INS / GPS tightly coupled integrated navigation enhancement method based on receiver clock bias

Abstract

The application is based on a receiver clock error INS / GPS tight coupling combined navigation enhancement method, which comprises the following steps: constructing a GPS receiver clock error prediction model; constructing a combined navigation system Kalman filter state vector with a strapdown inertial navigation system navigation solution error and calculating a strapdown inertial navigation system estimated pseudo-range and pseudo-range rate vector, a GPS receiver pseudo-range and pseudo-range rate, a GPS receiver clock error and clock drift calculation value or estimated value; when the number of visible satellites is greater than or equal to 4, when |ΔT|>T threshold , using the GPS receiver clock error and clock drift prediction value instead of the calculation value, otherwise using the calculation value to participate in the combined navigation Kalman filter operation; when the number of visible satellites is less than 4, directly using the GPS receiver clock error and clock drift prediction value to participate in the combined navigation Kalman filter operation; using the calculated INS navigation solution error correction value to correct the inertial navigation system error.

Description

Technical Field

[0001] This invention patent belongs to the field of INS / GPS integrated navigation, and specifically designs an enhancement method for tightly coupled INS / GPS integrated navigation based on receiver clock bias. Background Technology

[0002] INS / GPS tightly coupled integrated navigation is a conventional integrated navigation method. It involves constructing an integrated navigation model and using pseudorange and pseudorange rate measurements from navigation satellites obtained from a GPS receiver to calibrate the position, velocity, and attitude solutions calculated by the inertial navigation system, thereby obtaining a high-precision and high-update-rate integrated navigation solution. However, in real-world environments, GPS navigation satellite signals are easily blocked or interfered with, leading to significant errors in pseudorange and pseudorange rate measurements obtained by end users, or insufficient availability of satellites. This significantly reduces the reliability of conventional INS / GPS tightly coupled integrated navigation systems and the accuracy of the integrated navigation solution. Summary of the Invention

[0003] The purpose of this invention is to address the shortcomings of conventional INS / GPS tightly coupled integrated navigation systems in the prior art, where GPS navigation satellite signals are easily blocked or interfered with, resulting in large errors in the pseudorange and pseudorange rate measurements obtained by end users, or insufficient available satellite resources. The invention provides an INS / GPS tightly coupled integrated navigation enhancement method based on receiver clock bias.

[0004] In order to achieve the purpose of this invention, the following technical solution is adopted:

[0005] The present invention provides a receiver clock bias-based INS / GPS tightly coupled integrated navigation enhancement method, comprising the following steps:

[0006] (1) The GPS receiver receives satellite navigation signals and calculates the clock error of the GPS receiver over 60 seconds, which is used to construct the GPS receiver clock error model.

[0007] (2) Select at least three sets of GPS receiver clock difference values ​​and corresponding times from the above 60s GPS receiver measurements, and use x(t)=a0+a1t+ε x (t)---Formula (1) is used to construct a linear model of the GPS receiver clock bias for predicting the GPS receiver clock bias, where: a0 is the initial value deviation of the GPS receiver clock; a1 is the clock drift of the GPS receiver; ε x (t) represents the random variation of the GPS receiver clock bias; x(t) represents the predicted GPS receiver clock bias value. a0, a1, and ε are calculated using the least squares linear fitting method. x (t);

[0008] (3) The navigation solution error of the strapdown inertial navigation system (INS) includes: position measurement error vector. Velocity error measurement vector Attitude error measurement error vector Gyroscope error vector δw=[δw x δw y δw z ] and accelerometer error vector δf=[δf x δf y δf z The state vector of the Kalman filter of the integrated navigation system is constructed from the above five vectors. in: δλ is the longitude measurement error, δh is the latitude measurement error, and δv is the elevation measurement error. e For the eastward velocity measurement error, δv n For northbound velocity measurement error, δv u δp is the pitch angle measurement error, δr is the roll angle measurement error, δA is the yaw angle measurement error, and [δw] is the yaw angle measurement error. x δw y δw z [δf] represents the error vectors in the three directions of the gyroscope, respectively. x δf y δf z These are the error vectors in the three directions of the accelerometer;

[0009] (4) Estimate the pseudorange and pseudorange rate vectors using the strapdown inertial navigation system; calculate the pseudorange and pseudorange rate using GPS receiver information; and construct the measurement vectors of the Kalman filter of the integrated navigation system using the calculated or estimated values ​​of GPS receiver clock bias and clock drift. Where, ρ INS and The vectors are n×1 dimensional pseudorange and pseudorange rate, estimated from measurement information obtained by a strapdown inertial navigation system; ρ GPS and The vectors are n×1 dimensional pseudorange and pseudorange rate calculated from GPS measurement information; δt u and δd u These are the calculated or estimated values ​​of GPS receiver clock bias and clock drift, respectively. n is the number of GPS visible satellites in the current epoch; the calculated values ​​of GPS receiver clock bias and clock drift are obtained by formulas (3.138) and (3.139) in reference [1]; the estimated values ​​of GPS receiver clock bias and clock drift are obtained by formula (1) in step (2);

[0010] (5) Number of visible satellites monitored by the GPS receiver

[0011] (a) When the number of visible satellites is greater than or equal to 4

[0012] Calculate the error detection threshold T threshold and the detection sample ΔT, where

[0013] Where σ cal σ is the standard deviation of the calculated GPS receiver clock bias over a 60-second period. pre The standard deviation is the predicted clock bias of a GPS receiver over a 60-second period; the standard deviation is expressed as... The calculation yields x; where x i The calculated or predicted value of the clock difference per second in 60 seconds; The average of the calculated or predicted clock difference over 60 seconds;

[0014] ΔT=δt u(cal) -δt u(pre) , where δt u(cal) Current receiver clock bias calculation value, δt u(pre) The current receiver clock bias prediction value is calculated using formula (1);

[0015] When |ΔT|>T threshold If an error is found, the predicted values ​​of the GPS receiver clock bias and clock drift are used instead of the calculated values ​​to participate in the combined navigation Kalman filter operation in formula (2); otherwise, if no error is found, the calculated values ​​of the receiver clock bias and clock drift are used to participate in the combined navigation Kalman filter operation.

[0016] (b) When the number of visible satellites is less than 4

[0017] When there are fewer than 4 satellites, the predicted values ​​of GPS receiver clock bias and clock drift cannot be calculated. Therefore, the combined navigation Kalman filter operation of formula (2) is directly used.

[0018] (6) Use The error correction value of the INS navigation solution output by the Kalman filter of the integrated navigation system is calculated, where: X is the state vector constructed in step (3); Z is the measurement vector constructed in step (4); K is the Kalman gain; H is the measurement matrix; This is the error correction value for INS navigation.

[0019] (7) Use the INS navigation solution error correction value from step (6). Error correction is performed on the navigation solutions of the inertial navigation system, including position, velocity, and attitude. This is the bit error correction amount. This is the speed error correction amount. δω is the attitude error correction value. corr δf is the gyroscope error correction value. corr The accelerometer error measurement correction is used to correct the position measurement error of the inertial navigation solution through the formula. Make corrections, where r ins r represents the position measurement value obtained by the inertial navigation system. ins / gps The corrected position measurement value; the velocity measurement error of the inertial navigation solution is expressed by the formula... Make corrections, where v ins The velocity measurement value, v, is obtained by the inertial navigation system. ins / gps The corrected velocity measurement value; the attitude measurement error of the inertial navigation solution is expressed by the formula... Make corrections, where ε ins The attitude measurement value ε obtained by the inertial navigation system ins / gps These are the corrected attitude measurements.

[0020] The present invention provides a method for enhancing INS / GPS tightly coupled navigation based on receiver clock bias, wherein: in step (2), five sets of GPS receiver clock bias values ​​and corresponding times are selected.

[0021] The present invention provides a receiver clock bias-based INS / GPS tightly coupled navigation enhancement method, wherein: in step (4), the pseudorange and pseudorange rate vector predicted by the strapdown inertial navigation system can be calculated.

[0022] The present invention provides a method for enhancing INS / GPS tightly coupled navigation based on receiver clock bias, wherein: in step (4), the pseudorange and pseudorange rate can be calculated using GPS receiver information.

[0023] The present invention provides a method for enhancing navigation based on receiver clock bias INS / GPS tightly coupled combination, wherein: in step (4), the calculated values ​​of GPS receiver clock bias and clock drift can be obtained.

[0024] The receiver clock bias-based INS / GPS tightly coupled integrated navigation enhancement method of the present invention has the following beneficial effects;

[0025] Compared to conventional tightly coupled INS / GPS integrated navigation systems, the receiver clock bias-based INS / GPS tightly coupled integrated navigation enhancement system adds GPS receiver clock bias and clock drift values ​​as measurement inputs to the integrated navigation Kalman filter in the integrated navigation model, and predicts clock bias data by constructing a receiver clock bias model. When there is satellite observation redundancy, the system detects and corrects measurement errors of the GPS receiver clock bias, reducing the impact of satellite measurement data errors on the integrated navigation solution. When the number of GPS satellites is insufficient, the predicted values ​​of receiver clock bias and clock drift are used as measurement inputs to the integrated navigation system, reducing the error accumulation rate of inertial navigation measurements and improving the robustness of the integrated navigation system. Attached Figure Description

[0026] Figure 1 This is a flowchart of the receiver clock bias-based INS / GPS tightly coupled integrated navigation enhancement method of the present invention. Detailed Implementation

[0027] like Figure 1 As shown, the present invention provides a method for enhancing tightly coupled INS / GPS integrated navigation based on receiver clock bias, comprising the following steps:

[0028] (1) The GPS receiver receives satellite navigation signals and calculates the clock error of the GPS receiver over 60 seconds, which is used to construct the GPS receiver clock error model.

[0029] (2) Select five sets of GPS receiver clock difference values ​​and corresponding times from the above 60s GPS receiver measurements, and use x(t)=a0+a1t+ε x (t)---Formula (1) is used to construct a linear model of the GPS receiver clock bias for predicting the GPS receiver clock bias, where: a0 is the initial value deviation of the GPS receiver clock; a1 is the clock drift of the GPS receiver; ε x (t) represents the random variation of the GPS receiver clock bias; x(t) represents the predicted GPS receiver clock bias value. a0, a1, and ε are calculated using the least squares linear fitting method. x (t);

[0030] (3) The navigation solution error of the strapdown inertial navigation system (INS) includes: position measurement error vector. Velocity error measurement vector Attitude error measurement error vector Gyroscope error vector δw=[δw x δw y δw z ] and accelerometer error vector δf=[δfx δf y δf z The state vector of the Kalman filter of the integrated navigation system is constructed from the above five vectors. in: δλ is the longitude measurement error, δh is the latitude measurement error, and δv is the elevation measurement error. e For the eastward velocity measurement error, δv n For northbound velocity measurement error, δv u δp is the pitch angle measurement error, δr is the roll angle measurement error, δA is the yaw angle measurement error, and [δw] is the yaw angle measurement error. x δw y δw z [δf] represents the error vectors in the three directions of the gyroscope, respectively. x δf y δf z ] are the error vectors of the accelerometer in three directions, and the above vectors can be obtained from equation (7.39) in reference [1].

[0031] (4) Estimate the pseudorange and pseudorange rate vectors using the strapdown inertial navigation system; calculate the pseudorange and pseudorange rate using GPS receiver information; and construct the measurement vectors of the Kalman filter of the integrated navigation system using the calculated or estimated values ​​of GPS receiver clock bias and clock drift. Where, ρ INS and The n×1 dimensional pseudorange and pseudorange rate vectors are estimated from the measurement information obtained by the strapdown inertial navigation system (see Equations (8.47) and (8.67) in Section 8.5.2 of Reference 1); ρ GPS and The n×1 dimensional pseudorange and pseudorange rate vectors are calculated from GPS measurement information (see equations (8.46) and (8.66) in section 8.5.2 of reference 1); δt u and δd u These are the calculated or estimated values ​​of GPS receiver clock bias and clock drift, respectively. n is the number of GPS visible satellites in the current epoch; the calculated values ​​of GPS receiver clock bias and clock drift are obtained by formulas (3.138) and (3.139) in reference [1]; the estimated values ​​of GPS receiver clock bias and clock drift are obtained by formula (1) in step (2);

[0032] (5) Number of visible satellites monitored by the GPS receiver

[0033] (a) When the number of visible satellites is greater than or equal to 4

[0034] Calculate the error detection threshold T thresholdand the detection sample ΔT, where

[0035] Where σ cal σ is the standard deviation of the calculated GPS receiver clock bias over a 60-second period. pre The standard deviation is the predicted clock bias of a GPS receiver over a 60-second period; the standard deviation is expressed as... The calculation yields x; where x i The calculated or predicted value of the clock difference per second in 60s; the calculated values ​​of clock difference and clock drift are obtained by formulas (3.138) and (3.139) in reference [1]; the predicted value of the GPS receiver clock difference is obtained by formula (1) in step (2). The average of the calculated or predicted clock difference over 60 seconds;

[0036] ΔT=δt u(cal) -δt u(pre) , where δt u(cal) Current receiver clock bias calculation value, δt u(pre) The current receiver clock bias prediction value is calculated using formula (1);

[0037] When |ΔT|>T threshold If an error is found, the predicted values ​​of the GPS receiver clock bias and clock drift are used instead of the calculated values ​​to participate in the combined navigation Kalman filter operation in formula (2); otherwise, if no error is found, the calculated values ​​of the receiver clock bias and clock drift are used to participate in the combined navigation Kalman filter operation.

[0038] (b) When the number of visible satellites is less than 4

[0039] When there are fewer than 4 satellites, the predicted values ​​of GPS receiver clock bias and clock drift cannot be calculated. Therefore, the combined navigation Kalman filter operation of formula (2) is directly used.

[0040] (6) Use The error correction value of the INS navigation solution output by the Kalman filter of the integrated navigation system is calculated, where: X is the state vector constructed in step (3); Z is the measurement vector constructed in step (4); K is the Kalman gain (see Table 7.1 in Reference 1); H is the measurement matrix (see Table 7.1 in Reference 1); This is the error correction value for INS navigation solution;

[0041] (7) Use the INS navigation solution error correction value from step (6). Error correction is performed on the navigation solutions of the inertial navigation system, including position, velocity, and attitude. This is the bit error correction amount. This is the speed error correction amount. δω is the attitude error correction value. corr δf is the gyroscope error correction value. corr The accelerometer error measurement correction is used to correct the position measurement error of the inertial navigation solution through the formula. Make corrections, where r ins r represents the position measurement value obtained by the inertial navigation system. ins / gps The corrected position measurement value; the velocity measurement error of the inertial navigation solution is expressed by the formula... Make corrections, where v ins The velocity measurement value, v, is obtained by the inertial navigation system. ins / gps The corrected velocity measurement value; the attitude measurement error of the inertial navigation solution is expressed by the formula... Make corrections, where ε ins The attitude measurement value ε obtained by the inertial navigation system ins / gps These are the corrected attitude measurements.

[0042] Reference 1 is Noureldin A, Karamat TB, Georgy J. Fundamentals of Inertial Navigation, Satellite-based Positioning and their Integration [M]. 2013.

[0043] The above description only discloses specific embodiments of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or modifications that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for enhancing tightly coupled INS / GPS integrated navigation based on receiver clock bias, characterized in that: It includes the following steps: (1) The GPS receiver receives satellite navigation signals and calculates the clock error of the GPS receiver over 60 seconds, which is used to construct the GPS receiver clock error model; (2) Select at least three sets of GPS receiver clock difference values ​​and corresponding times from the above 60s GPS receiver measurements, and use... ---Formula (1) is used to construct a linear model of the GPS receiver clock bias for predicting the GPS receiver clock bias, where: This represents the initial value deviation of the GPS receiver clock. For clock drift in GPS receivers; This represents the random variation in the GPS receiver clock bias; To predict the clock bias value of the GPS receiver, the least squares linear fitting method was used to calculate... , and ; (3) The navigation solution error of the strapdown inertial navigation system (INS) includes: position measurement error vector. Velocity error measurement vector Attitude error measurement error vector gyroscope error vector and accelerometer error vector The state vector of the Kalman filter of the integrated navigation system is constructed from the above five vectors. ,in: For longitude measurement error, For latitude measurement error, For elevation measurement errors, For the measurement error of eastward velocity, For northbound velocity measurement error, For the measurement error of the celestial velocity, For pitch angle measurement error, For roll angle measurement error, For heading angle measurement error, These are the error vectors in the three directions of the gyroscope, These are the error vectors in the three directions of the accelerometer; (4) Estimate the pseudorange and pseudorange rate vectors using the strapdown inertial navigation system; calculate the pseudorange and pseudorange rate using GPS receiver information; and construct the measurement vectors of the Kalman filter of the integrated navigation system using the calculated or estimated values ​​of GPS receiver clock bias and clock drift. ---Formula (2), where, and The information obtained from measurements using a strapdown inertial navigation system is predicted. 3D pseudorange and pseudorange rate vector; and Calculated using GPS measurement information dimensional pseudorange and pseudorange rate vector; and These are the calculated or estimated values ​​of GPS receiver clock bias and clock drift, respectively. ; The number of GPS visible satellites at the current epoch. (5) Number of visible satellites monitored by the GPS receiver (a) When the number of visible satellites is greater than or equal to 4 Calculate the error detection threshold and test samples ,in ,in The standard deviation of the calculated clock bias values ​​for a 60-second GPS receiver; The standard deviation is the predicted clock bias of a GPS receiver over a 60-second period; the standard deviation is expressed as... ----Formula (3) is used to calculate the result; where x i The calculated or predicted value of the clock difference per second in 60 seconds; The average of the calculated or predicted clock difference over 60 seconds; ,in Current receiver clock bias calculation value, The current receiver clock bias prediction value is calculated using formula (1); when If an error is found, the predicted values ​​of the GPS receiver clock bias and clock drift are used instead of the calculated values ​​to participate in the combined navigation Kalman filter operation in formula (2); otherwise, if no error is found, the calculated values ​​of the receiver clock bias and clock drift are used to participate in the combined navigation Kalman filter operation. (b) When the number of visible satellites is less than 4 When there are fewer than 4 satellites, the predicted values ​​of GPS receiver clock bias and clock drift cannot be calculated. Therefore, the combined navigation Kalman filter operation in formula (2) is directly used. (6) Use The INS navigation solution error correction value of the Kalman filter output of the integrated navigation system is calculated, where: The state vector constructed in step (3); The measurement vector constructed in step (4); Kalman gain; For measurement matrix; This is the error correction value for INS navigation solution; (7) Use the INS navigation solution error correction value from step (6). Error correction is performed on the inertial navigation system, including the position, velocity, and attitude navigation solutions. This is the bit error correction amount. This is the speed error correction amount. This is the attitude error correction amount. This is the gyroscope error correction amount. The accelerometer error measurement correction is used to correct the position measurement error of the inertial navigation solution through the formula. Make corrections, among which... The position measurement value obtained by the inertial navigation system. The corrected position measurement value; the velocity measurement error of the inertial navigation solution is expressed by the formula... Make corrections, among which... The velocity measurement value is obtained from the inertial navigation system. The corrected velocity measurement value; the attitude measurement error of the inertial navigation solution is expressed by the formula... Make corrections, among which... Attitude measurements obtained from an inertial navigation system. These are the corrected attitude measurements.

2. The receiver clock bias-based INS / GPS tightly coupled integrated navigation enhancement method as described in claim 1, characterized in that: In step (2), select the clock difference values ​​and corresponding times of five GPS receivers.

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