CN112923805A - Pneumatic layout of small high-mobility missile - Google Patents
Pneumatic layout of small high-mobility missile Download PDFInfo
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- CN112923805A CN112923805A CN202110075751.XA CN202110075751A CN112923805A CN 112923805 A CN112923805 A CN 112923805A CN 202110075751 A CN202110075751 A CN 202110075751A CN 112923805 A CN112923805 A CN 112923805A
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- 241000272517 Anseriformes Species 0.000 claims abstract description 31
- 230000002093 peripheral effect Effects 0.000 claims abstract description 10
- 241000272525 Anas platyrhynchos Species 0.000 claims description 36
- 238000000034 method Methods 0.000 claims description 5
- 230000004323 axial length Effects 0.000 claims description 4
- 238000012545 processing Methods 0.000 claims description 2
- 238000013461 design Methods 0.000 abstract description 10
- 238000005096 rolling process Methods 0.000 abstract description 7
- 230000000694 effects Effects 0.000 abstract description 4
- 230000008859 change Effects 0.000 description 7
- 230000007123 defense Effects 0.000 description 2
- 230000004075 alteration Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 230000018109 developmental process Effects 0.000 description 1
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- 238000005516 engineering process Methods 0.000 description 1
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
- F42B10/60—Steering arrangements
- F42B10/62—Steering by movement of flight surfaces
- F42B10/64—Steering by movement of flight surfaces of fins
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
- F42B15/01—Arrangements thereon for guidance or control
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- Engineering & Computer Science (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Aviation & Aerospace Engineering (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
Abstract
本发明一种小型高机动导弹气动布局，属于导弹设计领域；包括导弹头部、弹身、一级鸭翼和尾翼，所述导弹头部为旋成体外形，同轴安装于导弹的前端；所述弹身为圆柱体结构；四个一级鸭翼沿周向均布于弹身外周面靠近头部的位置，并呈“×”字型；四个尾翼沿周向均布于弹身尾端的外周面，并呈“×”字型；还包括两个二级鸭翼，两个二级鸭翼以弹身的中心轴对称、设置于弹身外周面靠近一级鸭翼的位置，并分别位于弹身的正上方和正下方成“Ⅰ”字型。本发明增加二级鸭翼的控制效率更高，滚转控制效果好。二级鸭翼产生的滚转控制力矩为近似线性变化的，控制系统设计简单。
The invention discloses an aerodynamic layout of a small high-mobility missile, belonging to the field of missile design. It includes a missile head, a projectile body, a first-level canard and a tail. The bullet body is a cylindrical structure; four first-level canard wings are evenly distributed on the outer peripheral surface of the bullet body near the head in the circumferential direction, and are in the shape of "X"; four tail wings are uniformly distributed on the outer peripheral surface of the rear end of the bullet body It is in the shape of "X"; it also includes two secondary canards, which are symmetrical with the central axis of the projectile body and are arranged on the outer peripheral surface of the projectile body close to the first-level canards, and are respectively located on the projectile body. The top and bottom of it are in the shape of "I". The invention increases the control efficiency of the secondary canard wings, and has a good rolling control effect. The roll control torque generated by the secondary canard changes approximately linearly, and the control system is simple in design.
Description
Technical Field
The invention belongs to the field of missile design, and particularly relates to a pneumatic layout of a small high-mobility missile.
Background
Along with the development of national defense technology, the miniaturization of missiles is more and more emphasized, and various countries successively begin to research and develop microminiature guided munitions suitable for individual combat and the like. Meanwhile, the air defense requirements of targets such as unmanned planes, small airplanes and shells are met, and high requirements are provided for low-cost miniaturization and maneuvering performance of the missiles.
The maneuvering of the missile is mainly realized by changing the magnitude and direction of the force and moment acting on the missile. One way to achieve this is to implement direct force control by a vector engine, but this method has the disadvantages of difficult force control, large system size, and inability to operate for long periods of time. Another method is to vary the aerodynamic forces and moments acting on the missile by manipulating the deflection of the aerodynamic control surfaces. The required high maneuvering characteristics can be realized by reasonably designing the aerodynamic control surface of the missile, namely the aerodynamic layout of the missile.
Disclosure of Invention
The technical problem to be solved is as follows:
in order to avoid the defects of the prior art, the invention provides a pneumatic layout of a small high-mobility missile, which reduces the induced rolling influence of a duck-type missile by forming two-stage duck wings distributed in an I shape on the outer peripheral surface of the missile, solves the coupling problem of a rolling control channel and a pitching and yawing control channel on the basis, and reduces the design difficulty of a missile system.
The technical scheme of the invention is as follows: a small-sized high-mobility missile aerodynamic configuration comprises a missile head, a missile body, a first-level duck wing and a tail wing, wherein the missile head is in a shape of a rotating body and is coaxially arranged at the front end of a missile; the bullet body is of a cylindrical structure; the four first-stage duck wings are uniformly distributed on the outer peripheral surface of the missile body close to the head along the circumferential direction and are in an X shape; the four tail wings are uniformly distributed on the peripheral surface of the tail end of the bullet body along the circumferential direction and are shaped like an X; the method is characterized in that: still include two second grade duck wings, two second grade duck wings are with the central axis symmetry of bullet body, set up in the position that the bullet body outer peripheral face is close to the one-level duck wing to be located directly over the bullet body respectively and become "I" style of calligraphy under with.
The further technical scheme of the invention is as follows: the first-stage duck wing and the second-stage duck wing are both in trapezoidal or triangular flat wing structures, and chamfering processing is carried out on the outer edges of the duck wings.
The further technical scheme of the invention is as follows: the distance between the second-stage duck wing and the first-stage duck wing is 1-2 times of the span length of the first-stage duck wing.
The further technical scheme of the invention is as follows: the diameter of the small high maneuvering missile is 40 mm; the head of the missile is of a rotating body structure, a bus is a parabola, the axial length of the head is 86mm, and the maximum diameter of the head is 40 mm; the bullet body is cylindrical, the axial length is 578mm, and the diameter of the bullet body is 40 mm;
the root chord length of the first-stage duck wing is 40mm, the tip chord length is 20mm, the half-spread length is 30mm, the thickness of the wing surface is 2mm, and the distance from the hinge position to the top point of the bullet-guiding head is 196 mm; the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shafts of the steering engines, and the four first-stage duck wing surfaces are respectively connected with four independent output shafts of the steering engines; the hinge is the steering engine output shaft;
the root chord length of the second-stage duck wing is 20mm, the tip chord length is 10mm, the half span length is 20mm, the wing surface thickness is 2mm, and the distance between the hinge and the head is 246 mm; the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shaft of the steering engine, and 2 wing surfaces are respectively connected with 2 independent steering engine output shafts; the hinge is used for controlling the output shaft of the steering engine of the second-level duck wing.
The tail wing is 70mm in root chord length, 40mm in tip chord length, 40mm in half span length and 2mm in wing surface thickness, the tail wing rear edge is flush with the bottom of the missile, and the four tail wings are fixedly connected with the surface of the missile body.
The further technical scheme of the invention is as follows: the missile head is in a spherical structure or a parabolic structure.
The further technical scheme of the invention is as follows: the tail wing is of a trapezoidal flat wing structure, and the edge of the tail wing is subjected to chamfering treatment; the four tail wings are the same as the four first-level canard wings in layout and are opposite in position.
Advantageous effects
The invention has the beneficial effects that:
1. the control efficiency of the second-level canard is higher, and the rolling control effect is good.
