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

Abstract

The application discloses a kind of based on navigation array antenna and MEMS inertial navigation system's combined orientation, measure attitude system and method, including satellite navigation array antenna, MEMS inertial navigation, MEMS inertial navigation attitude solution module, combined attitude filter module, each single satellite navigation array antenna is integrated with a MEMS inertial navigation, satellite navigation array antenna is connected with combined attitude filter module, MEMS inertial navigation is connected with combined attitude filter module by MEMS inertial navigation attitude solution module.This application integrates MEMS inertial navigation in satellite navigation array antenna, on the one hand relative to the traditional use of optical fiber inertial navigation or laser inertial navigation and satellite navigation array antenna is combined in the way of orientation / measure attitude system, it will effectively reduce system volume, improve system integration, reduce system cost;On the other hand, the orientation of satellite navigation system, measure attitude data and the measure attitude data of MEMS inertial navigation are fused, can effectively improve satellite navigation array antenna orientation, measure attitude precision, improve the output frequency of orientation, attitude data, simultaneously can maintain the continuous working ability of orientation, measure attitude system in a certain time when satellite navigation signal is lost.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of orientation and attitude determination, and in particular to a combined orientation and attitude determination system and method based on a navigation array antenna and MEMS inertial navigation. BACKGROUND

[0002] An orientation / attitude determination system based on a satellite navigation system can determine two attitude angles of a carrier, i.e., an azimuth angle and a pitch angle (or a roll angle), if two antennas are used. If three or more antennas are used, the system can determine three attitude angles of the carrier, i.e., an azimuth angle, a pitch angle and a roll angle. The satellite orientation / attitude determination technology uses carrier phase observations of multiple satellites to obtain a baseline vector, and then obtains an angle of the baseline vector relative to a navigation coordinate system, i.e., an attitude angle of the carrier relative to the navigation coordinate system. The accuracy of the orientation / attitude determination method is closely related to the accuracy of the carrier phase observations, the stability of the antenna phase center and the length of the baseline.

[0003] In the prior art, the accuracy of the carrier phase observations provided by a satellite navigation chip has reached the millimeter level. To obtain better attitude angle accuracy, on the one hand, the design of the antenna is optimized, such as using a symmetric design for the antenna elements and using double or multi-feed technology for the elements to improve the stability of the antenna phase center. On the other hand, in the application scenario, the length between two or more antennas (i.e., the length of the baseline) is as long as possible, i.e., several meters or even tens of meters. The longer the antenna, the higher the accuracy obtained.

[0004] In some application scenarios, such as a vehicle-mounted satellite dynamic channel, the integration degree of the antenna is getting higher and higher, and multiple antenna elements are integrated into an antenna array panel. The distance between the elements on the antenna array panel is only tens of centimeters, and the optimization of the elements is limited due to the size limitation. The stability of the antenna phase center is deteriorated. Compared with the traditional multi-antenna separation design, in the satellite orientation / attitude determination application with an antenna distance of meters, the accuracy will decrease sharply, from the order of 0.1 degrees to several degrees. SUMMARY

[0005] To solve the above problems, the present application provides a combined orientation and attitude determination system and method based on a navigation array antenna and MEMS inertial navigation. The orientation / attitude determination data of the satellite navigation system and the attitude determination data of the MEMS inertial navigation are fused together through a combined attitude determination filter. The difference between the attitude error angles obtained by the MEMS inertial navigation and the attitude error angles obtained by the satellite navigation is converted into an attitude misalignment angle through a conversion matrix to obtain a measurement equation of the filter, and then the final orientation / attitude estimation is given through the filter.

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

[0007] A kind of combined orientation, attitude determination system based on navigation array antenna and MEMS inertial navigation, including satellite navigation array antenna, MEMS inertial navigation, MEMS inertial navigation attitude solution module, combined attitude determination filter module, each satellite navigation array antenna is integrated with a MEMS inertial navigation, satellite navigation array antenna is connected with combined attitude determination filter module, MEMS inertial navigation is connected with combined attitude determination filter module by MEMS inertial navigation attitude solution module.

