CN110081883A - Low cost integrated navigation system and method suitable for high speed rolling flight device - Google Patents

Abstract

The invention discloses a low-cost integrated navigation system and method for a high-speed rolling aircraft. In the system, the position information and speed information of the aircraft are obtained in real time through a satellite navigation module, and the three-axis acceleration and pitch of the aircraft are obtained in real time through an inertial navigation module. Angular velocity, yaw angular velocity and roll angular velocity; the required overload is calculated by the microprocessor, and the required overload is transmitted to the actuator, and the steering gear is controlled by the actuator; among them, the three-axis acceleration, pitch angular velocity and yaw The angular velocity of the flight can be directly measured, and the angular velocity of the roll can be obtained through real-time calculation, which can be directly used as the input of the microprocessor.

Description

Low-cost integrated navigation system and method suitable for high-speed rolling aircraft

technical field

The invention relates to the field of guided aircraft control, in particular to a low-cost integrated navigation system and method suitable for high-speed rolling aircraft.

Background technique

The navigation system of the aircraft obtains the flight parameters of the aircraft in real time, calculates the required overload according to the specific parameters, obtains enough flight parameters through the satellite system and the strapdown inertial navigation system, and then comprehensively uses the flight parameters to achieve The purpose of real-time control of the aircraft;

The traditional strapdown inertial navigation (SINS) system uses an accelerometer to measure the linear acceleration of the aircraft, a gyroscope to measure the angular velocity of the aircraft, and then calculates the speed, position and attitude angle of the aircraft through SISN. However, for an aircraft with a high rotational speed, its rotational speed may exceed the measurement range of the gyroscope, resulting in the failure to measure the roll angular velocity of the aircraft.

In addition, in the prior art, when the roll angle is measured by the gyroscope, due to the accumulation of cumulative errors, the roll angle directly measured by the gyroscope contains a relatively large cumulative error. When performing subsequent calculations based on the measurement results, timely Correct the cumulative error and verify it with other measurements;

Moreover, for the aircraft guidance system, the cost of the gyroscope is relatively high. If the number of gyroscopes can be reduced, the overall cost of the guidance system can be effectively reduced.

Based on the above-mentioned technical problems, the present inventor has done in-depth research on the setting positions of the gyroscopes and accelerometers on the existing aircraft, and proposed a corresponding solution method to solve the problem of low gyroscope accuracy by reducing the number of gyroscope settings. problems, and reduce the cost of setting up the aircraft, and obtain accurate guidance parameters through a reasonable calculation process to improve the reliability of the aircraft.

SUMMARY OF THE INVENTION

In order to overcome the above-mentioned problems, the inventor has carried out determined research and designed a low-cost integrated navigation system and method suitable for high-speed rolling aircraft. In this system, the position information and speed information of the aircraft are obtained in real time through the satellite navigation module. The navigation module obtains the three-axis acceleration, pitch angular velocity, yaw angular velocity and roll angular velocity of the aircraft in real time; calculates the required overload through the microprocessor, and transmits the required overload to the actuator, and controls the steering gear to work through the actuator Wherein, triaxial acceleration, pitch angular velocity and yaw angular velocity can all be directly measured, and roll angular velocity can be obtained by real-time calculation, which can be directly used as the input quantity of the microprocessor, thereby completing the present invention.

Specifically, the object of the present invention is to provide a low-cost integrated navigation system suitable for high-speed rolling aircraft, which system includes a satellite navigation module 1, an inertial navigation module 2, a microprocessor 3 and an actuator 4;

Wherein, the satellite navigation module 1 is used to obtain the position information and speed information of the aircraft in real time;

The inertial navigation module 2 is used to obtain the triaxial acceleration, pitch rate, yaw rate and roll rate of the aircraft in real time;

The microprocessor 3 is used to receive the information transmitted by the satellite navigation module 1 and the inertial navigation module 2 in real time, and calculate the required overload accordingly, and pass the required overload to the actuator 4;

The actuator 4 is used to control the rudder operation of the steering gear according to the received overload.

