CN111045437A - Anti-high-overload integrated guidance control system - Google Patents
Anti-high-overload integrated guidance control system Download PDFInfo
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- CN111045437A CN111045437A CN201811187337.2A CN201811187337A CN111045437A CN 111045437 A CN111045437 A CN 111045437A CN 201811187337 A CN201811187337 A CN 201811187337A CN 111045437 A CN111045437 A CN 111045437A
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- 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
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- 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
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
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- 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
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- 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
Abstract
The invention discloses an integrated guidance control system for resisting high overload, wherein a navigation system, a guidance system and a control system are integrally designed, data are transmitted to a microprocessor module for processing, and the influence of time lag of signal transmission among subsystems on aircraft control is avoided; in addition, because the line-of-sight angular rates obtained by the navigation module and the guidance module are not accurate enough and have errors, the microprocessor module performs navigation calculation by using the fused and calculated line-of-sight angular rates, and the hit precision can be further improved.
Description
Technical Field
The invention relates to a guidance control system of an aircraft, in particular to an integrated guidance control system capable of resisting high overload.
Background
The traditional guidance control system is generally composed of a space orientation gyro, a platform laser seeker, a pneumatic steering engine and the like, but the elements essentially determine that the designed high-dynamic aircraft cannot resist destructive influence of high overload on the system, a plurality of original pieces no longer have theoretical working performance under the condition of high dynamic, and the traditional guidance control system usually needs to design a navigation system, a guidance system and a control system respectively according to specific models, so that a large amount of time and energy are needed, and the problems of incompatibility and mismatching among subsystems exist. This conventional design makes the command transfer between the navigation system, guidance system and control system even more delayed, especially under high overload conditions, with serious consequences. Therefore, the general integrated navigation guidance control system capable of resisting high overload is designed, can be integrated into a whole and normally work under the flight conditions of high overload and large maneuver, and has important significance for improving the performance of a high-dynamic aircraft.
For the above reasons, the present inventors have made intensive studies on the existing guidance control system, and have awaited the design of a new guidance control system capable of solving the above-mentioned problems.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention makes a keen study and designs an integrated guidance control system for resisting high overload, wherein a navigation system, the guidance system and a control system are integrally designed, and data are transmitted to a microprocessor module for processing, so that the influence of time lag of signal transmission among subsystems on aircraft control is avoided, and a sheet-shaped high overload resisting antenna is specially designed, so that compared with a traditional conical antenna and an improved annular antenna, the sheet-shaped antenna not only has stronger satellite signal acceptance capability, but also has the characteristic of resisting high overload, and can stably work under the condition of high dynamic high overload; in addition, because the line-of-sight angular rates obtained by the navigation module and the guidance module are not accurate enough and have errors, the microprocessor module performs navigation calculation by using the fused and calculated line-of-sight angular rates, so that the hit precision can be further improved, and the method is completed.
In particular, the invention aims to provide an integrated guidance control system resistant to high overloads, which is capable of precisely controlling the attitude of an aircraft in the event of high overloads.
The system comprises an integrated microprocessor module 1, which is used for providing demand overload for accurately controlling the attitude of the aircraft and generating a rudder deflection instruction.
Wherein, this system still includes:
the navigation module 2 is used for acquiring the position and speed information of the aircraft in real time;
the guidance module 3 is used for acquiring the roll angle, the rotating speed and the line-of-sight angle information of the aircraft in real time;
and the control module 4 is used for executing the rudder deflection instruction, feeding the rudder deflection state back to the microprocessor module 1 in real time, and performing feedback compensation on the rudder deflection instruction to improve the rudder-striking precision.
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 5 is arranged on the outer wall of the aircraft, the high overload resistant antenna 21 is arranged in the accommodating groove 5, and a protective baffle 51 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 guidance module 3 comprises a geomagnetic sensor 31 and a strapdown laser seeker 32,
the geomagnetic sensor 31 is used for sensing the roll angle and the rotating speed of the aircraft in real time,
the strapdown laser seeker 32 is used to sense the line-of-sight angle of the aircraft in real time.
