Combined orientation and attitude measurement system and method based on navigation array antenna and MEMS inertial navigation

Abstract

The invention discloses a combined orientation and attitude measurement system and method based on navigation array antennas and MEMS inertial navigation, the system comprises satellite navigation array antennas, MEMS inertial navigation, an MEMS inertial navigation attitude resolving module and a combined attitude measurement filter module, one MEMS inertial navigation is integrated in each satellite navigation array antenna, the satellite navigation array antennas are connected with the combined attitude measurement filter module, and the MEMS inertial navigation attitude resolving module is connected with the combined attitude measurement filter module. And the MEMS inertial navigation is connected with the combined attitude measurement filter module through the MEMS inertial navigation attitude resolving module. The MEMS inertial navigation is integrated in the satellite navigation array antenna, on one hand, compared with a traditional combination mode of using optical fiber inertial navigation or laser inertial navigation and a satellite navigation orientation / attitude measurement system, the system size is effectively reduced, the system integration degree is improved, and the system cost is reduced; and on the other hand, the orientation and attitude measurement data of the satellite navigation system and the attitude measurement data of the MEMS inertial navigation system are fused, so that the orientation and attitude measurement precision of the satellite navigation system can be effectively improved, the output frequency of the orientation and attitude data can be improved, and meanwhile, the continuous working capability of the orientation and attitude measurement system can be maintained within a certain time when the satellite navigation signal is unlocked.

Description

Technical Field

[0001] The present invention relates to the technical field of orientation and attitude measurement, and in particular to a combined orientation and attitude measurement system and method based on a navigation array antenna and a MEMS inertial navigation system. Background Art

[0002] The orientation / attitude measurement system based on the satellite navigation system, if using dual antennas, can determine the two attitude angles of the carrier: azimuth and pitch (or roll). If three or more antennas are used, the three attitude angles of the carrier can be determined: azimuth, pitch and roll. Satellite orientation / attitude measurement technology uses the carrier phase observations of multiple satellites to obtain the baseline vector, and then obtains the angle of the baseline vector relative to the navigation coordinate system from the baseline vector, that is, the attitude angle of the carrier relative to the navigation coordinate system. The accuracy of the orientation / attitude measurement method is closely related to the accuracy of the carrier phase observation, the stability of the antenna phase center and the length of the baseline.

[0003] With existing technology, the carrier phase observation accuracy given by satellite navigation chips has reached the millimeter level. To obtain better attitude angle accuracy, on the one hand, the antenna design is optimized, such as using a symmetrical design for antenna array elements and using dual-feed or multi-feed technology for array elements to improve the stability of the antenna phase center. On the other hand, in application scenarios, the length between two or more antennas (i.e., baseline length) is made as long as possible to reach several meters or even tens of meters. The longer the antenna, the higher the accuracy of the solution.

[0004] In some application scenarios, such as vehicle-mounted satellite communications, the integration of antennas is getting higher and higher. Multiple antenna elements are integrated into one antenna array dish. The distance between elements on the antenna array dish is only tens of centimeters. In addition, the optimization of the elements is limited due to size constraints, and the stability of the phase center of the antenna elements has deteriorated. Compared with the traditional multi-antenna separation design, the accuracy will drop sharply in satellite orientation / attitude measurement applications with antenna distances of meters, from 0.1 degrees to several degrees of accuracy. Summary of the invention

[0005] In view of the above problems, the present invention provides a combined orientation and attitude measurement system and method based on a navigation array antenna and MEMS inertial navigation, which fuses the orientation / attitude measurement data of a satellite navigation system and the attitude measurement data of a MEMS inertial navigation system through a combined attitude measurement filter. The difference between the attitude error angle obtained by the MEMS inertial navigation and the attitude error angle obtained by the satellite navigation is converted into an attitude misalignment angle through a conversion matrix, and the measurement equation of the filter is obtained, and then the final orientation / attitude estimation is given through the filter.

[0006] To achieve the above object, the present invention adopts the following technical solutions:

[0007] A combined orientation and attitude measurement system based on a navigation array antenna and a MEMS inertial navigation system comprises a satellite navigation array antenna, a MEMS inertial navigation system, a MEMS inertial navigation attitude solution module and a combined attitude measurement filter module. A MEMS inertial navigation system is integrated in each single satellite navigation array antenna. The satellite navigation array antenna is connected to the combined attitude measurement filter module, and the MEMS inertial navigation system is connected to the combined attitude measurement filter module via the MEMS inertial navigation attitude solution module.

[0008] As a preferred embodiment of the above scheme, the satellite navigation array antenna includes a satellite navigation baseband chip and a hardware platform for running a MEMS inertial navigation attitude solution module and a combined attitude measurement filter module; the satellite navigation baseband chip is used to receive and process satellite signals from two or more arrays, solve one or more baseline vectors, and finally provide the attitude angle of satellite navigation orientation and attitude measurement solved by the satellite navigation; the MEMS inertial navigation attitude solution module is used to obtain the angular velocity and acceleration information provided by the MEMS inertial navigation, and solve the attitude information of the inertial navigation. When the inertial navigation is initially aligned, the azimuth angle is provided by the satellite navigation baseband chip; the combined attitude measurement filter module is used to fuse the orientation and attitude measurement data of the satellite navigation system with the MEMS inertial navigation attitude measurement data, and convert the difference between the attitude error angle obtained by the MEMS inertial navigation and the attitude error angle obtained by the satellite navigation through a conversion matrix, to obtain the measurement equation of the filter, and then provide the final orientation and attitude estimation through the filter.

[0009] As a preferred embodiment of the above scheme, the combined attitude measurement filter module adopts a Kalman filter in the form of a loose combination.

[0010] A combined orientation and attitude measurement method based on a navigation array antenna and MEMS inertial navigation comprises the following steps:

[0011] Step 1: Establish the combined orientation and attitude measurement equation:

[0012] Equation of state:

[0013] Measurement equation: Z = HX + V

[0014] In the above formula, X is the state variable, F is the system matrix, G is the noise driving matrix, W is the system noise, Z is the measurement vector, H is the measurement matrix, and V is the measurement noise;

[0015] The measurement equation is characterized by:

[0016]

[0017] In the formula, ΔA=A INS -A GNSS is the attitude error angle, A INS is the attitude angle updated by MEMS inertial navigation, AGNSS is the attitude angle obtained by satellite orientation and attitude measurement, is the attitude misalignment angle, ε b is the gyro drift along the three axes of the carrier system, is the accelerometer zero bias along the three axes of the load system, C t is the transformation matrix related to attitude misalignment angle and attitude error angle;

[0018] Step 2: Discretization of orientation and attitude measurement equations:

[0019]

[0020] One-step transfer matrix Φ k / k-1 In a shorter time [t k-1 ,t k ] is approximately Φ k / k-1 ≈I+F(t k-1 )T s , T s =T k -T k-1 , Γ k-1 W k-1 is the system noise term;

[0021] Step 3: Establish the Kalman filter equation:

[0022] State one-step prediction:

[0023] State Estimation:

[0024] One-step forecast mean square error:

[0025] Filter gain:

[0026] Estimated mean square error: P k =(IK k H k ) k / k-1

[0027] R k is the measurement noise variance matrix, Q k is the system noise variance matrix, Simplified to empirical constant value.

[0028] Due to the above structure, the beneficial effects of the present invention are:

[0029] The present application integrates MEMS inertial navigation into the satellite navigation array antenna. On the one hand, compared with the traditional combination of fiber-optic inertial navigation or laser inertial navigation and satellite navigation orientation / attitude measurement system, it will effectively reduce the system volume, improve the system integration and reduce the system cost. On the other hand, the orientation / attitude measurement data of the satellite navigation system is integrated with the attitude measurement data of the MEMS inertial navigation, which will effectively improve the orientation / attitude measurement accuracy of the satellite navigation and increase the output frequency of the orientation / attitude data. At the same time, when the satellite navigation signal is lost, the orientation / attitude measurement system can maintain its continuous working capability for a certain period of time. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings required for describing the embodiments are briefly introduced below.

[0031] Figure 1 It is a block diagram of the working principle of the system of the present invention;

[0032] Figure 2 It is a schematic diagram of the system structure of the present invention. DETAILED DESCRIPTION

[0033] The technical solution of the present invention will be described clearly and completely below in conjunction with the accompanying drawings of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in the field without creative work are within the scope of protection of the present invention.

[0034] This embodiment provides a combined orientation and attitude measurement system based on a navigation array antenna and a MEMS inertial navigation system. Figure 1 As shown, it includes a satellite navigation array antenna, a MEMS inertial navigation, a MEMS inertial navigation attitude solution module, and a combined attitude measurement filter module. A MEMS inertial navigation is integrated in each single satellite navigation array antenna. The satellite navigation array antenna is connected to the combined attitude measurement filter module, and the MEMS inertial navigation is connected to the combined attitude measurement filter module through the MEMS inertial navigation attitude solution module.

[0035] Specific:

[0036] The satellite navigation array antenna includes a satellite navigation baseband chip and a hardware platform for running a MEMS inertial navigation attitude solution module and a combined attitude measurement filter module;

[0037] The satellite navigation baseband chip is used to receive and process satellite signals from two or more arrays, calculate one or more baseline vectors, and finally provide the attitude angle of satellite navigation orientation and attitude measurement calculated by the satellite navigation;

[0038] The MEMS inertial navigation attitude calculation module runs on the hardware platform of the array antenna, and is used to obtain the angular velocity and acceleration information given by the MEMS inertial navigation, and calculate the attitude information of the inertial navigation. When the inertial navigation is initially aligned, the azimuth is given by the satellite navigation baseband chip;

[0039] The combined attitude filter module runs on the hardware platform of the array antenna, adopts the Kalman filter, and adopts the loose combination form to fuse the orientation and attitude data of the satellite navigation system and the MEMS inertial navigation attitude data together, and converts the difference between the attitude error angle obtained by the MEMS inertial navigation and the attitude error angle obtained by the satellite navigation into the attitude misalignment angle through the conversion matrix, obtains the measurement equation of the filter, and then gives the final orientation and attitude estimation through the filter. The filter consists of two sets of equations: state equation and measurement equation. The state equation is an error equation, and the correction adopts feedback correction, and the estimation result is fed back to the inertial navigation attitude solution to correct the inertial device error.

[0040] The system structure diagram is as follows: Figure 2 As shown, S10 is the antenna cover, S11 is the MEMS inertial navigation, S12, S14, S15 and S16 are the antenna arrays, the number of arrays is ≥2, and at least 2 arrays are used to receive satellite navigation signals. In the figure, S15 array 1 and S12 array 2 are used to receive satellite navigation signals. The line connecting the phase centers of the two arrays forms a baseline vector, which is parallel to one axis of the MEMS inertial navigation. When only two arrays are used, only two attitude angles of the carrier can be determined: azimuth and pitch (or roll). If three or more arrays are used, the three attitude angles of the carrier can be determined: azimuth, pitch and roll.

[0041] This embodiment also provides a combined orientation and attitude measurement method based on a navigation array antenna and a MEMS inertial navigation system, comprising the following steps:

[0042] Step 1: Establish the combined orientation and attitude measurement equation:

[0043] The state variables of the system are: attitude misalignment angle, gyroscope drift error and accelerometer bias error, a total of 9 dimensions, namely:

[0044]

[0045] The subscripts e, n, and u represent the east, north, and celestial directions of the geographic navigation coordinate system (n system), respectively. represents the misalignment angle, ε i (i=x,y,z) represents the gyro drift along the three axes of the carrier system, It represents the zero bias of the accelerometer along the three axes of the load system.

[0046] From the attitude error equation, considering that the errors of the gyroscope and accelerometer are first-order Markov processes, the state equation of the filter can be obtained:

[0047]

[0048] Among them, X is the state variable, W is the system noise, F is the system matrix, and G is the noise driving matrix. The specific forms of F and G are as follows:

[0049]

[0050] in is the attitude matrix of the carrier system (b system) relative to the navigation system (n system); the diagonal matrix β g =diag(1 / τ gx 1 / τ gy 1 / τ gz ) and β a =diag(1 / τ ax 1 / τ ay 1 / τ az ) are the anti-correlation time constants of the first-order Markov processes of the gyroscope and accelerometer, respectively. x ω y ω z a x a y a z ] T ,ω i 、a i (i=x, y, z) are the zero-mean normally distributed white noises of the gyroscope and accelerometer in the carrier coordinate system.

[0051] In the state equation, ε and It is modeled as a first-order Markov process, that is,

[0052]

[0053] Where i = x, y, z represents the three components of the rectangular coordinate system, w g and w a They are the gyroscope angular rate white noise and accelerometer relative white noise respectively.

[0054] The measurement equation is established as follows:

[0055] Z=HX+V (5)

[0056] in When the antenna array uses two elements, the roll angle error is ignored and replaced by 0. represents the attitude angle, where θ, γ and They represent the carrier's pitch angle, roll angle, and azimuth angle (north-west is considered positive); V is the measurement noise. Attitude (mathematical platform) misalignment angle The conversion relationship between the attitude error angle ΔA is: The attitude error angle ΔA can be defined as ΔA = A INS -A GNSS , A INS is the attitude angle updated by the inertial navigation, A GNSS It is the attitude angle obtained by satellite navigation and orientation.

[0057] Step 2: Discretization of orientation and attitude measurement equations:

[0058] The state equation is in continuous form. In order to facilitate numerical calculation, it is necessary to use a discrete form. The discrete form of can be expressed as:

[0059] X k =Φ k / k-1 X k-1 +Γ k-1 W k-1 (6)

[0060] One-step transfer matrix Φ k / k-1 In a shorter time [t k-1 ,t k ] is approximately Φ k / k-1 ≈I+F(t k-1 )T s , T s =T k -T k-1 Γ k-1 W k-1 In engineering practice, as a system noise term, the mean square error term P is predicted in one step. k / k-1 Use empirical constants in processing.

[0061] The measurement equation is discretized as:

[0062] Z k =HX k +V k (7)

[0063] Step 3: Establish the Kalman filter equation:

[0064] There are two types of filter updates: time update and measurement update. k Estimates Follow these steps to solve:

[0065] State one-step prediction:

[0066]

[0067] State Estimation:

[0068]

[0069] One-step forecast mean square error:

[0070]

[0071] Filter gain:

[0072]

[0073] Estimate the mean squared error:

[0074] P k =(IK k H k ) k / k-1 (12)

[0075] R k is the measurement noise variance matrix, Q k is the system noise variance matrix, Simplified to empirical constant value.

[0076] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A combined orientation and attitude measurement system based on navigation array antenna and MEMS inertial navigation, characterized in that: It includes a satellite navigation array antenna, a MEMS inertial navigation, a MEMS inertial navigation attitude solution module, and a combined attitude measurement filter module. A MEMS inertial navigation is integrated in each single satellite navigation array antenna. The satellite navigation array antenna is connected to the combined attitude measurement filter module, and the MEMS inertial navigation is connected to the combined attitude measurement filter module through the MEMS inertial navigation attitude solution module.

2. The combined orientation and attitude measurement system based on navigation array antenna and MEMS inertial navigation according to claim 1 is characterized in that: The satellite navigation array antenna includes a satellite navigation baseband chip and a hardware platform for running a MEMS inertial navigation attitude solution module and a combined attitude measurement filter module; the satellite navigation baseband chip is used to receive and process satellite signals from two or more arrays, solve one or more baseline vectors, and finally provide the attitude angle of satellite navigation orientation and attitude measurement solved by the satellite navigation; the MEMS inertial navigation attitude solution module is used to obtain the angular velocity and acceleration information provided by the MEMS inertial navigation, and solve the attitude information of the inertial navigation. When the inertial navigation is initially aligned, the azimuth angle is provided by the satellite navigation baseband chip; the combined attitude measurement filter module is used to fuse the orientation and attitude measurement data of the satellite navigation system with the MEMS inertial navigation attitude measurement data, and convert the difference between the attitude error angle obtained by the MEMS inertial navigation and the attitude error angle obtained by the satellite navigation through a conversion matrix, so as to obtain the measurement equation of the filter, and then provide the final orientation and attitude estimation through the filter.

3. The combined orientation and attitude measurement system based on navigation array antenna and MEMS inertial navigation according to claim 1 is characterized in that: The combined attitude measurement filter module adopts a Kalman filter in a loose combination form.

4. A combined orientation and attitude measurement method based on a navigation array antenna and MEMS inertial navigation according to any one of claims 1 to 3, characterized in that: The steps include: Step 1: Establish the combined orientation and attitude measurement equation: Equation of state: Measurement equation: Z = HX + V In the above formula, X is the state variable, F is the system matrix, G is the noise driving matrix, W is the system noise, Z is the measurement vector, H is the measurement matrix, and V is the measurement noise; The measurement equation is characterized by: In the formula, ΔA=A INS -A GNSS is the attitude error angle, A INS is the attitude angle updated by MEMS inertial navigation, A GNSS is the attitude angle obtained by satellite orientation and attitude measurement, is the attitude misalignment angle, ε b is the gyro drift along the three axes of the carrier system, is the accelerometer zero bias along the three axes of the load system, C t is the transformation matrix related to attitude misalignment angle and attitude error angle; Step 2: Discretization of orientation and attitude measurement equations: One-step transfer matrix Φ k / k-1 In a shorter time [t k-1 ,t k ] is approximately Φ k / k-1 ≈I+F(t k-1 )T s , T s =T k -T k-1 , Γ k-1 W k-1 is the system noise term; Step 3: Establish the Kalman filter equation: State one-step prediction: State Estimation: One-step forecast mean square error: Filter gain: Estimated mean square error: P k =(IK k H k ) k / k-1 R k is the measurement noise variance matrix, Q k is the system noise variance matrix, Simplified to empirical constant value.

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