Under the same flight condition and rudder deflection angle, the control surface area with the same size and the roll control torque generated by the two-stage canard wing configuration are 1.98 times of the roll torque generated by the conventional canard wing layout missile. Meanwhile, in practical application, the requirement on missile roll control torque is low, and the roll control surface with small size can be independently designed by the layout scheme of the second-stage canard wing, so that the same or even better roll control effect is achieved. For example, when the second-stage canard wing for roll control is designed independently in the design, the wing area of the canard wing is only one third of that of the original single-stage canard wing missile, and meanwhile, the roll control of the missile is carried out independently by the second-stage canard wing without participation of the first-stage canard wing, so that the design of a control system is simpler and more efficient, and the precision of the control system is improved.
2. The rolling control torque generated by the second-level canard is approximately linearly changed, and the control system is simple in design.
As shown in fig. 3, under the same sliding angle, the roll control moment coefficient generated by the second-level duck rudder is relatively stable with respect to the change of the attack angle. Except the curve of the state of the side sliding angle of 20 degrees, other broken lines are basically changed linearly. Therefore, the roll control moment generated by the second-level canard in the longitudinal symmetric plane of the missile basically has a linear rule to the change of the attack angle.
3. The two-stage canard design enables the rolling channel to be decoupled from the other two channels, and the design of a control system is simplified.
As shown in fig. 4 and 5, the variation of the missile pitch moment coefficient with the sideslip angle and the attack angle is respectively in the non-rudder deflection state and the two-stage canard differential 10 ° state. As shown in fig. 6 and 7, the change of the yaw moment coefficient is associated with each change. The comparison of the two conditions shows that the differential rudder deflection of the duck rudder increases the pitch moment coefficient and the yaw moment coefficient, but does not change the basic variation trend. The roll control channel and other control channels can be decoupled, and the method has important significance for pneumatic parameter calculation of pitching and yawing channels, control design and the like.
Drawings
FIG. 1 is a schematic view of a canard missile layout.
FIG. 2 is a schematic view of an "x" font layout.
FIG. 3 is a graph showing the variation of the roll control moment coefficient generated by the two-stage canard with respect to the angle of attack for different sideslip angles.
FIG. 4 shows the change rule of the pitching moment coefficient along with the attack angle when the second-stage canard wing has no rudder deflection under different sideslip angles.
FIG. 5 shows the variation of the pitching moment coefficient with the angle of attack for a two-stage canard differential motion of 10 degrees under different sideslip angles.
FIG. 6 shows the change rule of the yaw moment coefficient along with the angle of attack when the second-stage canard wing has no rudder deflection under different sideslip angles.
FIG. 7 shows the variation law of the yawing moment coefficient with the angle of attack when the two-stage canard differential motion is 10 degrees under different sideslip angles.
FIG. 8 is a schematic diagram of the aerodynamic layout of the small high maneuvering missile and the relative positions of the parts.
Description of reference numerals: the missile comprises a missile head 1, a missile body 2, a first-stage duck wing 3, a second-stage duck wing 4 and a tail wing 5.
Detailed Description
The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", and the like, indicate orientations and positional relationships based on those shown in the drawings, and are used only for convenience of description and simplicity of description, and do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be considered as limiting the present invention.
As shown in FIG. 8, the pneumatic layout of a small high maneuvering missile with the diameter of 40mm of the invention is implemented as follows:
the missile head 1 is in a rotating body structure, and a generatrix is a parabola. The length of the head is 86mm, and the maximum diameter of the head is 40 mm. The missile head 1 is connected with the front end of the missile body 2.
The bullet body 2 is cylindrical, and length is 578mm, and the bullet body diameter is 40 mm.
The first-stage duck wing 3 is a flat wing, 4 wing surfaces are arranged in an X shape, the root chord length is 40mm, the tip chord length is 20mm, the half-spread length is 30mm, the wing surfaces are 2mm thick, and the hinge position (a steering engine output shaft) is 196mm according to the head. And the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shaft of the steering engine. And 4 wing surfaces are respectively connected with 4 independent steering engine output shafts.
