CN111290427A - High-overload-resistant aircraft lateral deviation correction system - Google Patents

High-overload-resistant aircraft lateral deviation correction system Download PDF

Info

Publication number
CN111290427A
CN111290427A CN201811533712.4A CN201811533712A CN111290427A CN 111290427 A CN111290427 A CN 111290427A CN 201811533712 A CN201811533712 A CN 201811533712A CN 111290427 A CN111290427 A CN 111290427A
Authority
CN
China
Prior art keywords
aircraft
overload
module
navigation
antenna
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
CN201811533712.4A
Other languages
Chinese (zh)
Other versions
CN111290427B (en
Inventor
王伟
师兴伟
林德福
宁波
王江
裴培
林时尧
王雨辰
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Military Representative Office Of Pla In 844 Factory
Beijing Institute of Technology BIT
Original Assignee
Military Representative Office Of Pla In 844 Factory
Beijing Institute of Technology BIT
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Military Representative Office Of Pla In 844 Factory, Beijing Institute of Technology BIT filed Critical Military Representative Office Of Pla In 844 Factory
Publication of CN111290427A publication Critical patent/CN111290427A/en
Application granted granted Critical
Publication of CN111290427B publication Critical patent/CN111290427B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
    • G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
    • G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
    • G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B19/00—Programme-control systems
    • G05B19/02—Programme-control systems electric
    • G05B19/04—Programme control other than numerical control, i.e. in sequence controllers or logic controllers
    • G05B19/042—Programme control other than numerical control, i.e. in sequence controllers or logic controllers using digital processors
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B2219/00—Program-control systems
    • G05B2219/20—Pc systems
    • G05B2219/25—Pc structure of the system
    • G05B2219/25252—Microprocessor

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • General Physics & Mathematics (AREA)
  • Remote Sensing (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Health & Medical Sciences (AREA)
  • Artificial Intelligence (AREA)
  • Computer Vision & Pattern Recognition (AREA)
  • Evolutionary Computation (AREA)
  • Medical Informatics (AREA)
  • Software Systems (AREA)
  • Traffic Control Systems (AREA)
  • Position Fixing By Use Of Radio Waves (AREA)

Abstract

The invention discloses an anti-high-overload aircraft sideslip correction system, which comprises a microprocessor module, a navigation module and a navigation ratio output module, wherein the microprocessor module is used for calculating the required sideslip for the aircraft sideslip correction, the navigation module is used for providing the position and speed information of an aircraft for the microprocessor module, the navigation ratio output module is used for providing a real-time variable navigation ratio for the microprocessor module, the navigation module comprises an anti-high overload antenna and can provide stable guidance information under the condition of high overload, the navigation ratio output module is used for obtaining the real-time variable navigation ratio according to the information such as the total range, the real-time sideslip distance and the like when the aircraft starts to control, so that the guidance performance is improved, a target is ensured to enter the field range of a guidance head when the aircraft enters the final guidance, in addition, because the navigation ratio is continuously changed in a small amplitude, the large-amplitude vibration of the flight track can not be caused, the stable flight process is ensured, and the final hit precision is high.