Wherein, the satellite navigation module 1 includes an antenna 11, a receiver filter 12 and a satellite receiver 13;

The receiver filter 12 is used for filtering the signal received by the antenna 11 and delivering the filtered signal to the satellite receiver 13 .

Wherein, the inertial navigation module 2 includes an accelerometer and a gyroscope,

Measuring the three-axis acceleration of the aircraft through the accelerometer,

The pitch rate and yaw rate of the aircraft are measured by the gyroscope.

Wherein, the accelerometer includes a first accelerometer, a second accelerometer and a third accelerometer arranged at the position of the center of mass of the aircraft, wherein the measurement result f _ox =f ₁ of the first accelerometer represents the direction along the X _b axis Acceleration, the measurement result of the second accelerometer f _oy =f ₂ represents the acceleration along the Y _b -axis direction, and the measurement result of the third accelerometer f _oz =f ₃ represents the acceleration along the Z _b -axis direction;

Preferably, the accelerometer also includes _a fourth accelerometer and a fifth accelerometer arranged in a plane _OYbZb perpendicular to the axis of the aircraft,

Wherein, the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b -axis direction;

The fourth accelerometer is arranged on the Y _b axis, and the distance between the center of mass and the position of the aircraft is L, and its measurement result is f ₄ ;

The fifth accelerometer is arranged on the Z _b axis, and the distance between it and the center of mass of the aircraft is L, and its measurement result is f ₅ .

Wherein, the gyroscope includes a first gyroscope for measuring pitch rate, and a second gyroscope for measuring yaw rate;

Wherein, the first gyroscope measurement result ω _y represents the pitch angular velocity;

The second gyroscope measurement ω _z represents the yaw rate.

Wherein, the inertial navigation module 2 also includes a roll angular velocity calculation module 21,

Described roll angular velocity calculation module 21 solves the roll angular velocity of aircraft in real time by following formula (1);

Among them, _ωx represents the roll angular velocity; Indicates the roll angular acceleration; Δω _x indicates the increment of the roll angular velocity;

Preferably, the Δω _x is obtained by the following formula (2):

in, Indicates the square root output at time k+1, Indicates the square root output at time k.

Among them, the Obtained by the following formula (3), Obtained by the following formula (4):

Wherein, described microprocessor 3 obtains required overload by following formula (5):

Among them, a represents the need to use overload, N represents the proportional guidance law, v represents the speed of the aircraft, Indicates the bullet eye line of sight angular velocity.

The present invention also provides a low-cost integrated navigation method suitable for high-speed rolling aircraft, the method comprising:

measuring the acceleration along the _Xb -axis direction by the first accelerometer;

measuring the acceleration along the _Yb -axis direction by the second accelerometer;

Measuring the acceleration along the Z _b -axis direction by the third accelerometer;

Obtain its measurement result f ₄ by the fourth accelerometer measurement;

Obtain its measurement result f ₅ by the fifth accelerometer measurement;

measuring the pitch rate by the first gyroscope;

measuring the yaw rate through the second gyroscope;

Wherein, the fourth accelerometer and the fifth accelerometer are all arranged in the plane OY _b Z _b perpendicular to the aircraft axis, and the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b axis direction;

The fourth accelerometer is arranged on the Y _b axis, and the distance between the center of mass and the position of the aircraft is L;

The fifth accelerometer is arranged on the Z _b axis, and the distance between the fifth accelerometer and the center of mass of the aircraft is L.

Among them, the roll angular velocity calculation module receives the information measured by the accelerometer and the gyroscope, and calculates the roll angular velocity accordingly,

Then calculate the required overload according to the roll angular velocity, pitch angular velocity and yaw angular velocity, three-axis acceleration, position information of the aircraft, and velocity information of the aircraft.

The beneficial effects that the present invention has include:

(1) According to the low-cost integrated navigation system applicable to high-speed rolling aircraft provided by the present invention, the accelerometer is set to measure the acceleration, and then the roll rate is jointly calculated by the acceleration, the pitch rate, and the yaw rate, so as to obtain a high-speed roll rate at a high speed. Under the condition of rotating speed, obtain the rolling angular velocity in real time with the accuracy meeting the requirements of use;

(2) The low-cost integrated navigation system applicable to high-speed rolling aircraft provided by the present invention adopts MEMS inertial measurement devices and uses accelerometers instead of gyroscopes for measurement, which reduces the cost and improves the performance of the inertial navigation system.

