CN112556724A - Initial coarse alignment method for low-cost navigation system of micro aircraft in dynamic environment - Google Patents

Initial coarse alignment method for low-cost navigation system of micro aircraft in dynamic environment Download PDF

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CN112556724A
CN112556724A CN202011428960.XA CN202011428960A CN112556724A CN 112556724 A CN112556724 A CN 112556724A CN 202011428960 A CN202011428960 A CN 202011428960A CN 112556724 A CN112556724 A CN 112556724A
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coordinate system
aircraft
navigation
vector
initial
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张达
李康伟
潘芷纯
刘青
黄晓龙
张华君
许铠通
裴家涛
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Hubei Institute Of Aerospacecraft
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Hubei Institute Of Aerospacecraft
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    • 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
    • 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
    • G01C21/005—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 with correlation of navigation data from several sources, e.g. map or contour matching
    • 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
    • G01C21/165—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

Abstract

The initial coarse alignment method of the low-cost navigation system of the micro aircraft in the dynamic environment is characterized in that the micro aircraft is provided with sensor equipment such as a three-axis inertial device, a magnetometer and a satellite navigation system GNSS, the initial alignment is carried out by adopting the combined information of the inertial device, the magnetometer and the satellite navigation system GNSS, the three-axis inertial device and the magnetometer on the aircraft respectively measure the acceleration and the magnetic field intensity information under a coordinate system of the aircraft body, the GNSS of the satellite navigation system measures the current longitude and latitude of a carrier and the speed information relative to a geographic navigation coordinate system, a direction cosine matrix is converted from the coordinate system of the aircraft body to the initial attitude of the geographic navigation coordinate system through calculation, and then the initial attitude angle of the aircraft relative to the geographic navigation coordinate system is determined to complete the initial.

Description

Initial coarse alignment method for low-cost navigation system of micro aircraft in dynamic environment
Technical Field
The invention belongs to the technical field of aircraft navigation positioning, and particularly relates to an initial coarse alignment method of a low-cost navigation system of a miniature aircraft in a dynamic environment.
Background
In recent years, with the rapid development of unmanned aerial vehicles and micro aircrafts in military and civil industries, the expansion of the application field of the unmanned aerial vehicle industry becomes continuous power for promoting the development of the unmanned aerial vehicle industry. The unmanned micro aircraft used after ground launching or aerial release and then electrified is different from the traditional ground take-off and landing aircraft, has low requirements on fields and use environments, can be applied to the military fields of remote monitoring, investigation and striking and the like of the sea and enemy territory, and the civil fields of aerial photography, disaster rescue, mapping and the like during sea and forest fires. Positioning accuracy of a navigation system using a micro aircraft during air electrification becomes the first guarantee for successfully executing a flight task, and when a Micro Electro Mechanical System (MEMS) navigation system widely adopted by the micro aircraft is initially aligned, the initial attitude angles of an engine body relative to a geographic navigation coordinate system, namely an initial pitch angle and a roll angle, need to be determined, and the determination of the initial attitude angle becomes the key of subsequent positioning accuracy, so that the initial alignment method of the low-cost navigation system of the micro aircraft under the air dynamic environment has great application prospect.
The MEMS navigation positioning sensor widely adopted by the micro aircraft has the advantages of low cost, small size and the like, but also has the defects of low precision and the like, can not identify the rotation angular velocity of the earth, and is difficult to solve the problem of self-alignment. Meanwhile, aiming at the micro aircraft used after being launched on the ground or being charged after being thrown in the air, because the micro aircraft is in the dynamic environment in the air at the initial moment and is simultaneously influenced by aerodynamic lift, aerodynamic resistance, aircraft self power, gravity and the like, the acceleration caused by acting force except the gravity exists, and the method for determining the initial pitch angle and the roll angle only by the accelerometer measurement value in the static arrangement state on the ground adopted by the traditional micro aircraft at the initial alignment cannot be applied. Therefore, in order to obtain the initial coarse alignment posture of the micro-aircraft low-cost MEMS navigation system in the air dynamic environment, the alignment must be assisted by external information.
The navigation system of the traditional unmanned aerial vehicle widely adopts a magnetometer, a satellite navigation system GNSS and the like to carry out external information to assist a low-cost MEMS inertial device to carry out combined navigation, so that the task requirement is met only by modifying a software algorithm under the condition of utilizing the existing hardware and not increasing a measuring sensor, the cost can be greatly reduced, and the practicability and the economic benefit are improved. However, for the alignment problem in the air dynamic environment, since the method of providing the initial absolute reference information by using the gravity acceleration in the ground static state cannot be adopted, another kind of absolute reference information needs to be found. Meanwhile, the updating speed of the GNSS of the satellite navigation system is low, the sampling rate of the MEMS inertial navigation is high, and the difference of the sampling rates can cause information delay to generate errors, so that how to reasonably use the magnetometer, the GNSS of the satellite navigation system and the like to carry out auxiliary alignment of external information needs to be further solved.
Disclosure of Invention
The invention aims to solve the task application problems of low precision of a low-cost navigation system and alignment of an air dynamic environment, and provides an initial rough alignment method of a low-cost navigation system of a micro aircraft under the dynamic environment, wherein the micro aircraft is provided with sensor equipment such as a triaxial inertia device, a magnetometer and a satellite navigation system GNSS, the initial alignment is carried out by adopting the combined information of the inertia device, the magnetometer and the satellite navigation system GNSS, the airborne triaxial inertia device and the magnetometer respectively measure the acceleration and magnetic field intensity information under a body coordinate system, the satellite navigation system GNSS measures the current longitude and latitude of a carrier and the speed information relative to a geographical navigation coordinate system, converting the initial attitude of the coordinate system of the body to the coordinate system of the geographical navigation into a direction cosine matrix through calculation, and then determining an initial attitude angle of the aircraft relative to a geographical navigation coordinate system to complete initial coarse alignment.
Further, the method comprises the following steps:
s1, the magnetic field intensity vector of the real-time current position of the micro aircraft is obtained by inquiring an earth magnetic field distribution table according to the GNSS positioning data or directly calculated according to an earth magnetic field model,
the vector relationship between the measurement data of the airborne measuring magnetometer and the magnetic field intensity under the local geographic coordinate system is as follows:
Figure BDA0002825940590000031
in the formula, mnIs the magnetic field intensity vector under the local geographical navigation coordinate system,
Figure BDA0002825940590000032
the vector of the magnetic field intensity under the coordinate system of the machine body is measured by the magnetometer,
Figure BDA0002825940590000033
converting a direction cosine matrix from a body coordinate system to a geographical navigation coordinate system;
the origin of the geographic navigation coordinate system is located at the center of mass of the aircraft, and the three-dimensional coordinate axes respectively point to the fixed directions of the north, the east and the ground; the origin of the body coordinate system is positioned at the mass center of the aircraft, and the three-dimensional coordinate axes point to the reference directions of the carrier, namely the front, the right and the lower sides;
s2, for the accelerometer measuring data, since the accelerometer measures the specific force, the specific force is not the motion acceleration of the carrier, but the difference between the absolute acceleration of the carrier relative to the inertia space and the gravity acceleration of the earth, wherein the inertia space is the difference between the motion and the rest of the object described by newton' S law relative to a special reference frame, there is the following relationship:
Figure BDA0002825940590000034
in the formula (f)nFor local geography guidanceA specific force vector under the navigation coordinate system;
Figure BDA0002825940590000035
a specific force vector of a coordinate system of the body measured for the accelerometer;
aiming at the alignment problem in the air dynamic environment, because the method of providing initial absolute reference information by using the gravity acceleration under the static state of the ground cannot be used, another absolute reference information needs to be searched, and the specific force vector f under the local geographic coordinate systemnAccording to a gravity field data model and a micro aircraft carrier current motion velocity vector vnDetermining, based on the Coriolis acceleration theorem, a specific force f on a geographical coordinate systemnWith acceleration in a geographical coordinate system
Figure BDA0002825940590000036
And acceleration of gravity gnThe following relationship is satisfied:
Figure BDA0002825940590000041
in the formula (I), the compound is shown in the specification,
Figure BDA0002825940590000042
the rotation angular speed of the geographic coordinate system;
Figure BDA0002825940590000043
projecting the rotational angular velocity of the earth on a geographic coordinate system; for a carrier with a small moving speed, and the measurement accuracy limit of a low-cost MEMS navigation system,
Figure BDA0002825940590000044
negligible, so the following formula:
Figure BDA0002825940590000045
s3 speed information, namely the current motion speed vector of the micro-aircraft carrierQuantity vnThe measurement is obtained through a GNSS system;
s4 magnetic field intensity vector under assumed measurement vector machine body coordinate system
Figure BDA0002825940590000046
Specific force vector of body coordinate system
Figure BDA0002825940590000047
The vectors are not parallel, and if the two vectors are parallel, the measurement can be performed when the two vectors are not parallel, such as the next measurement time, and the like, by judging that the two vectors are parallel, so that a third vector equation is obtained by the vector cross product of the two vectors as follows:
Figure BDA0002825940590000048
wherein the content of the first and second substances,
Figure BDA0002825940590000049
is a posture conversion direction cosine matrix of the vector coordinate conversion from the body coordinate system to the geographic coordinate system;
s5, calculating a vector m in three known calculation vectorsn、fnAnd mn×fnAnd three measurement vectors
Figure BDA00028259405900000410
And
Figure BDA00028259405900000411
in combination with the above formula, the following formula is given:
Figure BDA00028259405900000412
the orientation transformation direction cosine matrix from the computer body coordinate system to the geographic coordinate system is as follows:
Figure BDA00028259405900000413
s6, rolling angle phi, pitching angle theta and yaw angle psi of the aircraft and attitude transformation direction cosine matrix from the body coordinate system to the geographic coordinate system
Figure BDA0002825940590000051
The relationship of (a) to (b) is as follows:
Figure BDA0002825940590000052
converting a direction cosine matrix according to attitude
Figure BDA0002825940590000053
The initial attitude angle (namely the roll angle phi) during the coarse alignment is calculated through an inverse trigonometric function0Angle of pitch theta0And yaw angle psi0The following were used:
Figure BDA0002825940590000054
in conclusion, the invention obtains the initial attitude angle in the air dynamic environment.
Further, step S1 is preceded by:
and S0, the micro aircraft is powered on after being launched from the ground or thrown in the air, and when the standby navigation sensor equipment is started to work normally after a period of time, the airborne inertial device, the magnetometer and the satellite navigation system GNSS work normally, and the micro aircraft navigation system starts to run.
Preferably, in step S4, since the speed information is obtained by measuring through the GNSS system, but the sampling rate of the MEMS inertial device is high, the sampling rate of the satellite navigation system is low, and time correction is required to be performed on the speed data to solve the problem of sampling rate mismatch, and by correcting the data compared at the sampling time, the error caused by the inconsistency of the sampling time between the GNSS system and the inertial device can be reduced. The method for time correction of the speed information measured by the GNSS system comprises the following steps:
if the output acceleration signal of the inertia device is at tk-1And tkObtained while the neighboring GNSS measurement speed information is
Figure BDA0002825940590000061
And
Figure BDA0002825940590000062
the velocity information of the GNSS measurement obtained above is
Figure BDA0002825940590000063
And
Figure BDA0002825940590000064
method for obtaining micro aircraft at time t by using linear interpolationk-1And tkUpper velocity estimation information vn(tk-1) And vn(tk) The details are as follows
Figure BDA0002825940590000065
Figure BDA0002825940590000066
According to the sampling data time interval of the MEMS inertial navigation sensor, namely the accelerometer, as delta t, the speed derivative on the geographic coordinate system, namely the acceleration information
Figure BDA0002825940590000067
Can be obtained by correcting the corrected speed information vn(tk-1) And vn(tk) Time difference obtaining is carried out, namely:
Figure BDA0002825940590000068
compared with the prior art, the technical scheme of the invention can obtain the following beneficial effects:
1. the initial alignment method designed by the invention obtains absolute reference information for determining the attitude angle of the initial carrier, namely the initial pitch angle and the roll angle, namely the absolute reference magnetic field intensity information of the current position is obtained by looking up a table of earth magnetic field distribution according to the current longitude and latitude, and the absolute reference acceleration information relative to the ground is obtained according to the differential value of the speed relative to the ground navigation coordinate system, so that the accuracy of the system is improved;
2. the initial alignment method designed by the invention solves the problem of mismatching of the sampling rates of the inertial device and the satellite navigation, reduces the alignment error, meets the continuous requirement of the sampling data by time correction of the acquired data, and improves the reliability of the system. .