Description

High-overload-resistant aircraft lateral deviation correction system
Technical Field
The invention relates to the field of guidance control of high overload aircrafts, in particular to a high overload resistant aircraft sideslip correction system.
Background
For a remote guidance aircraft, in order to improve the range of the aircraft, various measures are mostly adopted in the climbing section of a flight trajectory to enable the climbing height of the aircraft to be higher, such as rocket range extension, bottom row technology or high-power gunpowder, and the like, but the measures usually prolong the flight time of the climbing section of the aircraft, so that the starting and controlling time of the aircraft is generally set to be 50s after launching. The long flight time before starting control makes the aircraft unable to control the aircraft to fly to the target along the expected trajectory in the time, and the influence of the lateral wind, the magnus force generated by self rotation and the interference of the launching end often forces the aircraft to have a large lateral deviation distance during starting control, while even the general lateral guidance method and system can control the aircraft to fly to the target, when the aircraft enters the last guidance section, the general lateral guidance method and system often have difficulty in controlling the aircraft to make the target enter the field of view of the guidance head, and the evaluation standard of entering the field of view is as follows: and when the distance is 3km from the target, the lateral deviation is less than 600 m.
In addition, for a high-overload remote guidance aircraft, the difference from a common aircraft is large, under the action of high overload, the transmission instruction lag among a navigation system, the guidance system and a control system is more serious due to the design scheme of a space directional gyroscope, a platform laser guide head, a pneumatic steering engine and the like in the traditional guidance control system, and the information obtained by the navigation system is not accurate enough, can not be continuous and has large fluctuation, so that serious consequences are caused.
If the aircraft cannot enable the target to enter the field of view of the guide head when entering the final guide section, the aircraft cannot capture the target in the final guide section, and the target is probably missed finally; in the guidance control process of the aircraft, if guidance laws with large differences are adopted for different stages, the flight trajectory of the aircraft is inevitably vibrated greatly, and the stability of the aircraft is reduced;
for the above reasons, the present inventors have conducted extensive studies on existing high overload resistant aircraft control systems in order to devise a new high overload resistant aircraft yaw correction system that solves the above problems.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention carries out intensive research and designs an anti-high-overload aircraft sideslip correction system, which comprises a microprocessor module, a navigation module and a navigation ratio output module, wherein the microprocessor module is used for calculating the required sideslip overload required by the aircraft sideslip correction, the navigation module is used for providing the position and speed information of the aircraft for the microprocessor module, the navigation ratio output module is used for providing the real-time variable navigation ratio for the microprocessor module, the navigation module comprises an anti-high overload antenna and can provide stable guidance information under the condition of high overload, the navigation ratio output module brings the total range and the real-time sideslip distance of the aircraft during starting and controlling and the length of a connecting line between the aircraft and a target point projected on a transmitting point and a connecting line between the aircraft and the target point into consideration of a guidance algorithm to obtain the scientific and reasonable real-time variable navigation ratio, therefore, the guidance performance of the device is improved, the target is ensured to enter the field range of the seeker when the device enters the final guidance, in addition, the navigation ratio is continuously changed in a small amplitude, the large amplitude vibration of the flight track is not caused, the stable flight process is ensured, the final hit precision is high, and the invention is completed.
In particular, the invention aims to provide an aircraft sideslip correction system resistant to high overloads, comprising:
the microprocessor module 1 is used for calculating the required sideslip overload required by the aircraft sideslip correction;
and the navigation module 2 is used for acquiring the position and speed information of the aircraft in real time.
In the microprocessor module 1, the required sideslip overload is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the angular rate of the line of sight of the missile in the sideslip direction;
preferably, the lateral bias is obtained in real time by the following formula (one):
Figure BDA0001906347530000031
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
Figure BDA0001906347530000032
representing the angular rate of the aircraft's yaw direction line of sight.
The navigation module 2 comprises a high overload resistant antenna 21, an anti-interference sub-module 22 and a satellite guidance sub-module 23;
the high overload resistant antenna 21 is in the shape of a sheet, for receiving satellite signals at high overload,
the anti-interference submodule 22 is connected to the high overload resistant antenna 21, and is configured to perform filtering processing on the satellite signal,
the satellite guidance sub-module 23 receives the filtered satellite signals and calculates the position and speed information of the aircraft in real time according to the signals.
Wherein the high overload resistant antenna 21 is arranged on the outer wall of the aircraft,
preferably, an inwards concave accommodating groove 4 is formed in the outer wall of the aircraft, the high overload resistant antenna 21 is installed in the accommodating groove 4, and a protective baffle 41 is arranged outside the high overload resistant antenna 21.
The anti-high overload antenna 21 is provided with a plurality of pieces which are uniformly distributed around the aircraft, and preferably, the anti-high overload antenna 21 is provided with 4 pieces.
Wherein, the system is also provided with a navigation ratio output module 3 for calculating the navigation ratio;
the navigation ratio output module 3 outputs the lateral deviation distance z of the aircraft according to the control startingmSelects the corresponding navigation ratio N and transmits the navigation ratio N to the microprocessor module 1 in real time.
Wherein the offset distance z of the aircraft during the controlmWhen the lateral deviation is large,
when in use
Figure BDA0001906347530000041
When the temperature of the water is higher than the set temperature,
Figure BDA0001906347530000042
when in use
Figure BDA0001906347530000043
And xmWhen the speed is higher than 3km,
Figure BDA0001906347530000044