Description of drawings

Fig. 1 shows a low-cost integrated navigation system suitable for high-speed rolling aircraft and a logic diagram of the overall structure of the method according to a preferred embodiment of the present invention;

Fig. 2 shows a schematic structural diagram of the relative position relationship between the accelerometer and the angular rate gyroscope in the low-cost integrated navigation system and method suitable for high-speed rolling aircraft according to a preferred embodiment of the present invention;

Fig. 3 shows the structure schematic diagram of the antenna in the low-cost integrated navigation system and method suitable for high-speed rolling aircraft according to a preferred embodiment of the present invention;

Fig. 4 shows the flight track figure that experiment example 1 and experiment example 2 obtain in the low-cost integrated navigation system and method that are applicable to high-speed rolling aircraft according to a preferred embodiment of the present invention;

Fig. 5 shows the flight trajectory diagrams obtained in Experimental Example 3 and Experimental Example 4 in the low-cost integrated navigation system and method suitable for high-speed rolling aircraft according to a preferred embodiment of the present invention.

Explanation of reference numbers:

1-Satellite Navigation Module

2- Inertial Navigation Module

21-Roll angular velocity calculation module

3- Microprocessor

4- Executive body

11-antenna

12- Receiver filter

13-Satellite receiver

14-Accommodating slot

15-Protective baffle

Detailed ways

The present invention will be further described in detail through the drawings and examples below. Through these descriptions, the features and advantages of the present invention will become more apparent.

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 various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise indicated.

According to the low-cost integrated navigation system and method suitable for high-speed rolling aircraft provided by the present invention, as shown in Figure 1, when the rotational speed of the aircraft is above 10 revolutions per second, it can be called a high rotational speed. The maximum speed of the aircraft is generally 20 rpm;

In a preferred embodiment, as shown in Figure 1, the low-cost integrated navigation system includes a satellite navigation module 1, an inertial navigation module 2, a microprocessor 3 and an actuator 4;

Wherein, the satellite navigation module 1 is used to obtain the position information and speed information of the aircraft in real time;

The inertial navigation module 2 is used to obtain the triaxial acceleration, pitch angular velocity, yaw angular velocity and roll angular velocity of the aircraft in real time; wherein, the triaxial acceleration represents the acceleration along the X _b axis direction in _Fig . Acceleration and acceleration along the Z _b -axis direction.

The microprocessor 3 is used to receive the information that the satellite navigation module 1 and the inertial navigation module 2 transmit in real time, and solve the required overload accordingly, and pass the required overload to the actuator 4; It is made of existing chips in the art, and this application does not specifically limit the specific model selection, as long as it can meet the above functions;

The actuator 4 is used to control the steering of the steering gear according to the received overload, so as to complete the control work of the navigation system; the actuator 4 includes a pneumatic steering gear or an electric steering gear.

In a preferred embodiment, as shown in Figure 3, satellite navigation module 1 comprises antenna 11, receiver filter 12 and satellite receiver 13; Said satellite receiver 13 can be in the GPS/Beidou/GLONASS receiver One or more of the above; setting the above-mentioned multiple receivers can improve the accuracy and receiving ability of obtaining satellite information.

Wherein, the antenna 11 is arranged on the outer wall of the aircraft, through which the satellite signal is received, and the receiver filter 12 is used for filtering the signal received by the antenna 11, and passing the filtered signal to the satellite receiver 13. Analyze the satellite signal through the satellite receiver 13 to obtain the position information of the aircraft and the speed information of the aircraft.

Preferably, as shown in FIG. 3 , a concave accommodation groove 14 is provided on the outer wall of the aircraft, the antenna 11 is installed in the accommodation groove 14 , and the depth dimension of the accommodation groove 14 is larger than the The thickness dimension of the antenna, and a protective baffle 15 is provided outside the antenna 11 .

