CN103407570A - Eddy current generating device for controlling lateral force of large-incidence-angle slender body - Google Patents
Eddy current generating device for controlling lateral force of large-incidence-angle slender body Download PDFInfo
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- CN103407570A CN103407570A CN2013102940750A CN201310294075A CN103407570A CN 103407570 A CN103407570 A CN 103407570A CN 2013102940750 A CN2013102940750 A CN 2013102940750A CN 201310294075 A CN201310294075 A CN 201310294075A CN 103407570 A CN103407570 A CN 103407570A
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Abstract
The invention relates to an eddy current generating device for controlling a lateral force of a large-incidence-angle slender body. A rotating mechanism is installed inside a conical head part of an aircraft, and the axis of the rotating mechanism is overlapped with that of the aircraft; a retraction jack is installed inside the conical head part of the aircraft and positioned between the rotating mechanism and the top end of the conical head part of the aircraft; a micro triangular wing in the retraction jack is arranged on the top end outside the conical head part of the aircraft. The micro triangular wing is driven by an electric telescoping rod to do telescoping movement along the axis direction of the aircraft. Unsteady weak turbulent flow generated by the micro triangular wing under the effect of airflow is used for performing the proportional control on the lateral force within the range of an angle of attack being 25 to 60 degree, and the control range of the lateral force coefficient is -3 to 3. The micro triangular wing can be withdrawn into the conical head part of the aircraft under the situations of zero lateral force and small lateral force, no additional resistance is produced, and the eddy current generating device is applicable to the aircraft with a taper angle being more than 30 degrees. The eddy current generating device has the characteristics of simplicity in structure, reliability in performance and the like.
Description
Technical field
The present invention relates to aircraft Aerodynamic Design field, specifically a kind of be used to controlling the vortex generating means of side force of slender body at large incidence angle.
Background technology
Modern combat aircraft and guided missile are being done the large angle of attack when motor-driven, and head there will be the unexpected and uncertain sideslipping motion of direction, and this phenomenon is called as " phantom break away (Phantom yaw) ".This phenomenon be because elongated body of revolution under large angle of attack state, even without angle of side slip, leeward district also can produce complicated many vortex systems and flow, this flow brought a size almost with the side force of lift with magnitude, thereby induce very large driftage and rolling moment; And the direction of side force is uncertain.
In the Low Angle Of Attack scope (0 °≤α<10 °), the streaming for adhering to stable flowing of model, side force is zero; In middle Low Angle Of Attack scope (10 °<α≤20 °), model flows and separates, the symmetrical whirlpool that a pair of rotation direction of rolling in the leeward district of model is opposite, and the locus of vortex does not change in time, and side force is zero; In medium angle of attack scope (20 °<α≤30 °), flow for stable asymmetric back of the body whirlpool in the leeward district of model, due to vortex intensity a little less than, so side force is less; In large angle of attack scope (30 °<α≤70 °), flow for asymmetric back of the body vortex system in the leeward district of model, the vortex intensity of Asymmetric Vortex strengthens along with the increase of the angle of attack, but due to the reducing of main whirlpool affected zone, making the variation tendency of side force in this angle of attack scope is first to increase afterwards and reduce; In very large angle of attack scope (α > 70 °), rear portion, the leeward district of model is non-permanent separated flow, on body after major part, form the non-permanent whirlpool come off, and along with the zone in the non-permanent whirlpool alternately come off of the increase of the angle of attack is increasing, side force significantly reduces, until α=90 ° form karman vortex, the side force aviation value is entirely zero.
