing trailing edge suction surface and trailing edge flap suction surface; Described Plasma Actuator forms by being insulated the asymmetric layout electrode that material separates, and the upper surface electrode of insulating material is connected with pulse plasma power high-pressure side, the lower surface electrode ground connection of insulating material; Described Plasma Actuator lay position at leading edge and the trailing edge of separation point;
Step 2: when taking off, land while opening leading edge slat, trailing edge flap, send " opening " signal, open plasma exiter, the upper surface electrode of Plasma Actuator is connected with ground terminal with the high-pressure side of pulse plasma power respectively with lower surface electrode, the output voltage waveforms of power supply is nanosecond pulse, voltage is between 1～30kV, and discharge frequency is between 1～5kHz;
When aircraft is packed up leading edge slat, trailing edge flap, send " pass " signal, close Plasma Actuator.
2. high-lift wing system method for controlling plasma flow according to claim 1, is characterized in that: the upper surface electrode width d of described insulating material
1be 1～10mm, the lower surface electrode width d of insulating material
2be 1～10mm, the separation delta d of upper surface electrode and lower surface electrode is 0～2mm, the thickness h of upper surface electrode and lower surface electrode
ebe 0.001～0.035mm, insulation thickness h
dbe 0.2～2mm.
3. high-lift wing system method for controlling plasma flow according to claim 1, is characterized in that: described upper surface electrode and lower surface electrode all adopt copper electrode.
4. high-lift wing system method for controlling plasma flow according to claim 1, is characterized in that: described interelectrode insulating material adopts polyimide.
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Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN106184720A (en) * | 2016-08-08 | 2016-12-07 | 北京航空航天大学 | Lift-drag ratio enhancement mode wing based on Plasma Actuator and gurney flap |
| CN107037824A (en) * | 2017-06-09 | 2017-08-11 | 中国航空工业集团公司哈尔滨空气动力研究所 | A kind of all-wing aircraft model transverse control device and control method |
| CN107444614A (en) * | 2017-09-08 | 2017-12-08 | 中国民航大学 | Suitable for the aerofoil flexibility plasma drag reduction paster of small-sized Fixed Wing AirVehicle |
| CN107645822A (en) * | 2017-09-18 | 2018-01-30 | 中国人民解放军空军工程大学 | A kind of air intake duct shock wave control device and method based on the electric discharge of surface magnetic control arc |
| CN107651027A (en) * | 2017-10-30 | 2018-02-02 | 吉林大学 | A kind of automobile tail separation method of flow control and damping device based on plasma excitation |
| CN107914865A (en) * | 2017-11-27 | 2018-04-17 | 西北工业大学 | The virtual dynamic bionic apparatus and method of plasma for the leading edge of a wing |
| CN108928503A (en) * | 2018-07-27 | 2018-12-04 | 中国人民解放军空军工程大学 | Unmanned plane plasma flow control flight test TT&C system |
| CN110040235A (en) * | 2019-05-05 | 2019-07-23 | 中国人民解放军国防科技大学 | Active and passive combined flow control method and device |
| JP2019189045A (en) * | 2018-04-25 | 2019-10-31 | 株式会社Ｓｕｂａｒｕ | Wing structure, method for controlling wing structure, and aircraft |
| CN113120218A (en) * | 2021-05-25 | 2021-07-16 | 中国人民解放军空军工程大学 | Composite plasma excitation method for flow separation control of high-subsonic wing |
| CN113200141A (en) * | 2021-05-26 | 2021-08-03 | 西安理工大学 | Suction type lift increasing device based on Laval tubular plasma |
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Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106184720A (en) * | 2016-08-08 | 2016-12-07 | 北京航空航天大学 | Lift-drag ratio enhancement mode wing based on Plasma Actuator and gurney flap |
| CN107037824A (en) * | 2017-06-09 | 2017-08-11 | 中国航空工业集团公司哈尔滨空气动力研究所 | A kind of all-wing aircraft model transverse control device and control method |
| CN107037824B (en) * | 2017-06-09 | 2023-10-24 | 中国航空工业集团公司哈尔滨空气动力研究所 | Transverse control device and control method for flying wing model |
| CN107444614A (en) * | 2017-09-08 | 2017-12-08 | 中国民航大学 | Suitable for the aerofoil flexibility plasma drag reduction paster of small-sized Fixed Wing AirVehicle |
| CN107645822A (en) * | 2017-09-18 | 2018-01-30 | 中国人民解放军空军工程大学 | A kind of air intake duct shock wave control device and method based on the electric discharge of surface magnetic control arc |
| CN107651027A (en) * | 2017-10-30 | 2018-02-02 | 吉林大学 | A kind of automobile tail separation method of flow control and damping device based on plasma excitation |
| CN107914865B (en) * | 2017-11-27 | 2020-09-25 | 西北工业大学 | Plasma virtual dynamic bionic device and method for wing leading edge |
| CN107914865A (en) * | 2017-11-27 | 2018-04-17 | 西北工业大学 | The virtual dynamic bionic apparatus and method of plasma for the leading edge of a wing |
| JP2019189045A (en) * | 2018-04-25 | 2019-10-31 | 株式会社Ｓｕｂａｒｕ | Wing structure, method for controlling wing structure, and aircraft |
| CN108928503A (en) * | 2018-07-27 | 2018-12-04 | 中国人民解放军空军工程大学 | Unmanned plane plasma flow control flight test TT&C system |
| CN108928503B (en) * | 2018-07-27 | 2021-07-20 | 中国人民解放军空军工程大学 | Unmanned aerial vehicle plasma flow control flight test measurement and control system |
| CN110040235A (en) * | 2019-05-05 | 2019-07-23 | 中国人民解放军国防科技大学 | Active and passive combined flow control method and device |
| CN113120218A (en) * | 2021-05-25 | 2021-07-16 | 中国人民解放军空军工程大学 | Composite plasma excitation method for flow separation control of high-subsonic wing |
| CN113200141A (en) * | 2021-05-26 | 2021-08-03 | 西安理工大学 | Suction type lift increasing device based on Laval tubular plasma |
| CN113200141B (en) * | 2021-05-26 | 2022-05-31 | 西安理工大学 | Suction type lift increasing device based on Laval tubular plasma |
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Application publication date: 20140122 |