The antenna 11 is fixed on the bottom of the receiving groove 14. Preferably, the receiving groove can just accommodate the antenna 11, and the side wall of the receiving groove can provide a lateral limit for the antenna 11 to prevent the antenna 11 from moving in series. The protective baffle 15 It is fixed on the top of the holding tank, and it is completely placed inside the holding tank, so that the outer surface of the aircraft can be basically smooth. The outer shape of the protective baffle adapts to the outline of the aircraft, and can be curved or flat. The inside of the protective baffle abuts against the antenna 11 to fix the antenna 11 to ensure that the antenna 11 will not move or be damaged during acceleration.

The protective baffle 15 is used to protect the antenna 11 inside the aircraft during the acceleration phase, preventing the antenna 11 from being damaged during the acceleration process. When the aircraft enters the guidance phase, the protective baffle 15 is separated from the aircraft, so that The antenna 11 is exposed outside, so as to facilitate receiving satellite signals with the antenna 11 and avoid shielding/interfering satellite signals with the protective baffle 15 . Preferably, the antenna 11 is similar to the steering gear on the aircraft, and it needs to start work only in the guidance phase, so the protective baffle 15 and the external baffle of the aircraft steering gear can be synchronously controlled and disengaged.

The shape of the antenna 11 is a sheet shape, that is, the antenna 11 is a sheet antenna or a thin plate antenna. The antenna can be a rectangular flat plate or a curved plate with a radian. Contour setting, in this application, it is preferably a curved plate shape with a radian, which matches the outline of the aircraft, and during the rolling process of the aircraft, the arc plate antenna with a radian takes longer to receive satellite signals , the signal strength is better,

Preferably, the antenna 11 is provided with multiple pieces, which are evenly distributed around the aircraft. Preferably, the antenna 11 is provided with 4 pieces. In this application, the antenna 11 is preferably arranged along the circumferential direction of the aircraft roll, In order to ensure that the satellite signal receiving ability of the aircraft will not be weakened when the aircraft rolls at high speed.

Compared with the traditional conical antenna or loop antenna, the chip antenna 11 in this application is less susceptible to external noise or interference because the chip antenna occupies a small area, and the chip antenna is more integrated, and its satellite The signal acceptance ability is stronger.

Preferably, the sheet antenna 11 can be prepared with the same material as the traditional loop antenna or cone antenna, and the antenna 11 can reduce the thickness as much as possible on the basis of ensuring stability and physical strength, so as to reduce costs;

Preferably, the length of the antenna 11 is preferably 120-200 mm, the width of the antenna 11 is preferably 50-70 mm, and its thickness is 4-8 mm.

In a preferred embodiment, the inertial navigation module 2 includes an accelerometer and a gyroscope,

Measuring the three-axis acceleration of the aircraft through the accelerometer,

Measuring the pitch rate and yaw rate of the aircraft through the gyroscope;

The accelerometer described in the application can choose the existing accelerometer in the prior art, can realize the above-mentioned function, the gyroscope described in the application can choose the existing gyroscope in the prior art, can realize the above-mentioned function.

In a preferred embodiment, as shown in FIG. 2 , the accelerometers include a first accelerometer, a second accelerometer and a third accelerometer arranged at the center of mass of the aircraft, wherein the measurement of the first accelerometer The result f _ox =f ₁ represents the acceleration along the X _b axis, the measurement result of the second accelerometer f _oy =f ₂ represents the acceleration along the Y _b axis, and the measurement result of the third accelerometer f _oz =f ₃ represents the acceleration along the X b axis. The acceleration in Z _b -axis direction; said Fig. 2 is the coordinate system of the aircraft itself, also can be referred to as the projectile coordinate system, described projectile coordinate system can be converted between the ground coordinate system, can select for use in the prior art Conventional existing methods are used, and there is no special limitation in this application.