The control module 4 comprises a steering engine servo submodule 41 and an electric steering engine 42;
the steering engine servo submodule 41 is used for receiving a steering deviation instruction, converting the steering deviation instruction into a steering engine signal and controlling the electric steering engine 42 to steer;
the steering engine servo submodule 41 is also used for sensing the steering deflection state of the electric steering engine.
Wherein, in the microprocessor module 1, a first line-of-sight angular rate is obtained according to the position and speed information of the aircraft obtained from the navigation module 2
Then obtaining a second line-of-sight angular rate according to the line-of-sight angle of the aircraft obtained from the guidance module 3
Will be aligned with
And
final line-of-sight angular rate obtained after fusion processing
As the line-of-sight angular velocity of the aircraft.
The invention has the advantages that:
(1) according to the guidance control system, the geomagnetic sensor replaces a traditional space orientation gyroscope, the strapdown laser seeker replaces a platform seeker, and the electric steering engine replaces a pneumatic steering engine, so that the system has good high overload resistance, can be applied to high-overload and large-motor aircrafts, can provide accurate guidance data information, and can reduce damage risks;
(2) in the guidance control system, the navigation system, the guidance system and the control system are integrally designed, so that the influence of time lag of signal transmission among subsystems on aircraft control is avoided;
(3) the guidance control system uses the strapdown laser guide head to replace a platform guide head, and integrally designs the navigation system, the guidance system and the control system, thereby saving the loading space of the aircraft and improving the load capacity of the aircraft;
(4) the guidance control system has broad-spectrum adaptability, is suitable for aircrafts of various models, and is a general guidance control system scheme.
Drawings
FIG. 1 is a general block diagram of an integrated guidance control system for resisting high overload according to a preferred embodiment of the invention;
FIG. 2 is a schematic structural diagram of a high overload resisting antenna in the integrated guidance control system for resisting high overload according to a preferred embodiment of the invention;
fig. 3 shows the aircraft trajectory under three angular rate conditions in the experimental example.
The reference numbers illustrate:
1-microprocessor module
2-navigation module
21-high overload resistant antenna
22-anti-interference submodule
23-satellite guidance sub-module
3-guidance module
31-geomagnetic sensor
32-strapdown laser seeker
4-control Module
41-steering engine servo submodule
42-electric steering engine
5-holding tank
51-protective baffle
6-Power supply module
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 integrated guidance control system for resisting high overload, disclosed by the invention, as shown in the figure 1, the system can accurately control the attitude of an aircraft under the condition of high overload.
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 (generally, the flight condition with the normal acceleration of more than 10g is called large-maneuvering flight, and g represents gravity acceleration); generally, under the condition of high overload/high dynamic, sensitive devices on the aircraft, such as a space gyroscope, an inertial gyroscope, a platform type laser guide head and the like, lose the measurement reference, so that an accurate measurement result is difficult to obtain.
In a preferred embodiment, the system includes an integrated microprocessor module 1 for providing demand overload for precise control of the attitude of the aircraft and generating rudder deflection commands. The integration means that the calculation parts of the navigation module, the guidance module and the control module in the guidance control system are integrated, basic data are transmitted to the microprocessor module after being sensitively obtained, and the basic data are uniformly processed by the microprocessor module, so that the problem of delay of transmission instructions among the navigation system, the guidance system and the control system is avoided, and system noise caused by mutual interference of signals can be reduced;
preferably, as shown in fig. 1, the system further comprises: the navigation module 2, the guidance module 3 and the control module 4; wherein the content of the first and second substances,
the navigation module 2 is used for acquiring the position and speed information of the aircraft in real time;
the guidance module 3 is used for acquiring the roll angle, the rotating speed and the line-of-sight angle information of the aircraft in real time;
and the control module 4 is used for executing the rudder deflection instruction, feeding the rudder deflection state back to the microprocessor module 1 in real time, and performing feedback compensation on the rudder deflection instruction to improve the rudder-striking precision. The feedback compensation is that the real-time rudder deflection state is used as the input quantity of the control module 4, so that a closed-loop system is formed. By feedback compensation, the accuracy of the rudder deflection command can be improved. The feedback compensation in the present invention may be feedback compensation known in the art, and is not particularly limited.
In a further 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.