when x ismWhen the speed is less than or equal to 3km,N＝4
wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a value obtained by real-time measurement and calculation, and changes along with the position change of the aircraft; x is the number of*Representing the length, x, of the projection of the line between the aircraft point and the target point on the line between the launch point and the target point at the time of the take-off*Take a constant value during the calculation.
Wherein the offset distance z of the aircraft during the controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
Figure BDA0001906347530000045
when x ismWhen the length is less than or equal to 3km, N is 4.
Wherein the offset distance z of the aircraft during the controlmWhen the lateral deviation is small, the device can be used,
N＝4。
wherein the offset distance z of the aircraft during the takeoff controlmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral offset distance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation.
The invention has the advantages that:
according to the high-overload-resistant aircraft sidesway correction system provided by the invention, the radial range from a target when an aircraft starts controlling, the real-time sidesway distance and the projection length of a connecting line between the aircraft located point and a target point on the connecting line of a transmitting point and the target point are taken into consideration of a guidance algorithm, so that the navigation ratio can be adaptively adjusted according to the self sidesway condition and the flight condition of the aircraft, namely, the navigation ratio is increased when the sidesway is large, and the navigation ratio is reduced when the sidesway is small;
in addition, in the high-overload-resistant aircraft lateral deviation correction system, the change of the navigation ratio is smooth and continuous, so that the deflection failure of an actuating mechanism caused by the discontinuity of control quantity is avoided;
the stable position and speed information of the aircraft can be continuously provided under the condition of high overload by arranging the plurality of high overload resistant antennas, and the stability and accuracy of guidance control during high overload are ensured.
Drawings
FIG. 1 is a logic diagram of the overall structure of a high overload resistant aircraft sideslip correction system according to a preferred embodiment of the present invention;
FIG. 2 is a schematic structural diagram of a high overload resistant antenna in a high overload resistant aircraft sideslip correction system according to a preferred embodiment of the present invention;
FIG. 3 illustrates a schematic diagram of the location of the target point, the launch point and the aircraft in accordance with a preferred embodiment of the present invention;
FIG. 4 shows a trajectory graph related to lateral deviation and a shooting distance after control activation, namely a lateral trajectory graph after control activation, in a simulation experiment of the invention;
fig. 5 shows the trajectory profile of the present invention after the activation and before the final guide segment, which is related to the lateral deviation and the shooting distance, i.e. the lateral trajectory profile before entering the final guide segment.
Description of the reference numerals
1-microprocessor module
2-navigation module
21-high overload resistant antenna
22-anti-interference submodule
23-satellite guidance sub-module
3-navigation ratio output module
4-storage tank
41-protective baffle
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the invention, an aircraft sideslip correction system resistant to high overloads is provided, as shown in fig. 1, comprising: the system comprises a microprocessor module 1, a navigation module 2 and a navigation ratio output module 3; wherein the content of the first and second substances,
the microprocessor module 1 is used for calculating the required sideslip overload required by the aircraft sideslip correction;
the navigation module 2 is used for acquiring the position and speed information of the aircraft in real time.
The overload needing to be used is index data used for controlling the workload of a steering engine on the aircraft, and the steering engine on the aircraft performs steering operation according to the calculated overload needing to be used. The lateral bias requiring overload is the lateral overload that the steering engine needs to provide in order to eliminate the lateral bias.
The high overload in the invention means that the ratio of the resultant force of aerodynamic force and engine thrust acting on the aircraft to the gravity of the aircraft is 10000 or more; the high dynamic state means that the aircraft can carry out large-maneuvering flight and has large normal acceleration (the flight condition with the normal acceleration of more than 10g is generally called large-maneuvering flight, and g represents gravity acceleration). In a preferred embodiment, in the microprocessor module 1, the sidesway overload requirement is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the line-of-sight angular rate of the bullet;
preferably, the lateral bias is obtained in real time by the following formula (one):
Figure BDA0001906347530000071
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
Figure BDA0001906347530000072
representing the angular rate of the aircraft's yaw direction line of sight. Since the aim of the application is to study the correction of the lateral deviation, the angular rate of the visual line of the bullet eyes in the lateral deviation direction is abbreviated as the angular rate of the visual line of the bullet eyes, and the lateral deviation requiring overload can also be abbreviated as overload requiring.
The flight speed of the aircraft is measured in real time by a navigation module 2 on the aircraft, the line-of-sight angular rate of the missile can be measured in real time by a sensing element or can be obtained by calculation, and generally, the normal line-of-sight angular rate of the missile and the line-of-sight angular rate of the missile in the lateral deviation direction can be obtained by aircraft position information and target point position information which are solved by satellite signals in a middle guidance section; and directly measuring by a platform laser guide head during final guide section to obtain the normal line-of-sight angular rate of the bullet eyes and the lateral deviation direction line-of-sight angular rate of the bullet eyes, wherein the normal line-of-sight angular rate and the lateral deviation direction line-of-sight angular rate are not particularly limited in the application.
The overload needing to be used is a special term in the field, and in the guidance control process of the guidance aircraft, the overload needing to be used must be firstly solved and converted into an overload instruction, and then the steering engine is controlled to steer;
in a preferred embodiment, as shown in fig. 1, 2, the navigation module 2 comprises an anti-high overload antenna 21, an anti-jamming sub-module 22 and a satellite guidance sub-module 23;
the high overload resistant antenna 21 is in the shape of a sheet, for receiving satellite signals at high overload,