Preferably, the accelerometer also includes _a fourth accelerometer and a fifth accelerometer arranged in a plane _OYbZb perpendicular to the axis of the aircraft,

Wherein, the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b -axis direction;

The fourth accelerometer is arranged on the Y _b axis, the distance between the fourth accelerometer and the center of mass position of the aircraft is L, and its measurement result is f ₄ ;

The configuration parameters of the fourth accelerometer are as follows:

θ _y4 =0°; θ _z4 =90°; r _y4 =L; r _z4 =0

Wherein, θ _y4 represents the angle between the fourth accelerometer f ₄ and the Y _b axis; θ _z4 represents the angle between the fourth accelerometer f ₄ and the Z _b axis; _ry 4 represents the angle between the fourth accelerometer f ₄ and the Y _b axis Distance; r _z4 represents the distance between the fourth accelerometer f ₄ and the Y _b axis;

The fifth accelerometer is arranged on the Z _b axis, the distance between the fifth accelerometer and the center of mass of the aircraft is L, and its measurement result is f ₅ .

The configuration parameters of the fifth accelerometer are as follows:

θ _y5 =0°; θ _z5 =90°; r _y5 =0; r _z5 =L

Wherein, θ _y5 represents the angle between the _fifth accelerometer f5 and the Y _b axis; θ _z5 represents the angle between the _fifth accelerometer f5 and the Z _b axis; _ry5 represents the angle between the _fifth accelerometer f5 and the Y _b axis Distance; r _z5 represents the distance between the fifth accelerometer f ₅ and the Y _b axis;

In this application, by specifically selecting the location and measurement direction of the fourth accelerometer and the fifth accelerometer, this configuration can be easily installed and occupies less space), and finally a more accurate roll angle can be obtained through calculation.

In a preferred embodiment, the gyroscope includes a first gyroscope for measuring pitch rate, and a second gyroscope for measuring yaw rate;

Wherein, the first gyroscope measurement result ω _y represents the pitch angular velocity;

The second gyroscope measurement ω _z represents the yaw rate.

In a preferred embodiment, the inertial navigation module 2 also includes a roll angular velocity calculation module 21,

Described roll angular velocity calculation module 21 solves the roll angular velocity of aircraft in real time by following formula (1);

Among them, _ωx represents the roll angular velocity; Indicates the roll angular acceleration; Δω _x represents the increment of the roll angular velocity; the roll angular velocity is calculated in real time, specifically after the aircraft is controlled in real time, and the calculation frequency is 50-100Hz, preferably is 50Hz, at the initial moment, that is, when k=0, both the roll angular acceleration and the increment of the roll angular velocity are zero, and t represents the instantaneous moment, that is, the instantaneous moment at which ω _x is obtained from the solution, and the t≥k+ 1 means any non-zero moment after the start of control.

Preferably, the Δω _x is obtained by the following formula (2):

in, Indicates the square root output at time k+1, that is, the absolute value of the roll angular velocity obtained at time k+1, Indicates the square root output at time k, that is, the absolute value of the roll angular velocity obtained at time k.

In a preferred embodiment, the Obtained by the following formula (3), Obtained by the following formula (4):

Both the formula (3) and the formula (4) are reasonably derived according to information such as the setting position and setting direction of the fourth accelerometer and the fifth accelerometer,

Wherein, the output equation of the uniaxial accelerometer is referring to the following formula (6);

Among them, ω _x represents the roll angular velocity, ω _y represents the yaw angular velocity, ω _z represents the pitch angular velocity, θ _y and θ _z represent the sensitive vectors to the y and z axes respectively, and the sensitive vector means that the accelerometer is sensitive to this direction. , r _y , r _z represent the installation vector respectively, and the installation vector refers to the position and direction of the accelerometer from the coordinate axis when the accelerometer is installed; f ₁ , f ₂ , f ₃ represent the three-axis acceleration respectively.

Simplify formula (6) to get:

Substituting the corresponding sensitive vector and installation vector into the above two formulas respectively, formula (3) and formula (4) can be obtained;

In a preferred embodiment, the microprocessor 3 obtains the required overload through the following formula (5):

Among them, a represents the need to use overload, N represents the proportional guidance law, v represents the speed of the aircraft, Indicates the line-of-sight rate of bullets;

In a preferred embodiment, the system also includes a power supply module, the power supply module is connected to the thermal power source loaded on the aircraft, and integrates the input and output of the entire circuit to prevent the system from being burned due to problems such as short circuit ;The power supply module can provide the required rated voltage to each subsystem, such as the satellite navigation module and inertial navigation module, etc., to ensure the normal operation of the components; for the specific needs of some subsystems, the power supply module can provide it with a reset voltage signal .

