CN201712784U - Airplane main wing leading-edge separation vortex control mechanism - Google Patents

Abstract

The utility model relates to an airplane main wing leading-edge separation vortex control mechanism which comprises a separation vortex control wing plate rotating shaft and a separation vortex control wing plate; the basic shape of the separation vortex control wing plate is trapezoidal, the shape of the upper surface of the separation vortex control wing plate is matched with that of the upper wing surface, and the separation vortex control wing plate is fixed in the concave surface of the upper wing surface and constitutes a complete airfoil surface altogether with the upper wing surface; the bevel edge on the rear edge of the separation vortex control wing plate takes a wavy shape, the number N of periods of the wavy line is 6-8; the front edge of the separation vortex control wing plate is positioned in the position 10 percent of the local chord length of the wing, and the chordwise width of the separation vortex control wing plate along the wing is 40 percent of the local chord length; the separation vortex control wing plate is matched with the separation vortex control wing plate rotating shaft, and the opening and the closing control to the separation vortex control wing plate are realized through a fly-by-wire system; and the separation vortex control wing plate is opened during the post-stall maneuvering process of the airplane to inhibit the instability and the nonstationarity of wake flow of the wing, therefore, the control of the airplane under the post-stall maneuver condition is realized.

Description

A kind of aircraft host wing leading edge separation vortex control mechanism

Technical field

The present invention relates to aerospace field, is a kind of aircraft host wing leading edge separation vortex control mechanism.

Background technology

Modern opportunity of combat is used guided weapon in a large number, and guided weapon provides locking information by airborne radar, therefore, as early as possible the other side is included in one's own side's radar view or escapes from the key element that the other side's radar view is attacking and defending as early as possible.If aircraft can only carry out the flight under the conventional state, be to turn or climb all to need the long period to finish so, thereby be unfavorable for the locking of airborne radar and lock fixed.Therefore, in close air combat fistfight, aim at enemy plane rapidly and make and oneself be in adversary's aiming all the time beyond the invisible, promptly the enemy is implemented the fast speed attack effectively preserving under the own prerequisite, become the most important operation quality of modern opportunity of combat.Modern combat aircraft driving engine under moderate velocity can also keep very big surplus power, has good accelerating ability, therefore can the attitude of aircraft be made the fast speed adjustment, thereby reach the purpose that moment changes enemy and we's situation surpassing under the big state of angle of attack of self stalling incidence.

The fighter plane angle of attack substantially exceeds its stalling incidence, and under low-speed condition, still can control aspect, thereby the tactical maneuver that changes the sensing of its flying speed and head rapidly is called " post stall maneuver ", be also referred to as " super motor-driven ", at first propose in early 1980s by He Baisite.Behind the big quantity research, He Baisite thinks, carry out super motor-drivenly, aircraft must satisfy so a series of conditions: (1) aircraft should have enough maneuvering abilities at pitching, driftage and three passages of lift-over, and is low to 0.1, still can keep higher driving efficiency when the angle of attack reaches 70 ° at Mach number; (2) aircraft need adopt closed loop control and advanced aerodynamic arrangement so that have fabulous low speed, big angle of attack stability; (3) aircraft should be able to change soon, acceleration and deceleration are fast, promptly has the ability of the instantaneous angular velocity that generation is very big in very short time.Obtain the advantage that changes own attitude apace by this post stall maneuver, become the design objective of opportunity of combat of new generation, and be regarded as the 4th generation fighter plane one of feature.

At present, existing a lot of country is studied super motor-driven opportunity of combat.The U.S. and German cooperation research and development a series of proof machine, be used for verifying the super maneuverability of aircraft.These aircrafts have all flown into the fault speed district, all have good manoevreability and agility.For example, the X-29A maximum angle of attack reaches 68 °, and X-31A and F-18HARV have flown to 70 ° of angles of attack, and has finished some maneuvers.New-type fighter planes such as " fitful wind " of the 4th generation opportunity of combat F-22 that comprises the U.S. that on engineering, is applied, France, Muscovite Soviet Union-35.These opportunities of combat are just paid much attention to the post stall maneuver design at the beginning of design, and have super maneuverability, and wherein typical opportunity of combat F-22 has reached the super motor-driven warfighting capabilities of the angle of attack more than 60 °.So super maneuverability also must become one of the main performance of the opportunity of combat indispensability of future generation of China.

