CN116817895A - Satellite guidance system containing guidance instrument and guidance method - Google Patents
Satellite guidance system containing guidance instrument and guidance method Download PDFInfo
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Abstract
本发明公开了一种含有制导仪的卫星制导系统及制导方法，系统包括制导仪和弹上卫星制导组件；制导仪获取目标相对制导仪的位置并将该信息发送给弹上卫星制导组件，并与弹上卫星制导组件进行差分定位；弹上卫星制导组件根据制导仪发送的星历信息和制导仪自身位置信息进行差分定位，解算导弹自身相对制导仪的位置，再根据制导仪发送的目标相对制导仪的位置信息解算导弹相对目标的位置和速度，并根据导弹相对目标的运动信息生成控制指令，控制导弹飞向目标。本发明能够对移动目标进行高精度打击。
The invention discloses a satellite guidance system and a guidance method containing a guidance instrument. The system includes a guidance instrument and a satellite guidance assembly on the missile; the guidance instrument obtains the position of the target relative to the guidance instrument and sends the information to the satellite guidance assembly on the missile, and Differential positioning is performed with the satellite guidance component on the missile; the satellite guidance component on the missile performs differential positioning based on the ephemeris information sent by the guidance instrument and the position information of the guidance instrument itself, and calculates the position of the missile relative to the guidance instrument, and then based on the target sent by the guidance instrument The position information of the relative guidance instrument calculates the position and speed of the missile relative to the target, and generates control instructions based on the movement information of the missile relative to the target to control the missile to fly to the target. The invention can carry out high-precision strikes against moving targets.
Description
Technical Field
The invention relates to the technical field of guidance control, in particular to a satellite guidance system with a guidance instrument and a guidance method.
Background
Satellite guidance is a low-cost guidance mode, and is widely applied to the fields of unmanned aerial vehicles, missiles and the like, and the working principle of the guidance mode is as follows: installing a satellite receiver on an aircraft, and acquiring coordinates (precision, latitude and altitude) of the position of the aircraft in real time by using the satellite receiver; the coordinate information of the emission point and the target point is stored in the flight control computer in advance; in the flight process of the aircraft, the flight control computer calculates the relative position of the aircraft and the target in real time, calculates a control instruction through a guidance control algorithm, and controls the aircraft to fly to the target. The positioning accuracy of the traditional satellite receiver is about 10 m.
Differential satellite navigation is a satellite positioning technology which utilizes a differential satellite receiving reference station with known accurate three-dimensional coordinates to obtain a pseudo-range correction amount or a position correction amount, and then sends the pseudo-range correction amount or the position correction amount to a user (other satellite receivers) in real time or in a future to correct measurement data of the user so as to improve satellite positioning accuracy. The key of the technology is that a differential satellite receiving reference table with known three-dimensional coordinates is provided, a satellite receiver is arranged on the reference table, and the pseudo-range error of the area near the reference table is corrected by comparing satellite information received by the satellite receiver with known accurate coordinate information of the reference table, so that the positioning accuracy of the satellite receiver is improved (the positioning can be improved to be within 0.1 m).
The satellite guidance has the advantages of low cost and no matter what the satellite guidance is after being launched, but the guidance mode also has limitations in use, firstly, the precision is insufficient, the hit precision of the traditional satellite guidance aircraft is about 10m, and the traditional satellite guidance aircraft can only be used for attacking large-scale buildings such as clustered targets or bridges, and the accurate attack on single-point targets is difficult to realize; secondly, the moving target is difficult to attack, the target coordinates of the satellite guided aircraft to be attacked are set in a flight control computer before attack, the satellite guided aircraft is usually a fixed target or a target with known motion rules and weaker maneuverability (such as a large ship), and for targets with certain maneuverability such as ground moving vehicles, the satellite guided aircraft is difficult to attack accurately. Differential satellite positioning requires accurate three-dimensional coordinates of a reference station although positioning accuracy is high, and accurate positioning of a certain reference station is difficult to achieve in a complex battlefield environment.
In summary, the existing satellite guidance system has the problems that high-precision hitting of a target is difficult to achieve, moving targets are difficult to hit, and the like. Therefore, in order to complement the gap in the application capability of the existing satellite guidance system, it is necessary to develop a satellite guidance system having a high-precision striking capability for a moving object.
Disclosure of Invention
In view of the above, the present invention provides a satellite guidance system and a satellite guidance method including a guidance instrument, which can strike a moving target with high accuracy.
The technical scheme adopted by the invention is as follows:
a satellite guidance system with a guidance instrument comprises the guidance instrument and an on-bullet satellite guidance assembly;
the guidance instrument comprises a control cabin, a satellite receiver I, a laser range finder, an angular rate gyroscope, a cradle head, a signal transmitter and an observation device; the satellite receiver I is used for receiving ephemeris information sent by the navigation satellite and calculating the position information and the movement speed of the guidance instrument according to the ephemeris information; the system comprises a laser range finder, an angular rate gyroscope and an observing and sighting device, wherein the laser range finder is used for measuring the distance between a target and the guidance instrument; the control cabin is used for providing a navigation coordinate system reference and a horizontal reference, and the control cabin calculates the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser ranging machine relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument; the signal transmitter transmits the position information, ephemeris information, relative position and movement speed of the target relative to the guidance instrument to the satellite guidance assembly on the bullet;
the on-board satellite guidance assembly comprises a satellite receiver II, a signal receiver and a flight control computer; the satellite receiver II is used for receiving ephemeris information sent by the navigation satellite; the signal receiver is used for receiving the position information, the ephemeris information, the relative position and the movement speed information of the target relative to the guidance instrument of the guidance instrument sent by the signal transmitter on the guidance instrument; the flight control computer is used for calculating the position and the speed of the guided missile relative to the guidance instrument according to the ephemeris information received by the satellite receiver II, the position information and the ephemeris information of the guidance instrument received by the signal receiver, and calculating the position and the speed of the guided missile relative to the target in real time by combining the relative position and the speed information of the target relative to the guidance instrument received by the signal receiver, and calculating the flight control instruction according to the position and the speed.
Further, the guidance instrument further comprises an angular position sensor, wherein the angular position sensor is fixed at the rotating shaft of the two-degree-of-freedom cradle head and is used for measuring the azimuth angle and the elevation angle of the target relative to the guidance instrument.
Further, the angular rate gyroscope is replaced by an angular position sensor, and the angular position sensor is fixed at a rotating shaft of the two-degree-of-freedom cradle head and is used for measuring the azimuth angle and the altitude angle of a target relative to the guidance instrument; the angular velocity of the target relative to the guidance instrument is calculated from the azimuth angle and the elevation angle.
Further, the control cabin comprises an electronic compass, a level meter and a guidance computer;
the electronic compass is used for providing a navigation coordinate system reference; the level gauge is used for providing a level reference; the guidance computer is used for calculating the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser range finder relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument.
Further, the sight direction of the sight device is parallel to the light outlet direction of the laser range finder.
Further, the flight control computer calculates the flight control instruction by adopting a proportional guidance law or a speed tracking guidance law.
The satellite guidance method with the guidance instrument adopts the satellite guidance system, and the guidance method comprises the following steps:
leveling a guidance instrument, enabling an angular rate gyro to finish initial alignment, and determining the initial orientation of a laser range finder in a navigation coordinate system; the satellite receiver I starts self-checking, searches for satellites and receives ephemeris information sent by the navigation satellites;
step two, a shooter searches and locks a target through the sighting device, and simultaneously rotates the sighting device and a laser range finder fixedly connected with the sighting device, so that a laser beam emitted by the laser range finder can irradiate the target;
step three, the laser range finder sends the measured distance from the target to the guidance instrument to the control cabin; the angular rate gyroscope sends the high-low angle and azimuth angle rate of the laser range finder to the control cabin, calculates the high-low angle and azimuth angle of the laser range finder according to the high-low angle and azimuth angle rate and sends the high-low angle and azimuth angle to the control cabin;
step four, the control cabin calculates the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser range finder relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument;
step five, launching a missile loaded with an onboard satellite guidance assembly;
step six, the signal transmitter transmits the relative position and movement speed information of the target calculated by the control cabin relative to the guidance instrument, ephemeris information received by the satellite receiver I on the guidance instrument and position information of the guidance instrument itself calculated by the control cabin to the satellite guidance assembly on the bullet;
step seven, a signal receiver in the missile-borne satellite guidance assembly receives the information sent by the signal transmitter and transmits the information to the flight control computer; the flight control computer calculates the relative distance and speed between the missile and the target by combining the ephemeris information received by the satellite receiver II and the information received by the signal receiver, so as to calculate the flight control instruction and control the missile to fly to the target.
Further, the guidance instrument further comprises an angular position sensor, and in the third step, the angular position sensor sends the measured high-low angle and azimuth angle of the laser range finder to the control cabin.
Further, the flight control computer in the step seven calculates the flight control instruction by adopting a proportional guidance law or a speed tracking guidance law.
The beneficial effects are that:
the satellite guidance system of the invention comprises a guidance instrument and an on-bullet satellite guidance assembly. The guidance instrument can acquire the position of the target relative to the guidance instrument, send the information to the satellite guidance assembly on the bullet, and perform differential positioning with the satellite guidance assembly on the bullet; the satellite guidance assembly on the missile can conduct differential positioning according to ephemeris information sent by the guidance instrument and position information of the guidance instrument, the position of the missile relative to the guidance instrument is calculated, the position and the speed of the missile relative to the target are calculated according to the position information of the target relative to the guidance instrument sent by the guidance instrument, and a control instruction is generated according to the movement information of the missile relative to the target so as to control the missile to fly to the target. Therefore, the invention has high-precision striking capability on the moving target, overcomes the limitation that the traditional satellite guided weapon can only attack the stationary cluster target, and improves the hit precision of the satellite guided weapon from 10m to within 1.5 m.
Secondly, compared with other guidance systems achieving the same hit precision (1.5 m), such as laser semi-active guidance, image guidance, infrared guidance and the like, which need to be provided with a tens of thousands of elements or even hundreds of thousands of elements of guide heads, the invention adopts satellite guidance, and can achieve the hit precision the same as the guidance system by only needing to be provided with a satellite receiver with thousands of elements, and the cost is lower.
Drawings
FIG. 1 is a schematic diagram of the working principle of the present invention;
FIG. 2 is a flow chart of the operation of the method of the present invention;
FIG. 3 is a schematic view of the structure of the guidance instrument of the present invention;
the system comprises a 1-control cabin, a 2-azimuth position sensor, a 3-cradle head, a 4-high-low angular position sensor, a 5-signal transmitter, a 6-satellite receiver I, a 7-laser range finder and an 8-sighting device.
Detailed Description
The invention will now be described in detail by way of example with reference to the accompanying drawings.
The invention provides a satellite guidance system with a guidance instrument, which is shown in figure 1 and comprises the guidance instrument and an on-bullet satellite guidance assembly; the guidance instrument can acquire the position of the target relative to the guidance instrument, send the information to the satellite guidance assembly on the bullet, and perform differential positioning with the satellite guidance assembly on the bullet; the satellite guidance assembly on the missile can conduct differential positioning according to ephemeris information sent by the guidance instrument and position information of the guidance instrument, the position of the missile relative to the guidance instrument is calculated, the position and the speed of the missile relative to the target are calculated according to the position information of the target relative to the guidance instrument sent by the guidance instrument, and a control instruction is generated according to the movement information of the missile relative to the target so as to control the missile to fly to the target.
As shown in fig. 3, the guidance instrument in this embodiment includes a control cabin 1, an azimuth position sensor 2, a cradle head 3, a high-low angular position sensor 4, a triaxial angular rate gyro, a signal transmitter 5, a satellite receiver i 6, a laser range finder 7, and an observation device 8. The guidance instrument control cabin 1 comprises a guidance computer, a level instrument and an electronic compass, wherein the electronic compass is used for providing a navigation coordinate system reference, and the level instrument is used for providing a level reference.
If the guidance instrument comprises an angular rate gyroscope and does not comprise an angular position sensor, the azimuth angle and the high-low angle of the target relative to the guidance instrument can be obtained by the calculation of the angular rate integration; if the guidance instrument comprises an angular position sensor and does not comprise an angular rate gyro, the angular rate of the target relative to the guidance instrument is obtained by the azimuth angle and the high-low angle micro-resolution calculation.
The control cabin 1 can be placed on a tripod or carried on a vehicle such as an airplane. The guidance computer can receive the azimuth position information of the target measured by the azimuth position sensor 2 relative to the guidance instrument, the height angle position information of the target measured by the height angle position sensor 4 relative to the guidance instrument, the distance information of the target measured by the laser range finder 7 to the guidance instrument, the height angle speed and azimuth speed information of the target measured by the triaxial angular rate gyro relative to the guidance instrument, and the azimuth information of the initial pointing direction of the laser range finder 7 in the navigation coordinate system output by the electronic compass in the control cabin 1. And the guidance computer calculates the relative position and the movement speed of the target relative to the guidance instrument through a target movement calculation algorithm according to the information.
The target motion calculation algorithm specifically comprises the following steps:
constructing a navigation coordinate system: the position of the guidance instrument is taken as the origin of coordinates, the forward eastern direction is taken as the X axis of the navigation coordinate system, the north direction is taken as the Y axis of the navigation coordinate system, and the sky direction is taken as the Z axis of the navigation coordinate system.
Defining the azimuth lambda of the target relative to the guidance instrument T The north-to-east direction is positive for the included angle between the connecting line of the target and the guidance instrument and the north direction; let the initial orientation azimuth of the laser rangefinder 7 measured by the electronic compass be lambda 0 The azimuth angle of the laser range finder 7, which is measured by the azimuth angle position sensor 2, relative to the initial orientation when pointing at the target is Δλ, then the azimuth angle of the target relative to the guidance instrument in the navigation coordinate system is:
λ T ＝λ 0 +Δλ (1)
defining the height angle theta of the target relative to the guidance instrument T The included angle between the connecting line of the target and the guidance instrument and the horizontal plane is the positive direction of the connecting line of the target and the guidance instrument above the horizontal plane; after the level gauge is adjusted to the zero position, the high and low angle output by the high and low angle position sensor 4 is delta theta, and then the high and low angle of the target relative to the guidance gauge under the navigation coordinate system is as follows:
θ T ＝Δθ (2)
let the distance between the target measured by the laser range finder 7 and the guidance instrument be L T Then, in the navigation coordinate system, the position coordinates of the target relative to the guidance instrument are:
let the azimuth angle rate measured by the triaxial angular rate gyro in the laser range finder 7 be ω λ High and low angular rate omega θ The movement speed of the target relative to the guidance instrument is:
wherein,,for the derivative of the target-to-guidance distance, the target-to-guidance distance L can be measured by the laser rangefinder 7 T And (5) obtaining differentiation.
The cradle head 3 is arranged on the guidance instrument control cabin 1, the cradle head 3 and the guidance instrument control cabin 1 can rotate in a single degree of freedom in the azimuth direction, and an azimuth position sensor 2 is arranged at the joint of the cradle head 3 and the guidance instrument control cabin 1 and is used for measuring the rotating azimuth angle between the cradle head 3 and the guidance instrument control cabin 1. The cradle head 3 is provided with a high-low angle position sensor 4, a signal transmitter 5, a satellite receiver I6, a laser range finder 7 and an observing and aiming device 8. The signal transmitter 5 and the satellite receiver I6 are fixedly connected to the cradle head 3, and the laser range finder 7 and the sighting device 8 can rotate in a single degree of freedom relative to the cradle head 3.
The laser range finder 7 is fixedly connected with the triaxial angular rate gyroscope and the sighting device 8. The sight direction of the sight device 8 is parallel to the light outlet direction of the laser range finder 7 so as to ensure that the sight direction of a shooter is consistent with the laser emission direction. The laser range finder 7 is arranged on the cradle head 3 and can rotate in a single degree of freedom in a high-low angle direction relative to the cradle head 3. The rotating shaft of the laser range finder 7 connected with the cradle head 3 is provided with a high-low angle position sensor 4 for measuring the rotating angle of the laser range finder 7 relative to the cradle head 3. The laser rangefinder 7 may emit a laser beam toward the target and calculate the distance of the guidance instrument to the target based on the length of time that the reflected laser beam is received from the target.
The satellite receiver I6 can receive ephemeris information sent by the navigation satellite and calculate the coordinate position of the guidance instrument. The signal transmitter 5 can transmit the motion information (position and speed) of the target relative to the guidance instrument, the coordinate position of the guidance instrument itself and the ephemeris information received by the satellite receiver I6 on the guidance instrument to the signal receiver in the satellite guidance assembly on the bullet.
The missile-borne satellite guidance assembly is carried on a missile for attacking a target, and can be carried on a rotor or a fixed-wing aircraft besides the missile. The missile-borne satellite guidance assembly comprises a satellite receiver II, a signal receiver and a flight control computer. The satellite receiver II can receive the ephemeris information sent by the navigation satellite and transmit the ephemeris information to the flight control computer. The signal receiver can receive the motion information (position and speed) of the target relative to the guidance instrument, which is sent by the signal transmitter 5 on the guidance instrument, the coordinate position of the guidance instrument itself and the ephemeris information received by the satellite receiver I6. And the flight control computer calculates the relative distance between the missile and the target in real time through a missile relative motion calculation algorithm according to the information received by the satellite receiver II and the signal receiver, so as to form a flight control instruction and control the missile to fly to the target.
The bullet relative motion calculation algorithm is as follows:
the coordinates of the guidance instrument in the geodetic coordinate system are set asThe missile has the following coordinates in the geodetic coordinate systemThe j navigation satellite coordinates are +.>Let the relative coordinates between missile and guidance instrument under the geodetic coordinate system be (x) E y E z E ) And->
By usingAnd->The pseudo range of the j navigation satellite received by the satellite receiver I6 on the guidance instrument and the satellite receiver II on the missile satellite guidance assembly are respectively shown, and the pseudo range difference is +.>
The actual distance from the guidance instrument to the jth navigation satellite is set asThe expression is as follows:
let c be the light velocity, dT t Receiver clock difference dT for satellite receiver I6 on guidance instrument m For the receiver clock error of the satellite receiver ii in the satellite guidance assembly on the missile, the difference d=c (dT m -dT t )。
Construction state variables X ＝[x E y E z E d] T Wherein (x) E y E z E ) And d is the distance difference between the guided vehicle and the guided vehicle caused by clock error.
The number of receivable navigation satellites is set as N, an N multiplied by 1 dimension vector Z is constructed, the content of the vector Z is the difference between the guidance instrument and the pseudo range of each navigation satellite guided by the missile, and the expression of the vector Z is as follows:
constructing an Nx1-dimensional vector H, wherein the expression is as follows:
the solving expression of the state variable X is:
X＝(H T H) -1 H T Z (7)
solving the solution (7) by using a least square method to obtain each parameter value in the state variable X, and further solving the coordinate (X E y E z E ). According to the coordinate of missile relative guidance instrument under the geodetic coordinate system, calculating navigationCoordinates of the missile relative to the guidance instrument under the coordinate system:
wherein alpha is g Longitude information, beta, calculated for satellite receiver II in satellite guidance assembly on bullet g Latitude information is calculated for a satellite receiver II in the satellite guidance assembly on the bullet.
And then solving the relative line-of-sight angle of the missile and the target:
after the relative line of sight angle of the missile and the target is calculated, the flight control computer can calculate the flight control instruction by adopting a proportional guidance law or a speed tracking guidance law.
The working process is as follows:
as shown in fig. 2, after the shooter enters the ground, the guidance instrument is initialized first: electrifying the guidance instrument, adjusting the guidance instrument to be horizontal, and ensuring that the indication of the level instrument is in a zero position; standing for a plurality of seconds to enable the angular rate gyro to finish initial alignment, and determining the initial orientation of the laser range finder 7 in a navigation coordinate system; the satellite receiver I6 starts self-checking, searches for satellites and receives ephemeris information transmitted by the navigation satellites.
The shooter searches and locks the target through the sighting device 8, and simultaneously rotates the sighting device 8 and the laser range finder 7 fixedly connected with the sighting device 8, so that the laser beam emitted by the laser range finder 7 can irradiate the target.
The laser distance measuring machine 7 sends the distance from the measured target to the guidance instrument to the guidance computer; the angular rate gyro sends the high-low angle and azimuth angle rate of the laser range finder 7 to the guidance computer; the azimuth position sensor 2 and the elevation position sensor 4 transmit the azimuth and elevation angles of the laser range finder 7 to the guidance computer, respectively.
The guidance computer calculates the position and velocity of the target relative to the guidance aid.
The shooter launches a missile carrying an onboard satellite guidance assembly. The signal transmitter 5 transmits the position and speed information of the target calculated by the guidance computer relative to the guidance instrument, the ephemeris information received by the satellite receiver I6 on the guidance instrument and the position information of the calculated guidance instrument to the satellite guidance assembly on the bullet. The signal receiver in the missile-borne satellite guidance assembly receives the information sent by the signal transmitter 5 and transmits the information to the flight control computer; the flight control computer calculates the relative distance and speed between the missile and the target by combining the ephemeris information received by the satellite receiver II and the information sent by the guidance instrument received by the signal receiver, so as to calculate the flight control instruction and control the missile to fly to the target.
In summary, the above embodiments are only preferred embodiments of the present invention, and are not intended to limit the scope of the present invention. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (9)
1. The satellite guidance system with the guidance instrument is characterized by comprising the guidance instrument and an on-bullet satellite guidance assembly;
the guidance instrument comprises a control cabin, a satellite receiver I, a laser range finder, an angular rate gyroscope, a cradle head, a signal transmitter and an observation device; the satellite receiver I is used for receiving ephemeris information sent by the navigation satellite and calculating the position information and the movement speed of the guidance instrument according to the ephemeris information; the system comprises a laser range finder, an angular rate gyroscope and an observing and sighting device, wherein the laser range finder is used for measuring the distance between a target and the guidance instrument; the control cabin is used for providing a navigation coordinate system reference and a horizontal reference, and the control cabin calculates the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser ranging machine relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument; the signal transmitter transmits the position information, ephemeris information, relative position and movement speed of the target relative to the guidance instrument to the satellite guidance assembly on the bullet;
the on-board satellite guidance assembly comprises a satellite receiver II, a signal receiver and a flight control computer; the satellite receiver II is used for receiving ephemeris information sent by the navigation satellite; the signal receiver is used for receiving the position information, the ephemeris information, the relative position and the movement speed information of the target relative to the guidance instrument of the guidance instrument sent by the signal transmitter on the guidance instrument; the flight control computer is used for calculating the position and the speed of the guided missile relative to the guidance instrument according to the ephemeris information received by the satellite receiver II, the position information and the ephemeris information of the guidance instrument received by the signal receiver, and calculating the position and the speed of the guided missile relative to the target in real time by combining the relative position and the speed information of the target relative to the guidance instrument received by the signal receiver, and calculating the flight control instruction according to the position and the speed.
2. The guidance system of claim 1, further comprising an angular position sensor fixed to the rotational axis of the two-degree-of-freedom cradle head for measuring the azimuth and elevation angle of the target relative to the guidance system.
3. The satellite guidance system with guidance according to claim 1, wherein the angular rate gyro is replaced with an angular position sensor fixed at the rotation axis of the two-degree-of-freedom cradle head for measuring the azimuth and elevation angle of the target relative to the guidance; the angular velocity of the target relative to the guidance instrument is calculated from the azimuth angle and the elevation angle.
4. The guidance system of claim 1, wherein the control pod comprises an electronic compass, a level, and a guidance computer;
the electronic compass is used for providing a navigation coordinate system reference; the level gauge is used for providing a level reference; the guidance computer is used for calculating the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser range finder relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument.
5. A satellite guidance system incorporating a guidance instrument according to any of claims 1-4, wherein the line of sight of the sighting device is parallel to the direction of the light outlet of the laser rangefinder.
6. The guidance system of claim 5, wherein the flight control computer calculates the flight control instructions using a proportional guidance law or a velocity tracking guidance law.
7. A satellite guidance method comprising a guidance instrument, characterized in that a satellite guidance system according to claim 1 is used, the guidance method comprising the steps of:
leveling a guidance instrument, enabling an angular rate gyro to finish initial alignment, and determining the initial orientation of a laser range finder in a navigation coordinate system; the satellite receiver I starts self-checking, searches for satellites and receives ephemeris information sent by the navigation satellites;
step two, a shooter searches and locks a target through the sighting device, and simultaneously rotates the sighting device and a laser range finder fixedly connected with the sighting device, so that a laser beam emitted by the laser range finder can irradiate the target;
step three, the laser range finder sends the measured distance from the target to the guidance instrument to the control cabin; the angular rate gyroscope sends the high-low angle and azimuth angle rate of the laser range finder to the control cabin, calculates the high-low angle and azimuth angle of the laser range finder according to the high-low angle and azimuth angle rate and sends the high-low angle and azimuth angle to the control cabin;
step four, the control cabin calculates the relative position and the movement speed of the target relative to the guidance instrument according to the distance of the target measured by the laser range finder relative to the guidance instrument and the angular velocity of the target measured by the angular velocity gyro relative to the guidance instrument;
step five, launching a missile loaded with an onboard satellite guidance assembly;
step six, the signal transmitter transmits the relative position and movement speed information of the target calculated by the control cabin relative to the guidance instrument, ephemeris information received by the satellite receiver I on the guidance instrument and position information of the guidance instrument itself calculated by the control cabin to the satellite guidance assembly on the bullet;
step seven, a signal receiver in the missile-borne satellite guidance assembly receives the information sent by the signal transmitter and transmits the information to the flight control computer; the flight control computer calculates the relative distance and speed between the missile and the target by combining the ephemeris information received by the satellite receiver II and the information received by the signal receiver, so as to calculate the flight control instruction and control the missile to fly to the target.
8. The method of claim 7, further comprising an angular position sensor, wherein in step three, the angular position sensor transmits the measured elevation and azimuth of the laser range finder to the control pod.
9. The method for guiding a satellite having a guidance instrument according to claim 7, wherein the flight control computer calculates the flight control command using a proportional guidance law or a velocity tracking guidance law in the seventh step.
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