CN113218390B - Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion - Google Patents
Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion Download PDFInfo
- Publication number
- CN113218390B CN113218390B CN202110585391.8A CN202110585391A CN113218390B CN 113218390 B CN113218390 B CN 113218390B CN 202110585391 A CN202110585391 A CN 202110585391A CN 113218390 B CN113218390 B CN 113218390B
- Authority
- CN
- China
- Prior art keywords
- navigation system
- attitude
- matrix
- star
- integrated navigation
- 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.)
- Active
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
- G01C21/20—Instruments for performing navigational calculations
-
- 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
- G01S13/00—Systems using the reflection or reradiation of radio waves, e.g. radar systems; Analogous systems using reflection or reradiation of waves whose nature or wavelength is irrelevant or unspecified
- G01S13/86—Combinations of radar systems with non-radar systems, e.g. sonar, direction finder
-
- 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
- G01S13/00—Systems using the reflection or reradiation of radio waves, e.g. radar systems; Analogous systems using reflection or reradiation of waves whose nature or wavelength is irrelevant or unspecified
- G01S13/88—Radar or analogous systems specially adapted for specific applications
- G01S13/882—Radar or analogous systems specially adapted for specific applications for altimeters
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F17/00—Digital computing or data processing equipment or methods, specially adapted for specific functions
- G06F17/10—Complex mathematical operations
- G06F17/16—Matrix or vector computation, e.g. matrix-matrix or matrix-vector multiplication, matrix factorization
Abstract
The invention discloses a rotary inertia astronomical combined navigation method based on the fusion of attitude and star altitude angle, which is characterized in that an inertia device and a star sensor are coincidently installed and are fixedly connected with a rotary mechanism; when the motor of the rotating mechanism drives the inertia device and the star sensor to rotate according to a single-shaft four-position rotation scheme, measuring data of the combined system under the condition; when the rotating mechanism is static, the star sensor starts to work, and data of the combined system under the condition are measured; sending the data of the two times into a single-axis rotation strapdown inertia/astronomical integrated navigation Kalman filter for filtering calculation to obtain an estimated value of the state error of the integrated navigation system; and finally, carrying out error correction on the integrated navigation system in real time through the estimated value of the state error of the integrated navigation system to obtain the high-precision attitude speed position information of the integrated navigation system. The invention can realize high-precision autonomous navigation of the aircraft.
Description
Technical Field
The invention belongs to the field of aircraft navigation, and particularly relates to a rotary inertia astronomical combined navigation method based on attitude and star altitude angle fusion.
Background
In order to enhance the autonomy of a navigation system of the hypersonic aircraft in future operations, an passivity navigation system should be selected as far as possible to meet the characteristics when the navigation system is selected. If the combined navigation system of the hypersonic aircraft in the near space depends heavily on a GNSS system, stable and reliable autonomous navigation is difficult to realize. It is reported that iran successfully captured the RQ-170 of the unmanned aerial vehicle in the united states in 2012 by using GPS spoofing, which indicates the importance of strong interference immunity when navigating an aircraft, particularly a hypersonic aircraft.
At present, an inertial navigation system is a passive navigation system, the anti-interference capability is strong, the quick information updating capability of the inertial navigation system can also provide information support for a quick maneuvering near space aircraft, and the inertial navigation system is the best choice of the near space aircraft navigation system, but the characteristic that errors are accumulated along with time does not meet the long-term flight requirement of the aircraft. Therefore, an auxiliary navigation system is needed to be matched for error correction, the current general solution is to match a global satellite navigation system for correction, and although the satellite navigation has high precision and complete information, the satellite navigation system is easy to be interfered, and the shortcoming of poor wartime usability exists.
Disclosure of Invention
Aiming at the defects in the prior art, the rotating inertia astronomical combined navigation method based on the fusion of the attitude and the star altitude angle solves the problems of poor autonomy and low precision of a navigation system.