When the fighter plane angle of attack increases and meets or exceeds stalling incidence, the body-shedding vortex that its big swept back wing leading edge forms can break, form time-dependent complex flow, like this mobile has certain randomness, in a single day fighter plane enters its angle-of-attack range, the efficient of conventional primary control surface will become very low even lose efficacy, and because speed is low, lift is little, thereby cause fighter plane to enter tailspin and crash in-flight at post stall maneuver.

At present, the common Thrust Vectoring Technology that adopts of the 4th generation fighter plane can substitute the controlsurface of former aircraft or strengthen the operating function of aircraft with the thrust component that engine thrust produces by jet pipe or tail jet deflection, the technology that fighter flight is controlled in real time, thus realize super motor-driven.Yet the vectored thrust technology can not solve the non-permanent problem in flow field of breaking and causing owing to vortexs such as aircraft wing leading edge body-shedding vortexs under the stalling incidence, thereby also can't solve horizontal stroke, the fore-and-aft stability problem that stalling incidence is got off the plane.So have important use to be worth to the design of China's the 4th generation fighter plane with the advanced Pneumatic Control Technology of driving engine vector control technology reasonable combination.

The practical Pneumatic Control Technology of modern combat aircraft can be divided into Passive Control and ACTIVE CONTROL two big classes basically, the Passive Control technology comprises that mainly installing fixing commentaries on classics of boundary 1ayer additional twists band, head fixedly edge strip, dorsal fin or abdomeinal fin, use different technology such as fighter plane nose shape, the big angle of attack is evaded and suppressed in the asymmetric whirlpool of appearance down, or reduce and eliminate the randomness of asymmetric whirlpool variation.Active control technology mainly contain at aircraft precursor head blow/air-breathing, install technology such as movable edge strip additional, can local input less energy just can obtain non local or overall mobile variation when the characteristics of ACTIVE CONTROL need to be, aircraft performance is improved significantly; Close when not required, also can not have influence on the aeroperformance of aircraft.

Summary of the invention

Stable defect of insufficient that the super maneuvering condition lower-pilot rudder face that prior art exists lost efficacy in order to overcome, the control ability under the aircraft post stall maneuver flight situation reduces and stalling incidence is got off the plane the present invention proposes a kind of aircraft host wing leading edge separation vortex control mechanism

The present invention includes separation vortex control wing rotating shaft and separation vortex control wing, and the basic configuration of this separation vortex control wing is trapezoidal.Separation vortex control wing upper surface shape with to cooperate upper surface of the airfoil shape, separation vortex control wing lower surface be angular plane surface, closely cooperate with the upper surface of the airfoil concave surface.The hypotenuse of separation vortex control wing trailing edge is a waveform, wave total number of cycles N=6～8.

When definite separation vortex control wing trailing edge waveform hypotenuse, to connect line between the separation vortex control wing trailing edge wave Origin And Destination as the x axle, the y axle is positioned at separation vortex control wing trailing edge wave starting point place, perpendicular to the x axle and point to aeroplane nose, set up plane right-angle coordinate; Each point coordinate that forms wave is determined by following equation:

$y=\frac{(N+1-m)}{N}\frac{C}{10}\sin \left(2\mathrm{\π}\frac{x-\mathrm{ml}+l}{l}\right)$

Wherein: the wave wavelength

N is an integer, and expression wave total number of cycles must satisfy following formula

$\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.2C}\&le;N\&le;\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.1C},$ Wherein:

B is that the wing exhibition is long; C is a wing root chord length; M is a variable, is determined that by the x coordinate of putting on the wave m equals to be not less than

Smallest positive integral.

The leading edge of described separation vortex control wing is positioned at 10% place of the local chord length of wing, and this separation vortex control wing is 40% of the local chord length of wing along airfoil chord to width.

At the upper surface of wing, the installation concave surface of separation vortex control wing is arranged from the leading edge of a wing 10% chord length place to the leading edge of a wing 50% place, this concave surface is the inclined-plane along chordwise direction; The degree of depth of this concave surface at leading edge of a wing starting point place is 30% of local wing thickness, and is 0 to the transition of trailing edge directional smoothing, until joining with aerofoil surface is smooth.

