CN-121994219-A - Carrier rocket autonomous combined navigation equipment and method based on star sensor

CN121994219ACN 121994219 ACN121994219 ACN 121994219ACN-121994219-A

Abstract

The scheme provides a carrier rocket autonomous combined navigation device and a carrier rocket autonomous combined navigation method based on a star sensor. The device is composed of an inertial measurement unit, a star sensor and an MEMS backup inertial navigation unit. The method comprises the steps of completing space-time alignment before emission, performing auxiliary correction by taking inertial navigation as a main body in flight, performing fitting correction on a sliding section in the atmosphere by using a least square method through low-frequency observation, switching to a high-frequency observation mode at the end of flight, performing real-time optimal estimation and correction on navigation parameters and errors of inertial devices through Kalman filtering by combining GNSS, and ensuring high-precision in-orbit. When the primary inertial navigation precision is abnormal, the system is automatically switched to the backup unit. According to the scheme, the reliability, the adaptability and the overall accuracy of the system are remarkably improved through the hardware redundancy design of the main/standby dual-inertial navigation system, the combination of front and rear intelligent segmentation correction and inertial measurement, and the combined navigation fusion algorithm of the star sensor and the GNSS.

Inventors

ZHOU XIN
WANG ZHIJUN
LI XUN
ZOU YANBING
LI XINRAN
ZENG ZIYI
LU HANG
HE FENG
ZHANG CHUNLAN

Assignees

航天科工火箭技术有限公司

Dates

Publication Date: 20260508
Application Date: 20251223

Claims (10)

1. The carrier rocket autonomous integrated navigation device based on the star sensor is characterized by comprising an inertial measurement assembly, the star sensor and a backup inertial navigation unit, wherein the inertial measurement assembly comprises a gyroscope and an accelerometer, navigation in a carrier rocket flight stage takes the inertial measurement assembly as a main component, the star sensor performs auxiliary correction, and after the inertial measurement assembly has abnormal precision reduction, the inertial measurement assembly is switched to the backup inertial navigation unit to continue navigation tasks.
2. The satellite sensor-based launch vehicle autonomous integrated navigation device of claim 1, wherein the backup inertial navigation unit is a MEMS component.
3. The satellite sensor-based carrier rocket autonomous integrated navigation method is applied to the satellite sensor-based carrier rocket autonomous integrated navigation device according to any one of claims 1 or 2, and is characterized by comprising the following steps: The preparation stage before launch of the carrier rocket, the time stamp of the inertial measurement unit and the star sensor is calibrated on the ground, so that time synchronization is ensured; initial azimuth acquisition, adopting static base acquisition alignment or moving base alignment acquisition: the static base alignment comprises the steps of observing the star azimuth through the star sensor after the carrier rocket is erected, measuring the gravity vector by combining with the inertia measurement combination, and calculating an initial gesture matrix; In the rocket flight stage, inertial navigation is taken as a leading, and the inertial measurement combines high-frequency output angular velocity and acceleration to perform navigation calculation in real time; The early stage of the rocket flight stage has complex environment in the atmosphere, a star sensor performs low-frequency star observation, and the error generated during inertial navigation is fitted through a least square method, so that correction is performed in a sliding section; the method comprises the steps that the interference factors at the end of a rocket flight stage are few, the star sensor is switched to a high-frequency correction mode, the absolute attitude angle is obtained at high frequency to conduct multi-source data fusion, the error of inertial navigation is optimally estimated based on Kalman filtering, the zero offset error of the gyroscope and the accelerometer is added in a generated state equation except for the position, the speed and the attitude deviation, the measurement equation is the deviation of the attitude angle calculated by inertial navigation relative to the absolute attitude angle measured by the star sensor, and the inertial navigation is accurately corrected before the inertial navigation is in orbit, so that high-precision orbit entering is achieved.
4. The satellite sensor-based carrier rocket autonomous integrated navigation method according to claim 3, wherein in a flight phase, after the satellite sensor detects that the inertial measurement unit has abnormal precision drop, the satellite sensor is switched to the backup inertial navigation unit to realize redundant design of a system.
5. The autonomous integrated navigation method of carrier rocket based on star sensor according to claim 3, wherein the error generated during inertial navigation is corrected in the sliding section by least square fitting, specifically, the star sensor can observe a plurality of stars at the same time, and the information of the plurality of stars is utilized by least square fitting, and the calculation steps are as follows: the star sensor obtains the direction of unit vectors of a plurality of stars in an arrow system; obtaining unit vector directions of a plurality of stars in an inertial system according to ephemeris information; Calculating the directions of the unit vectors corresponding to three axes of the arrow system in the inertial system by using a least square method based on the directions of the unit vectors in the arrow system and the directions of the unit vectors of the multiple stars in the inertial system; according to the directions of the unit vectors corresponding to the three axes of the rocket system in the inertial system, calculating the posture of the carrier rocket in the inertial system, and directly replacing the posture result of the current inertial navigation to realize inertial navigation correction.
6. The satellite sensor-based launch vehicle autonomous integrated navigation method of claim 3, wherein the specific data sources of the multi-source data fusion are the inertial measurement unit, the satellite sensor and the GNSS.
7. The satellite sensor-based carrier rocket autonomous integrated navigation method according to claim 6, wherein the Kalman filtering is used for optimally estimating inertial navigation errors, zero offset errors of the gyroscope and the accelerometer are added in a generated state equation except position, speed and attitude deviations, and a measurement equation is the deviation of an attitude angle calculated by inertial navigation relative to an absolute attitude angle measured by the satellite sensor, specifically, the inertial measurement combination, the satellite sensor and the GNSS are used for Kalman filtering, system state quantity is firstly set, a state matrix of a system is acquired, the state equation of the system is calculated, a measurement vector and a measurement matrix are acquired, and the measurement equation is calculated.
8. The satellite sensor-based carrier rocket autonomous integrated navigation method according to claim 7, wherein the integrated navigation algorithm flow is as follows: s1, setting a state initial value, predicting a mean square error initial value, and setting a covariance matrix; s2, inertial navigation is carried out, a conversion matrix from an emission inertial system to an arrow system is calculated, and the position and the speed of the inertial navigation are obtained; s3, updating the state matrix and the state transition matrix of the system; S4, judging whether new navigation information of the GNSS or the star sensor is received, if yes, entering a step S5, otherwise, jumping to a step S6; s5, the navigation information of the GNSS or the star sensor is received, and the measurement vector is constructed according to the navigation information; converting the position and the speed obtained by the GNSS from an earth fixedly connected coordinate system to a transmitting inertial coordinate system to obtain the position and the speed of the GNSS under the transmitting inertial system, and obtaining the position error and the speed error of inertial navigation by differentiating the position and the speed of inertial navigation with the position and the speed of the GNSS; Then, the difference between the inertial navigation gesture and the gesture obtained by the star sensor is obtained, namely, the misalignment angle is obtained; The measurement vector is composed of the difference between inertial navigation and the position and speed of the GNSS under the emission inertial coordinate system and the misalignment angle; Updating a filter according to the measurement vector, and predicting to obtain the state quantity at the current moment; S6, not receiving new navigation information of the GNSS or the star sensor, wherein the measurement vector is not updated, and the prediction update is performed based on the state quantity at the last moment of the filter; S7, according to the state quantity obtained by updating the filter, inertial navigation error correction is carried out, the inertial navigation position and speed error are corrected firstly, then the misalignment angle is utilized, the inertial navigation posture is corrected, and finally the state quantity of the filter is set to zero; s8, returning to the step S2.
9. A computer readable storage medium, on which a computer program is stored, characterized in that the program, when being executed by a processor, implements the steps of the method according to any of claims 3-8.
10. A computer device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, characterized in that the processor implements the steps of the method according to any of claims 3-8 when the program is executed by the processor.

Description

Carrier rocket autonomous combined navigation equipment and method based on star sensor Technical Field The invention relates to the technical field of carrier rocket navigation, in particular to a carrier rocket autonomous combined navigation device and method based on a star sensor. Background The carrier rocket navigation system generally mainly depends on inertial measurement combination navigation, but measurement errors of the inertial measurement combination are accumulated along with working time and are difficult to independently work for a long time, the satellite navigation system is easily affected by interference such as atmosphere, electromagnetic interference and man-made, the astronomical navigation data updating frequency is low, and continuous navigation is difficult to carry out. The single navigation means is difficult to meet the requirement of high-precision navigation, and in a special application scene, in order to improve the reliability of rocket flight, the navigation systems with high autonomy and strong anti-interference capability should be preferentially selected for combination, so that the requirement on the precision of the sub-navigation system is reduced while the navigation precision is ensured, and the cost is reduced. Disclosure of Invention Based on the above, the invention aims to provide the carrier rocket navigation equipment and the carrier rocket navigation method with high autonomy and strong anti-interference capability, which can reduce the requirement on the accuracy of a sub-navigation system while guaranteeing the navigation accuracy. In order to achieve the above purpose, the invention adopts the following technical scheme: According to the first aspect of the invention, a carrier rocket autonomous combined navigation device based on a star sensor is provided, and comprises an inertial measurement unit, a star sensor and a backup inertial navigation unit, wherein the inertial measurement unit comprises a gyroscope and an accelerometer, navigation in a carrier rocket flight stage takes the inertial measurement unit as a main component, the star sensor performs auxiliary correction, and after the inertial measurement unit has abnormal precision reduction, the inertial measurement unit is switched to the backup inertial navigation unit to continue navigation tasks. As a preferable scheme of the carrier rocket autonomous integrated navigation device based on the star sensor, the backup inertial navigation unit is an MEMS component. The invention provides a satellite sensor-based carrier rocket autonomous integrated navigation method, which is applied to any satellite sensor-based carrier rocket autonomous integrated navigation device, and specifically comprises the following steps: The preparation stage before launch of the carrier rocket, the time stamp of the inertial measurement unit and the star sensor is calibrated on the ground, so that time synchronization is ensured; initial azimuth acquisition, adopting static base acquisition alignment or moving base alignment acquisition: the static base alignment comprises the steps of observing the star azimuth through the star sensor after the carrier rocket is erected, measuring the gravity vector by combining with the inertia measurement combination, and calculating an initial gesture matrix; In the rocket flight stage, inertial navigation is taken as a leading, and the inertial measurement combines high-frequency output angular velocity and acceleration to perform navigation calculation in real time; The early stage of the rocket flight stage has complex environment in the atmosphere, a star sensor performs low-frequency star observation, and the error generated during inertial navigation is fitted through a least square method, so that correction is performed in a sliding section; the method comprises the steps that the interference factors at the end of a rocket flight stage are few, the star sensor is switched to a high-frequency correction mode, the absolute attitude angle is obtained at high frequency to conduct multi-source data fusion, the error of inertial navigation is optimally estimated based on Kalman filtering, the zero offset error of the gyroscope and the accelerometer is added in a generated state equation except for the position, the speed and the attitude deviation, the measurement equation is the deviation of the attitude angle calculated by inertial navigation relative to the absolute attitude angle measured by the star sensor, and the inertial navigation is accurately corrected before the inertial navigation is in orbit, so that high-precision orbit entering is achieved. As a preferable scheme of the carrier rocket autonomous integrated navigation method based on the star sensor, in the flight stage, after the star sensor detects that the inertial measurement unit has abnormal precision reduction, the star sensor is switched to the backup inertial navigation unit, so that the redundant des