Research is found, the faint disturbance of the asymmetric back of the body vortex pair of the large angle of attack of elongated body of revolution head is very responsive, by the body of revolution head being applied to certain faint disturbance, can realize eliminating and even control large angle of attack side force, mainly contain following several measure: 1) at the model head, the film flag is installed, under certain wind speed, flag generation autoexcitation, its wake flow is controlled body of revolution back of the body whirlpool, this measure advantage is not need to input energy, mechanism is simple, and shortcoming is that flag need to reach enough oscillation frequency, and its wake flow just can be eliminated asymmetric side force fully; 2) at the model head, lay the metal flag, flag is shimmy under the drive of motor, shimmy wake flow is controlled side force, from result by references, see that this kind mode not only can eliminate side force, and side force is linear change within the specific limits with the circumferential angle of shimmy balance, but shimmy mechanism more complicated, need the energy continued input; 3) in the perforate of model head, nose cone face single hole, porous are arranged and along several methods such as model axis perforates, by pulsation, blow/the air-breathing control that can realize side force, its shortcoming is that required the blowing of control Asymmetric Vortex/aspirated volume size is closely related with free incoming flow, be that aircraft speed increases, blow/aspirated volume also needs corresponding increase, and the added weight of equipment can be very large; 4) at the model head, movable edge strip is installed, the method had been carried out checking on aircraft, can realize the control of side force, and shortcoming is that edge strip is stressed larger, and required rotating machine power is also larger, and sleeve mechanism need to be paid larger weight cost; 5) repack the model head into movable nose cone, the deflection by nose cone or rotation to be to control, and mechanism's complexity of obvious this mode is larger to the change of model.
In the patent No., be in the innovation and creation of 200810226310.X, a kind of At High Angle of Attack Asymmetric Vortex Single hole microblowing disturbance active control method and device thereof are disclosed, proposition is arranging gas hole from 3mm position, model head tip, by changing air-blowing quantity and the circumferential angle of gas hole, can realize the control of side force, but this device needs the rotating model head, larger to model adaptation, and need source of the gas and other accessory equipment, complex structure.
In the patent No. is 201010013712.9 innovation and creation, a kind of device of eliminating flying lateral force of aircraft at high angle of attack is disclosed, this device is settled film directly over before Vehicle nose, film autoexcitation under airflow function during the large angle of attack, produce non-permanent flow-disturbing, can effectively eliminate side force, this apparatus structure is simple, but can only eliminate side force.
In the patent No. is 201110319551.0 innovation and creation, the non-permanent microvariations control setup of a kind of large angle of attack Asymmetric Vortex synthesizing jet-flow is disclosed, this device arranges the synthesizing jet-flow outlet at the model head, by changing the disturbance control frequency, can realize the control of large angle of attack Asymmetric Vortex.
In the patent No. is 201110319552.5 innovation and creation, a kind of large attack angle asymmetric vortex/lateral force closed-loop active control device is disclosed, in model head dead ahead, the microvariations sheet is set, by changing microvariations sheet frequency and the circumferential angle of disturbance balance, can realize controlling side force, in the process of controlling side force, this device needs sustained oscillation, consumes energy more.
In the patent No. is 201120218920.2 innovation and creation, a kind of asymmetrical vortex controlling device for aircraft forebody is disclosed, in both sides, the leeward district of aircraft precursor, puff port and oscillating jet device are set, by changing jet direction, can realize the control of Asymmetric Vortex, this device is more complicated and need the energy continued to input.
Summary of the invention
In order to eliminate the deficiency of the slender body aircraft such as fighter plane, tactical missile produce due to Asymmetric Vortex in large angles-of-attack situation side force, the present invention proposes a kind of be used to controlling the vortex generating means of side force of slender body at large incidence angle.
The present invention includes jack and rotating mechanism; Wherein, described rotating mechanism is arranged in the aircraft cone head part, and the dead in line of the axis of this rotating mechanism and aircraft; Jack is installed in the aircraft cone head part, and between described rotating mechanism and aircraft cone head part top; Miniature delta wing in jack is positioned at the outer top end of aircraft cone head part, and is connected by the electric expansion bar leading screw in pole and jack, under the drive of electric expansion bar along aircraft axis direction fore and aft motion.
Described jack comprises miniature delta wing, pole, strut support, electric expansion bar, electric expansion shaft bar; Electric expansion bar in jack is connected with the output shaft of described rotating machine; Miniature delta wing in jack is connected by the electric expansion bar leading screw in pole and jack.There is the via hole of pole at the strut support center, the axis of this via hole and aircraft dead in line.The electric expansion bar is arranged in the centre hole of the electric expansion shaft bar be comprised of front frame and after poppet; Axis and aircraft dead in line at described electric expansion bar.
Described miniature delta wing is the right angle trigonometry wing, is fixed in the pole top, and the length of described miniature delta wing is 0.12 times of aircraft maximum outside diameter, and sweepback angle is 45 °, without dihedral.