At the concave surface starting point place of upper surface of the airfoil, the separation vortex control wing rotating shaft is installed along wing tangential; The separation vortex control wing rotating shaft is fixed in the concave surface of upper surface of the airfoil, and has formed complete airfoil surface jointly by separation vortex control wing and aerofoil surface.

Described separation vortex control wing is fixed on rotating shaft on the wing strengthening rib rotating opening that makes progress around two ends.

At separation vortex control wing leading edge place, the axis hole that cooperates with the separation vortex control wing rotating shaft along the Zhan Xiangyou of separation vortex control wing.

The present invention is based on and separate the whirlpool in Field Flow Numerical Simulation with lee face hole and object plane undaform separation edge, average pressure drag of object (Temporal) that resulting separation whirlpool causes and instantaneous differential pressure resistance (MovingAverage), time dependent numerical result as shown in Figure 7.

Among Fig. 7 shown in a curve, when object does not have the lee face hole, its average pressure drag is fluctuation in the middle of 0.3 to 0.6, and the fluctuation range of instantaneous differential pressure resistance is 0.2.Among Fig. 7 shown in the b curve, when same object includes the lee face hole, cavity depth is 0.5 times of lee face height, and its average pressure drag fluctuates in 0.28 ± 0.5 scope, and the fluctuation range of instantaneous differential pressure resistance approximately is 0.1.Among Fig. 7 shown in the c curve, the edge, hole that separates when its body-shedding vortex of same object is trimmed to waveform, the Cycle Length of undaform is 3 times of lee face height, the average pressure drag of object fluctuates in 0.23 ± 0.1 scope, can be approximated to be constant, and the fluctuating range of instantaneous differential pressure resistance is less than 0.05.Therefore, the lee face hole can reduce average pressure drag, also can reduce the fluctuating range of instantaneous differential pressure resistance simultaneously; After the edge, hole that body-shedding vortex is separated was trimmed to waveform, the pressure drag of object can further reduce, and reach 55% with the amplitude that reduces that does not have lee face hole situation to compare average pressure drag, and the fluctuating range of instantaneous differential pressure resistance is very little.

The present invention installs separation vortex control wing on the wing aerofoil, wing plate length equates substantially with span length, the wing plate outer is trimmed to waveform, can improve under the super motor-driven angle of attack because fighter plane horizontal stroke, fore-and-aft stability problem that vortexs such as aircraft precursor body-shedding vortex, leading edge of a wing body-shedding vortex break and cause.Separation vortex control wing is installed in the wing top airfoil, with the distance of leading edge be 10% chord length.This wing plate closure when fighter plane carries out orthodox flight, top airfoil keeps original airfoil shape, and aircraft has good conventional aeroperformance.When aircraft carries out post stall maneuver flight, air-flow forms bigger separation whirlpool on tapered wing, thereby the fugitiveness and the non-stationarity of aircraft wake have been increased, control the rotating shaft that separation vortex control wing leans around the nearly leading edge of a wing and open this moment, the open angle of host wing top airfoil increases with the increase of flying angle relatively, near the top airfoil leading edge, form a local hole, the control leading edge is separated the whirlpool and is generated and a large amount of fragmentations takes place in the outer rim of this separation vortex control wing, thereby the fugitiveness and the non-stationarity that suppress the wing wake flow, make the primary control surface efficient of aircraft under the post-stall flight situation, the flicon ability, horizontal and vertical stability improves, and has realized the control of aircraft under the post-stall flight condition.

Description of drawings

Accompanying drawing 1 is the scheme drawing of aircraft layout;

Accompanying drawing 2 is scheme drawings of aircraft one side wing and separation vortex control wing;

Accompanying drawing 3 is aircraft wing A-A cutaway views, and separation vortex control wing installation site and relative size are described;

Accompanying drawing 4 is separation vortex control wing trailing edge wave coordinate scheme drawings;

Accompanying drawing 5 is separation vortex control wing when not opening, and upper surface of the airfoil separates the whirlpool scheme drawing; Wherein U ∞ represents speed of incoming flow, and α represents the angle of attack;

Accompanying drawing 6 is after separation vortex control wing is opened, and the separation vortex control wing wing is to the principle schematic of separation vortex control;

Accompanying drawing 7 is distribution schematic diagrams of instantaneous differential pressure resistance.Wherein:

1. wing 2. separation vortex control wing rotating shafts 3. separation vortex control wing 4. leading edges of a wing 5. trailing edges

The specific embodiment

Embodiment one

Present embodiment is a kind of mechanism that is used for the aircraft post-stall manipulation control, comprises wing 1, separation vortex control wing rotating shaft 2 and separation vortex control wing 3.

Wing 1 is trapezoidal, and its root string is 6 than λ; Wing setting α is 50 °, and the long b of wing exhibition is 4390mm, and root chord length C is 6282mm; Airfoil is elected NACA0006 as.At the upper surface of both sides wing 1, be symmetrically installed with separation vortex control wing 3; Present embodiment is that example is described with aircraft one side wing.

As shown in Figure 3.At the upper surface of both sides wing 1, the installation concave surface of separation vortex control wing 3 is arranged from wing 1 leading edge 10% chord length place to wing 1 leading edge 50% place, this concave surface is the inclined-plane along chordwise direction; This concave surface is 30% of a local wing thickness in the degree of depth at wing 1 leading edge starting point place, and is 0 to the transition of trailing edge directional smoothing, until joining with aerofoil surface is smooth.At the concave surface starting point place of wing 1 upper surface, separation vortex control wing rotating shaft 2 is installed along wing tangential.

As shown in Figure 2, the basic configuration of separation vortex control wing 3 is trapezoidal.The leading edge of separation vortex control wing 3 is positioned at 10% place of the local chord length of wing, and this separation vortex control wing 3 is 40% of the local chord length of wing along airfoil chord to width.The hypotenuse of separation vortex control wing 3 trailing edges is trimmed to waveform.

As shown in Figure 4, when definite separation vortex control wing 3 trailing edge hypotenuses, to connect line between the separation vortex control wing 3 trailing edge wave Origin And Destinations as the x axle, the y axle is positioned at separation vortex control wing 3 trailing edge wave starting point places, perpendicular to the x axle and point to aeroplane nose, set up plane right-angle coordinate.Each point coordinate that forms wave is determined by following equation:

$y=\frac{(N+1-m)}{N}\frac{C}{10}\sin \left(2\mathrm{\&pi;}\frac{x-\mathrm{ml}+l}{l}\right)$

Wherein: the wave wavelength

N is an integer, and expression wave total number of cycles need satisfy following formula

$\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.2C}\&le;N\&le;\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.1C},$

In the present embodiment, wave total number of cycles N=8; B is that the wing exhibition is long; C is a wing root chord length; M is a variable, is determined that by the x coordinate of putting on the wave m equals to be not less than

Smallest positive integral.

As shown in Figure 3, separation vortex control wing upper surface shape is with the upper surface shape of cooperate aerofoil profile; The separation vortex control wing lower surface is an angular plane surface, closely cooperates with the upper surface of the airfoil concave surface.After separation vortex control wing is installed in upper surface of the airfoil, formed complete airfoil surface jointly by separation vortex control wing and aerofoil surface.

As shown in Figure 3.Separation vortex control wing rotating shaft 2 is positioned at the leading edge of a wing, is fixed in the concave surface of upper surface of the airfoil, and its particular location is at the local chord length of the leading edge of a wing 10% place.The two ends of separation vortex control wing rotating shaft 2 are fixed on the strengthening rib of wing.At separation vortex control wing 3 leading edge places, the axis hole that cooperates with separation vortex control wing rotating shaft 2 along the Zhan Xiangyou of separation vortex control wing.Flight By Wire mechanism is positioned at wing, realizes opening and closed control separation vortex control wing 3 by this Flight By Wire mechanism.

When aircraft carried out orthodox flight, separation vortex control wing 3 was closed fully, as accompanying drawing 5.The upper surface of separation vortex control wing 3 and wing 1 has been formed complete smooth aerofoil surface jointly, makes aircraft have good conventional aeroperformance.When surpassing maneuvering flight, separation vortex control wing 3 is opened certain angle around separation vortex control wing rotating shaft 2 to the air speed direction, makes separation vortex control wing 3 and wing aerofoil form local hole at the top airfoil of aircraft wing 1 jointly, as shown in Figure 6.At this moment, the disengaged position of body-shedding vortex when the separation vortex control wing taper just becomes the fault speed high-angle-of-attack flight separates breaking of whirlpool with control, forms comparatively pulsation-free wake flow.Break process is carried out to the separation eddy field under the big angle of attack in the corrugated outer of wing plate, suppressing the generation in big separation whirlpool, thereby reduces the side force that can not estimate and rolling moment that fighter plane occurs under the big angle of attack, increases the stalling incidence of fighter plane.Make rudder face driving efficiency, flicon ability, the horizontal and vertical stability of aircraft under the post-stall flight situation improve.

Embodiment two

Present embodiment is a kind of mechanism that is used for the aircraft post-stall manipulation control, is used for aviette, comprises wing 1, separation vortex control wing rotating shaft 2 and separation vortex control wing 3.

Wing 1 is trapezoidal, and its root string is 5 than λ; Wing setting α is 45 °, and the long b of wing exhibition is 3000mm, and root chord length C is 3750mm; Airfoil is elected NACA0009 as.At the upper surface of both sides wing 1, be symmetrically installed with separation vortex control wing 3; Present embodiment is that example is described with aircraft one side wing.

As shown in Figure 3.At the upper surface of both sides wing 1, the installation concave surface of separation vortex control wing 3 is arranged from wing 1 leading edge 10% chord length place to wing 1 leading edge 50% place, this concave surface is the inclined-plane along chordwise direction; This concave surface is 30% of a local wing thickness in the degree of depth at wing 1 leading edge starting point place, and is 0 to the transition of trailing edge directional smoothing, until joining with aerofoil surface is smooth.At the concave surface starting point place of wing 1 upper surface, separation vortex control wing rotating shaft 2 is installed along wing tangential.

As shown in Figure 2, the basic configuration of separation vortex control wing 3 is trapezoidal.The leading edge of separation vortex control wing 3 is positioned at 10% place of the local chord length of wing, and this separation vortex control wing 3 is 40% of the local chord length of wing along airfoil chord to width.The hypotenuse of separation vortex control wing 3 trailing edges is trimmed to waveform.

As shown in Figure 4, when definite separation vortex control wing 3 trailing edge hypotenuses, to connect line between the separation vortex control wing 3 trailing edge wave Origin And Destinations as the x axle, the y axle is positioned at separation vortex control wing 3 trailing edge wave starting point places, perpendicular to the x axle and point to aeroplane nose, set up plane right-angle coordinate.Each point coordinate that forms wave is determined by following equation:

$y=\frac{(N+1-m)}{N}\frac{C}{10}\sin \left(2\mathrm{\&pi;}\frac{x-\mathrm{ml}+l}{l}\right)$

Wherein: the wave wavelength

N is an integer, and expression wave total number of cycles need satisfy following formula

$\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.2C}\&le;N\&le;\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.1C},$

In the present embodiment, wave total number of cycles N=6; B is that the wing exhibition is long; C is a wing root chord length; M is a variable, is determined that by the x coordinate of putting on the wave m equals to be not less than

Smallest positive integral.

As shown in Figure 3, separation vortex control wing upper surface shape is with the upper surface shape of cooperate aerofoil profile; The separation vortex control wing lower surface is an angular plane surface, closely cooperates with the upper surface of the airfoil concave surface.After separation vortex control wing is installed in upper surface of the airfoil, formed complete airfoil surface jointly by separation vortex control wing and aerofoil surface.

As shown in Figure 3, separation vortex control wing rotating shaft 2 is positioned at the leading edge of a wing, is fixed in the concave surface of upper surface of the airfoil, and its particular location is at the local chord length of the leading edge of a wing 10% place.The two ends of separation vortex control wing rotating shaft 2 are fixed on the strengthening rib of wing.At separation vortex control wing 3 leading edge places, the axis hole that cooperates with separation vortex control wing rotating shaft 2 along the Zhan Xiangyou of separation vortex control wing.Flight By Wire mechanism is positioned at wing, realizes opening and closed control separation vortex control wing 3 by this Flight By Wire mechanism.

When aircraft carried out orthodox flight, separation vortex control wing 3 was closed fully, as accompanying drawing 5.The upper surface of separation vortex control wing 3 and wing 1 has been formed complete smooth aerofoil surface jointly, makes aircraft have good conventional aeroperformance.When surpassing maneuvering flight, separation vortex control wing 3 is opened certain angle around separation vortex control wing rotating shaft 2 to the air speed direction, makes separation vortex control wing 3 and wing aerofoil form local hole at the top airfoil of aircraft wing 1 jointly, as shown in Figure 6.At this moment, the disengaged position of body-shedding vortex when the separation vortex control wing taper just becomes the fault speed high-angle-of-attack flight separates breaking of whirlpool with control, forms comparatively pulsation-free wake flow.Break process is carried out to the separation eddy field under the big angle of attack in the corrugated outer of wing plate, suppressing the generation in big separation whirlpool, thereby reduces the side force that can not estimate and rolling moment that fighter plane occurs under the big angle of attack, increases the stalling incidence of fighter plane.Make rudder face driving efficiency, flicon ability, the horizontal and vertical stability of aircraft under the post-stall flight situation improve.

Embodiment three

Present embodiment is a kind of mechanism that is used for the aircraft post-stall manipulation control, comprises wing 1, separation vortex control wing rotating shaft 2 and separation vortex control wing 3.

Wing 1 is trapezoidal, and its root string is 7 than λ; Wing setting α is 45 °, and the long b of wing exhibition is 4500mm, and root chord length C is 5250mm; Airfoil is elected NACA0012 as.At the upper surface of both sides wing 1, be symmetrically installed with separation vortex control wing 3; Present embodiment is that example is described with aircraft one side wing.

As shown in Figure 3, at the upper surface of both sides wing 1, the installation concave surface of separation vortex control wing 3 is arranged from wing 1 leading edge 10% chord length place to wing 1 leading edge 50% place, this concave surface is the inclined-plane along chordwise direction; This concave surface is 30% of a local wing thickness in the degree of depth at wing 1 leading edge starting point place, and is 0 to the transition of trailing edge directional smoothing, until joining with aerofoil surface is smooth.At the concave surface starting point place of wing 1 upper surface, separation vortex control wing rotating shaft 2 is installed along wing tangential.

As shown in Figure 2, the basic configuration of separation vortex control wing 3 is trapezoidal.The leading edge of separation vortex control wing 3 is positioned at 10% place of the local chord length of wing, and this separation vortex control wing 3 is 40% of the local chord length of wing along airfoil chord to width.The hypotenuse of separation vortex control wing 3 trailing edges is trimmed to waveform.

As shown in Figure 4, when definite separation vortex control wing 3 trailing edge hypotenuses, to connect line between the separation vortex control wing 3 trailing edge wave Origin And Destinations as the x axle, the y axle is positioned at separation vortex control wing 3 trailing edge wave starting point places, perpendicular to the x axle and point to aeroplane nose, set up plane right-angle coordinate.Each point coordinate that forms wave is determined by following equation:

$y=\frac{(N+1-m)}{\mathrm{<figure-callout\; id="10"\; label="N\; C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">N</figure-callout>}}\frac{C}{}\frac{}{10}\sin \left(2\mathrm{\&pi;}\frac{x-\mathrm{ml}+l}{l}\right)$

Wherein: the wave wavelength

N is an integer, and expression wave total number of cycles need satisfy following formula

$\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.2C}\&le;N\&le;\frac{\sqrt{{\mathrm{<figure-callout\; id="2"\; label="b"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">b</figure-callout>}}^2+{(\frac{C}{2}-\frac{\mathrm{<figure-callout\; id="2"\; label="C"\; filenames="HSA00000109562900021.png"\; state="\{\{state\}\}">C</figure-callout>}}{2\mathrm{\&lambda;}})}^2}}{0.1C},$

In the present embodiment, wave total number of cycles N=7; B is that the wing exhibition is long; C is a wing root chord length; M is a variable, is determined that by the x coordinate of putting on the wave m equals to be not less than

Smallest positive integral.

As shown in Figure 3, separation vortex control wing upper surface shape is with the upper surface shape of cooperate aerofoil profile; The separation vortex control wing lower surface is an angular plane surface, closely cooperates with the upper surface of the airfoil concave surface.After separation vortex control wing is installed in upper surface of the airfoil, formed complete airfoil surface jointly by separation vortex control wing and aerofoil surface.

As shown in Figure 3, separation vortex control wing rotating shaft 2 is positioned at the leading edge of a wing, is fixed in the concave surface of upper surface of the airfoil, and its particular location is at the local chord length of the leading edge of a wing 10% place.The two ends of separation vortex control wing rotating shaft 2 are fixed on the strengthening rib of wing.At separation vortex control wing 3 leading edge places, the axis hole that cooperates with separation vortex control wing rotating shaft 2 along the Zhan Xiangyou of separation vortex control wing.Flight By Wire mechanism is positioned at wing, realizes opening and closed control separation vortex control wing 3 by this Flight By Wire mechanism.

When aircraft carried out orthodox flight, separation vortex control wing 3 was closed fully, as accompanying drawing 5.The upper surface of separation vortex control wing 3 and wing 1 has been formed complete smooth aerofoil surface jointly, makes aircraft have good conventional aeroperformance.When surpassing maneuvering flight, separation vortex control wing 3 is opened certain angle around separation vortex control wing rotating shaft 2 to the air speed direction, makes separation vortex control wing 3 and wing aerofoil form local hole at the top airfoil of aircraft wing 1 jointly, as shown in Figure 6.At this moment, the disengaged position of body-shedding vortex when the separation vortex control wing taper just becomes the fault speed high-angle-of-attack flight separates breaking of whirlpool with control, forms comparatively pulsation-free wake flow.Break process is carried out to the separation eddy field under the big angle of attack in the corrugated outer of wing plate, suppressing the generation in big separation whirlpool, thereby reduces the side force that can not estimate and rolling moment that fighter plane occurs under the big angle of attack, increases the stalling incidence of fighter plane.Make rudder face driving efficiency, flicon ability, the horizontal and vertical stability of aircraft under the post-stall flight situation improve.

Claims (3)

1. an aircraft host wing leading edge separation vortex control mechanism is characterized in that the described wing plate that is used for the aircraft post-stall manipulation control comprises separation vortex control wing rotating shaft (2) and separation vortex control wing (3); And

The basic configuration of I, separation vortex control wing (3) is trapezoidal; Separation vortex control wing (3) upper surface shape with to cooperate wing (1) upper surface shape, separation vortex control wing lower surface be angular plane surface, closely cooperate with the upper surface of the airfoil concave surface; The hypotenuse of separation vortex control wing (3) trailing edge is a waveform, wave total number of cycles N=6～8; When definite separation vortex control wing (3) trailing edge waveform hypotenuse, to connect line between separation vortex control wing (3) the trailing edge wave Origin And Destination as the x axle, the y axle is positioned at separation vortex control wing (3) trailing edge wave starting point place, perpendicular to the x axle and point to aeroplane nose, set up plane right-angle coordinate; Each point coordinate that forms wave is determined by following equation:

$y=\frac{(N+1-m)}{N}\frac{C}{10}\sin \left(2\mathrm{\&pi;}\frac{x-\mathrm{ml}+l}{l}\right)$

Wherein: the wave wavelength

N is an integer, and expression wave total number of cycles need satisfy following formula

$\frac{\sqrt{b^2+{(\frac{C}{2}-\frac{C}{2\mathrm{\&lambda;}})}^2}}{0.2C}\&le;N\&le;\frac{\sqrt{b^2+{(\frac{C}{2}-\frac{C}{2\mathrm{\&lambda;}})}^2}}{0.1C},$ Wherein:

B is that the wing exhibition is long; C is a wing root chord length; M is a variable, is determined that by the x coordinate of putting on the wave m equals to be not less than

Smallest positive integral;

The leading edge of II, separation vortex control wing (3) is positioned at 10% place of the local chord length of wing, and this separation vortex control wing (3) is 40% of the local chord length of wing along airfoil chord to width;

III, at the upper surface of wing (1), to wing (1) leading edge 50% place one concave surface is arranged from wing (1) leading edge 10% chord length place, this concave surface is the inclined-plane along chordwise direction; This concave surface is 30% of a local wing thickness in the degree of depth at wing (1) leading edge starting point place, and is 0 to the transition of trailing edge directional smoothing, until joining with aerofoil surface is smooth.

2. a kind of according to claim 1 aircraft host wing leading edge separation vortex control mechanism is characterized in that, at the concave surface starting point place of wing (1) upper surface, along wing tangential separation vortex control wing rotating shaft (2) is installed; Separation vortex control wing rotating shaft (2) is fixed in the concave surface of upper surface of the airfoil, and has formed complete airfoil surface jointly by separation vortex control wing and aerofoil surface.

3. as a kind of aircraft host wing leading edge separation vortex control mechanism as described in the claim 2, it is characterized in that separation vortex control wing is fixed on rotating shaft (2) rotating opening upwards on the wing strengthening rib around two ends.

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