The present invention also provides a low-cost integrated navigation method suitable for high-speed rolling aircraft, in which first an accelerometer and a gyroscope are installed on a specific position of the high-speed rolling aircraft;

measuring the acceleration along the _Xb -axis direction by the first accelerometer;

measuring the acceleration along the _Yb -axis direction by the second accelerometer;

Measuring the acceleration along the Z _b -axis direction by the third accelerometer;

Obtain its measurement result f ₄ by the fourth accelerometer measurement;

Obtain its measurement result f ₅ by the fifth accelerometer measurement;

measuring the pitch rate by the first gyroscope;

measuring the yaw rate through the second gyroscope;

Wherein, the fourth accelerometer and the fifth accelerometer are all arranged in the plane OY _b Z _b perpendicular to the aircraft axis, and the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b axis direction;

The fourth accelerometer is arranged on the Y _b axis, and the distance between the center of mass and the position of the aircraft is L;

The fifth accelerometer is arranged on the Z _b axis, and the distance between the fifth accelerometer and the center of mass of the aircraft is L.

Preferably, the roll angular velocity calculation module receives the information measured by the accelerometer and the gyroscope, and calculates the roll angular velocity accordingly,

Then calculate the required overload according to the roll angular velocity, pitch angular velocity and yaw angular velocity, three-axis acceleration, position information of the aircraft, and velocity information of the aircraft;

Among them, the roll angular velocity calculation module calculates the roll angle through the following formulas (1), (2), (3) and (4),

ω _x represents the roll angular velocity; Indicates the roll angular acceleration; Δω _x represents the increment of the roll angular velocity; the roll angular velocity is calculated in real time, specifically after the aircraft is controlled in real time, and the calculation frequency is 50-100Hz, preferably is 50Hz, at the initial moment, that is, when k=0, both the roll angular acceleration and the increment of the roll angular velocity are zero, and t represents the instantaneous moment, that is, the instantaneous moment at which ω _x is obtained from the solution, and the t≥k+ 1 means any non-zero moment after starting the control;

in, Indicates the square root output at time k+1, that is, the absolute value of the roll angular velocity obtained at time k+1, Indicates the square root output at time k, that is, the absolute value of the roll angular velocity obtained at time k;

Preferably, the required overload is calculated by the following formula (5);

In this application, a represents the need to use overload, N represents the proportional guidance law, and v represents the speed of the aircraft, specifically the relative speed between the aircraft and the target, Indicates the bullet eye line of sight angular velocity.

x _r =x _T -x _M , y _r =y _T -y _M , v _rx =v _Mx -v _Tx , v _ry =v _My -v _Ty

N represents the navigation ratio, which is generally selected as 2 to 4, and the preferred value in this application is 3;

x _T represents the position of the target along the x-axis in the ground coordinate system;

y _T represents the position of the target along the y-axis in the ground coordinate system;

x _M represents the position of the aircraft along the x-axis in the ground coordinate system;

y _M represents the position of the aircraft along the y-axis in the ground coordinate system;

x _r represents the relative distance between the aircraft and the target along the x-axis in the ground coordinate system;

y _r represents the relative distance between the aircraft and the target along the y-axis in the ground coordinate system;

v _Mx represents the speed of the aircraft along the x-axis in the ground coordinate system;

v _My represents the speed of the aircraft along the y-axis in the ground coordinate system;

v _Tx represents the speed of the target along the x-axis in the ground coordinate system;

v _Ty represents the speed of the target along the y-axis in the ground coordinate system;

v _rx represents the relative velocity between the aircraft and the target along the x-axis in the ground coordinate system;

V _ry represents the relative velocity between the aircraft and the target along the y-axis in the ground coordinate system; for the ground coordinate system, the origin of the coordinates is usually taken as the launch point, that is, the take-off point of the aircraft, and the x-axis direction is the direction from the launch point to the target point direction, 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 speed of the aircraft itself are determined by the Obtained by the satellite navigation module and the inertial navigation module; after calculating the required overload, it is also necessary to adjust the rudder timing according to the roll angular velocity, so as to control the rudder work of the steering gear.

Experimental example:

Experimental example 1:

In the simulation experiment, the target point at a distance of 20km is taken as the target, and a rolling aircraft with a rotational speed of 5 rpm is selected to hit the target. The rolling aircraft is equipped with a satellite navigation module for obtaining the position information and speed information of the aircraft in real time. , is also provided with an inertial navigation module, wherein the inertial navigation module includes an accelerometer, a gyroscope and a rotational angular velocity calculation module, the three-axis acceleration of the aircraft is measured by the accelerometer, and the pitch angular velocity and yaw of the aircraft are measured by the gyroscope Angular velocity, solve the roll angular velocity of aircraft in real time by the roll angular velocity calculation module; :

The flight trajectory of the aircraft is shown by the solid line in FIG. 4 , and it can be known from the trajectory that the aircraft finally hits the target.

Experimental example 2:

In the simulation experiment, with the target point at a distance of 20 km as the target, a rolling aircraft with a rotating speed of 5 rpm was selected to strike the target. The rolling aircraft is provided with a navigation system that is basically the same as that in Embodiment 1. In Experimental Example 2, the navigation system uses the gyroscope to directly measure the roll angle of the aircraft, and does not use the angular velocity calculation module to calculate it in real time;

The flight trajectory of the aircraft is shown by the dotted line in FIG. 4 , and it can be seen from the trajectory that the aircraft finally hits the target.

Experimental example 3:

In the simulation experiment, the target point at a distance of 20 km was selected as the target, and the rolling aircraft whose rotation speed was basically the same as that in Experimental Example 1 was selected to hit the target, the only difference being that the rotation speed of the aircraft was 15 rpm;

The flight trajectory of the aircraft is shown by the solid line in FIG. 5 , and it can be known from the trajectory that the aircraft finally hits the target.

Experimental example 4:

In the simulation experiment, the target point at a distance of 20km was taken as the target, and a rolling aircraft with a rotational speed of 15 rpm was selected to hit the target. The rolling aircraft was basically the same as the rolling aircraft in Experimental Example 3. In the inertial navigation module of the navigation system in Example 4, a gyroscope for directly measuring the roll angle of the aircraft is installed instead of the roll angular velocity calculation module. Get the roll angle of the aircraft;

The flight trajectory of the aircraft is shown by the dotted line in FIG. 5 , and it can be known from the trajectory that the aircraft finally misses the target.

It can be seen from the above experimental example that when the rolling aircraft as given in Experimental Examples 1 and 3 is used to hit the target, the aircraft can effectively hit the target when the rotational speed is low or high; when using such as Experimental Example 2, When the rolling aircraft given in 4 strikes the target, if the aircraft maintains a low speed, it can still hit the target, but if the aircraft maintains a high speed, it cannot hit the target.

The present invention has been described above in conjunction with preferred embodiments, but these embodiments are only exemplary and serve as illustrations only. On this basis, various replacements and improvements can be made to the present invention, all of which fall within the protection scope of the present invention.

Claims (10)

1. A low-cost integrated navigation system suitable for high-speed rolling aircraft, characterized in that the system includes satellite navigation module (1), inertial navigation module (2), microprocessor (3) and executive agency (4) ; Wherein, the satellite navigation module (1) is used to obtain the position information and speed information of the aircraft in real time; The inertial navigation module (2) is used to obtain the triaxial acceleration, pitch angular velocity, yaw angular velocity and roll angular velocity of the aircraft in real time; The microprocessor (3) is used to receive the information transmitted by the satellite navigation module (1) and the inertial navigation module (2) in real time, and calculate the required overload accordingly, and transmit the required overload to the actuator (4); The actuator (4) is used to control the rudder operation of the steering gear according to the received overload. 2. the low-cost integrated navigation system applicable to high-speed rolling aircraft according to claim 1, characterized in that, The satellite navigation module (1) includes an antenna (11), a receiver filter (12) and a satellite receiver (13); The receiver filter (12) is used for filtering the signal received by the antenna (11), and delivering the signal obtained by the filtering process to the satellite receiver (13). 3. the low-cost integrated navigation system applicable to high-speed rolling aircraft according to claim 1, characterized in that, The inertial navigation module (2) includes an accelerometer and a gyroscope, Measuring the three-axis acceleration of the aircraft through the accelerometer, The pitch rate and yaw rate of the aircraft are measured by the gyroscope. 4. the low-cost integrated navigation system applicable to high-speed rolling aircraft according to claim 3, characterized in that, The accelerometer comprises a first accelerometer, a second accelerometer and a third accelerometer arranged at the position of the center of mass of the aircraft, wherein the measurement result f _ox =f ₁ of the first accelerometer represents the acceleration along the X _b axis direction, The measurement result f _oy = f ₂ of the second accelerometer represents the acceleration along the Y _b axis direction, and the measurement result f _oz = f ₃ of the third accelerometer represents the acceleration along the Z _b axis direction; Preferably, the accelerometer also includes _a fourth accelerometer and a fifth accelerometer arranged in a plane _OYbZb perpendicular to the axis of the aircraft, Wherein, the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b -axis direction; The fourth accelerometer is arranged on the Y _b axis, and the distance between the center of mass and the position of the aircraft is L, and its measurement result is f ₄ ; The fifth accelerometer is arranged on the Z _b axis, and the distance between it and the center of mass of the aircraft is L, and its measurement result is f ₅ . 5. the low-cost integrated navigation system applicable to high-speed rolling aircraft according to claim 3, characterized in that, The gyroscope includes a first gyroscope for measuring pitch rate, and a second gyroscope for measuring yaw rate; Wherein, the first gyroscope measurement result ω _y represents the pitch angular velocity; The second gyroscope measurement ω _z represents the yaw rate. 6. The low-cost integrated navigation system suitable for high-speed rolling aircraft according to one of claims 3-5, characterized in that, The inertial navigation module (2) also includes a roll angular velocity calculation module (21), Described roll angular velocity calculation module (21) solves the roll angular velocity of aircraft in real time by following formula (1); Among them, _ωx represents the roll angular velocity; Indicates the roll angular acceleration; Δω _x indicates the increment of the roll angular velocity; Preferably, the Δω _x is obtained by the following formula (2): in, Indicates the square root output at time k+1, Indicates the square root output at time k. 7. The low-cost integrated navigation system suitable for high-speed rolling aircraft according to claim 6, characterized in that, said Obtained by the following formula (3), Obtained by the following formula (4): 8. The low-cost integrated navigation system suitable for high-speed rolling aircraft according to claim 1, characterized in that, Described microprocessor (3) obtains required overload by following formula (5): Among them, a represents the need to use overload, N represents the proportional guidance law, v represents the speed of the aircraft, Indicates the bullet eye line of sight angular velocity. 9. A low-cost integrated navigation method suitable for high-speed rolling aircraft, characterized in that the method comprises, measuring the acceleration along the _Xb -axis direction by the first accelerometer; measuring the acceleration along the _Yb -axis direction by the second accelerometer; Measuring the acceleration along the Z _b -axis direction by the third accelerometer; Obtain its measurement result f ₄ by the fourth accelerometer measurement; Obtain its measurement result f ₅ by the fifth accelerometer measurement; measuring the pitch rate by the first gyroscope; measuring the yaw rate through the second gyroscope; Wherein, the fourth accelerometer and the fifth accelerometer are all arranged in the plane OY _b Z _b perpendicular to the aircraft axis, and the measurement directions of the fourth accelerometer and the fifth accelerometer are all towards the Y _b axis direction; The fourth accelerometer is arranged on the Y _b axis, and the distance between the center of mass and the position of the aircraft is L; The fifth accelerometer is arranged on the Z _b axis, and the distance between the fifth accelerometer and the center of mass of the aircraft is L. 10. the low-cost integrated navigation method applicable to high-speed rolling aircraft according to claim 9, characterized in that, Receive the information measured by the accelerometer and gyroscope through the roll angular velocity calculation module, and calculate the roll angular velocity accordingly, Then calculate the required overload according to the roll angular velocity, pitch angular velocity and yaw angular velocity, three-axis acceleration, position information of the aircraft, and velocity information of the aircraft.