On the aircraft cone head part top of described installation vortex generating means, open the expansion joint of a level, this expansion joint extends along the aircraft cone head part to the fuselage direction, until the length of this expansion joint on the aircraft cone head part is slightly larger than the length on miniature delta wing base, flexible with the aircraft cone head part that meets miniature delta wing edge.
When aircraft is in Low Angle Of Attack flight, miniature delta wing closes at aircraft conehead inside, when aircraft is in large angle of attack state, chaufeur or internal processes can be controlled the elongation of electric expansion bar, miniature delta wing is pushed into to aircraft cone head part outer end, and the adjusting rotating machine, make miniature delta wing be in suitable circumferential angle.Under the effect of air-flow, miniature delta wing produces non-permanent weak flow-disturbing, realizes the control to Asymmetric Vortex.
The non-permanent weak flow-disturbing that the present invention utilizes miniature delta wing to produce under airflow function, reach the purpose of controlling the slender body side force in the At High Angle of Attack situation, according to experimental result, this device can be realized proportional control to side force in 25 °～60 ° scopes of the angle of attack, the sideway force coefficient range of control is-3～3.Miniature delta wing is recoverable in zero side force and small side force situation can not produce extra resistance to aircraft conehead inside, is applicable to the aircraft that cone angle is greater than 30 °.The characteristics such as that the present invention has is simple in structure, dependable performance.
The accompanying drawing explanation
Fig. 1 is the birds-eye view of structure of the present invention;
Fig. 2 is that miniature delta wing and expansion joint are at aircraft cone head part position view;
Fig. 3 is the left view of aircraft conehead;
Fig. 4 is the birds-eye view of miniature delta wing in aircraft conehead installation site;
Fig. 5 is the A-A view of Fig. 1;
Fig. 6 is the B-B view of Fig. 1 and ball bearing of main shaft, electric expansion bar partial enlarged drawing;
Fig. 7 is the C-C view of Fig. 1;
Fig. 8 is the D-D view of Fig. 1.
Wherein:
1. miniature delta wing 2. pole 3. strut support 4. electric expansion bar front frames
5. electric expansion bar after poppet 6. rotating machine support 7. tightening screw hole 8. rotating machines
9. stepping motor 10. electric expansion bar shell 11. aircraft cone head part 12. leading screws
13. ball bearing of main shaft 14. feed screw nut 15. expansion joints
Specific embodiments
The present embodiment is a kind of be used to controlling the vortex generating means of side force of slender body at large incidence angle, comprises jack and rotating mechanism.Wherein, rotating mechanism comprises rotating machine 8 and rotating machine support 6.Described rotating machine 8 is positioned at the aircraft cone head part, and is fixed on the aircraft fuselage frame by rotating machine support 6, the axis of this rotating mechanism and the dead in line of aircraft.Jack is arranged in the aircraft cone head part by electric expansion bar front frame 4 and electric expansion bar after poppet 5, and between described rotating mechanism and aircraft cone head part top.The output shaft of described rotating machine 8 is connected with the electric expansion bar shell 10 in jack.The adjacent end face of feed screw nut in jack and strut support to the distance on aircraft cone head part top should make described miniature delta wing can stretch out the aircraft cone head part fully, in the present embodiment, the adjacent end face of the feed screw nut in jack and strut support to the distance on aircraft cone head part top is 140mm.
Rotating machine and electric expansion bar all join with the aircraft power supply, can be controlled by flight control system by signal wire (SW).
For convenience of miniature delta wing 1, along the aircraft axis direction, carry out fore and aft motion under the drive of electric expansion bar, on described aircraft flight device cone head part top, open the expansion joint 15 of a level, this expansion joint extends along the aircraft cone head part to the fuselage direction, until the length of this expansion joint on the aircraft cone head part is slightly larger than the length on miniature delta wing base, flexible along aircraft cone head part axis to meet miniature delta wing.In the present embodiment, the long 28mm of this expansion joint.
Described jack comprises miniature delta wing 1, pole 2, strut support 3, electric expansion bar, electric expansion bar front frame 4 and electric expansion bar after poppet 5.
As shown in Figure 1, miniature delta wing 1 is positioned at the outer top end of aircraft cone head part, and miniature delta wing 1 launches and regains under the drive of electric expansion bar leading screw 12.The expansion of described miniature delta wing 1 is that this miniature delta wing extends to outside the aircraft cone head part, and the withdrawal of described miniature delta wing 1 refers to that this miniature delta wing is contracted in the aircraft cone head part fully.
The right-angled triangle that described miniature delta wing 1 is made for aluminum alloy, be fixed in pole 2 tops.The length of described miniature delta wing is 0.12 times of aircraft maximum outside diameter, and the thickness of miniature delta wing must meet the rigidity needs of this miniature delta wing, and in the present embodiment, the length of described delta wing is 24mm, and thickness is 0.5mm.The sweepback angle of described delta wing is 45 °, without dihedral.
As shown in Figures 2 and 3, for convenience of miniature delta wing 1, along the aircraft axis direction, carry out fore and aft motion under the drive of electric expansion bar, on described aircraft flight device cone head part top, open the expansion joint 15 of a level, this expansion joint extends along the aircraft cone head part to the fuselage direction, until the length of this expansion joint on the aircraft cone head part is slightly larger than the length on miniature delta wing base, flexible along aircraft cone head part axis to meet miniature delta wing.In the present embodiment, the long 28mm of this expansion joint.
As shown in Figure 5, strut support 3 is " X " type support, " X " type support and the riveted joint of aircraft cone head part fuselage frame, there is through hole at strut support 3 centers, through-bore axis and aircraft dead in line, and the through hole internal diameter is identical with the diameter of pole 2, pole 2 can be slided along through hole, strut support 3 should guarantee that apart from the distance in Vehicle nose the place ahead miniature delta wing 1 can folding and unfolding, and in the present embodiment, strut support 3 is 64mm apart from the distance in Vehicle nose the place ahead.
As shown in Figure 6 and Figure 7, the electric expansion shaft bar is comprised of front frame 4 and after poppet 5.Described front frame 4 and after poppet 5 are " X " type support, are positioned at the aircraft cone head part, and rivet with aircraft cone head part fuselage frame.At the center of described front frame 4 and after poppet 5, a through hole is arranged respectively, be used to the electric expansion bar is installed; Through-bore axis and aircraft dead in line.Between through-hole wall and electric expansion bar shell 10, ball bearing of main shaft 13 is installed, under the drive of rotating machine 8, the electric expansion bar can rotate.
As shown in Figure 8, rotating machine support 6 is " X " type support, with the aircraft fuselage frame, rivet, described rotating machine carriage center has a through hole, through-bore axis and aircraft dead in line, and through-hole diameter is identical with rotating machine 8 diameters, on support, along the aircraft axis, two screw holes 7 are arranged, screw hole 7 diameter 5mm, by tightening screw, be connected rotating machine 8 and support 6.
Claims (7)
1. one kind be used to controlling the vortex generating means of side force of slender body at large incidence angle, it is characterized in that, comprises jack and rotating mechanism; Wherein, described rotating mechanism is arranged in the aircraft cone head part, and the dead in line of the axis of this rotating mechanism and aircraft; Jack is installed in the aircraft cone head part, and between described rotating mechanism and aircraft cone head part top; Miniature delta wing in jack is positioned at the outer top end of aircraft cone head part, and is connected by the electric expansion bar leading screw in pole and jack, under the drive of electric expansion bar along aircraft axis direction fore and aft motion.
2. a kind ofly as claimed in claim 1 be used to controlling the vortex generating means of side force of slender body at large incidence angle, it is characterized in that, described jack comprises miniature delta wing, pole, strut support, electric expansion bar, electric expansion shaft bar; Electric expansion bar in jack is connected with the output shaft of described rotating machine.
3. a kind of be used to controlling the vortex generating means of side force of slender body at large incidence angle as claimed in claim 1, it is characterized in that, described miniature delta wing is the right angle trigonometry wing, be fixed in the pole top, the length of described miniature delta wing is 0.12 times of aircraft maximum outside diameter, sweepback angle is 45 °, without dihedral; Miniature delta wing is connected with electric expansion bar leading screw by pole.
4. a kind ofly as claimed in claim 1 be used to controlling the vortex generating means of side force of slender body at large incidence angle, it is characterized in that, there is the via hole of pole at the strut support center, the axis of this via hole and aircraft dead in line.
5. a kind ofly as claimed in claim 1 be used to controlling the vortex generating means of side force of slender body at large incidence angle, it is characterized in that, the electric expansion bar is arranged in the centre hole of the electric expansion shaft bar be comprised of front frame and after poppet; Axis and aircraft dead in line at described electric expansion bar.
6. a kind ofly as claimed in claim 1 be used to controlling the vortex generating means of side force of slender body at large incidence angle, it is characterized in that the axis in rotating machine carriage center hole and aircraft dead in line.
7. a kind of be used to controlling the vortex generating means of side force of slender body at large incidence angle as claimed in claim 1, it is characterized in that, on the aircraft cone head part top of described installation vortex generating means, open the expansion joint of a level, this expansion joint extends along the aircraft cone head part to the fuselage direction, until the length of this expansion joint on the aircraft cone head part is slightly larger than the length on miniature delta wing base, flexible with the aircraft cone head part that meets miniature delta wing edge.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN2013102940750A CN103407570A (en) | 2013-07-12 | 2013-07-12 | Eddy current generating device for controlling lateral force of large-incidence-angle slender body |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN2013102940750A CN103407570A (en) | 2013-07-12 | 2013-07-12 | Eddy current generating device for controlling lateral force of large-incidence-angle slender body |
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| CN103407570A true CN103407570A (en) | 2013-11-27 |
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| CN2013102940750A Pending CN103407570A (en) | 2013-07-12 | 2013-07-12 | Eddy current generating device for controlling lateral force of large-incidence-angle slender body |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111846192A (en) * | 2020-06-04 | 2020-10-30 | 中国人民解放军国防科技大学 | Flight verification simulation cabin section for online identification of aircraft parameters |
| CN113525669A (en) * | 2021-05-29 | 2021-10-22 | 北京航空航天大学宁波创新研究院 | Large-attack-angle lateral force control method based on combined disturbance |
| WO2021212707A1 (en) * | 2020-04-20 | 2021-10-28 | 中国民用航空飞行学院 | Wing vortex-breaking structure, wing, and aircraft |
| CN114671005A (en) * | 2022-04-28 | 2022-06-28 | 威海光晟航天航空科技有限公司 | Aircraft nose cone structure and preparation method thereof |
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| US5050819A (en) * | 1990-08-10 | 1991-09-24 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Rotatable non-circular forebody flow controller |
| US5083724A (en) * | 1986-02-27 | 1992-01-28 | Messerschmitt-Bolkow-Blohm Gmbh | Device for controlling aerodynamic bodies |
| US5201829A (en) * | 1991-12-19 | 1993-04-13 | General Dynamics Corporation | Flight control device to provide directional control |
| CN101767648A (en) * | 2010-01-14 | 2010-07-07 | 西北工业大学 | Device for eliminating side force of slender body at large incidence angle |
| CN102390525A (en) * | 2011-10-20 | 2012-03-28 | 南京航空航天大学 | Large attack angle asymmetric vortex/lateral force closed-loop active control device |
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2013
- 2013-07-12 CN CN2013102940750A patent/CN103407570A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4015800A (en) * | 1975-04-22 | 1977-04-05 | Bihrle Jr William | Aerodynamic spin control device for aircraft |
| US5083724A (en) * | 1986-02-27 | 1992-01-28 | Messerschmitt-Bolkow-Blohm Gmbh | Device for controlling aerodynamic bodies |
| US5050819A (en) * | 1990-08-10 | 1991-09-24 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Rotatable non-circular forebody flow controller |
| US5201829A (en) * | 1991-12-19 | 1993-04-13 | General Dynamics Corporation | Flight control device to provide directional control |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2021212707A1 (en) * | 2020-04-20 | 2021-10-28 | 中国民用航空飞行学院 | Wing vortex-breaking structure, wing, and aircraft |
| CN111846192A (en) * | 2020-06-04 | 2020-10-30 | 中国人民解放军国防科技大学 | Flight verification simulation cabin section for online identification of aircraft parameters |
| CN113525669A (en) * | 2021-05-29 | 2021-10-22 | 北京航空航天大学宁波创新研究院 | Large-attack-angle lateral force control method based on combined disturbance |
| CN114671005A (en) * | 2022-04-28 | 2022-06-28 | 威海光晟航天航空科技有限公司 | Aircraft nose cone structure and preparation method thereof |
| CN114671005B (en) * | 2022-04-28 | 2024-04-12 | 威海光晟航天航空科技有限公司 | Aircraft nose cone structure and preparation method thereof |
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Application publication date: 20131127 |