Drawings
FIG. 1 is a diagram of the relationship between the combined information of an inertial device, a magnetometer and a GNSS of a satellite navigation system in accordance with the present invention;
FIG. 2 is a graph of the inertial device and time correction of satellite navigation sample data in accordance with the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
Example 1
The initial coarse alignment method of the low-cost navigation system of the miniature aircraft in the dynamic environment is characterized in that the miniature aircraft is provided with sensor equipment such as a triaxial inertia device, a magnetometer and a satellite navigation system GNSS, the initial alignment is carried out by adopting the combined information of the inertia device, the magnetometer and the satellite navigation system GNSS, the airborne triaxial inertia device and the magnetometer respectively measure the acceleration and magnetic field intensity information under a body coordinate system, the satellite navigation system GNSS measures the current longitude and latitude of a carrier and the speed information relative to a geographical navigation coordinate system, the initial attitude transformation direction cosine matrix from the body coordinate system to the geographical navigation coordinate system is obtained through calculation, and then the initial attitude angle of the aircraft relative to the geographical navigation coordinate system is determined to complete the initial coarse alignment.
The method comprises the following steps:
s0, the micro aircraft is powered on after being launched from the ground or thrown in the air, when a period of time passes and the standby navigation sensor equipment is started to work normally, the airborne inertial device, the magnetometer and the satellite navigation system GNSS work normally, and the micro aircraft navigation system starts to run;
s1, because the earth magnetic field data model is quite perfect, the magnetic field intensity vector of the real-time current position of the micro aircraft is obtained by inquiring the earth magnetic field distribution table according to the GNSS positioning data or directly calculated according to the earth magnetic field model,
the vector relationship between the measurement data of the airborne measuring magnetometer and the magnetic field intensity under the local geographic coordinate system is as follows:
Figure BDA0002825940590000081
in the formula, mnIs the magnetic field intensity vector under the local geographical navigation coordinate system,
Figure BDA0002825940590000082
the vector of the magnetic field intensity under the coordinate system of the machine body is measured by the magnetometer,
Figure BDA0002825940590000083
converting a direction cosine matrix from a body coordinate system to a geographical navigation coordinate system;
the origin of the geographic navigation coordinate system is located at the center of mass of the aircraft, and the three-dimensional coordinate axes respectively point to the fixed directions of the north, the east and the ground; the origin of the body coordinate system is positioned at the mass center of the aircraft, and the three-dimensional coordinate axes point to the reference directions of the carrier, namely the front, the right and the lower sides;
s2, for the accelerometer measuring data, since the accelerometer measures the specific force, the specific force is not the motion acceleration of the carrier, but the difference between the absolute acceleration of the carrier relative to the inertia space and the gravity acceleration of the earth, wherein the inertia space is the difference between the motion and the rest of the object described by newton' S law relative to a special reference frame, there is the following relationship:
Figure BDA0002825940590000084
in the formula (f)nA specific force vector under a local geographical navigation coordinate system;
Figure BDA0002825940590000085
a specific force vector of a coordinate system of the body measured for the accelerometer;
aiming at the alignment problem in the air dynamic environment, because the method of providing initial absolute reference information by using the gravity acceleration under the static state of the ground cannot be used, another absolute reference information needs to be searched, and the specific force vector f under the local geographic coordinate systemnAccording to a gravity field data model and a micro aircraft carrier current motion velocity vector vnDetermining, based on the Coriolis acceleration theorem, a specific force f on a geographical coordinate systemnWith acceleration in a geographical coordinate system
Figure BDA0002825940590000086
And acceleration of gravity gnThe following relationship is satisfied:
Figure BDA0002825940590000091
in the formula (I), the compound is shown in the specification,
Figure BDA0002825940590000092
the rotation angular speed of the geographic coordinate system;
Figure BDA0002825940590000093
projecting the rotational angular velocity of the earth on a geographic coordinate system; measurement for carrier with small motion speed and low-cost MEMS navigation systemThe limitation of the precision is that the precision is limited,
Figure BDA0002825940590000094
negligible, so the following formula:
Figure BDA0002825940590000095
s3, velocity information, namely the current motion velocity vector v of the micro aircraft carriernThe measurement is obtained through a GNSS system;
s4 magnetic field intensity vector under assumed measurement vector machine body coordinate system
Figure BDA0002825940590000096
Specific force vector of body coordinate system
Figure BDA0002825940590000097
The vectors are not parallel, and if the two vectors are parallel, the measurement can be performed when the two vectors are not parallel, such as the next measurement time, and the like, by judging that the two vectors are parallel, so that a third vector equation is obtained by the vector cross product of the two vectors as follows:
Figure BDA0002825940590000098
wherein the content of the first and second substances,
Figure BDA0002825940590000099
is a posture conversion direction cosine matrix of the vector coordinate conversion from the body coordinate system to the geographic coordinate system;
s5, calculating a vector m in three known calculation vectorsn、fnAnd mn×fnAnd three measurement vectors
Figure BDA00028259405900000910
And
Figure BDA00028259405900000911
in combination with the above formula, the following formula is given:
Figure BDA00028259405900000912
the orientation transformation direction cosine matrix from the computer body coordinate system to the geographic coordinate system is as follows:
Figure BDA00028259405900000913
s6, rolling angle phi, pitching angle theta and yaw angle psi of the aircraft and attitude transformation direction cosine matrix from the body coordinate system to the geographic coordinate system
Figure BDA0002825940590000101
The relationship of (a) to (b) is as follows:
Figure BDA0002825940590000102
converting a direction cosine matrix according to attitude
Figure BDA0002825940590000103
The initial attitude angle (namely the roll angle phi) during the coarse alignment is calculated through an inverse trigonometric function0Angle of pitch theta0And yaw angle psi0The following were used:
Figure BDA0002825940590000104
in conclusion, the invention obtains the initial attitude angle in the air dynamic environment.
The method obtains the initial attitude angle in the air dynamic environment, can provide calculation initial information for a subsequent navigation system, and meets the requirement of initial coarse alignment of a low-cost navigation system of a micro aircraft in the air dynamic environment.
Example 2
The present embodiment is different from embodiment 1 in that:
in the step S4, since the speed information is obtained by GNSS system measurement, but the sampling rate of the MEMS inertial device is fast, and the sampling rate of the GNSS navigation system is slow, time correction is required for the speed data to solve the problem of sampling rate mismatch, and by correcting the comparison data at the sampling time, the error caused by inconsistency of the sampling times of the GNSS system and the inertial device can be reduced. The method for time correction of the speed information measured by the GNSS system comprises the following steps:
if the output acceleration signal of the inertia device is at tk-1And tkObtained while the neighboring GNSS measurement speed information is
Figure BDA0002825940590000111
And
Figure BDA0002825940590000112
the velocity information of the GNSS measurement obtained above is
Figure BDA0002825940590000113
And
Figure BDA0002825940590000114
method for obtaining micro aircraft at time t by using linear interpolationk-1And tkUpper velocity estimation information vn(tk-1) And vn(tk) The method comprises the following steps:
Figure BDA0002825940590000115
Figure BDA0002825940590000116
time interval of sampled data according to MEMS inertial navigation sensor, namely accelerometer, is delta t, velocity derivative on geographic coordinate system
Figure BDA0002825940590000117
Can be obtained by correcting the corrected speed information vn(tk-1) And vn(tk) Time difference obtaining is carried out, namely:
Figure BDA0002825940590000118
the rest is the same as in example 1.

Claims (4)

1. The initial coarse alignment method of the low-cost navigation system of the micro aircraft in the dynamic environment is characterized in that the micro aircraft is provided with sensor equipment such as a three-axis inertial device, a magnetometer and a satellite navigation system GNSS, the initial alignment is carried out by adopting the combined information of the inertial device, the magnetometer and the satellite navigation system GNSS, the three-axis inertial device and the magnetometer on the aircraft respectively measure the acceleration and the magnetic field intensity information under a coordinate system of the aircraft body, the GNSS of the satellite navigation system measures the current longitude and latitude of a carrier and the speed information relative to a geographic navigation coordinate system, a direction cosine matrix is converted from the coordinate system of the aircraft body to the initial attitude of the geographic navigation coordinate system through calculation, and then the initial attitude angle of the aircraft relative to the geographic navigation coordinate system is determined to complete the initial.
2. The method for initial coarse alignment of a low-cost navigation system of a micro-aircraft in a dynamic environment according to claim 1, comprising the steps of:
s1, the magnetic field intensity vector of the real-time current position of the micro aircraft is obtained by inquiring an earth magnetic field distribution table according to the GNSS positioning data or directly calculated according to an earth magnetic field model,
the vector relationship between the measurement data of the airborne measuring magnetometer and the magnetic field intensity under the local geographic coordinate system is as follows:
Figure FDA0002825940580000011
in the formula, mnIs the magnetic field intensity vector under the local geographical navigation coordinate system,
Figure FDA0002825940580000012
the vector of the magnetic field intensity under the coordinate system of the machine body is measured by the magnetometer,
Figure FDA0002825940580000013
converting a direction cosine matrix from a body coordinate system to a geographical navigation coordinate system;
the origin of the geographic navigation coordinate system is located at the center of mass of the aircraft, and the three-dimensional coordinate axes respectively point to the fixed directions of the north, the east and the ground; the origin of the body coordinate system is positioned at the mass center of the aircraft, and the three-dimensional coordinate axes point to the reference directions of the carrier, namely the front, the right and the lower sides;
s2, for the accelerometer measuring data, since the accelerometer measures the specific force, the specific force is not the motion acceleration of the carrier, but the difference between the absolute acceleration of the carrier relative to the inertia space and the gravity acceleration of the earth, wherein the inertia space is the difference between the motion and the rest of the object described by newton' S law relative to a special reference frame, there is the following relationship:
Figure FDA0002825940580000021
in the formula (f)nA specific force vector under a local geographical navigation coordinate system;
Figure FDA0002825940580000022
a specific force vector of a coordinate system of the body measured for the accelerometer;
aiming at the alignment problem in the air dynamic environment, because the method of providing initial absolute reference information by using the gravity acceleration under the static state of the ground cannot be used, another absolute reference information needs to be searched, and the specific force vector f under the local geographic coordinate systemnAccording to a gravity field data model and a micro aircraft carrier current motion velocity vector vnDetermination based on Coriolis accelerationTheorem, determining specific force f on geographical coordinate systemnWith acceleration in a geographical coordinate system
Figure FDA0002825940580000023
And acceleration of gravity gnThe following relationship is satisfied:
Figure FDA0002825940580000024
in the formula (I), the compound is shown in the specification,
Figure FDA0002825940580000025
the rotation angular speed of the geographic coordinate system;
Figure FDA0002825940580000026
projecting the rotational angular velocity of the earth on a geographic coordinate system; for a carrier with a small moving speed, and the measurement accuracy limit of a low-cost MEMS navigation system,
Figure FDA0002825940580000027
negligible, so the following formula:
Figure FDA0002825940580000028
s3, velocity information, namely the current motion velocity vector v of the micro aircraft carriernThe measurement is obtained through a GNSS system;
s4 magnetic field intensity vector under assumed measurement vector machine body coordinate system
Figure FDA0002825940580000029
Specific force vector of body coordinate system
Figure FDA00028259405800000210
Non-parallel, if the two are judged to be in fact in the air because the carrier continuously moves in the airWhen the vectors are parallel, the measurement can be carried out again when the two vectors are not parallel, such as the next measurement time, and therefore, a third vector equation is obtained through the vector cross product of the two vectors as follows:
Figure FDA00028259405800000211
wherein the content of the first and second substances,
Figure FDA00028259405800000212
is a posture conversion direction cosine matrix of the vector coordinate conversion from the body coordinate system to the geographic coordinate system;
s5, calculating a vector m in three known calculation vectorsn、fnAnd mn×fnAnd three measurement vectors
Figure FDA0002825940580000031
Figure FDA0002825940580000032
And
Figure FDA0002825940580000033
in combination with the above formula, the following formula is given:
Figure FDA0002825940580000034
the orientation transformation direction cosine matrix from the computer body coordinate system to the geographic coordinate system is as follows:
Figure FDA0002825940580000035
s6, rolling angle phi, pitching angle theta and yaw angle psi of the aircraft and attitude transformation direction cosine matrix from the body coordinate system to the geographic coordinate system
Figure FDA0002825940580000039
The relationship of (a) to (b) is as follows:
Figure FDA0002825940580000036
converting a direction cosine matrix according to attitude
Figure FDA0002825940580000037
The initial attitude angle (namely the roll angle phi) during the coarse alignment is calculated through an inverse trigonometric function0Angle of pitch theta0And yaw angle psi0The following were used:
Figure FDA0002825940580000038
in conclusion, the invention obtains the initial attitude angle in the air dynamic environment.
3. The method for initial coarse alignment of a low-cost navigation system of a micro-aircraft in a dynamic environment according to claim 2, wherein said step S1 is preceded by the step of:
and S0, the micro aircraft is powered on after being launched from the ground or thrown in the air, and after a period of time and the standby navigation sensor equipment is started to work normally, the airborne inertial device, the magnetometer and the satellite navigation system GNSS work normally, and the micro aircraft navigation system starts to run.
4. The initial coarse alignment method for the low-cost navigation system of a micro-aircraft under the dynamic environment according to any one of claims 1 to 3, wherein in the step S4, the velocity information measured by the GNSS system is time-corrected by:
if the output acceleration signal of the inertia device is at tk-1And tkObtained while the neighboring GNSS measurement speed information is
Figure FDA0002825940580000041
And
Figure FDA0002825940580000042
the velocity information of the GNSS measurement obtained above is
Figure FDA0002825940580000043
And
Figure FDA0002825940580000044
method for obtaining micro aircraft at time t by using linear interpolationk-1And tkUpper velocity estimation information vn(tk-1) And vn(tk) The method comprises the following steps:
Figure FDA0002825940580000045
Figure FDA0002825940580000046
according to the sampling data time interval of the MEMS inertial navigation sensor, namely the accelerometer, as delta t, the speed derivative on the geographic coordinate system, namely the acceleration information
Figure FDA0002825940580000047
Can be obtained by correcting the corrected speed information vn(tk-1) And vn(tk) Time difference obtaining is carried out, namely:
Figure FDA0002825940580000048
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113218423A (en) * 2021-05-25 2021-08-06 上海机电工程研究所 Aerial coarse alignment method without reference attitude information during transmitting
CN113740890A (en) * 2021-08-31 2021-12-03 普宙科技(深圳)有限公司 Course angle correction method, system, computer equipment and storage medium
CN114018255A (en) * 2021-11-03 2022-02-08 湖南国天电子科技有限公司 Intelligent integrated navigation method, system, equipment and medium for underwater glider

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2002090173A (en) * 2000-09-18 2002-03-27 Toshiba Corp Inertia navigation system and its initial alignment method
CN103217174A (en) * 2013-04-10 2013-07-24 哈尔滨工程大学 Initial alignment method of strap-down inertial navigation system based on low-precision micro electro mechanical system
US20130245984A1 (en) * 2010-11-17 2013-09-19 Hillcrest Laboratories, Inc. Apparatuses and methods for magnetometer alignment calibration without prior knowledge of the local magnetic field
CN103900608A (en) * 2014-03-21 2014-07-02 哈尔滨工程大学 Low-precision inertial navigation initial alignment method based on quaternion CKF
CN104374405A (en) * 2014-11-06 2015-02-25 哈尔滨工程大学 MEMS strapdown inertial navigation initial alignment method based on adaptive central difference Kalman filtering
CN109556632A (en) * 2018-11-26 2019-04-02 北方工业大学 INS/GNSS/polarization/geomagnetic integrated navigation alignment method based on Kalman filtering
CN109752000A (en) * 2018-12-12 2019-05-14 哈尔滨工程大学 A kind of MEMS dual-axis rotation modulation type strapdown compass Initial Alignment Method
WO2019168735A1 (en) * 2018-02-28 2019-09-06 Idhl Holdings, Inc. Methods and apparatus for planar magnetometer calibration, heading determination, gyroscope assisted magnetometer amplitude calibration, magnetometer amplitude and alignment calibration, magnetometer mapping, and sensor fusion

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2002090173A (en) * 2000-09-18 2002-03-27 Toshiba Corp Inertia navigation system and its initial alignment method
US20130245984A1 (en) * 2010-11-17 2013-09-19 Hillcrest Laboratories, Inc. Apparatuses and methods for magnetometer alignment calibration without prior knowledge of the local magnetic field
CN103217174A (en) * 2013-04-10 2013-07-24 哈尔滨工程大学 Initial alignment method of strap-down inertial navigation system based on low-precision micro electro mechanical system
CN103900608A (en) * 2014-03-21 2014-07-02 哈尔滨工程大学 Low-precision inertial navigation initial alignment method based on quaternion CKF
CN104374405A (en) * 2014-11-06 2015-02-25 哈尔滨工程大学 MEMS strapdown inertial navigation initial alignment method based on adaptive central difference Kalman filtering
WO2019168735A1 (en) * 2018-02-28 2019-09-06 Idhl Holdings, Inc. Methods and apparatus for planar magnetometer calibration, heading determination, gyroscope assisted magnetometer amplitude calibration, magnetometer amplitude and alignment calibration, magnetometer mapping, and sensor fusion
CN109556632A (en) * 2018-11-26 2019-04-02 北方工业大学 INS/GNSS/polarization/geomagnetic integrated navigation alignment method based on Kalman filtering
CN109752000A (en) * 2018-12-12 2019-05-14 哈尔滨工程大学 A kind of MEMS dual-axis rotation modulation type strapdown compass Initial Alignment Method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
范蓓蓓: "外部信息辅助无陀螺捷联惯导初始对准仿真研究", 《电子测量与仪器学报》 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113218423A (en) * 2021-05-25 2021-08-06 上海机电工程研究所 Aerial coarse alignment method without reference attitude information during transmitting
CN113740890A (en) * 2021-08-31 2021-12-03 普宙科技(深圳)有限公司 Course angle correction method, system, computer equipment and storage medium
CN114018255A (en) * 2021-11-03 2022-02-08 湖南国天电子科技有限公司 Intelligent integrated navigation method, system, equipment and medium for underwater glider
CN114018255B (en) * 2021-11-03 2023-06-27 湖南国天电子科技有限公司 Intelligent integrated navigation method, system, equipment and medium of underwater glider

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