In order to achieve the purpose of the invention, the invention adopts the technical scheme that:
the rotation inertia astronomical combined navigation method based on the attitude and star altitude angle fusion is provided, and comprises the following steps:
s1, overlapping and installing the inertial device and the star sensor, and fixedly connecting the inertial device and the star sensor with the rotating mechanism to form a combined navigation system; the carrier where the integrated navigation system is located comprises an aircraft;
s2, the inertial device and the star sensor are driven to rotate according to a single-shaft four-position rotation scheme by a motor of the rotating mechanism;
s3, measuring attitude angular velocity and specific force information of the integrated navigation system through an inertial device;
s4, acquiring and obtaining a transformation matrix from the rotating coordinate system to the aircraft coordinate system according to the rotating angular velocity of the integrated navigation system;
s5, combining the attitude angular velocity and specific force information measured by the inertial device with the conversion matrix to calculate a strapdown inertial navigation algorithm, and obtaining attitude, velocity and position information of the aircraft under the rotation condition of the rotating mechanism;
s6, when the rotating mechanism is static, attitude information of the aircraft relative to the geocentric inertial system is obtained through the star sensor; calculating and solving altitude angle information of the fixed star by a spherical triangle method; measuring to obtain the altitude information of the aircraft through a radar altimeter; acquiring a state transition matrix of a state error equation of the integrated navigation system;
s7, sending the information obtained in the step S6 to a single-axis rotation strapdown inertia/astronomical integrated navigation Kalman filter for filtering calculation to obtain an estimated value of the state error of the integrated navigation system;
and S8, correcting the error of the information obtained in the step S5 in real time according to the estimated value of the state error of the integrated navigation system, and obtaining the high-precision attitude speed position information of the integrated navigation system.
Further, the specific method of step S4 is:
according to the formula:
obtaining a transformation matrix
Where ω is the angular velocity of rotation of the rotary mechanism and t is the time of rotation of the rotary mechanism.
Further, the specific method of step S6 is:
s6-1, obtaining a vector of the fixed star in an image space coordinate system through star map centroid subdivision positioning, obtaining a vector of the fixed star in a geocentric inertial coordinate system through star map matching identification, and obtaining attitude information of the combined navigation system relative to the geocentric inertial system through least square calculation;
s6-2, according to the formula:
sin(h star )＝sin(L)sin(δA)+cos(L)cos(δA)cos(λ+t g )
obtaining altitude angle information of stars, wherein h star Is star height angle, delta A is star declination, lambda is longitude, L is latitude, t g Greenwich mean time;
s6-3, according to the formula:
acquiring attitude information: attitude combination measurement value Z φ Attitude measurement matrix H φ And altitude information of stars: fixed star altitude angle measurement value Z E Altitude angle measurement matrix H E Altitude information of the aircraft: radar altimeter measurement Z h And a height measurement matrix H h ；Z k Measuring values of the integrated navigation system, including attitude integrated measuring values, star altitude angle measuring values and radar altimeter measuring values; v φ 、V E 、V h All are the measurement noise of the sensor; x is a state variable, wherein the state variable comprises three-axis attitude misalignment angle information of the integrated navigation system, three-axis speed error of the integrated navigation system, position error of the integrated navigation system, constant deviation of a gyroscope and installation error of a star sensor and an inertial device; δ λ is the longitude error of the aircraft; δ L is the latitude error of the aircraft; z is a linear or branched member dcm For constructed measurements, φ is the three-axis attitude misalignment angle information for the integrated navigation system, δ p is the position error for the integrated navigation system, which includes longitude, latitude, and altitude,
is the attitude matrix, V, of the aircraft relative to the integrated navigation system s For angle error, (. cndot.) x and [. cndot. ]]X is an antisymmetric matrix, L is latitude, λ is longitude, and h is altitude; i is an identity matrix, C δP Is a position matrix, α i To identify the sidereal Chijing, beta i Identified sidereal declination;
s6-4, according to the formula:
obtaining a state transition matrix Φ:
wherein:
f is the oblateness of the earth, R e Is the semi-major axis, omega, of the earth ie Is the rotational angular velocity, v, of the earth E East velocity, v N Is the north speed; f 0 、F 1 、F 2 、F 3 、F 4 、R M And R N Are all intermediate parameters; deltav is the three-axis velocity error of the integrated navigation system, epsilon is the constant deviation of the gyroscope,
is the constant deviation of the accelerometer, mu is the installation error of the star sensor and the inertial device,
differentiating the three-axis attitude misalignment angle information of the integrated navigation system,
for the three-axis velocity error differential of the integrated navigation system,
the position error differential for the integrated navigation system, which includes longitude, latitude and altitude,
is a constant offset derivative of the gyroscope,
for the accelerometer constant offset differential to be,
is the differential of the installation error of the star sensor and the inertial device,
is an antisymmetric matrix of the angular velocity of the navigational coordinate system relative to the earth's center inertial system,
antisymmetric matrix of specific force information measured for an accelerometer, v n X is the anti-symmetric matrix of the velocity,
is the angular velocity of the earth's rotation,
for the angular velocity, w, of the navigation coordinate system relative to the terrestrial coordinate system g For random drift of the gyroscope, w a Is the random drift of the accelerometer.