The second-stage canard wing 4 is a flat wing, 2 wing surfaces are in I-shaped layout, and the second-stage canard wing is located in a longitudinal symmetrical plane of the missile. The root chord length is 20mm, the tip chord length is 10mm, the half span length is 20mm, the airfoil thickness is 2mm, and the hinge (steering engine output shaft) is 246mm according to the head position. And the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shaft of the steering engine. And 2 wing surfaces are respectively connected with 2 independent steering engine output shafts.
The tail wing 5 is a flat plate wing, 4 wing surfaces are arranged in an X shape, the root chord length is 70mm, the tip chord length is 40mm, the half span length is 40mm, the wing surface thickness is 2mm, and the tail edge of the tail wing is flush with the bottom of the missile. 4 empennages are fixedly connected with the surface of the bullet body.
When the missile flies, 4 first-level canard wings execute the control of pitching and yawing channels, and 2 second-level canard wings execute the control of rolling channels. And the control system respectively adjusts and controls the deflection angle of each control surface according to the flying requirement of the missile.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made in the above embodiments by those of ordinary skill in the art without departing from the principle and spirit of the present invention.
Claims (6)
1. A small-sized high-mobility missile aerodynamic configuration comprises a missile head, a missile body, a first-level duck wing and a tail wing, wherein the missile head is in a shape of a rotating body and is coaxially arranged at the front end of a missile; the bullet body is of a cylindrical structure; the four first-stage duck wings are uniformly distributed on the outer peripheral surface of the missile body close to the head along the circumferential direction and are in an X shape; the four tail wings are uniformly distributed on the peripheral surface of the tail end of the bullet body along the circumferential direction and are shaped like an X; the method is characterized in that: still include two second grade duck wings, two second grade duck wings are with the central axis symmetry of bullet body, set up in the position that the bullet body outer peripheral face is close to the one-level duck wing to be located directly over the bullet body respectively and become "I" style of calligraphy under with.
2. The small high maneuver missile aerodynamic configuration of claim 1, wherein: the first-stage duck wing and the second-stage duck wing are both in trapezoidal or triangular flat wing structures, and chamfering processing is carried out on the outer edges of the duck wings.
3. The small high maneuver missile aerodynamic configuration of claim 1, wherein: the distance between the second-stage duck wing and the first-stage duck wing is 1-2 times of the span length of the first-stage duck wing.
4. The small high maneuver missile aerodynamic configuration of claim 1, wherein: the diameter of the small high maneuvering missile is 40 mm; the head of the missile is of a rotating body structure, a bus is a parabola, the axial length of the head is 86mm, and the maximum diameter of the head is 40 mm; the bullet body is cylindrical, the axial length is 578mm, and the diameter of the bullet body is 40 mm;
the root chord length of the first-stage duck wing is 40mm, the tip chord length is 20mm, the half-spread length is 30mm, the thickness of the wing surface is 2mm, and the distance from the hinge position to the top point of the bullet-guiding head is 196 mm; the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shafts of the steering engines, and the four first-stage duck wing surfaces are respectively connected with four independent output shafts of the steering engines; the hinge is the steering engine output shaft;
the root chord length of the second-stage duck wing is 20mm, the tip chord length is 10mm, the half span length is 20mm, the wing surface thickness is 2mm, and the distance between the hinge and the head is 246 mm; the symmetrical center of the side surface of the wing root is provided with a connecting hole position for connecting the output shaft of the steering engine, and 2 wing surfaces are respectively connected with 2 independent steering engine output shafts; the hinge is used for controlling the output shaft of the steering engine of the second-level duck wing.
The tail wing is 70mm in root chord length, 40mm in tip chord length, 40mm in half span length and 2mm in wing surface thickness, the tail wing rear edge is flush with the bottom of the missile, and the four tail wings are fixedly connected with the surface of the missile body.
5. The small high maneuver missile aerodynamic configuration of claim 1, wherein: the missile head is in a spherical structure or a parabolic structure.
6. The small high maneuver missile aerodynamic configuration of claim 1, wherein: the tail wing is of a trapezoidal flat wing structure, and the edge of the tail wing is subjected to chamfering treatment; the four tail wings are the same as the four first-level canard wings in layout and are opposite in position.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110075751.XA CN112923805A (en) | 2021-01-20 | 2021-01-20 | Pneumatic layout of small high-mobility missile |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202110075751.XA CN112923805A (en) | 2021-01-20 | 2021-01-20 | Pneumatic layout of small high-mobility missile |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114234734A (en) * | 2021-12-24 | 2022-03-25 | 中国工程物理研究院总体工程研究所 | Pneumatic layout of microminiature missile |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CA1341555C (en) * | 1983-09-19 | 2007-10-09 | Albert Dunn | Steering of missiles |
| US20080061188A1 (en) * | 2005-09-09 | 2008-03-13 | General Dynamics Ordnance And Tactical Systems, Inc. | Projectile trajectory control system |
| US20110297783A1 (en) * | 2009-10-26 | 2011-12-08 | Marcelo Edgardo Martinez | Rolling projectile with extending and retracting canards |
| US8552349B1 (en) * | 2010-12-22 | 2013-10-08 | Interstate Electronics Corporation | Projectile guidance kit |
| KR101421127B1 (en) * | 2014-01-15 | 2014-07-22 | 국방과학연구소 | Course correction fuze |
| CN106871742A (en) * | 2017-01-05 | 2017-06-20 | 北京航天自动控制研究所 | A kind of control system being arranged on body |
| CN106886625A (en) * | 2017-01-05 | 2017-06-23 | 北京航天自动控制研究所 | A kind of Design of Aerodynamic Configuration method of the dual-spin stabilization bullet based on fixed-wing duck rudder |
| CN110307759A (en) * | 2019-06-24 | 2019-10-08 | 中国航天空气动力技术研究院 | It is a kind of to be quickly laid out from overturning guided missile |
-
2021
- 2021-01-20 CN CN202110075751.XA patent/CN112923805A/en active Pending
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CA1341555C (en) * | 1983-09-19 | 2007-10-09 | Albert Dunn | Steering of missiles |
| US20080061188A1 (en) * | 2005-09-09 | 2008-03-13 | General Dynamics Ordnance And Tactical Systems, Inc. | Projectile trajectory control system |
| US20110297783A1 (en) * | 2009-10-26 | 2011-12-08 | Marcelo Edgardo Martinez | Rolling projectile with extending and retracting canards |
| US8552349B1 (en) * | 2010-12-22 | 2013-10-08 | Interstate Electronics Corporation | Projectile guidance kit |
| KR101421127B1 (en) * | 2014-01-15 | 2014-07-22 | 국방과학연구소 | Course correction fuze |
| CN106871742A (en) * | 2017-01-05 | 2017-06-20 | 北京航天自动控制研究所 | A kind of control system being arranged on body |
| CN106886625A (en) * | 2017-01-05 | 2017-06-23 | 北京航天自动控制研究所 | A kind of Design of Aerodynamic Configuration method of the dual-spin stabilization bullet based on fixed-wing duck rudder |
| CN110307759A (en) * | 2019-06-24 | 2019-10-08 | 中国航天空气动力技术研究院 | It is a kind of to be quickly laid out from overturning guided missile |
Non-Patent Citations (1)
| Title |
|---|
| 吴涛: "更锋利的格斗短刃 世界第四代近距空空导弹评析", 现代兵器, no. 11, pages 38 - 42 * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114234734A (en) * | 2021-12-24 | 2022-03-25 | 中国工程物理研究院总体工程研究所 | Pneumatic layout of microminiature missile |
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Application publication date: 20210608 |