[0008] As the preferred of above-mentioned scheme, the satellite navigation array antenna includes satellite navigation baseband chip and hardware platform for running MEMS inertial navigation attitude solution module and combined attitude determination filter module;The satellite navigation baseband chip is used to receive and process two or more array satellite signals, solve one or more baseline vectors, and finally give the attitude angle of satellite navigation orientation, attitude determination;The MEMS inertial navigation attitude solution module is used to obtain the angular velocity and acceleration information given by the MEMS inertial navigation, and solve the attitude information of the inertial navigation, and the azimuth angle is given by the satellite navigation baseband chip during the initial alignment of the inertial navigation;The combined attitude determination filter module is used to fuse the satellite navigation system orientation, attitude determination data and MEMS inertial navigation attitude determination data together, 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 into attitude misalignment angle through conversion matrix, obtain the measurement equation of the filter, and then give the final orientation, attitude estimation through the filter.

[0009] As the preferred of above-mentioned scheme, the combined attitude determination filter module adopts Kalman filter in the form of loose combination.

[0010] A kind of combined orientation, attitude determination method based on navigation array antenna and MEMS inertial navigation, comprising the following steps:

[0011] Step 1, establish combined orientation, attitude determination equation:

[0012] State equation:

[0013] Measurement equation: Z=HX+V

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

[0015] Wherein, the feature of measurement equation is:

[0016]

[0017] In the formula, ΔA=A INS -A GNSS Attitude error angle, A INS Attitude angle obtained by updating MEMS inertial navigation, AGNSS The attitude angles obtained from satellite orientation and attitude measurement. For the attitude misalignment angle, ε b For gyroscope drift along the three axes of the load system, To achieve zero bias of the accelerometers along the three axes of the load system, C t This is the transformation matrix related to the attitude misalignment angle and attitude error angle;

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

[0019]

[0020] One step of the transition matrix Φ k / k-1 In a relatively short time [t k-1 ,t k The inner approximation is Φ. k / k-1 ≈I+F(t k-1 )T s T s =T k -T k-1 , Γ k-1 W k-1 This refers to the system noise term.

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

[0022] State prediction in one step:

[0023] State estimation:

[0024] One-step prediction mean square error:

[0025] Filter gain:

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

[0027] R k To measure the noise variance matrix, Q k Let V be the system noise variance matrix. Simplify to empirical constant value processing.

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

[0029] The application integrates MEMS inertial navigation in a satellite navigation array antenna, which can effectively reduce the system volume, improve the system integration and reduce the system cost compared with the traditional combination of optical fiber inertial navigation or laser inertial navigation and satellite navigation orientation / attitude determination system; the combination of the satellite navigation system orientation / attitude determination data and the MEMS inertial navigation attitude determination data can effectively improve the satellite navigation orientation / attitude determination accuracy, improve the output frequency of the orientation / attitude data, and maintain the continuous working ability of the orientation / attitude determination system within a certain time when the satellite navigation signal is lost. BRIEF DESCRIPTION OF DRAWINGS

[0030] In order to more clearly illustrate the technical solutions in the embodiments of the application, the drawings needed in the embodiment description will be briefly introduced.

[0031] Figure 1 The system working principle block diagram of the application is shown in the figure.

[0032] Figure 2 The system structure schematic diagram of the application is shown in the figure. DETAILED DESCRIPTION

[0033] The technical solutions of the application will be described clearly and completely in combination with the drawings of the application. All other embodiments obtained by those skilled in the art based on the embodiments in the application without creative labor are within the protection scope of the application.

[0034] The embodiment provides a combined orientation and attitude determination system based on a navigation array antenna and MEMS inertial navigation, as shown in the figure. Figure 1 The combined orientation and attitude determination system includes a satellite navigation array antenna, a MEMS inertial navigation, a MEMS inertial navigation attitude determination module and a combined attitude determination filter module, one MEMS inertial navigation is integrated in each satellite navigation array antenna, the satellite navigation array antenna is connected with the combined attitude determination filter module, and the MEMS inertial navigation is connected with the combined attitude determination filter module through the MEMS inertial navigation attitude determination module.

[0035] Specifically,

[0036] The satellite navigation array antenna includes a satellite navigation baseband chip and a hardware platform for running the MEMS inertial navigation attitude determination module and the combined attitude determination filter module.

[0037] The satellite navigation baseband chip is used for receiving and processing satellite signals of two or more arrays, determining one or more baseline vectors, and finally giving the attitude angle of the satellite navigation orientation and attitude determination.

[0038] The MEMS inertial navigation attitude resolving module runs on the hardware platform of the array antenna, and is used for obtaining the angular velocity and acceleration information given by the MEMS inertial navigation, and resolving the attitude information of the inertial navigation. When the inertial navigation is initially aligned, the azimuth angle is given by the satellite navigation baseband chip.

[0039] The combined attitude resolving filter module runs on the hardware platform of the array antenna, adopts the Kalman filter, adopts the loose combination form, and is used for fusing the directional and attitude resolving data of the satellite navigation system and the attitude resolving data of the MEMS inertial navigation together, and converting the difference between the attitude error angles obtained by the MEMS inertial navigation and the attitude error angles obtained by the satellite navigation into the attitude misalignment angle through the conversion matrix, so as to obtain the measurement equation of the filter, and then giving the final directional and attitude estimation through the filter. The filter is composed of two equations: the state equation and the measurement equation. The state equation is the error equation, and the correction adopts the feedback correction, and the estimation result is fed back to the inertial navigation attitude resolving to correct the inertial device error.

[0040] The system structure schematic diagram is shown in Figure 2 The figure shows that S10 is the antenna cover, S11 is the MEMS inertial navigation, S12, S14, S15 and S16 are the array elements of the antenna, the number of the array elements is greater than or equal to 2, and at least two array elements are used for receiving the satellite navigation signal. In the figure, the array element 1 S15 and the array element 2 S12 are used for receiving the satellite navigation signal, the connecting line of the phase centers of the two array elements forms the baseline vector, and is parallel to the axis of the MEMS inertial navigation. When only two array elements are used, only two attitude angles of the carrier can be determined: the azimuth angle and the pitch angle (or the roll angle). If three or more array elements are used, three attitude angles of the carrier can be determined: the azimuth angle, the pitch angle and the roll angle.

[0041] The embodiment also provides a combined directional and attitude resolving method based on the navigation array antenna and the MEMS inertial navigation, which comprises the following steps:

[0042] Step 1, establishing the combined directional and attitude resolving equation:

[0043] The state variables of the system are selected as the attitude misalignment angle, the gyro drift error and the accelerometer zero offset error, and there are 9 dimensions, namely:

[0044]

[0045] Wherein, the subscripts e, n and u respectively represent the east, north and sky directions of the geographic navigation coordinate system (n system), represents the misalignment angle, and ε i represents the gyro drift along the three axes of the carrier system, represents the accelerometer zero offset along the three axes of the carrier system.

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

[0047]

[0048] where 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] where 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 / tau ay 1 / tau az ) are the inverse correlation time constants of the first-order Markov processes of the gyroscope and the accelerometer, respectively. W = [ω x ω y ω z a x a y a z ] T , ω i , a i (i = x, y, z) are white noises of the gyroscope and the accelerometer in the carrier coordinate system, respectively, with zero mean and normal distribution.

[0051] In the state equation, ε and are modeled as first-order Markov processes, i.e.

[0052]

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

[0054] The measurement equation is established as follows:

[0055] Z = HX + V (5)

[0056] where When the antenna array uses 2 elements, the roll angle error is ignored, and 0 is used instead, represents the attitude angle, where θ, γ, and denote the carrier pitch, roll and yaw angles (north and west positive) respectively; V is the measurement noise. The misalignment angle of the attitude (mathematical platform) and the conversion relationship of 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 obtained by inertial navigation update, and A GNSS is the attitude angle obtained by satellite navigation orientation.

[0057] Step 2, orientation, and discretization of the attitude equation:

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

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

[0060] Where the one-step transition matrix Φ k / k-1 is approximated as Φ k / k-1 ≈ I + F (t k-1 ) T s , T s = T k -T k-1 in a short time [t k-1 , t k ]. Γ k-1 W k-1 is used as a system noise term in engineering practice, and is treated as an empirical constant in the one-step prediction mean square error term P k / k-1 .

[0061] The measurement equation is discretized as:

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

[0063] Step 3, establish Kalman filter equation:

[0064] Filter update is divided into two kinds: time update and measurement update. The estimate k of the state quantity X is solved according to the following steps:

[0065] State one-step prediction:

[0066]

[0067] State estimation:

[0068]

[0069] One-step prediction mean square error:

[0070]

[0071] Filter gain:

[0072]

[0073] Estimated mean square error:

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

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

[0076] The above only is the preferred embodiment of the present application, and does not limit the present application, for those skilled in the art, the present application can have various modifications and changes. Any modification, equivalent replacement, improvement, etc. within the spirit and principle of the present application, should be included in the protection scope of the present application.

Claims

1. A combined orientation and attitude measurement method based on a navigation array antenna and a MEMS inertial navigation system, characterized in that, It includes a satellite navigation array antenna, a MEMS inertial navigation system, a MEMS inertial navigation attitude calculation module, and a combined attitude measurement filter module. Each individual satellite navigation array antenna integrates a MEMS inertial navigation system. 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 through the MEMS inertial navigation attitude calculation module. The method includes the following steps: Step 1: Establish combined orientation and attitude measurement equations: Equations 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 Let A be the attitude error angle. INS For the attitude angles obtained from the MEMS inertial navigation system update, A GNSS The attitude angles obtained from satellite orientation and attitude measurement. For the attitude misalignment angle, ε b For gyroscope drift along the three axes of the load system, To achieve zero bias of the accelerometers along the three axes of the load system, C t This is the transformation matrix related to the attitude misalignment angle and attitude error angle; Step 2: Discretization of orientation and attitude measurement equations: One step of the transition matrix Φ k / k-1 In a relatively short time [t k-1 ,t k The inner approximation is Φ. k / k-1 ≈I+F(t k-1 )T s T s =T k -T k-1 , Γ k-1 W k-1 This refers to the system noise term. Step 3: Establish the Kalman filter equation: State prediction in one step: State estimation: One-step prediction mean square error: Filter gain: Estimate mean square error: P k =(IK k H k )P k / k-1 R k To measure the noise variance matrix, Q k Let V be the system noise variance matrix. Simplify to empirical constant value processing.

2. The orientation and attitude measurement method based on a combination of navigation array antenna and MEMS inertial navigation according to claim 1, 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 calculation module and a combined attitude measurement filter module. The satellite navigation baseband chip receives and processes satellite signals from two or more arrays, calculates one or more baseline vectors, and finally provides the orientation and attitude angles calculated by the satellite navigation system. The MEMS inertial navigation attitude calculation module acquires the angular velocity and acceleration information provided by the MEMS inertial navigation system, calculates the attitude information of the inertial navigation system, and provides the azimuth angle during initial alignment by the satellite navigation baseband chip. The combined attitude measurement filter module fuses the orientation and attitude measurement data of the satellite navigation system and the attitude measurement data of the MEMS inertial navigation system, and converts the difference between the attitude error angle obtained by the MEMS inertial navigation system and the attitude error angle obtained by the satellite navigation system into the attitude misalignment angle through a transformation matrix, obtains the measurement equation of the filter, and then provides the final orientation and attitude estimation through the filter.

3. The orientation and attitude measurement method based on a combination of navigation array antenna and MEMS inertial navigation according to claim 1, characterized in that: The combined attitude measurement filter module uses a Kalman filter and is implemented in a loosely combined manner.

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