Further, the specific method of step S7 is:
s7-1, according to the formula:
obtaining a state one-step prediction value
Wherein phi k/k-1 Is the state transition matrix at the last moment in time,
the state optimal estimated value at the last moment is obtained;
s7-2, according to the formula:
obtaining a one-step prediction error variance P k/k-1 (ii) a Wherein P is k-1 Is the error variance matrix of the previous time (·) T Being a transpose of a matrix, Γ k-1 For the system noise interference matrix, Q k-1 Is a process noise matrix;
s7-3, according to the formula:
obtaining a filter gain matrix K k (ii) a Wherein H k Is a measurement matrix having a value of
R k Measuring a noise variance matrix;
s7-4, according to the formula:
obtaining the optimal state error estimation of the integrated navigation system at the moment
The invention has the beneficial effects that: on the basis of traditional inertial astronomical combined navigation, observation of star height angles is added, the error of an inertial device is delayed by using a rotation technology, and the navigation precision can be improved while the autonomy is enhanced.
Drawings
FIG. 1 is a block diagram of a rotational inertial astronomical combination system based on attitude and star altitude fusion;
FIG. 2 is a flow chart of a rotational inertial astronomical integrated navigation method based on attitude and star altitude fusion;
FIG. 3 is a plot of attitude error contrast;
FIG. 4 is a graph comparing velocity errors;
fig. 5 is a position error comparison diagram.
Detailed Description
The following description of the embodiments of the present invention is provided to facilitate the understanding of the present invention by those skilled in the art, but it should be understood that the present invention is not limited to the scope of the embodiments, and it will be apparent to those skilled in the art that various changes may be made without departing from the spirit and scope of the invention as defined and defined in the appended claims, and all matters produced by the invention using the inventive concept are protected.
As shown in fig. 1 and 2, the rotational inertial astronomical combined navigation method based on attitude and sidereal altitude fusion comprises the following steps:
s1, overlapping and installing the inertial device and the star sensor, and fixedly connecting the inertial device and the star sensor with the rotating mechanism to form a combined navigation system; the carrier where the integrated navigation system is located comprises an aircraft;
s2, the inertial device and the star sensor are driven to rotate according to a single-shaft four-position rotation scheme by a motor of the rotating mechanism;
s3, measuring attitude angular velocity and specific force information of the integrated navigation system through an inertial device;
s4, acquiring and obtaining a transformation matrix from the rotating coordinate system to the aircraft coordinate system according to the rotating angular velocity of the integrated navigation system;
s5, combining the attitude angular velocity and specific force information measured by the inertial device with the conversion matrix to calculate a strapdown inertial navigation algorithm, and obtaining attitude, velocity and position information of the aircraft under the rotation condition of the rotating mechanism;
s6, when the rotating mechanism is static, attitude information of the aircraft relative to the geocentric inertial system is obtained through the star sensor; calculating and solving altitude angle information of the fixed star by a spherical triangle method; measuring to obtain the altitude information of the aircraft through a radar altimeter; acquiring a state transition matrix of a state error equation of the integrated navigation system;
s7, sending the information obtained in the step S6 to a single-axis rotation strapdown inertia/astronomical integrated navigation Kalman filter for filtering calculation to obtain an estimated value of the state error of the integrated navigation system;
and S8, correcting the error of the information obtained in the step S5 in real time according to the estimated value of the state error of the integrated navigation system, and obtaining the high-precision attitude speed position information of the integrated navigation system.
The specific method of step S4 is:
according to the formula:
obtaining a transformation matrix
Where ω is the angular velocity of rotation of the rotary mechanism and t is the time of rotation of the rotary mechanism.
The specific method of step S6 is:
s6-1, obtaining a vector of the fixed star in an image space coordinate system through star map centroid subdivision positioning, obtaining a vector of the fixed star in a geocentric inertial coordinate system through star map matching identification, and obtaining attitude information of the combined navigation system relative to the geocentric inertial system through least square calculation;
s6-2, according to the formula:
sin(h star )＝sin(L)sin(δA)+cos(L)cos(δA)cos(λ+t g )
obtaining elevation angle information of stars, wherein h star Is star height angle, delta A is star declination, lambda is longitude, L is latitude, t g Greenwich mean time;
s6-3, according to the formula:
acquiring attitude information: attitude combination measurement value Z φ Attitude measurement matrix H φ Altitude information of stars: fixed star altitude angle measurement value Z E Height angle measurement matrix H E Altitude information of the aircraft: radar altimeter measurement Z h And a height measurement matrix H h ；Z k Measuring values of the integrated navigation system, including attitude integrated measuring values, star altitude angle measuring values and radar altimeter measuring values; v φ 、V E 、V h All are the measurement noise of the sensor; x is a state variable, wherein the state variable comprises three-axis attitude misalignment angle information of the integrated navigation system, three-axis speed error of the integrated navigation system, position error of the integrated navigation system, constant deviation of a gyroscope and installation error of a star sensor and an inertial device; δ λ is the longitude error of the aircraft; δ L is the latitude error of the aircraft; z dcm For constructed measurements, φ is three of the integrated navigation systemThe shaft attitude misalignment angle information, δ p is the position error of the integrated navigation system, which includes longitude, latitude and altitude,
is the attitude matrix, V, of the aircraft relative to the integrated navigation system s For angle error, (. times.and [. times.]X is an antisymmetric matrix, L is latitude, λ is longitude, and h is altitude; i is an identity matrix, C δP Is a position matrix, α i To identify the sidereal right channel, beta i Identified star declination;
s6-4, according to the formula:
obtaining a state transition matrix Φ:
wherein:
f is the oblateness of the earth, R e Is the semi-major axis, omega, of the earth ie Is the rotational angular velocity, v, of the earth E East speed, v N Is the north velocity; f 0 、F 1 、F 2 、F 3 、F 4 、R M And R N Are all intermediate parameters; deltav is the three-axis velocity error of the integrated navigation system, epsilon is the constant deviation of the gyroscope,
for constant deviation of accelerometerMu is the installation error of the star sensor and the inertial device,
differentiating the three-axis attitude misalignment angle information of the integrated navigation system,
for the three-axis velocity error differential of the integrated navigation system,
for the position error differential of the integrated navigation system, which includes longitude, latitude and altitude,
is the constant offset derivative of the gyroscope,
for the accelerometer constant offset differential to be,
is the differential of the installation error of the star sensor and the inertial device,
is an antisymmetric matrix of the angular velocity of the navigational coordinate system relative to the earth's center inertial system,
antisymmetric matrix of specific force information measured for an accelerometer, v n X is the anti-symmetric matrix of the velocity,
is the angular velocity of the earth's rotation,
for the angular velocity, w, of the navigation coordinate system relative to the terrestrial coordinate system g For random drift of the gyroscope, w a Is the random drift of the accelerometer.
The specific method of step S7 is:
s7-1, according to the formula:
obtaining a state one-step prediction value
Wherein phi k/k-1 For the state transition matrix at the last moment in time,
the state optimal estimated value at the last moment is obtained;
s7-2, according to the formula:
obtaining a one-step prediction error variance P k/k-1 (ii) a Wherein P is k-1 Is the error variance matrix of the previous moment, (. DEG) T Being a transpose of a matrix, Γ k-1 For the system noise interference matrix, Q k-1 Is a process noise matrix;
s7-3, according to the formula:
obtaining a filter gain matrix K k (ii) a Wherein H k Is a measurement matrix having a value of
R k Measuring a noise variance matrix;
s7-4, according to the formula:
obtaining the optimal state error estimation of the integrated navigation system at the moment
In one embodiment of the present invention, the comparison result between the digital simulation results of the conventional inertial/astronomical integrated navigation system and the present invention is:
table 1 initial conditions settings
TABLE 2 inertial navigation error injection
As shown in fig. 3, 4 and 5, in order to compare the simulation results, the conventional inertial/astronomical combined navigation and the rotation inertial astronomical combined navigation result based on the attitude and sidereal altitude fusion in the invention are drawn together.
Compared with the traditional inertial astronomical combined navigation system, the latitude error of the rotary inertial astronomical combined navigation based on the fusion of the attitude angle and the star altitude angle is obviously reduced from 1570m to 355.6m by about 77%, and the horizontal position error of the aircraft is reduced from 1654m to 756.7m by about 54%. Compared with two groups of simulation results, the improved single-axis rotation inertia astronomical combined navigation precision is obviously superior to the traditional inertia astronomical combined navigation precision, the fluctuation is small, and the integral performance of navigation is good.
The invention adds the observation of the star height angle on the basis of the traditional inertial astronomical combined navigation, delays the error of an inertial device by using a rotation technology, and can improve the navigation precision while enhancing the autonomy.
Claims (4)
1. A rotation inertia astronomical combined navigation method based on attitude and star altitude angle fusion is characterized by comprising the following steps:
s1, overlapping and installing the inertial device and the star sensor, and fixedly connecting the inertial device and the star sensor with the rotating mechanism to form a combined navigation system; the carrier where the integrated navigation system is located comprises an aircraft; the pointing direction of the star sensor is the positive direction of the z axis of the rotating mechanism;
s2, the inertial device and the star sensor are driven to rotate according to a single-shaft four-position rotation scheme by a motor of the rotating mechanism;
s3, measuring attitude angular velocity and specific force information of the integrated navigation system through an inertial device;
s4, acquiring and obtaining a transformation matrix from the rotating coordinate system to the aircraft coordinate system according to the rotating angular velocity of the integrated navigation system;
s5, combining the attitude angular velocity and specific force information measured by the inertial device with the conversion matrix to calculate a strapdown inertial navigation algorithm, and obtaining attitude, velocity and position information of the aircraft under the rotation condition of the rotating mechanism;
s6, when the rotating mechanism is static, attitude information of the aircraft relative to the geocentric inertial system is obtained through the star sensor; calculating and solving altitude angle information of the fixed star by a spherical triangle method; measuring to obtain the altitude information of the aircraft through a radar altimeter; acquiring a state transition matrix of a state error equation of the integrated navigation system;
s7, sending the information obtained in the step S6 to a single-axis rotation strapdown inertia/astronomical integrated navigation Kalman filter for filtering calculation to obtain an estimated value of the state error of the integrated navigation system;
and S8, correcting the error of the information obtained in the step S5 in real time according to the estimated value of the state error of the integrated navigation system, and obtaining the high-precision attitude speed position information of the integrated navigation system.
2. The method for rotational inertial astronomical combined navigation based on attitude and sidereal altitude fusion according to claim 1, wherein the specific method of step S4 is as follows:
according to the formula:
obtaining a transformation matrix
Where ω is the angular velocity of rotation of the rotary mechanism and t is the time of rotation of the rotary mechanism.
3. The method for rotational inertial astronomical combined navigation based on attitude and sidereal altitude fusion according to claim 1, wherein the specific method of step S6 is as follows:
s6-1, obtaining a vector of a fixed star in an image space coordinate system through star map centroid subdivision positioning, obtaining a vector of the fixed star in a geocentric inertial coordinate system through star map matching identification, and obtaining attitude information of the combined navigation system relative to the geocentric inertial system through calculation by using a least square method;
s6-2, according to the formula:
sin(h star )＝sin(L)sin(δA)+cos(L)cos(δA)cos(λ+t g )
obtaining elevation angle information of stars, wherein h star Is star height angle, delta A is star declination, lambda is longitude, L is latitude, t g Greenwich mean time;
s6-3, according to the formula:
acquiring attitude information: attitude combination measurement value Z φ Attitude measurement matrix H φ Altitude information of stars: fixed star altitude angle measurement value Z E Altitude angle measurement matrix H E Altitude information of the aircraft: radar altimeter measurement Z h And a height measurement matrix H h ；Z k Measuring values of the integrated navigation system comprise attitude integrated measuring values, star altitude angle measuring values and radar altimeter measuring values; v φ 、V E 、V h All are the measurement noise of the sensor; x is a state variable, wherein the state variable comprises three-axis attitude misalignment angle information of the integrated navigation system, three-axis speed error of the integrated navigation system, position error of the integrated navigation system, constant value deviation of a gyroscope and installation error of a star sensor and an inertial device; δ λ is the longitude error of the aircraft; δ L is the latitude error of the aircraft; z dcm Phi is the three-axis attitude misalignment angle information of the integrated navigation system, δ p is the position error of the integrated navigation system, including longitude, latitude and altitude,
is the attitude matrix, V, of the aircraft relative to the integrated navigation system s For angle error, (. times.and [. times.]X is an antisymmetric matrix, L is latitude, λ is longitude, and h is altitude; i is an identity matrix, C δP Is a position matrix, α i To identify the sidereal right channel, beta i Identified sidereal declination;
s6-4, according to the formula:
obtaining a state transition matrix Φ:
wherein:
f is the oblateness of the earth, R e Is the semi-major axis of the earth, omega ie Is the rotational angular velocity, v, of the earth E East speed, v N Is the north speed; f 0 、F 1 、F 2 、F 3 、F 4 、R M And R N Are all intermediate parameters; deltav is the three-axis velocity error of the integrated navigation system, epsilon is the constant deviation of the gyroscope,
is the constant deviation of the accelerometer, mu is the installation error of the star sensor and the inertial device,
for the three-axis attitude misalignment angle information differentiation of the integrated navigation system,
for the three-axis velocity error differential of the integrated navigation system,
for the position error differential of the integrated navigation system, which includes longitude, latitude and altitude,
is a constant offset derivative of the gyroscope,
for the accelerometer constant offset differential to be,
is the differential of the installation error of the star sensor and the inertial device,
is an antisymmetric matrix of the angular velocity of the navigational coordinate system relative to the earth's center inertial system,
antisymmetric matrix of specific force information measured for an accelerometer, v n X is the anti-symmetric matrix of the velocity,
is the angular velocity of the earth's rotation,
for the angular velocity, w, of the navigation coordinate system relative to the terrestrial coordinate system g For random drift of the gyroscope, w a Is the random drift of the accelerometer.
4. The method for rotational inertial astronomical combined navigation based on attitude and sidereal altitude fusion according to claim 3, wherein the specific method of step S7 is as follows:
s7-1, according to the formula:
obtaining a state one-step prediction value
Wherein phi k/k-1 For the state transition matrix at the last moment in time,
the state optimal estimated value at the last moment is obtained;
s7-2, according to the formula:
obtaining a one-step prediction error variance P k/k-1 (ii) a Wherein P is k-1 Is the error variance matrix of the previous moment, (. DEG) T Being a transpose of a matrix, Γ k-1 For the system noise interference matrix, Q k-1 Is a process noise matrix;
s7-3, according to the formula:
obtaining a filter gain matrix K k (ii) a Wherein H k Is a measurement matrix having a value of
R k Measuring a noise variance matrix;
s7-4, according to the formula:
obtaining the optimal state error estimation of the integrated navigation system at the moment
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110585391.8A CN113218390B (en) | 2021-05-27 | 2021-05-27 | Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110585391.8A CN113218390B (en) | 2021-05-27 | 2021-05-27 | Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113218390A CN113218390A (en) | 2021-08-06 |
| CN113218390B true CN113218390B (en) | 2022-09-27 |
Family
ID=77098863
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110585391.8A Active CN113218390B (en) | 2021-05-27 | 2021-05-27 | Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113218390B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116952206B (en) * | 2023-09-19 | 2024-02-02 | 中国科学院光电技术研究所 | Space target detection method based on starlight attitude measurement |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1118663A (en) * | 1965-06-03 | 1968-07-03 | North American Aviation Inc | Inertial navigation system error correcting methods |
| CN101788296A (en) * | 2010-01-26 | 2010-07-28 | 北京航空航天大学 | SINS/CNS deep integrated navigation system and realization method thereof |
| CN104165640A (en) * | 2014-08-11 | 2014-11-26 | 东南大学 | Near-space missile-borne strap-down inertial navigation system transfer alignment method based on star sensor |
| CN108871326A (en) * | 2018-07-09 | 2018-11-23 | 北京航空航天大学 | A kind of single-shaft-rotation modulation inertia-astronomy deep integrated navigation method |
| CN111121766A (en) * | 2019-12-17 | 2020-05-08 | 南京理工大学 | Astronomical and inertial integrated navigation method based on starlight vector |
| CN111964671A (en) * | 2020-07-14 | 2020-11-20 | 西北工业大学 | Inertial astronomical integrated navigation system and method based on double-axis rotation modulation |
| CN112577519A (en) * | 2020-11-23 | 2021-03-30 | 南京航空航天大学 | Aerospace vehicle star sensor installation error online calibration method |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110986934B (en) * | 2019-12-12 | 2022-10-18 | 北京自动化控制设备研究所 | Navigation method and system of integrated double-shaft rotation inertial navigation astronomical integrated navigation system |
-
2021
- 2021-05-27 CN CN202110585391.8A patent/CN113218390B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1118663A (en) * | 1965-06-03 | 1968-07-03 | North American Aviation Inc | Inertial navigation system error correcting methods |
| CN101788296A (en) * | 2010-01-26 | 2010-07-28 | 北京航空航天大学 | SINS/CNS deep integrated navigation system and realization method thereof |
| CN104165640A (en) * | 2014-08-11 | 2014-11-26 | 东南大学 | Near-space missile-borne strap-down inertial navigation system transfer alignment method based on star sensor |
| CN108871326A (en) * | 2018-07-09 | 2018-11-23 | 北京航空航天大学 | A kind of single-shaft-rotation modulation inertia-astronomy deep integrated navigation method |
| CN111121766A (en) * | 2019-12-17 | 2020-05-08 | 南京理工大学 | Astronomical and inertial integrated navigation method based on starlight vector |
| CN111964671A (en) * | 2020-07-14 | 2020-11-20 | 西北工业大学 | Inertial astronomical integrated navigation system and method based on double-axis rotation modulation |
| CN112577519A (en) * | 2020-11-23 | 2021-03-30 | 南京航空航天大学 | Aerospace vehicle star sensor installation error online calibration method |
Non-Patent Citations (1)
| Title |
|---|
| 孙建业等.基于旋转式惯性/天文组合导航系统信息融合算法研究.《战术导弹技术》.2019,(第4期),第55-61页. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113218390A (en) | 2021-08-06 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN109459044B (en) | GNSS dual-antenna assisted vehicle-mounted MEMS inertial navigation combined navigation method | |
| CN110764119B (en) | Satellite antenna autonomous measurement and control method and system independent of satellite navigation information | |
| CN108508918B (en) | High-precision real-time ground pointing control method for data transmission antenna of static orbit remote sensing satellite | |
| CN108731674B (en) | Inertial astronomical integrated navigation system based on single-axis rotation modulation and calculation method | |
| CN111121766B (en) | Astronomical and inertial integrated navigation method based on starlight vector | |
| CN108088443B (en) | Speed compensation method for positioning and orienting equipment | |
| CN105716615A (en) | Method for autonomously navigating Mars probes | |
| CN110398242B (en) | Attitude angle determination method for high-rotation-height overload condition aircraft | |
| CN112880669B (en) | Spacecraft starlight refraction and single-axis rotation modulation inertial integrated navigation method | |
| CN113108783B (en) | inertial/Doppler combined navigation method for unmanned underwater vehicle | |
| CN113503892B (en) | Inertial navigation system moving base initial alignment method based on odometer and retrospective navigation | |
| CN103604428A (en) | Star sensor positioning method based on high-precision horizon reference | |
| CN110986934B (en) | Navigation method and system of integrated double-shaft rotation inertial navigation astronomical integrated navigation system | |
| CN113218390B (en) | Rotation inertia astronomy combined navigation method based on attitude and star altitude angle fusion | |
| JP2002158525A (en) | Antenna controller for satellite tracking | |
| CN116105730A (en) | Angle measurement-only optical combination navigation method based on cooperative target satellite very short arc observation | |
| CN112833878B (en) | Near-ground multi-source astronomical autonomous navigation method | |
| CN107807375B (en) | Unmanned aerial vehicle attitude tracking method and system based on multiple GPS receivers | |
| CN112525204A (en) | Spacecraft inertia and solar Doppler velocity combined navigation method | |
| CN110388942B (en) | Vehicle-mounted posture fine alignment system based on angle and speed increment | |
| CN115542363B (en) | Attitude measurement method suitable for vertical downward-looking aviation pod | |
| CN116222551A (en) | Underwater navigation method and device integrating multiple data | |
| CN111964671B (en) | Inertial astronomical integrated navigation system and method based on double-axis rotation modulation | |
| CN109099911B (en) | Navigation positioning method and system for aviation system | |
| CN113985464A (en) | Zero calibration method and system for vehicle-mounted satellite communication-in-motion |
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 | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |