CN114440865A - Integrated navigation method based on MEMS inertial measurement system - Google Patents

Integrated navigation method based on MEMS inertial measurement system Download PDF

Info

Publication number
CN114440865A
CN114440865A CN202111499808.5A CN202111499808A CN114440865A CN 114440865 A CN114440865 A CN 114440865A CN 202111499808 A CN202111499808 A CN 202111499808A CN 114440865 A CN114440865 A CN 114440865A
Authority
CN
China
Prior art keywords
navigation
navigation system
satellite
error
inertial
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
CN202111499808.5A
Other languages
Chinese (zh)
Inventor
易文俊
陈明玥
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Nanjing University of Science and Technology
Original Assignee
Nanjing University of Science and Technology
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Nanjing University of Science and Technology filed Critical Nanjing University of Science and Technology
Priority to CN202111499808.5A priority Critical patent/CN114440865A/en
Publication of CN114440865A publication Critical patent/CN114440865A/en
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/04—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by terrestrial means
    • G01C21/08—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by terrestrial means involving use of the magnetic field of the earth
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/18—Stabilised platforms, e.g. by gyroscope
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C23/00—Combined instruments indicating more than one navigational value, e.g. for aircraft; Combined measuring devices for measuring two or more variables of movement, e.g. distance, speed or acceleration
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/13—Receivers
    • G01S19/14—Receivers specially adapted for specific applications
    • G01S19/18—Military applications
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/40—Correcting position, velocity or attitude
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/42—Determining position
    • G01S19/48—Determining position by combining or switching between position solutions derived from the satellite radio beacon positioning system and position solutions derived from a further system
    • G01S19/49—Determining position by combining or switching between position solutions derived from the satellite radio beacon positioning system and position solutions derived from a further system whereby the further system is an inertial position system, e.g. loosely-coupled
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/53—Determining attitude

Landscapes

  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Computer Networks & Wireless Communication (AREA)
  • Automation & Control Theory (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • General Life Sciences & Earth Sciences (AREA)
  • Geology (AREA)
  • Environmental & Geological Engineering (AREA)
  • Manufacturing & Machinery (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Navigation (AREA)

Abstract

The invention relates to a combined navigation method based on an MEMS inertial measurement system. The method comprises the following steps: (1): collecting data; (2) aligning and correcting a platform error angle: aligning the guided missile by adopting a chip of the integrated navigation system, and correcting a platform error angle according to satellite information; (3) data processing and integrated navigation: using the position error, the speed error and the attitude error of the MEMS inertial measurement system as state variables of the integrated navigation system; taking a position difference value, a speed difference value and an attitude difference value output by the MEMS inertial measurement system and output by the satellite navigation system and the magnetic sensor as the measurement values of the system; the coarse alignment and the fine alignment of the MEMS inertial measurement system data and the correction of the INS navigation result are realized through the magnetic sensor and the satellite, and the navigation result is finally output. The invention provides a scheme of a combined navigation system formed by an MEMS inertial magnetizing sensor and satellite navigation, and the navigation precision is improved by utilizing the autonomy of MEMS and the reliability of magnetic sensors and satellite navigation.

Description

Integrated navigation method based on MEMS inertial measurement system
Technical Field
The invention belongs to the field of guidance control, and particularly relates to a combined navigation method based on an MEMS (micro-electromechanical system) inertia measurement system.
Background
With the rapid development of science and technology, modern war forms tend to be intelligent and informationized. However, conventional artillery still achieves tactical goals in a "mass-winning" manner, and cannot meet the requirements for weapon precision and efficiency under modern war conditions, so that accurate guidance of the weapon arises. The hitting precision of the weapon can be greatly improved through the guidance transformation of the conventional ammunition, so that the weapon can effectively destroy targets, and therefore, the accurate guided weapon adopting the flight control technology to adjust the trajectory and the aerodynamic characteristics becomes the research focus of various countries. During the launching and flying processes of the guided missile and arrow, the state and environmental characteristic quantity of the guided missile and arrow are continuously collected, logically analyzed and resolved, and decided, and even information interaction is carried out with a weapon system, so that the flight trajectory is adjusted, and the purposes of range extension, accurate striking or information confrontation and the like are achieved. Compared with other weapons, the guided projectile has low maintenance cost and strong maneuverability, and can be launched by using the traditional artillery platform. The guidance modes of the guided projectile are mainly divided into three modes of laser guidance, millimeter guidance and combined guidance, because the launching condition of the guided projectile is very severe, the requirement on an inertial sensor is greatly limited, and therefore the inertial sensor based on a micro-electro-mechanical system is generally adopted.
The MEMS is a micro-electromechanical system, and the interior thereof is mainly composed of a micro-mechanical structure, a micro-signal sensor and a micro-actuator, and in addition, circuit interfaces, communication, power supply and Digital Signal Processing (DSP) circuits can also be designed in the system. The MEMS sensor is a novel sensor manufactured by adopting micro-electronics and micro-machining technologies, and compared with the traditional sensor, the MEMS sensor has the characteristics of small volume, light weight, low cost, low power consumption, high overload resistance, high reliability, starting from batch production, easiness in integration and intelligentization realization. Therefore, the MEMS inertial sensing technology has wide application in the fields of aerospace, military and the like.
MEMS inertial sensors mainly include accelerometers (or acceleration sensors) and angular rate sensors (gyros) and their single, double, and three-axis combination IMU (inertial measurement unit), AHRS (attitude reference system including magnetic sensors). The working principle of the method mainly depends on Newton's law, under the condition of giving initial conditions of motion, angular motion parameters and linear motion parameters of the carrier relative to an inertial space are measured, and after the angular motion parameters and the linear motion parameters are solved by a navigation computer, the speed, the position, the attitude and the heading of the carrier are solved. The MEMS inertial sensor is a key component of the guided rocket, measures information such as spatial position, speed, acceleration, attitude angle, attitude angular velocity and the like of the rocket in real time in the flying process, and provides basic information for a control system.
In the last decade, the research and development of the MEMS inertial sensor in China is very hot, and particularly after 2010, the MEMS sensor and the pop-up application thereof are researched and developed by the top research institution in China. The main research methods are as follows: the study of the MEMS-INS/GPS double-star combined positioning method comprises the steps of Wangpeng, modern navigation, volume 6, phase 6, pages 487 to 491, and 2015. "inertial and starlight combined navigation method based on online estimation and correction of gyro error under inertial system", Zhao Hui, war institute of war, Vol.37, No. 12, pp.2259-2267, 2016. "multi-constellation GNSS/INS tightly coupled method", volume 23, No. 1, pages 38-42, 2015, Tao, China inertial technology report. "remote rocket projectile combination navigation method based on ultra-tight coupling GPS/INS", Zhao HUNTO, test science and instrument, Vol.6, No. 2, pages 153-160, 2015.
In the existing navigation mode, the MEMS inertial navigation is the most common navigation mode, but the inertial navigation also has the following disadvantages in the missile application: the small-volume rocket has poor overload resistance, is easy to disperse for a long time, has complex algorithm, and is not suitable for the rocket with high overload, limited volume and high rolling speed. Therefore, research on methods and technologies with high overload resistance, high precision, low cost, micro power consumption and integration capability has become the key point of domestic and foreign research.
Disclosure of Invention
The invention aims to provide a combined navigation method based on an MEMS inertial measurement system.
The technical solution for realizing the purpose of the invention is as follows: an integrated navigation method based on an MEMS inertial measurement system, which adopts the MEMS inertial measurement system as a basis, and a magnetic sensor and a satellite are used for assisting the integrated navigation system, comprises the following steps:
step (1): data acquisition: reading the output data of a gyroscope and an accelerometer of the MEMS inertial measurement system and the output data of a magnetic sensor, and receiving satellite data;
step (2): aligning and correcting a platform error angle: aligning the missile by adopting a chip of the integrated navigation system, and correcting a platform error angle according to satellite information;
and (3): data processing and integrated navigation: using the position error, the speed error and the attitude error of the MEMS inertial measurement system as state variables of the integrated navigation system; taking a position difference value, a speed difference value and an attitude difference value output by the MEMS inertial measurement system and output by the satellite navigation system and the magnetic sensor as the measurement values of the system; the coarse alignment and the fine alignment of the MEMS inertial measurement system data and the correction of the INS navigation result are realized through the magnetic sensor and the satellite, and the navigation result is finally output.
Further, the algorithm adopted in the step (3) "data processing, combined navigation" is specifically as follows:
on the basis of an MEMS inertial measurement system state equation, data of a magnetic sensor and a satellite are considered, and a state equation and a measurement equation of the combined navigation system are obtained, wherein the state equation is as follows:
Figure BDA0003401067870000031
wherein the state variable X (t) is represented by the following formula:
Figure BDA0003401067870000032
where δ L, δ λ, δ h are the position errors of the inertial navigation system, δ VE，δVN，δVUIs the velocity error of the inertial navigation system,
Figure BDA0003401067870000033
is the attitude error, ε, of an inertial navigation systemx，εy，εzIs the drift error of the MEMS sensor;
f is a transfer matrix, G(t)System excitation input matrix, W (t) excitation noise
The measurement equation is as follows:
Zk＝HkXk+Vk
wherein the state variable ZkComprises the following steps:
Figure BDA0003401067870000034
Figure BDA0003401067870000035
wherein, Δ L, Δ λ, Δ h are position differences between the inertial navigation system and the satellite navigation output; Δ Ve，ΔVn，ΔVuIs the speed difference between the inertial navigation system and the satellite navigation output;
Figure BDA0003401067870000036
is the misalignment angle of the inertial navigation system with respect to the geomagnetic sensor output; d is a misalignment angle; n is a radical ofE，NN，NUA deviation value of the position information received for satellite navigation; mE，MN，MUThe satellite navigation receives a speed deviation value, R is the earth radius, and L represents longitude.
Further, the transition matrix F in the state equation is:
Figure BDA0003401067870000037
in the formula:
Figure BDA0003401067870000038
Figure BDA0003401067870000041
Figure BDA0003401067870000042
Figure BDA0003401067870000043
Figure BDA0003401067870000044
Figure BDA0003401067870000045
Figure BDA0003401067870000051
wherein, L, lambda and h respectively represent longitude, latitude and height; ve、Vn、VuRespectively measuring the speed of the projectile in the northeast direction by the inertial navigation system; f. ofe、fu、fnOutputting the accelerometer of the inertial navigation system in the northeast direction; tau isGDx、τGDy、τGDzIs the correlation time; omegaieIs the rotational angular velocity of the earth; r is the radius of the earth.
Further, the excitation noise w (t) is represented by the following formula:
W(t)＝[σx σy σz εx εy εz εGDWx εGDWy εGDWz]
wherein σx、σy、σzIs the accelerometer noise error; epsilonx、εy、εzIs the gyroscope drift error; epsilonGDWx、εGDWy、εGDWzIs the satellite navigation system drift error.
Further, the system stimulates the input matrix G(t)Comprises the following steps:
Figure BDA0003401067870000052
wherein
Figure BDA0003401067870000053
Is the attitude transformation matrix, and I is the identity matrix.
Further, on the basis of the state equation of the MEMS inertial measurement system, the specific process of obtaining the measurement equation of the integrated navigation system in consideration of the data of the magnetic sensor and the satellite is as follows:
in order to add the attitude information obtained by the geomagnetic sensor into the integrated navigation, the attitude angle measured by the inertial navigation system and the attitude angle measured by the geomagnetic sensor are differentiated:
Figure BDA0003401067870000054
the difference is then converted into the form of the misalignment angle, which can be represented by a matrix transformation:
Figure BDA0003401067870000055
wherein, thetas、γG、ψsMeasuring an attitude angle by a geomagnetic sensor; thetaI、γI、ψIMeasuring an attitude angle by an inertial navigation system;
Figure BDA0003401067870000056
the difference value of the attitude angle is obtained by measuring the inertial navigation system relative to the geomagnetic sensor;
the measurement equation expression composed of the inertial navigation system, the satellite and the geomagnetic acquired position, speed and attitude difference value is as follows:
Figure BDA0003401067870000061
wherein N isE、NN、NUThe deviation value of the position information measured by the satellite navigation system and an actual value is obtained; mE、MN、MUIs the velocity bias value of the satellite navigation system.
Compared with the prior art, the invention has the remarkable advantages that:
the method of the invention provides a scheme of an integrated navigation system composed of an MEMS inertial magnetizing sensor and satellite navigation; the method mainly researches a system combination mode of inertia/geomagnetism/satellite combination, takes an MEMS inertia measurement system as a basis, uses a position error, a speed error and an attitude error of an inertial navigation system as state variables of the system, uses a geomagnetic sensor to measure an attitude angle and satellite navigation to obtain position and speed information to optimize the system, and uses a Kalman filtering algorithm to design an optimal filter, thereby realizing the improvement of MEMS inertial navigation precision.
The implementation of the scheme can effectively solve the problem that the error is dispersed along with time and the navigation effect is lost when an inertial measurement system is used independently, the system precision is improved, and the method provides guarantee for realizing accurate target strike.
Drawings
FIG. 1 is a schematic diagram of a data acquisition module according to the present invention.
FIG. 2 is a schematic diagram of the INS navigation algorithm of the present invention.
FIG. 3 is a block diagram of a data fusion module according to the present invention.
FIG. 4 is a schematic flow chart of the guidance control system based on the MEMS inertial measurement system.
Detailed Description
The present invention is described in further detail below with reference to the attached drawing figures.
The invention aims to provide an integrated navigation control system based on an MEMS inertial measurement system. The method provides a scheme of a combined navigation system formed by an MEMS inertial magnetizing sensor and satellite navigation, and improves the navigation precision by utilizing the autonomy of the MEMS and the reliability of the magnetic sensor and the satellite navigation. The method mainly researches a system combination mode of inertia/geomagnetism/satellite combination, takes an MEMS inertia measurement system as a basis, uses a position error, a speed error and an attitude error of an inertial navigation system as state variables of the system, uses a geomagnetic sensor to measure an attitude angle and satellite navigation to obtain position and speed information to optimize the system, and uses a Kalman filtering algorithm to design an optimal filter, thereby realizing the improvement of MEMS inertial navigation precision. The system flow is shown in fig. 4. The implementation of this scheme can effectual solution independent use inertial measurement system can lead to the error to disperse along with time and thereby lead to losing the problem of navigation effect, improves system's precision, provides the guarantee for realizing hitting the accuracy of target. The successful development of the invention has important significance for developing high-precision guided weapons. The specific technical scheme is as follows:
(1) data acquisition
The data acquisition module mainly has the functions that the central processing unit reads output data of the gyroscope, the accelerometer and the magnetic sensor; reading magnetoresistive sensor output data; data of a GNSS is received. The specific flow is shown in figure 1.
(2) Navigation algorithm design
The INS navigation algorithm mainly realizes the initialization of the system and the navigation updating and resolving of the MEMS-IMU through the acquisition and processing of data of the gyroscope, the accelerometer and the magnetoresistive sensor, and outputs the attitude, the speed and the position information of the carrier in real time. The INS navigation algorithm flow is shown in fig. 2. The navigation algorithm firstly needs to establish a system model, and the system model mainly comprises a system state equation and a measurement equation. Because the invention is a combined navigation system formed by an MEMS inertial magnetizing sensor and satellite navigation, when a navigation algorithm is designed, the position error, the speed error and the attitude error of the inertial navigation system are used as state variables of the system, and the position difference, the speed difference and the attitude difference output by the inertial navigation system and the satellite navigation system and the magnetic sensor are used as the measurement of the system.
The state equation of the system is
Figure BDA0003401067870000071
Wherein the state variable X (t) is represented by the following formula:
Figure BDA0003401067870000072
where δ L, δ λ, δ h are the position errors of the inertial navigation system, δ VE，δVN，δVUIs the velocity error of the inertial navigation system,
Figure BDA0003401067870000073
is the attitude error, ε, of an inertial navigation systemx，εy，εzIs the drift error of the MEMS sensor.
From the velocity position error propagation equation of the inertial navigation system, its transfer matrix can be listed:
Figure BDA0003401067870000074
in the formula:
Figure BDA0003401067870000075
Figure BDA0003401067870000081
Figure BDA0003401067870000082
Figure BDA0003401067870000083
Figure BDA0003401067870000084
Figure BDA0003401067870000085
Figure BDA0003401067870000091
wherein, L, lambda and h respectively represent longitude, latitude and height; ve、Vn、VuRespectively measuring the speed of the projectile in the northeast direction by the inertial navigation system; f. ofe、fu、fnOutputting the accelerometer of the inertial navigation system in the northeast direction; tau isGDx、τGDy、τGDzIs the correlation time; omegaieIs the rotational angular velocity of the earth; r is the radius of the earth.
Excitation noise w (t) is represented by:
W(t)＝[σxσyσzεxεyεzεGDWxεGDWyεGDWz]
wherein sigmax、σy、σzIs the accelerometer noise error; epsilonx、εy、εzIs the gyroscope drift error; epsilonGDWx、εGDWy、εGDWzIs the satellite navigation system drift error.
System excitation input matrix G(t)Comprises the following steps:
Figure BDA0003401067870000092
wherein
Figure BDA0003401067870000093
Is a postureAnd converting the matrix.
The measurement equation of the combined system is Zk＝HkXk+VkWherein the state variables are:
Figure BDA0003401067870000094
Figure BDA0003401067870000095
wherein, Δ L, Δ λ, Δ h are position differences between the inertial navigation system and the satellite navigation output; Δ Ve，ΔVn，ΔVuIs the speed difference between the inertial navigation system and the satellite navigation output;
Figure BDA0003401067870000096
is the misalignment angle of the inertial navigation system with respect to the geomagnetic sensor output; d is a misalignment angle; n is a radical ofE，NN，NUA deviation value of the position information received for satellite navigation; mE，MN，MUAnd the speed deviation value received by the satellite navigation.
In order to add the attitude information obtained by the geomagnetic sensor into the integrated navigation, the attitude angle measured by the inertial navigation system is different from the attitude angle measured by the geomagnetic sensor, and the following steps are obtained:
Figure BDA0003401067870000097
the difference is then converted into the form of the misalignment angle, which can be represented by a matrix transformation:
Figure BDA0003401067870000101
wherein, thetas、γG、ψsMeasuring an attitude angle by a geomagnetic sensor; thetaI、γI、ψIIs inertial navigationThe system measures an attitude angle;
Figure BDA0003401067870000102
the difference of the attitude angle measured by the inertial navigation system relative to the geomagnetic sensor.
The inertial navigation system and the satellite, geomagnetic obtained position, speed and attitude difference form an expression as follows:
Figure BDA0003401067870000103
wherein N isE、NN、NUThe deviation value of the position information measured by the satellite navigation system and an actual value is obtained; mE、MN、MUIs the velocity bias value of the satellite navigation system.
(3) Data fusion
The data fusion module performs data fusion on the carrier attitude, speed, position and other information obtained by the calculation of the INS and the position and speed information output by the GPS, so as to realize the functions of initial coarse alignment and fine alignment of the INS and correction of the INS navigation result, and finally outputs the navigation result, as shown in fig. 3.

Claims (6)

1. An integrated navigation method based on an MEMS inertial measurement system is characterized in that the integrated navigation system which takes the MEMS inertial measurement system as a basis and is assisted by a magnetic sensor and a satellite comprises the following steps:
step (1): data acquisition: reading the output data of a gyroscope and an accelerometer of the MEMS inertial measurement system and the output data of a magnetic sensor, and receiving satellite data;
step (2): aligning and correcting a platform error angle: aligning the missile by adopting a chip of the integrated navigation system, and correcting a platform error angle according to satellite information;
and (3): data processing and integrated navigation: using the position error, the speed error and the attitude error of the MEMS inertial measurement system as state variables of the integrated navigation system; taking a position difference value, a speed difference value and an attitude difference value output by the MEMS inertial measurement system and output by the satellite navigation system and the magnetic sensor as the measurement values of the system; the coarse alignment and the fine alignment of the MEMS inertial measurement system data and the correction of the INS navigation result are realized through the magnetic sensor and the satellite, and the navigation result is finally output.
2. The method according to claim 1, wherein the algorithm adopted in the step (3) "data processing, combined navigation" is specifically as follows:
on the basis of an MEMS inertial measurement system state equation, data of a magnetic sensor and a satellite are considered, and a state equation and a measurement equation of the combined navigation system are obtained, wherein the state equation is as follows:
Figure FDA0003401067860000011
wherein the state variable X (t) is represented by the following formula:
Figure FDA0003401067860000012
where δ L, δ λ, δ h are the position errors of the inertial navigation system, δ VE，δVN，δVUIs the speed error of the inertial navigation system,
Figure FDA0003401067860000013
is the attitude error, ε, of an inertial navigation systemx，εy，εzIs the drift error of the MEMS sensor;
f is a transfer matrix, G(t)System excitation input matrix, W (t) excitation noise
The measurement equation is as follows:
Zk＝HkXk+Vk
wherein the state variable ZkComprises the following steps:
Figure FDA0003401067860000014
Figure FDA0003401067860000021
wherein, Δ L, Δ λ, Δ h are position differences between the inertial navigation system and the satellite navigation output; Δ Ve，ΔVn，ΔVuIs the speed difference between the inertial navigation system and the satellite navigation output;
Figure FDA0003401067860000022
is the misalignment angle of the inertial navigation system with respect to the geomagnetic sensor output; d is a misalignment angle; n is a radical ofE，NN，NUA deviation value of the position information received for satellite navigation; mE，MN，MUThe satellite navigation receives a speed deviation value, R is the earth radius, and L represents the longitude.
3. The method of claim 2, wherein the transition matrix F in the state equation is:
Figure FDA0003401067860000023
in the formula:
Figure FDA0003401067860000024
Figure FDA0003401067860000025
Figure FDA0003401067860000031
Figure FDA0003401067860000032
Figure FDA0003401067860000033
Figure FDA0003401067860000034
Figure FDA0003401067860000041
wherein, L, lambda and h respectively represent longitude, latitude and height; ve、Vn、VuRespectively measuring the speed of the projectile body in the northeast direction by an inertial navigation system; f. ofe、fu、fnOutputting the accelerometer of the inertial navigation system in the northeast direction; tau isGDx、τGDy、τGDzIs the correlation time; omegaieIs the rotational angular velocity of the earth; r is the radius of the earth.
4. The method of claim 3, wherein the excitation noise W (t) is represented by the following equation:
W(t)＝[σx σy σz εx εy εz εGDWx εGDWy εGDWz]
wherein sigmax、σy、σzIs the accelerometer noise error; epsilonx、εy、εzIs the gyroscope drift error; epsilonGDWx、εGDWy、εGDWzIs the satellite navigation system drift error.
5. The method of claim 4Method, characterized in that the system excites an input matrix G(t)Comprises the following steps:
Figure FDA0003401067860000042
wherein
Figure FDA0003401067860000043
Is the attitude transformation matrix, and I is the identity matrix.
6. The method of claim 5, wherein the measurement equation of the integrated navigation system is obtained by considering the data of the magnetic sensor and the satellite based on the state equation of the MEMS inertial measurement system as follows:
in order to add the attitude information obtained by the geomagnetic sensor into the integrated navigation, the attitude angle measured by the inertial navigation system and the attitude angle measured by the geomagnetic sensor are differentiated:
Figure FDA0003401067860000044
the difference is then converted into the form of the misalignment angle, which can be represented by a matrix transformation:
Figure FDA0003401067860000045
wherein, thetas、γG、ψsMeasuring an attitude angle by a geomagnetic sensor; thetaI、γI、ψIMeasuring an attitude angle by an inertial navigation system;
Figure FDA0003401067860000051
the difference value of the attitude angle is obtained by measuring the inertial navigation system relative to the geomagnetic sensor;
the measurement equation expression composed of the inertial navigation system, the satellite and the geomagnetic acquired position, speed and attitude difference value is as follows:
Figure FDA0003401067860000052
wherein N isE、NN、NUThe deviation value of the position information measured by the satellite navigation system and an actual value is obtained; mE、MN、MUIs the velocity bias value of the satellite navigation system.
CN202111499808.5A 2021-12-09 2021-12-09 Integrated navigation method based on MEMS inertial measurement system Withdrawn CN114440865A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202111499808.5A CN114440865A (en) 2021-12-09 2021-12-09 Integrated navigation method based on MEMS inertial measurement system

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202111499808.5A CN114440865A (en) 2021-12-09 2021-12-09 Integrated navigation method based on MEMS inertial measurement system

Publications (1)

Publication Number Publication Date
CN114440865A true CN114440865A (en) 2022-05-06

Family

ID=81363899

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202111499808.5A Withdrawn CN114440865A (en) 2021-12-09 2021-12-09 Integrated navigation method based on MEMS inertial measurement system

Country Status (1)

Country Link
CN (1) CN114440865A (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115507849A (en) * 2022-11-22 2022-12-23 北京理工大学前沿技术研究院 Magnetic sensor correction method and system based on INS/GNSS combined navigation assistance
CN117146810A (en) * 2023-08-29 2023-12-01 广东精天科技有限公司 Combined positioning system based on satellite navigation and MEMS inertial navigation

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115507849A (en) * 2022-11-22 2022-12-23 北京理工大学前沿技术研究院 Magnetic sensor correction method and system based on INS/GNSS combined navigation assistance
CN117146810A (en) * 2023-08-29 2023-12-01 广东精天科技有限公司 Combined positioning system based on satellite navigation and MEMS inertial navigation
CN117146810B (en) * 2023-08-29 2024-02-13 广东精天科技有限公司 Combined positioning system based on satellite navigation and MEMS inertial navigation

Similar Documents

Publication Publication Date Title
CN111426318B (en) Low-cost AHRS course angle compensation method based on quaternion-extended Kalman filtering
CN107314718B (en) High speed rotation bullet Attitude estimation method based on magnetic survey rolling angular rate information
CN105606094B (en) A kind of information condition matched filtering method of estimation based on MEMS/GPS combined systems
TW468035B (en) Micro inertial measurement unit
CN105607106B (en) A kind of low-cost and high-precision BD/MEMS fusions attitude measurement method
CN114440865A (en) Integrated navigation method based on MEMS inertial measurement system
CN109596018B (en) High-precision estimation method for flight attitude of spinning projectile based on magnetic roll angle rate information
CN106342284B (en) A kind of flight carrier attitude is determined method
CN107655476A (en) Pedestrian's high accuracy foot navigation algorithm based on Multi-information acquisition compensation
CN102445200A (en) Microminiature personal combined navigation system as well as navigating and positioning method thereof
CN109373832B (en) Method for measuring initial parameters of rotating projectile muzzle based on magnetic rolling
CN111044075B (en) SINS error online correction method based on satellite pseudo-range/relative measurement information assistance
Bezick et al. Inertial navigation for guided missile systems
CN112504275A (en) Water surface ship horizontal attitude measurement method based on cascade Kalman filtering algorithm
CN104864874B (en) A kind of inexpensive single gyro dead reckoning navigation method and system
CN113503892B (en) Inertial navigation system moving base initial alignment method based on odometer and retrospective navigation
CN110243362A (en) A kind of high and medium ultrasonic target air navigation aid
Yang et al. A novel method for fast stationary initial alignment based on extended measurement information
CN113701752A (en) Full strapdown attitude measurement device and method for spinning projectile
Minor et al. Utilization of GPS/MEMS-IMU for measurement of dynamics for range testing of missiles and rockets
CN111337056A (en) Optimization-based LiDAR motion compensation position and attitude system alignment method
CN115371707A (en) Coarse alignment method for Doppler radar assisted strapdown inertial navigation motion base under large azimuth misalignment angle
CN114578857A (en) Guidance aircraft autonomous control method, device and system based on full trajectory information
CN110986926B (en) Flight projectile body rotation attitude measurement method based on geomagnetic elements
CN110220534B (en) Online calibration method applied to on-missile inertial measurement unit

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
WW01 Invention patent application withdrawn after publication

Application publication date: 20220506

WW01 Invention patent application withdrawn after publication