In a preferred embodiment, as shown in fig. 1, the guidance module 3 comprises a geomagnetic sensor 31 and a strapdown laser seeker 32,
the geomagnetic sensor 31 is used for sensing the roll angle and the rotating speed of the aircraft in real time and transmitting the sensing information to the microprocessor module 1, and compared with a traditional space orientation gyroscope, the geomagnetic sensor is not limited by the frame angle, so that the geomagnetic sensor can normally work under the condition of high overload; the sensitivity referred to in the present invention is understood to be a measurement or a sensing for obtaining dynamic information of the aircraft.
The geomagnetic sensor 31 in the present invention is a commonly used geomagnetic sensor in the art, and does not have any special requirement for this, and can satisfy the above functions;
the strapdown laser seeker 32 is a commonly used strapdown laser seeker in the field, does not need to be specially required, and can meet the functions;
the strapdown laser seeker 32 is used for sensing the sight angle of the aircraft in real time and transmitting the measurement information to the microprocessor module, compared with a traditional platform seeker, the strapdown seeker does not need to be installed on the platform and can be directly and fixedly connected to the aircraft, the loading space of the aircraft is saved, the working performance is good under the condition of high overload, and the integrated implementation of a navigation guidance control system is facilitated.
In a preferred embodiment, as shown in fig. 1, the control module 4 comprises a steering engine servo sub-module 41 and an electric steering engine 42;
the steering engine servo submodule 41 is used for receiving a steering deviation instruction, converting the steering deviation instruction into a steering engine signal and controlling the electric steering engine 42 to steer; the steering engine 42 is preferably a proportional electric steering engine, has better high overload resistance compared with a traditional pneumatic steering engine, and can realize accurate control of the attitude of the aircraft particularly under the condition of large maneuvering;
the steering engine servo submodule 41 is also used for sensing the steering deflection state of the electric steering engine.
In a preferred embodiment, in the microprocessor module 1, a first line-of-sight angular rate is derived from the position and speed information of the aircraft obtained from the navigation module 2
Then obtaining a second line-of-sight angular velocity according to the line-of-sight angle of the aircraft obtained from the guidance module 3Rate of change
Will be aligned with
And
final line-of-sight angular rate obtained after fusion processing
As the line-of-sight angular velocity of the aircraft.
In particular, the present invention relates to a method for producing,
obtained by the following formula (one):
wherein x isr＝xT-xM，yr＝yT-yM，vrx＝vMx-vTx，vry＝vMy-vTy
xTRepresenting the position of the target along the x-axis direction under the ground coordinate system;
yTrepresenting the position of the target along the y-axis direction under the ground coordinate system;
xMrepresenting the position of the aircraft along the x-axis direction under a ground coordinate system;
yMrepresenting the position of the aircraft along the y-axis direction under a ground coordinate system;
xrrepresenting the relative distance between the aircraft and the target along the x-axis direction under a ground coordinate system;
yrrepresenting the relative distance between the aircraft and the target along the y-axis direction under a ground coordinate system;
vMxrepresenting the aircraft in the x-axis direction under the ground coordinate systemThe speed of (d);
vMyrepresenting the speed of the aircraft in the direction of the y axis under a ground coordinate system;
vTxrepresenting the speed of the target along the x-axis direction under the ground coordinate system;
vTyrepresenting the speed of the target along the y-axis direction under the ground coordinate system;
vrxrepresenting the relative speed of the aircraft and the target along the x-axis direction under a ground coordinate system;
vryrepresenting the relative speed of the aircraft and the target along the y-axis direction under a ground coordinate system; regarding the ground coordinate system, the origin of coordinates is usually taken as the emission point, the x-axis direction is the direction from the emission point to the target point, and the y-axis direction is perpendicular to the x-axis and vertically upward; when the target is a static target, the speed of the target is 0, the position of the target is pre-filled on the aircraft, and the position and the speed of the aircraft are sensitively obtained by devices on the aircraft.
Generally, a navigation system inevitably has time lag due to the fact that satellite signals are received, and the line-of-sight angular rate is calculated according to the calculated speed and position information, so that the finally obtained line-of-sight angular rate is not accurate enough and has errors; the line-of-sight angle measured by the strapdown laser head can obtain the line-of-sight angular rate after being calculated by a differentiator, but the result calculated by the differentiator usually has larger error. Therefore, the microprocessor module performs data fusion on the line-of-sight angular rate calculated by the navigation module and the strapdown laser seeker, and improves the accuracy of the finally obtained line-of-sight angular rate, so that a more accurate rudder deflection instruction is calculated, and the hit accuracy of the aircraft is improved.
Line of sight angle q2Can be directly obtained by sensing the strapdown laser seeker 32, and can obtain a second line-of-sight angular rate after being calculated by a differentiator
The fusion calculation was performed by the following formula (two):
in the invention, the microprocessor module obtains the final visual angle rate through data fusion
Then, based on the angular rate of the line of sight
The required overload is calculated and, in particular,
generally speaking, the navigation ratio is 2-4.
In a preferred embodiment, the guidance control system further comprises a power supply module, wherein the power supply module is connected to a thermal power supply loaded on the aircraft and integrates input and output of the whole circuit, so that the system is prevented from being burnt out due to short circuit and the like; the power supply module can provide rated voltage required by each module, and ensures that each element works normally; for the particular needs of a molecular system, the power supply module may provide a reset voltage signal thereto.
The invention also provides a guidance control method for resisting high overload, wherein the guidance control system for resisting high overload is adopted to control the aircraft. In particular, in the method, the first step,
satellite signals are received by the high overload resistant antenna 21 under high overload conditions,
the satellite signals are filtered by the anti-jamming sub-module 22,
the satellite guidance sub-module 23 is used for solving the position and speed information of the aircraft in real time according to the signals;
the roll angle and the rotating speed of the aircraft are sensed in real time through the geomagnetic sensor 31,
sensing the line-of-sight angle of the aircraft in real time by the strapdown laser seeker 32;
the overload required is resolved through the microprocessor module 1, and a rudder deflection instruction is generated,
the steering wheel deflection instruction is converted into a steering wheel signal through the steering wheel servo submodule 41,
the electric steering engine is used for steering, so that the attitude of the aircraft is accurately controlled.
Wherein the microprocessor module first obtains a final line-of-sight angular rate through data fusion
And then the overload is resolved.
Experimental example:
launching a plurality of aircrafts with the same model to the same target position at the same launching place, wherein for each aircraft, the target point is within the range, the rotating speed of the aircraft in the advancing process is controlled to be 6-10 r/s, the overload on each aircraft is over 10000g, and the flight track of each aircraft is mapped to obtain a graph 3;
wherein a first aircraft is equipped with an anti-high overload antenna as shown in fig. 2, a satellite signal is received through the anti-high overload antenna, and the line-of-sight angular rate is calculated in a microprocessor module of the aircraft through a data fusion algorithm
Then, the overload is calculated according to the calculated overload, wherein, in the calculation process, the visual angle rate is obtained by the following formula (two)
Wherein the content of the first and second substances,
q2is directly obtained by the direct sensitivity of a strapdown laser seeker,
by making a pair q2Do differential operationObtaining;
xr＝xT-xM，yr＝yT-yM，vrx＝vMx-vTx，vry＝vMy-vTy
xTthe value is 15000, yTA value of 0, xM、yM、xr、yr、vMx、vMyThe values of (A) are all obtained sensitively by the implementation of components on board the aircraft and are variable in real time, generally speaking, xrIs less than 15000, yrThe value of (a) is between-200 and 200, vMxIs between 0 and 1000, vTxAnd vTyAll take the value 0. The trajectory curve of the aircraft is shown in FIG. 3
Where the length is in meters and the velocity is in meters per second.
Compared with the first aircraft, the antenna on the second aircraft is a spiral antenna or a straight antenna in the prior art, and the microprocessor module on the second aircraft performs navigation guidance by using the line-of-sight angular rate obtained by the navigation module, and overload is needed for resolving; the trajectory curve of the aircraft is shown in FIG. 3
And (4) showing.
Compared with the first aircraft, the antenna of the third aircraft is a spiral antenna or a straight antenna in the prior art, and the microprocessor module on the third aircraft performs navigation guidance by using the line-of-sight angular rate obtained by the guidance module and solves the overload; the trajectory curve of the aircraft is shown in FIG. 3
And (4) showing.
As can be seen from FIG. 3, the line-of-sight angular rate generated through data fusion
The more accurate rudder deflection instruction can be provided, the aircraft can accurately hit a target under the action of the rudder deflection instruction generated by the line-of-sight angular rate, and the circular probability error can be controlled within 15 meters; and under the action of rudder deviation commands generated by the other two kinds of line-of-sight angular rates, the probability error of the circle is about 100 meters generally, and the final result has 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 integrated guidance control system for resisting high overload is characterized in that the system can accurately control the attitude of an aircraft under the condition of high overload.
2. The guidance control system of claim 1,
the system comprises an integrated microprocessor module (1) for providing a demand overload for the precise control of the attitude of the aircraft and generating a rudder deflection command.
3. The guidance control system according to claim 2,
the system further comprises:
the navigation module (2) is used for acquiring the position and speed information of the aircraft in real time;
the guidance module (3) is used for acquiring roll angle, rotating speed and line-of-sight angle information of the aircraft in real time;
and the control module (4) is used for executing the rudder deflection instruction, feeding back the rudder deflection state to the microprocessor module (1) in real time, and performing feedback compensation on the rudder deflection instruction to improve the rudder-turning precision.
4. The guidance control system of claim 3,
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.
5. The guidance control system of claim 4,
the high overload resistant antenna (21) is arranged on the outer wall of the aircraft,
preferably, an inwards concave accommodating groove (5) is formed in the outer wall of the aircraft, the high-altitude overload-resistant antenna (21) is installed in the accommodating groove (5), and a protective baffle (51) is arranged outside the high-altitude overload-resistant antenna (21).
6. The guidance control system of claim 5,
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.
7. The guidance control system of claim 3,
the guidance module (3) comprises a geomagnetic sensor (31) and a strapdown laser seeker (32),
the geomagnetic sensor (31) is used for sensing the roll angle and the rotating speed of the aircraft in real time,
the strapdown laser seeker (32) is used for sensing the line-of-sight angle of the aircraft in real time.
8. The guidance control system of claim 3,
the control module (4) comprises a steering engine servo submodule (41) and an electric steering engine (42);
the steering engine servo sub-module (41) is used for receiving a steering deviation instruction, converting the steering deviation instruction into a steering engine signal and controlling an electric steering engine (42) to steer;
the steering engine servo submodule (41) is also used for sensing the steering deviation state of the electric steering engine.
9. The guidance control system according to claim 2,
in the microprocessor module (1), a first line-of-sight angular rate is obtained from the position and speed information of the aircraft obtained from the navigation module (2)
Then obtaining a second line-of-sight angular rate from the line-of-sight angle of the aircraft obtained from the guidance module (3)
Will be aligned with
And
final line-of-sight angular rate obtained after fusion processing
As the line-of-sight angular velocity of the aircraft.
10. A guidance control method for resisting high overload, which is characterized in that the guidance control system for resisting high overload of any one of claims 1-9 is used for controlling an aircraft.
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| CN201811187337.2A CN111045437A (en) | 2018-10-12 | 2018-10-12 | Anti-high-overload integrated guidance control system |
| PCT/CN2019/092205 WO2020073682A1 (en) | 2018-10-12 | 2019-06-21 | Anti-high overload integrated guidance and control system |
| JP2021502935A JP7262845B2 (en) | 2018-10-12 | 2019-06-21 | Integrated guidance and control system with high load capacity |
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| CN201811187337.2A CN111045437A (en) | 2018-10-12 | 2018-10-12 | Anti-high-overload integrated guidance control system |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN113885334A (en) * | 2021-10-25 | 2022-01-04 | 北京无线电测量研究所 | Robust terminal guidance law calculation method and device based on flight target interception |
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| CN116700306B (en) * | 2023-06-02 | 2024-02-27 | 北京理工大学 | Integrated guidance control method for strapdown guided aircraft |
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| JP7262845B2 (en) | 2023-04-24 |
| JP2022539926A (en) | 2022-09-14 |
| WO2020073682A1 (en) | 2020-04-16 |
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