CN102135621A

CN102135621A - Fault recognition method for multi-constellation integrated navigation system

Info

Publication number: CN102135621A
Application number: CN2010106172501A
Authority: CN
Inventors: 陆伟宁; 王千喜; 翟羽佳; 杨晓昆; 胡强; 李秋凤; 刘岩
Original assignee: CHINA AEROSPACE SCIENCE & INDUSTRY ACADEMY OF INFORMATION TECHNOLOGY
Current assignee: CHINA AEROSPACE SCIENCE & INDUSTRY ACADEMY OF INFORMATION TECHNOLOGY
Priority date: 2010-12-31
Filing date: 2010-12-31
Publication date: 2011-07-27
Anticipated expiration: 2030-12-31
Also published as: CN102135621B

Abstract

The invention discloses a fault recognition method for a multi-constellation integrated navigation system. The method comprises the following steps of: unifying time and space coordinates of the multi-constellation system; determining fault detection usability; determining a fault; recognizing and removing faulty satellites and the like. Faults of various satellites can be monitored and recognized simultaneously, the success rate of recognizing and removing the faulty satellites is greatly improved due to unification of the time and space coordinates of the multi-constellation system and repeated detection and judgment, and the navigation positioning accuracy, positioning performance and reliability of the multi-constellation integrated navigation system are improved; therefore, the navigation service performance is improved.

Description

Fault identification method of multi-constellation combined navigation system

Technical Field

The invention relates to a fault identification method of a navigation system, in particular to a fault identification method of a multi-constellation combined navigation system.

Background

The GPS (global navigation satellite system) has been widely used in the world, and in the GPS, a multi-constellation combined navigation system has high navigation positioning accuracy and good positioning performance and reliability, so that the navigation service provided by the multi-constellation combined navigation system can reach the level that the enhancement system must be used before.

However, when the error of the multi-constellation combined navigation system exceeds the allowable limit due to various reasons such as a failed satellite or a long-time false lock, the multi-constellation combined navigation system cannot be qualified for navigation work. At present, the positioning of the satellite navigation system is failed mainly due to the following reasons:

1) the positioning accuracy of satellite navigation is influenced by the number of satellites and the geometric distribution of the satellites, and the performance of a GPS system is deteriorated in areas with a small number of satellites and poor positioning geometric distribution;

2) the satellite navigation system is large and complex, and the software and hardware faults of the system can also increase the satellite navigation positioning error, so that the flight safety of the airplane is influenced;

3) the owning country of the GPS system adopts some measures for limiting the positioning precision for the safety benefit of the own country, and is difficult to ensure that the owning country does not adopt similar measures for the military benefit under the condition of military conflict and war at present;

4) electromagnetic waves in the external environment, ionosphere changes, natural interference and man-made interference, particularly hostile interference, can also affect the reliability of satellite navigation.

The fault identification and integrity enhancement mainly adopts monitoring technology to provide real-time alarm capability. Methods for achieving system integrity can be divided into two categories, one being internal and the other being external. The internal method uses receiver internal sensor information to achieve integrity monitoring, RAIM.

The invention patent application with the application number of CN 200610165465.8 discloses a GNSS receiver autonomous integrity monitoring method based on multi-satellite fault recognition, which comprises the following steps: a. performing an availability analysis of autonomous integrity; b. and (3) judging single-star and multi-star faults: comparing the detection statistic with a single-star or multi-star detection threshold, if the detection statistic does not exceed any threshold, indicating that no fault exists at present, and continuing to monitor; if the threshold is exceeded, determining to enter a corresponding fault identification step; c. single-star and multi-star fault identification is carried out: carrying out fault identification on single-satellite time by using a characteristic deviation line method, and carrying out fault identification on multi-satellite time by using a hypothesis verification method; d. and (4) carrying out exclusion verification: removing the fault satellite from the selected satellite combination, and repeating the step a and the step b; if no fault is found, indicating that the step c is correct and the fault satellite is eliminated; if a new fault is found, it indicates that step c fails, and the specific conditions need to be analyzed: if the single-satellite fault is failed to be eliminated, the fault is probably a multi-satellite fault; and if the multi-satellite fault removal fails, determining that the measured data at the current moment cannot complete integrity monitoring. The GNSS receiver autonomous integrity monitoring method based on multi-satellite fault identification has the serious defect of low success rate of identifying and removing the fault satellite.

Disclosure of Invention

The invention aims to overcome the defects in the prior art and provide a fault identification method of a multi-constellation combined navigation system, which has reliable performance and high success rate of identifying and eliminating fault satellites.

In order to solve the problems in the prior art, the invention discloses a fault identification method of a multi-constellation combined navigation system, which comprises the following steps:

1) unification of time and space coordinates for multi-constellation systems

1.1) unification in a multi-constellation system: according to a fixed conversion relation between the system time of each constellation and UTC (universal time), roughly resolving the system times of different constellations, and unifying the system times of all the constellation systems; clock difference between system time of each constellation and UTC

The following pseudo-range observation equation gives:

wherein, the superscript g represents a system serial number, i is an observation serial number (i is 1, 2, 3, 4, 5); (X, y, z) is the coordinates of the receiver in the selected coordinate system, (X)^g，Y^g，Z^g) Converting the observation satellites in each constellation to coordinates under a selected coordinate system by using a coordinate conversion formula;

is the clock difference (r ═ 1, 2.. N) between the receiver and each constellation,

is a pseudo range observation value of each visible satellite, and c is the speed of light;

for N constellations, the receiver has N +3 unknowns to be measured, and needs N +3 pseudo-range observation equations to solve to obtain the clock error between the system time and the UTC of each constellation

1.2) unifying the spatial coordinates of multiple constellations

The unification of the space coordinates is completed by adopting the following coordinate conversion formula:

wherein (Δ x, Δ y, Δ z) is the earth center offset, θ_x，θ_y，θ_zRotation angle of coordinate axis, m is scale factor, (x, y, z)_sysTo be the coordinates in the coordinate system to be transformed, (x, y, z)_sys1Coordinates of a target coordinate system;

2) fault detection availability determination

2.1) judging the number of the visible satellites, if the number of the visible satellites is less than N +3, indicating that fault detection cannot be carried out, carrying out integrity alarm by the system, and otherwise, continuing;

2.2) availability determination:

firstly, a fault detection threshold value sigma is calculated_TThe formula is as follows:

wherein σ₀Is the variance value of the pseudo-range measurement error; n is the number of visible satellites; the threshold value T is determined by:

<math><mrow><mi>Pr</mi><mrow><mo>(</mo><mi>SSE</mi><mo>/</mo><msubsup><mi>σ</mi><mn>0</mn><mn>2</mn></msubsup><mo><</mo><msup><mi>T</mi><mn>2</mn></msup><mo>)</mo></mrow><mo>=</mo><msubsup><mo>&Integral;</mo><mn>0</mn><msup><mi>T</mi><mn>2</mn></msup></msubsup><msub><mi>f</mi><msubsup><mi>χ</mi><mrow><mo>(</mo><mi>n</mi><mo>-</mo><mn>4</mn><mo>,</mo><mi>λ</mi><mo>)</mo></mrow><mn>2</mn></msubsup></msub><mrow><mo>(</mo><mi>x</mi><mo>)</mo></mrow><mi>dx</mi><mo>=</mo><mn>1</mn><mo>-</mo><msub><mi>P</mi><mi>FA</mi></msub></mrow></math>

wherein, P_FATolerable false alarm probabilities;

(x) Is a chi with a degree of freedom of n-4²A probability density function of the distribution;

statistical summary of residual statistical sum of squares less than threshold TRate;

in the formula:

v＝(I-G(G^TWG)^-1G^TW)ε

g is a linearization matrix formed by cosine vectors from each satellite to the receiver, epsilon is a pseudo-range error vector, W is an n multiplied by n dimensional observation pseudo-range weight matrix, and I is a unit matrix;

suppose that the ith satellite has a fault with a bias of b_iSSE obeys a decentralization of x with a degree of freedom of n-4²The distribution, non-center parameter λ, can be obtained by:

<math><mrow><mi>λ</mi><mo>=</mo><mi>E</mi><mrow><mo>(</mo><msup><mi>v</mi><mi>T</mi></msup><mi>Wv</mi><mo>)</mo></mrow><mo>/</mo><msubsup><mi>σ</mi><mn>0</mn><mn>2</mn></msubsup><mo>=</mo><msubsup><mi>RPE</mi><mi>i</mi><mn>2</mn></msubsup><mo>/</mo><msubsup><mi>σ</mi><mn>0</mn><mn>2</mn></msubsup><msubsup><mi>δHDOP</mi><mi>i</mi><mn>2</mn></msubsup></mrow></math>

wherein,

{RPE}_{i} = (A_{1 i}^{2} + A_{2 i}^{2}) W_{ii}^{2} b_{i}

A＝(G^TWG)^-1G^T

Q_v＝W^-1-G(G^TWG)^-1G^T

HDOP denotes the horizontal positioning accuracy factor, HDOP, of all satellites in view_iRepresenting the horizontal positioning accuracy factor after the ith satellite is removed;

calculating delta HDOP_max；

Before fault detection, HDOP corresponding to each satellite is calculated in real time_iRemoving and taking the maximum value of delta HDOP_max；

Calculating a horizontal positioning error protection limit value HPL, wherein the formula is as follows;

<math><mrow><mi>HPL</mi><mo>=</mo><msub><mi>δHDOP</mi><mi>max</mi></msub><mo>×</mo><msub><mi>σ</mi><mn>0</mn></msub><mo>×</mo><msqrt><mi>λ</mi></msqrt></mrow></math>

comparing the HPL with a horizontal error protection limit value HAL, if the HPL is out of limit, the fault detection is unavailable, and the system gives an integrity alarm, otherwise, the system continues to give an alarm;

3) fault determination

By error of actual observed pseudo-rangeVariance (variance)

With the fault detection threshold value sigma obtained in step 2)_TCompare if, if

If the fault is detected, the work is continued, otherwise, the work is finished;

4) identifying and troubleshooting satellites

4.1) judging the number of the visible satellites, if the number of the visible satellites is less than N +4, fault identification cannot be carried out, and the system carries out integrity alarm, otherwise, the system continues to carry out;

4.2) identifying a fault satellite: the identification of the fault satellite adopts a characteristic line deviation method, firstly, an observation coefficient matrix G is subjected to QR decomposition to obtain a matrix Q^T：

Q^{T} = [\begin{matrix} Q_{X} \\ Q_{P} \end{matrix}]

Wherein Q is_XIs Q^TFirst 4 rows, Q_PIs Q^TThe remaining n-4 rows;

the parity residual vector p is p ═ Q_Py＝Q_P(Gx+ε)＝Q_Pε

Calculating the characteristic deviation line K of each satellite_cha：

K_{cha} = \frac{Q_{p} (1, i)}{Q_{p} (2, i)}

Computing a characteristic deviation slope K for an odd-even space vector_p：

K_p＝p₁/p₂

Wherein p is₁And p₂Is an element of the parity vector p if K of the ith satellite_chaAnd K_pVery close, the ith satellite is identified as the failed satellite;

4.3) troubleshooting the satellite: removing the fault satellite identified in the substep 4.2);

5) repeating the steps 1) to 4);

6) and (3) repeating the step 2) and the step 3), if no fault exists, the satellite with the fault is successfully eliminated, the work is finished, otherwise, the measured data cannot be tested for the self-integrity, and the system gives an integrity alarm.

The fault identification method of the multi-constellation combined navigation system can monitor and identify faults of a plurality of satellites at the same time, greatly improves the success rate of identifying and eliminating the fault satellites through unified and multiple detection and judgment of time and space coordinates of the multi-constellation system, and enhances the navigation positioning precision, positioning performance and reliability of the multi-constellation combined navigation system, thereby improving the navigation service performance.

Drawings

Fig. 1 is a general flowchart of a fault identification method of a multi-constellation combined navigation system according to the present invention.

Detailed Description

FIG. 1 is a general flowchart of a fault identification method of a multi-constellation combined navigation system according to the present invention

As shown in fig. 1, the method comprises the steps of:

step S1, unifying the time and space coordinates of the multi-constellation system, and unifying the time and space coordinates of the multi-constellation system for the observation data from a plurality of different constellations, so as to ensure the accuracy of the system in resolving the data and ensure the success rate of identifying and removing the fault satellite. The method comprises the following two sub-steps:

and step S1.1, unifying the multi-constellation system, roughly resolving the system time of different constellations according to the fixed conversion relation between the system time of each constellation and the UTC, and unifying the system time of all the constellation systems.

Clock difference between system time of each constellation and UTC

The following pseudo-range observation equation gives:

is the clock difference (r ═ 1, 2.. N) between the receiver and each constellation,is a pseudo range observation value of each visible satellite, and c is the speed of light;

And step S1.2, unifying the space coordinates of multiple constellations, wherein a unified coordinate system is needed for simultaneously processing the observation data of different constellations. Coordinates of all constellations are unified under a coordinate system, and coordinate conversion is needed, in the embodiment, the coordinate conversion adopts a classical burst-Wolf model, and the formula is as follows:

wherein (Δ x, Δ y, Δ z) is the earth center offset, θ_x，θ_y，θ_zRotation angle of coordinate axis, m is scale factor, (x, y, z)_sysTo be the coordinates in the coordinate system to be transformed, (x, y, z)_sys1Are the coordinates of the target coordinate system.

Through the step S1, all the observation data can be unified to the same time and space coordinate system, thereby shielding the multiple constellation differences.

Step S2, determining availability for fault detection, wherein the determination of availability for fault detection before fault determination is required after the system completion and the spatial coordinates unification of step S1, and the method includes the following two sub-steps:

step S2.1, firstly, judging the number of the visible satellites, if the number of the visible satellites is less than N +3, indicating that fault detection cannot be carried out, carrying out integrity alarm by the system, and if the number of the visible satellites is more than or equal to N +3, continuously executing step S2.2;

step S2.2, availability determination, failure detection availability determination, first of all, the failure detection threshold value sigma is found_TThe formula is as follows:

wherein, P_FATolerable false alarm probabilities;

is a chi with a degree of freedom of n-4²A probability density function of the distribution;

is the statistical probability that the sum of the squared residuals is less than the threshold value T.

In the formula:

v＝(I-G(G^TWG)^-1G^TW)ε

wherein G is a linearization matrix formed by cosine vectors of directions from each satellite to the receiver, epsilon is a pseudo-range error vector, W is an n multiplied by n dimensional observation pseudo-range weight matrix, and I is a unit matrix.

wherein,

{RPE}_{i} = (A_{1 i}^{2} + A_{2 i}^{2}) W_{ii}^{2} b_{i}

A＝(G^TWG)^-1G^T

Q_v＝W^-1-G(G^TWG)^-1G^T

HDOP denotes the horizontal positioning accuracy factor, HDOP, of all satellites in view_iIndicating the horizontal positioning accuracy factor after the ith satellite is removed.

Calculating delta HDOP_max；

Before fault detection, HDOP corresponding to each satellite is calculated in real time_iRemoving and taking the maximum value of delta HDOP_max。

the horizontal positioning error protection limit HPL is compared with the horizontal error protection limit HAL, and if the limit is exceeded, the fault detection is not available and the system gives an integrity alarm, otherwise it continues with step S3.

Step S3, failure determination, variance of error of actual observed pseudo-range

σ is performed with the failure detection threshold value obtained in step S2_TCompare if, if

It indicates that a fault is detected and execution continues at step S4, otherwise the operation ends.

Step S4, identifying and eliminating the failed satellite, when the existence of the failed satellite is detected in step S3, the identification and elimination of the failed satellite are carried out, and the method comprises the following 3 substeps:

and S4.1, firstly, judging the number of the visible satellites, if the number of the visible satellites is less than N +4, failure identification cannot be carried out, carrying out integrity alarm by the system, and if the number of the visible satellites is more than or equal to N +4, continuously executing the step S4.2.

Step S4.2, identifying the fault satellite by using a characteristic line deviation method, and firstly, carrying out QR decomposition on an observation coefficient matrix G to obtain a matrix Q^T：

Q^{T} = [\begin{matrix} Q_{X} \\ Q_{P} \end{matrix}]

Wherein Q is_XIs Q^TFirst 4 rows, Q_PIs Q^TThe remaining n-4 rows;

the parity residual vector p is p ═ Q_Py＝Q_P(Gx+ε)＝Q_Pε

Each time of calculationCharacteristic deviation line K of particle satellite_cha：

K_{cha} = \frac{Q_{p} (1, i)}{Q_{p} (2, i)}

Computing a characteristic deviation slope K for an odd-even space vector_p：

K_p＝p₁/p₂

Wherein p is₁And p₂Is an element of the parity vector p if K of the ith satellite_chaAnd K_pVery close, the ith satellite is identified as the failed satellite.

And (4.3) removing the fault satellite: the faulty satellite identified in substep S4.2 is rejected.

In order to ensure the success rate of identifying and eliminating the fault satellite of the fault identification method of the multi-constellation combined navigation system, multiple identifications and judgments are needed for eliminating multiple fault satellites.

And step S5, repeating the steps S1 to S4, and identifying and eliminating the fault satellite still existing. The step S1 is repeated to prevent clock difference between the system time of each constellation and UTC

And when new changes exist, unifying the re-read observation data of each constellation to a new same time and space coordinate system, and realizing shielding of the difference of the plurality of constellations again.

Repeating the step S2 and the step S3, if the judgment result shows that no fault exists, the fault star is proved to be successfully eliminated, and the work is finished; if there is still a fault, step S4 is continued to identify and eliminate the faulty satellite.

And S6, repeating the steps S2 and S3, if no fault is judged, proving that the fault satellite is successfully eliminated and the work is finished, otherwise, the fault satellite cannot be eliminated again due to the limitation of the prior art, so that the measured data cannot be subjected to the autonomous integrity test, and the system gives an integrity alarm.

In summary, the embodiments of the present invention disclose preferred embodiments thereof, but are not limited thereto. Those skilled in the art can easily appreciate the spirit of the present invention from the above-mentioned embodiments, and make various extensions and changes, which are within the protection scope of the present invention, without departing from the spirit of the present invention.

Claims

1. A fault identification method of a multi-constellation combined navigation system is characterized by comprising the following steps:

1) unification of time and space coordinates for multi-constellation systems

1.1) unification in a multi-constellation system: according to the fixed conversion relation between the system time of each constellation and the UTC, roughly resolving the system times of different constellations, and unifying the system times of all the constellation systems;

clock difference between system time of each constellation and UTC

The following pseudo-range observation equation gives:

wherein, the superscript g represents a system serial number, i is an observation serial number (i is 1, 2, 3, 4, 5); (X, y, z) is the coordinates of the receiver in the selected coordinate system, (X)^g，Y^g，Z^g) For transforming the observation satellites in each constellation to the selected coordinate system by using a coordinate transformation formulaCoordinates;