the anti-interference submodule 22 is connected to the high overload resistant antenna 21, and is configured to perform filtering processing on the satellite signal,
the satellite guidance sub-module 23 receives the filtered satellite signals and calculates the position and speed information of the aircraft in real time according to the signals.
Wherein the high overload resistant antenna 21 is arranged on the outer wall of the aircraft,
preferably, as shown in fig. 2, an inward concave accommodating groove 5 is provided on the outer wall of the aircraft, the high overload resistant antenna 21 is installed in the accommodating groove 5, the depth dimension of the accommodating groove 5 is greater than the thickness dimension of the antenna, and a protective baffle 51 is provided outside the high overload resistant antenna 21.
Anti high antenna 21 that transships is fixed in the bottom of holding tank 5, preferably, the holding tank just can hold anti high antenna 21 that transships, and the lateral wall of holding tank can provide the side direction spacing for anti high antenna 21 that transships, prevents that anti high antenna 21 that transships from moving, guard flap 51 is fixed at the top of holding tank, and inside its self was arranged the holding tank completely, can make aircraft surface smooth basically, guard flap external shape suits with the appearance profile of aircraft, can be the arc, also can be dull and stereotyped shape, guard flap inboard and anti high antenna 21 looks butt that transships for fixed anti high antenna 21 that transships, guarantee that anti high antenna 21 that transships can not remove and destroy in acceleration process.
The protective baffle 51 is used for protecting the high-altitude overload antenna 21 on the inner side of the protective baffle in the acceleration stage of the aircraft, and preventing the high-altitude overload antenna 21 from being damaged in the acceleration process, when the aircraft enters the guidance stage, the protective baffle 51 is separated from the aircraft, so that the high-altitude overload antenna 21 is exposed outside, satellite signals can be conveniently received by the high-altitude overload antenna 21, and the protective baffle 51 is prevented from shielding/interfering the satellite signals. Preferably, the high overload resistant antenna 21 is similar to steering engines on an aircraft and needs to be started in the guidance stage, so that the protective baffle 51 and the baffle outside the steering engine of the aircraft can be synchronously controlled and synchronously separated.
The shape of the high overload resistant antenna 21 is a sheet shape, that is, the high overload resistant antenna 21 is a sheet antenna or a thin plate antenna, the antenna can be a rectangular flat plate shape, and also can be an arc plate shape with a radian, and can be arranged according to the outline of the aircraft, in this application, the arc plate shape with the radian is preferred, and is matched with the outline of the aircraft, and in the rolling process of the aircraft, the time for receiving satellite signals by the arc plate antenna with the radian is longer, the signal intensity is better,
preferably, the high overload resistant antenna 21 is provided with a plurality of pieces which are uniformly distributed around the aircraft, preferably, the high overload resistant antenna 21 is provided with 4 pieces, and preferably, the high overload resistant antenna 21 is arranged along the circumferential direction of the rolling of the aircraft in the application, so that the satellite signal receiving capability of the aircraft is not weakened when the aircraft rolls at a high speed.
The anti high antenna 21 that transships of slice in this application compares traditional cone antenna or loop antenna, because slice antenna occupation space area is little, is difficult for receiving external noise or the influence of interference, and slice antenna integrated level is higher moreover, and its satellite signal receptivity is stronger.
Preferably, the sheet-shaped high-overload-resistant antenna 21 can be prepared from the same material as that of a traditional loop antenna or a traditional cone antenna, and the thickness of the high-overload-resistant antenna 21 can be reduced as much as possible on the basis of ensuring stability and physical strength so as to reduce cost;
preferably, the length dimension of the high overload resistant antenna 21 is preferably 120-200 mm, the width dimension of the high overload resistant antenna 21 is preferably 50-70 mm, and the thickness of the high overload resistant antenna is 4-8 mm.
Preferably, the satellite guidance sub-module 23 includes a GPS receiver, a beidou receiver and a GLONASS receiver, which are configured to improve the accuracy and receptivity of acquiring satellite information.
The above equation (a) is also an overload demand calculation equation which is the most widely applied proportion guidance law in the field, but the guidance law in the prior art generally takes a fixed value, and the navigation ratio in the guidance law is adjusted by the navigation ratio output module 3 to give different overload demands.
The navigation ratio output module 3 outputs the lateral deviation distance z of the aircraft according to the control startingmSelects the corresponding navigation ratio N and transmits the navigation ratio N to the microprocessor module 1 in real time.
In the invention, the position of the aircraft, the target position and the launching position are all regarded as one point, namely the position of the aircraft, the target point and the launching point are obtained;
the offset distance zmAs shown in fig. 3, the target point and the launching point are connected by a straight line, and the distance between the point where the aircraft is located and the straight line is the offset distance; to refer to the extent to which the aircraft is sailing off in the lateral direction.
The starting control point is a time node in the flight process of the aircraft, the aircraft flies in an uncontrolled inertia mode before the starting control point, and when the aircraft passes through the time node, a guidance control system on the aircraft starts to work, so that the flight direction of the aircraft is adjusted, the flight deviation is corrected, and the aircraft can finally hit a target.
In a preferred embodiment, the yaw distance z of the aircraft is determined as a function of the departure controlmSelects the corresponding navigation ratio N to calculate the yaw overload.
Wherein preferably the offset z of the aircraft at the time of takeoff controlmWhen the lateral deviation is large,
when in use
Figure BDA0001906347530000101
When the temperature of the water is higher than the set temperature,
Figure BDA0001906347530000102
when in use
Figure BDA0001906347530000103
And xmWhen the speed is higher than 3km,
Figure BDA0001906347530000104
when x ismWhen the length is less than or equal to 3km, N is 4
Wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a variation value obtained by real-time measurement and calculation; as the position of the aircraft changes; x is the number of*Representing the length, x, of the projection of the line between the aircraft point and the target point on the line between the launch point and the target point at the time of the take-off*Taking a constant value in the calculation process; x is the number ofm、x*And zmCan be seen in
The schematic shown in FIG. 3;
according to the above calculation formula, when
Figure BDA0001906347530000105
During the process, the calculation formula of the navigation ratio N is changed, but the value of N is gradually changed along the curve all the time, no abrupt change point exists, the N is smooth and continuous, the aircraft can only provide continuous and stable overload, and larger instantaneous overload is not needed to be provided due to the abrupt change of the navigation ratio, so that the deflection failure of an actuating mechanism caused by the discontinuity of the control quantity is avoided.
In a preferred embodiment, the offset z of the aircraft is measured during the takeoff controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
Figure BDA0001906347530000111
when x ismWhen the length is less than or equal to 3km, N is 4.
At xmWhen the distance between the aircraft and the target is less than or equal to 3km, the aircraft enters a final guide section, and the lateral deviation is corrected to be within an allowable range, so that a guide head on the aircraft can capture the target, and the target is guided by adopting a proportional guide law, wherein the guide head can be a laser guide head and the like.
In a preferred embodiment, the offset z of the aircraft is measured during the takeoff controlmWhen the lateral deviation is small, the device can be used,
n is 4; namely, only fixed navigation ratio is needed to be used for guidance calculation when the vehicle is deflected to a small side.
In a preferred embodiment, the offset z of the aircraft is the distance of the aircraft during the takeoff controlmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral offset distance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation. Corresponding navigation ratio calculation formulas are selected according to different lateral deviation amounts, so that ammunition under different lateral deviation amounts can enable a target point to enter a visual field area before a final guide sectionI.e. the seeker captures the target.
In a preferred embodiment, said xmAnd zmAll are obtained by real-time solution, and the solution process comprises
Pre-stored longitude and latitude coordinates of the launching point and the longitude and latitude coordinates of the target point are called,
the longitude and latitude coordinates of the position of the aircraft are solved in real time through a navigation module,
then x is calculated according to the real-time position relation among the position of the aircraft, the launching point and the target pointmAnd zmThe calculation relationship may be as shown in fig. 3, and a specific calculation method may be a method known in the art, which is not particularly limited in this application.
The aircraft adopts a proportion guidance law based on the gradual change of the navigation ratio of the satellite signals to guide before the final guidance segment, and can capture the laser signals during the final guidance segment, thereby switching to laser guidance at the final guidance segment and greatly improving the hit precision.
In a preferred embodiment, since the present invention is directed to a method and a system for correcting aircraft lateral deviation, during the research process, all points need to be projected onto the same plane for research, so all points involved in the present invention, such as an aircraft point, an emission point, a target point, a start control point, and the like, refer to the projected point of the point on the same horizontal plane.
Examples of the experiments
In order to verify that the high-overload-resistant aircraft sideslip correction system has better sideslip correction capability and can improve the hit rate compared with a traditional guidance control system, simulation is carried out in a simulation verification mode;
setting the shooting distance between the starting control time of the aircraft and the target to be 20km and the lateral deviation to be 3 km; the lateral deviation is required to be within 600m at a position 3km away from a target, namely, the target can be captured by a guide head when entering a final guide section, the flying speed of an aircraft is 300m/s, and the flying direction is parallel to a connecting line from a launching point to a target point; for this example, by ballistic simulationThe ballistic curves in fig. 4 and 5 are obtained in stages, wherein the first scheme (solid line) represents the ballistic curve obtained by using the high overload resistant aircraft lateral deviation correction system provided in the present application, the second scheme (dotted line) represents the ballistic curve obtained by using the conventional proportional guidance algorithm,
Figure BDA0001906347530000121
where N is 4, the resulting ballistic curve.
FIG. 4 shows a diagram of the lateral ballistic trajectory of the aircraft after takeoff; fig. 5 shows lateral ballistic trajectory diagrams before the aircraft enters the final section in both scenarios, i.e., fig. 4 and 5 are not complete lateral ballistic trajectory diagrams, but are partial phase lateral ballistic trajectory diagrams.
The shooting distance in the invention refers to: calculating from the starting control time of the aircraft, and projecting the flight distance of the aircraft on the connecting line of the emission point and the target point; in the experimental example, the shooting distance when starting control is 0, and the shooting distance when just hitting a target is 20 km;
as can be seen from fig. 4, the trajectory correction condition obtained by the aircraft yaw correction system for resisting high overload provided by the present application is obviously due to the trajectory correction condition obtained by the conventional proportional guidance algorithm, and under the same large yaw condition, that is, the yaw is 3km, the aircraft yaw correction system for resisting high overload provided by the present application can effectively control the aircraft to fly to the target, whereas the conventional proportional guidance algorithm finally has a miss distance of about 300m and cannot accurately hit the target.
As can be seen from FIG. 5, the high overload resistant aircraft lateral deviation correction system provided by the present application can be used as desired at xmCorrecting the lateral deviation to be within 600m when the lateral deviation is 3km, and accurately obtaining the lateral deviation to be about 500 m; the traditional proportional guidance algorithm can not complete the task index, and is in xmAbout 850 meters is still left when the lateral deviation is 3 km;
therefore, the comparison can show that the high overload resistant aircraft sideslip correction system provided by the application can effectively correct the sideslip under the condition of high overload and reduce the miss distance.
The present invention has been described above in connection with preferred embodiments, but these embodiments are merely exemplary and merely illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.

Claims (10)

1. An aircraft yaw correction system that is resistant to high overloads, the system comprising:
a microprocessor module (1) for calculating the yaw demand overload required for the correction of the yaw of the aircraft;
and the navigation module (2) is used for acquiring the position and speed information of the aircraft in real time.
2. The system of claim 1,
in the microprocessor module (1), the required sideslip overload is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the angular rate of the line of sight of the missile in the sideslip direction;
preferably, the lateral bias is obtained in real time by the following formula (one):
Figure FDA0001906347520000011
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
Figure FDA0001906347520000012
representing the angular rate of the aircraft's yaw direction line of sight.
3. The system of claim 1,
the navigation module (2) comprises a high overload resistant antenna (21), an anti-interference sub-module (22) and a satellite guidance sub-module (23);
the shape of the high overload resistant antenna (21) is a sheet shape for receiving satellite signals in high overload,
the anti-interference submodule (22) is connected with the anti-high overload antenna (21) and is used for filtering the satellite signals,
and the satellite guidance sub-module (23) receives the satellite signals subjected to filtering processing and calculates the position and speed information of the aircraft in real time according to the signals.
4. The system of claim 3,
the high overload resistant antenna (21) is arranged on the outer wall of the aircraft,
preferably, an inwards concave accommodating groove (4) is formed in the outer wall of the aircraft, the anti-high overload antenna (21) is installed in the accommodating groove (4), and a protective baffle (41) is arranged outside the anti-high overload antenna (21).
5. The system of claim 4,
the anti-high overload antenna (21) is provided with a plurality of pieces which are uniformly distributed around the aircraft, and preferably, the anti-high overload antenna (21) is provided with 4 pieces.
6. The system of claim 2,
the system is also provided with a navigation ratio output module (3) for calculating a navigation ratio;
the navigation ratio output module (3) outputs the lateral deviation distance z of the aircraft during control startingmSelects the corresponding navigation ratio N and transmits the navigation ratio N to the microprocessor module (1) in real time.
7. The system of claim 6,
offset distance z of aircraft during takeoff and controlmWhen the lateral deviation is large,
when in use
Figure FDA0001906347520000021
When the temperature of the water is higher than the set temperature,
Figure FDA0001906347520000022
when in use
Figure FDA0001906347520000023
And xmWhen the speed is higher than 3km,
Figure FDA0001906347520000024
when x ismWhen the length is less than or equal to 3km, N is 4
Wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a value obtained by real-time measurement and calculation, and changes along with the position change of the aircraft; x is the number of*And the length of a connecting line between the aircraft located point and the target point projected on the connecting line between the emission point and the target point at the starting and controlling moment is represented.
8. The system of claim 6,
offset distance z of aircraft during takeoff and controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
Figure FDA0001906347520000025
when x ismWhen the length is less than or equal to 3km, N is 4.
9. The system of claim 6,
offset distance z of aircraft during takeoff and controlmWhen the lateral deviation is small, the device can be used,
N＝4。
10. the system of claim 7, 8, 9,
offset distance z of aircraft when taking off controlmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral deviation isDistance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation.
CN201811533712.4A 2018-12-06 2018-12-14 High-overload-resistant aircraft lateral deviation correction system Active CN111290427B (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CN201811487456 2018-12-06
CN201811487456X 2018-12-06

Publications (2)

Publication Number Publication Date
CN111290427A true CN111290427A (en) 2020-06-16
CN111290427B CN111290427B (en) 2021-07-09

Family

ID=71025080

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201811533712.4A Active CN111290427B (en) 2018-12-06 2018-12-14 High-overload-resistant aircraft lateral deviation correction system

Country Status (1)

Country Link
CN (1) CN111290427B (en)

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4492352A (en) * 1982-09-22 1985-01-08 General Dynamics, Pomona Division Noise-adaptive, predictive proportional navigation (NAPPN) guidance scheme
US4494202A (en) * 1982-09-22 1985-01-15 General Dynamics, Pomona Division Fourth order predictive, augmented proportional navigation system terminal guidance design with missile/target decoupling
EP1352206B1 (en) * 2000-12-13 2005-02-09 Saab Ab Method for controlling a missile
CN103245256A (en) * 2013-04-25 2013-08-14 北京理工大学 Multi-missile cooperative attack guidance law designing method
CN103954179A (en) * 2014-04-30 2014-07-30 北京理工大学 System for evaluating disturbance rejection rate parasitical loop of strap down infrared seeker
CN104035335A (en) * 2014-05-27 2014-09-10 北京航空航天大学 High accuracy longitudinal and cross range analytical prediction method based smooth gliding reentry guidance method
CN105043171A (en) * 2015-06-30 2015-11-11 北京航天长征飞行器研究所 Longitudinal guidance method of rocket projectile with inclined-angle restraining
CN207116690U (en) * 2017-09-13 2018-03-16 武汉雷可达科技有限公司 Missile-borne conformal antenna and missile-borne conformal antenna system

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4492352A (en) * 1982-09-22 1985-01-08 General Dynamics, Pomona Division Noise-adaptive, predictive proportional navigation (NAPPN) guidance scheme
US4494202A (en) * 1982-09-22 1985-01-15 General Dynamics, Pomona Division Fourth order predictive, augmented proportional navigation system terminal guidance design with missile/target decoupling
EP1352206B1 (en) * 2000-12-13 2005-02-09 Saab Ab Method for controlling a missile
CN103245256A (en) * 2013-04-25 2013-08-14 北京理工大学 Multi-missile cooperative attack guidance law designing method
CN103954179A (en) * 2014-04-30 2014-07-30 北京理工大学 System for evaluating disturbance rejection rate parasitical loop of strap down infrared seeker
CN104035335A (en) * 2014-05-27 2014-09-10 北京航空航天大学 High accuracy longitudinal and cross range analytical prediction method based smooth gliding reentry guidance method
CN105043171A (en) * 2015-06-30 2015-11-11 北京航天长征飞行器研究所 Longitudinal guidance method of rocket projectile with inclined-angle restraining
CN207116690U (en) * 2017-09-13 2018-03-16 武汉雷可达科技有限公司 Missile-borne conformal antenna and missile-borne conformal antenna system

Non-Patent Citations (10)

* Cited by examiner, † Cited by third party
Title
TALAAT IBRAHIM，ETC.: ""Optimum Dynamic Navigation Ratio for Launch Vehicles"", 《2014 IEEE AEROSPACE CONFERENCE》 *
YUE MENG，等: ""Research on the visual/inertial integrated carrier landing guidance algorithm"", 《INTERNATIONAL JOURNAL OF ADVANCED ROBOTIC SYSTEMS》 *
孟克子，等: ""多约束条件下的最优中制导律设计"", 《系统工程与电子技术》 *
常超，等: ""战术导弹GPS制导控制系统设计"", 《导箭与制导学报》 *
徐平，等: ""基于无人机平台制导控制半实物仿真系统研究"", 《中北大学学报（自然科学版）》 *
李辕，等: ""拦截高速机动目标偏置比例制导律研究"", 《装备学院学报》 *
林琳: ""导引头与制导控制律参数综合优化方法研究"", 《中国优秀硕士学位论文全文数据库 工程科技II辑》 *
毕永涛，等: ""直接侧向力与气动力复合控制导弹脱靶量分析"", 《哈尔滨工业大学学报》 *
王嘉鑫，等: ""引入参考目标的比例导引制导律研究"", 《航天控制》 *
臧路尧，等: ""一种适用于红外制导弹药的变增益比例导引率"", 《红外与激光工程》 *

Also Published As

Publication number Publication date
CN111290427B (en) 2021-07-09

Similar Documents

Publication Publication Date Title
CN111351401B (en) Anti-sideslip guidance method applied to strapdown seeker guidance aircraft
EP3043164B1 (en) Aero-wave instrument for the measurement of the optical wavefront disturbances in the airflow around airborne systems
CN111692919B (en) Precise guidance control method for aircraft with ultra-close range
CN114502465B (en) Determination of attitude by pulsed beacons and low cost inertial measurement units
US20050045761A1 (en) Proactive optical trajectory following system
CN108931155B (en) Autonomous guidance system independent of satellite navigation extended-range guidance ammunition
CN111434586B (en) Aircraft guidance control system
CN111290002B (en) Aircraft lateral deviation correction system applied to satellite signal unstable area
US8433460B1 (en) Onboard sensor suite for determining projectile velocity
CN111221348B (en) Sideslip correction method applied to remote guidance aircraft
CN111412793B (en) Anti-sideslip full-range coverage control system applied to remote guidance aircraft
CN111380405B (en) Guidance control system of high-dynamic aircraft with strapdown seeker
CN111290427B (en) High-overload-resistant aircraft lateral deviation correction system
JP7262845B2 (en) Integrated guidance and control system with high load capacity
CN112180971A (en) Multi-mode guidance method and system for multi-rotor aircraft
CN112445230A (en) High-dynamic aircraft multi-mode guidance system and guidance method under large-span complex environment
US8237095B2 (en) Spot leading target laser guidance for engaging moving targets
CN111273682B (en) Sideslip correction method based on virtual target point
US4160250A (en) Active radar missile launch envelope computation system
CN111284690B (en) Composite range-extending aircraft capable of correcting lateral deviation
CN114153226A (en) Unmanned aerial vehicle view field keeping and guiding method and system assisted by dynamic sight line information
CN115617063A (en) Aircraft guidance control device with falling angle constraint and method
CN116203849B (en) Falling angle constraint control system applied to remote composite guidance aircraft
CN113759954B (en) Composite guidance method for maneuvering target
RU2129696C1 (en) Sighting system

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant