Flight calibration method and system for structural parameters of star sensor and inertial unit

Abstract

The invention relates to a flight calibration method and system for structural parameters of a star sensor and an inertial unit, and the method comprises the following steps: setting parameters and environmental parameters of the star sensor and the inertial unit, and setting an initial state of a structural parameter deviation estimator; after the inertial unit and the star sensor output data, performing time propagation on estimation values and estimation covariance of constant zero offset, zero offset drift and structural parameter deviation; calculating attitude increments of the inertial unit and the star sensor by using the angular velocity and star observation information, and constructing a measurement equation of structural parameter deviation; solving a corresponding regularization least square problem by using a regularization robust filtering framework, and completing estimation value updating and estimation covariance updating of constant value zero offset, zero offset drift and structure parameter deviation; and after the stable estimated value of the structure parameter deviation is obtained, the structure parameters are compensated. According to the method, high-precision structural parameter calibration of the star sensor and the inertial unit can be realized, and complicated astronomical-ground coordinate conversion can be carried out without the aid of time service positioning information.

Description

Technical Field

[0001] The invention relates to the technical field of combined application of a star sensor and an inertial unit, and in particular to a flight calibration method and system for structural parameters of a star sensor and an inertial unit. Background Art

[0002] Inertial navigation system is an important navigation system in aviation. It obtains the position, velocity and attitude information of the carrier in real time by integrating the measurement values ​​of the inertial unit. It has been widely used because of its strong autonomy and good anti-interference ability. In the real flight environment, complex factors such as overload, vibration, temperature change, etc. will act on the inertial unit in a comprehensive manner, resulting in the actual performance of the inertial navigation system being inferior to the nominal performance. Therefore, it is necessary to test the performance of the inertial navigation system in a real flight environment. Star sensor is an astronomical navigation device that provides the attitude information of the carrier by observing stars. Since the star sensor has the characteristics of high precision and no drift, it can be used as the attitude reference of the inertial navigation system after being strapdown-installed with the inertial navigation system, thereby realizing the in-flight attitude accuracy assessment of the inertial navigation system. In addition, the star sensor and the inertial unit can also form a combined navigation system. However, the star sensor and the inertial unit have independent measurement coordinate systems. Whether it is in-flight attitude accuracy assessment or combined navigation, it is necessary to align the measurement coordinate system first, that is, it is necessary to calibrate the structural parameters between the star sensor and the inertial unit.

[0003] After the strapdown system of the star sensor and the inertial unit is assembled, its structural parameters can be calibrated by experimental methods. However, factors such as impact during transportation, release of installation preload, and vibration and temperature changes in the working environment may cause structural changes between the star sensor and the inertial measurement unit, so it is necessary to calibrate the structural parameters in real time online. The existing online calibration method of structural parameters uses the attitude data output by the inertial navigation system and the structural parameters pre-calibrated in the laboratory to infer the direction of the stars observed by the star sensor, and then uses the difference between the inferred star direction and the star direction actually observed by the star sensor to establish the measurement equation of the structural parameter deviation. Finally, the Kalman filter is used to calibrate the structural parameters based on the measurement equation.

[0004] Existing methods use the attitude information of the inertial navigation system to establish the measurement equation of the structural parameter deviation, but the cumulative error contained in the attitude information has a significant impact on the calibration process. Existing methods use the process data inside the inertial navigation system to perform complex recursion and compensation for the attitude cumulative error, but these process data cannot be obtained in the application context of the inertial navigation system attitude accuracy assessment. At the same time, when observing the stars in the atmosphere, the atmospheric refraction causes the direction of the starlight to be deflected, so that the star observation information of the star sensor is biased. However, the existing methods do not provide a means to overcome this problem. Therefore, when the existing methods are used to calibrate the structural parameters in the atmosphere, the results will be significantly affected by the atmospheric refraction of starlight. Therefore, a more effective flight calibration method for the structural parameters of star sensors and inertial units is currently needed to solve the above problems. Summary of the invention

[0005] In order to solve the problem of insufficient calibration accuracy of structural parameters of existing star sensors and inertial units, the purpose of the present invention is to provide a method and system for flight calibration of structural parameters of star sensors and inertial units. The method of the present invention is a method for flight calibration of structural parameters based on attitude increment measurement. The method obtains the attitude increment of the inertial unit by integrating the angular velocity information in a short time; and calculates the attitude increment of the star sensor using the stargazing information at the adjacent time. A measurement equation for structural parameter deviation is established based on the attitude increment of the inertial unit and the star sensor, and then a structural parameter deviation estimator is designed using a regularized robust filtering framework to calibrate the structural parameters. The present invention avoids the introduction of attitude accumulation error and starlight atmospheric refraction, so it can achieve high-precision flight calibration of structural parameters.

[0006] A first aspect of the present invention provides a method for flight calibration of star sensor and inertial unit structural parameters, comprising:

[0007] Step 1: Set the sensor accuracy, seeing, attitude increment integration time, and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data manual and the weather conditions during calibration, and set the initial state of the structural parameter deviation estimator according to the above parameters;

[0008] Step 2: after the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information, the estimated values ​​and estimated covariances of the gyroscope's constant bias, bias drift, and structural parameter deviations are time-propagated according to the parameter model;

[0009] Step 3: Use the angular velocity measurement values ​​and star observation information obtained in step 2 to construct the attitude increments of the inertial unit and the star sensor, and then construct the measurement equation of the structural parameter deviation and the measurement noise covariance;

[0010] Step 4: Based on the measurement equation described in step 3, the regularized robust filtering framework is used to solve the corresponding regularized least squares problem, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation;

[0011] Step 5: After obtaining the estimated value of the stable structural parameter deviation, the structural parameters are compensated.

[0012] Furthermore, among the parameters that need to be set in step 1, the sensor accuracy includes the zero bias repeatability σ of the three-axis gyroscope in the inertial unit bx ,σ by ,σ bz , angle random walk N Gx ,N Gy ,N Gz , rate random walk K Gx ,K Gy ,K Gz 、The star measurement accuracy of the star sensor σ w ; Seeing degree σ η Determined by the half-width of the seeing disk, which is 2.355σ under Gaussian distribution. η ; The attitude increment integration time is represented by the symbol Δk;

[0013] The structural parameters refer to the installation structural parameters of the star sensor and the inertial unit, which are represented by an orthogonal matrix S, which describes the coordinate transformation from the star sensor coordinate system to the inertial unit coordinate system. The pre-calibrated value of this matrix is ​​S o , the deviation between the pre-calibrated value and the actual value is the structural parameter deviation matrix M, and their relationship is S = MS o , the structural parameter deviation M represents a small angle rotation, which is approximated by the corresponding rotation vector θ:

[0014]

[0015] in,[·] × Represents the antisymmetric matrix of the vector, and the uncertainty of the structural parameters is the standard deviation of θ1, θ2, θ3, expressed as σ θx ,σ θy ,σ θz .

[0016] Furthermore, in step 2, the parameter model includes a model of the constant bias, bias drift and structural parameter deviation of the gyroscope, and the bias drift d of the gyroscope is k Modeled as a first-order Markov process d k =d k-1 +ε k-1 , where ε k-1 is the rate random walk noise, with a constant zero bias bk and structural parameter deviation θ k It can be modeled as a constant quantity to be estimated. Assuming that at time k, the inertial unit outputs a new angular velocity measurement and the star sensor outputs new star observation information, the estimated value and covariance of the state quantity composed of the gyroscope's constant zero bias, zero bias drift, and structural parameter deviation are propagated from time k-1 to time k as follows:

[0017]

[0018] P x,k|k-1 =P x,k-1|k-1 +Q k-1

[0019] in, and P x,k-1|k-1 They represent the posterior estimate and posterior covariance of the state quantity at time k-1 respectively, and P x,k|k-1 They represent the prior estimate and prior covariance of the state quantity at time k, Q k-1 The covariance matrix of the process noise is Where T s is the sampling interval.

[0020] Furthermore, in step 3, the construction of the attitude increments of the inertial unit and the star sensor includes the following steps:

[0021] (1) The starting time of the calculated attitude increment is set to The termination time is set to k;

[0022] (2) Let the angular velocity measurement value of the inertial unit at time k be ω gm,k , then from time i-1 to time i The rotation vector of the inertial unit coordinate system can be estimated as:

[0023]

[0024] in, is the a posteriori estimate of the zero bias drift at time i, and the matrix Λ contains the scale error λ x ,λ y ,λ z and the non-orthogonality error δ xy , δ xz , δ yz , has the following form:

[0025]

[0026] Rotation Vector Can be converted into rotation matrix form Therefore, the attitude increment of the inertial unit can be estimated as

[0027] (3) Let the direction vector of the i-th star observed by the star sensor at time k in the star sensor coordinate system be w i,k At the adjacent time k and time l, the star sensor observes several common stars, and the collection of these stars is recorded as Star sensor attitude increment Use the QUEST algorithm or SVD decomposition method to solve the following Wahba problem:

[0028]

[0029] Furthermore, in step 4, the regularized least squares problem to be solved is:

[0030]

[0031] in, H k and W k are the measured values, sensitivity matrix and measurement noise covariance of the measurement equation established in step 3, respectively, ||·|| W represents the weighted Euclidean norm of the vector, δH k Denotes the sensitivity matrix H k Possible deviations come from possible disturbances of the matrix Λ during operation. Λ Δ Λ E Λ , the matrix δH k It can be expressed as:

[0032]

[0033] Δ Λ =diag(Rλ x ,Rδ xy ,Rδ xz ,Rλ y ,Rδ yz ,Rλ z )

[0034]

[0035] Among them, if λ' is used x Represents λ x The disturbance amount, then Uλ x Represents λ' x The limit of Rλ x =λ' x / Uλ x, other variables are expressed in the same way.

[0036] A second aspect of the present invention provides a flight calibration system for star sensor and inertial unit structural parameters, comprising:

[0037] A setting module is used to set the sensor accuracy, seeing, attitude increment integration time and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data manual and the weather conditions during calibration, and to set the initial state of the structural parameter deviation estimator according to the above parameters;

[0038] A time propagation module is used to perform time propagation of the estimated values ​​and covariance of the gyroscope's constant bias, bias drift, and structural parameter deviation according to the parameter model after the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information;

[0039] A construction module is used to construct the attitude increments of the inertial unit and the star sensor using the obtained angular velocity measurements and star observation information, and then construct the measurement equations of the structural parameter deviations and the measurement noise covariance;

[0040] An iterative estimation module is used to solve the corresponding regularized least squares problem based on the measurement equation using a regularized robust filtering framework, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation;

[0041] The compensation module is used to compensate the structural parameters after obtaining the estimated value of the stable structural parameter deviation.

[0042] The technical solution of the present invention can achieve the following beneficial technical effects:

[0043] The present invention establishes a measurement equation for structural parameter deviation based on the attitude increments of the inertial unit and the star sensor. Due to the short integration time, the attitude increment of the inertial unit contains only a small cumulative error; since the atmospheric refraction at adjacent moments is approximately constant, the starlight atmospheric refraction can be offset in the calculation process of the star sensor attitude increment. The method of the present invention avoids the introduction of attitude cumulative errors and starlight atmospheric refraction, so it can achieve high-precision structural parameter calibration. In addition, each of the above-mentioned attitude increments is based on its own measurement coordinate system, so the method of the present invention does not need to use timing and positioning information to perform complex astronomical-ground coordinate conversion. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 It is a flow chart of steps of the flight calibration method of the star sensor and the inertial unit structural parameters of the present invention;

[0045] Figure 2 It is the structural parameter deviation estimation result curve of the method of the present invention. DETAILED DESCRIPTION

[0046] In order to enable those skilled in the art to better understand the scheme of the present invention, the technical scheme in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of the present invention.

[0047] It should be noted that the terms "first", "second", etc. in the specification and claims of the present invention and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the data used in this way can be interchanged where appropriate, so that the embodiments of the present invention described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "including" and "having" and any variations thereof are intended to cover non-exclusive inclusions, for example, a process, method, system, product or device that includes a series of steps or units is not necessarily limited to those steps or units that are clearly listed, but may include other steps or units that are not clearly listed or inherent to these processes, methods, products or devices.

[0048] According to an embodiment of the present invention, an embodiment of an on-orbit multi-load reference deviation measurement method and system is provided. It should be noted that the steps shown in the flowchart of the accompanying drawings can be executed in a computer system such as a set of computer executable instructions, and although a logical order is shown in the flowchart, in some cases, the steps shown or described can be executed in an order different from that shown here.

[0049] The present invention is described in detail below with reference to the accompanying drawings and embodiments.

[0050] A first aspect of the present invention provides a method for flight calibration of a star sensor and an inertial unit structural parameters, specifically comprising the following steps:

[0051] Step 1: Set the sensor accuracy, seeing, attitude increment integration time, and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data sheet and the weather conditions during calibration, and set the initial state of the structural parameter deviation estimator according to the above parameters.

[0052] Specifically, among the parameters that need to be set, the sensor accuracy includes the zero bias repeatability σ of the three-axis gyroscope in the inertial unit bx ,σ by ,σ bz , angle random walk N Gx ,N Gy ,NGz , rate random walk K Gx ,K Gy ,K Gz 、The star measurement accuracy of the star sensor σ w ; Seeing degree σ η Determined by the half-width of the seeing disk, which is 2.355σ under Gaussian distribution. η ; The attitude increment integration time is represented by the symbol Δk.

[0053] Specifically, the structural parameters refer to the installation structural parameters of the star sensor and the inertial unit, which are represented by an orthogonal matrix S, which describes the coordinate transformation from the star sensor coordinate system to the inertial unit coordinate system. The pre-calibrated value of this matrix is ​​S o , the deviation between the pre-calibrated value and the actual value is the structural parameter deviation matrix M, and their relationship is S = MS o The structural parameter deviation M usually represents a small angle rotation and can therefore be approximated by the corresponding rotation vector θ:

[0054]

[0055] in,[·] × Represents the antisymmetric matrix of the vector. The uncertainty of the structural parameters is the standard deviation of θ1, θ2, θ3, expressed as σ θx ,σ θy ,σ θz .

[0056] Specifically, the state quantity of the structural parameter deviation estimator includes three variables: the constant zero bias b output by the gyro k , zero bias drift d k and the structural parameter deviation vector θ between the star sensor and the inertial measurement unit k , where the subscript k is the time variable, that is, the state vector is The initial state of the structural parameter deviation estimator that needs to be set is the posterior estimate of the state quantity at time 0 and the posterior covariance

[0057] Step 2: After the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information, the estimated values ​​and estimated covariances of the gyroscope's constant bias, bias drift, and structural parameter deviations are time-propagated according to the parameter model.

[0058] Specifically, the parameter model includes the model of the constant bias, bias drift and structural parameter deviation of the gyroscope. k It can be modeled as a first-order Markov process d k =d k-1 +εk-1 , where ε k-1 is the rate random walk noise, with a constant zero bias b k and structural parameter deviation θ k It can be modeled as a constant quantity to be estimated. Assuming that at time k, the inertial unit outputs a new angular velocity measurement and the star sensor outputs new star observation information, the estimated value and covariance of the state quantity composed of the gyroscope's constant zero bias, zero bias drift, and structural parameter deviation are propagated from time k-1 to time k as follows:

[0059]

[0060] P x,k|k-1 =P x,k-1|k-1 +Q k-1

[0061] in, and P x,k-1|k-1 They represent the posterior estimate and posterior covariance of the state quantity at time k-1 respectively, and P x,k|k-1 They represent the prior estimate and prior covariance of the state quantity at time k, Q k-1 The covariance matrix of the process noise is Where T s is the sampling interval.

[0062] Step 3: Use the angular velocity measurements obtained in step 2 and the star observation information to construct the attitude increments of the inertial unit and the star sensor, and then construct the measurement equations of the structural parameter deviations and the measurement noise covariance.

[0063] Specifically, the construction of the attitude increment of the inertial unit and the star sensor includes the following steps:

[0064] (1) The starting time of the calculated attitude increment is set to The end time is set to k.

[0065] (2) Let the angular velocity measurement value of the inertial unit at time k be ω gm,k , then from time i-1 to time i The rotation vector of the inertial unit coordinate system can be estimated as:

[0066]

[0067] in, is the a posteriori estimate of the zero bias drift at time i, and the matrix Λ contains the scale error λ x ,λ y ,λ z and the non-orthogonality error δ xy , δ xz, δ yz , has the following form:

[0068]

[0069] Rotation Vector Can be converted into rotation matrix form Therefore, the attitude increment of the inertial unit can be estimated as

[0070] (3) Let the direction vector of the i-th star observed by the star sensor at time k in the star sensor coordinate system be w i,k . At the adjacent k moments and At the moment, the star sensor can often observe several common stars, and the collection of these stars is recorded as Therefore, the attitude increment of the star sensor is The following Wahba problem can be solved using QUEST algorithm or SVD decomposition method.

[0071]

[0072] Specifically, the construction of the measurement equation of the structural parameter deviation and the measurement noise covariance includes the following steps:

[0073] (1) From the above-mentioned inertial unit attitude increment Star sensor attitude increment and pre-calibrated values ​​S of structural parameters o , which can form a quantity close to the unit matrix Its deviation from the identity matrix can be used to establish the measurement quantity The relationship is

[0074] (2) Measurement quantity With the state quantity x k The relationship between is the measurement equation, which is in, represents the measurement noise, H k is the sensitivity matrix, and its expression is:

[0075]

[0076] Among them, J i is the left-multiplied Baker-Campbell-Hausdorff approximate Jacobian matrix at time i, given by:

[0077]

[0078] (3) Measurement noise The covariance matrix ofk , expressed as:

[0079]

[0080] Among them, P d,i|i is the a posteriori estimated covariance of the zero bias drift at time i, and the covariance matrix of the star sensor attitude increment and the angular random walk noise covariance P of the three-axis gyroscope Δω They are expressed as:

[0081]

[0082] Step 4: Based on the measurement equation in step 3, the regularized robust filtering framework is used to solve the corresponding regularized least squares problem, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation.

[0083] Specifically, the regularized least squares problem that needs to be solved is:

[0084]

[0085] in, H k and W k are the measured values, sensitivity matrix and measurement noise covariance of the measurement equation established in step 3, respectively, ||·|| W represents the weighted Euclidean norm of the vector, δH k Denotes the sensitivity matrix H k Possible deviations come from possible disturbances of the matrix Λ during operation. Λ Δ Λ E Λ , the matrix δH k It can be expressed as

[0086]

[0087] Δ Λ =diag(Rλ x ,Rδ xy ,Rδ xz ,Rλ y ,Rδ yz ,Rλ z )

[0088]

[0089] Among them, if λ' is used x Represents λ x The disturbance amount, then Uλ x Represents λ'x The limit of Rλ x =λ' x / Uλ x , other variables are expressed in the same way.

[0090] Specifically, we assume that a robust version of the regularized least squares problem can be described as:

[0091]

[0092] Among them, x is the value to be estimated, A and y are known matrix parameters and vector parameters respectively, Q and W are symmetric and semi-positive definite weight matrices, δA and δy represent the uncertainty of parameters A and y, have the same dimensions as A and y respectively, and can be decomposed as follows:

[0093] [δAδy]=CΔ[E a E y ]

[0094] Among them, C, E a , E y are preset quantities with appropriate dimensions, Δ is an arbitrary contraction matrix that satisfies ||Δ||≤1, and for matrices, ||·|| represents the 2-induced norm. Then the unique solution to this problem can be expressed as

[0095]

[0096] The new weight matrix It is modified from the original weight matrix {Q, W} as follows

[0097]

[0098] mark represents the pseudoinverse of a matrix, rather than a negative scalar parameter Determine by the following optimization problem

[0099]

[0100] The function G(β) is defined as

[0101]

[0102] Here is

[0103]

[0104] Specifically, after making the following substitutions for the regularized least squares problem to be solved, the above solution method can be used to obtain the posterior estimate of the state quantity at time k:

[0105]

[0106] At the same time, the posterior estimated covariance of the state quantity at time k can be updated as:

[0107]

[0108] Step 5: After obtaining the estimated value of the stable structural parameter deviation, the structural parameters are compensated.

[0109] Specifically, after obtaining the estimated value of the stable structural parameter deviation, the stable value is recorded as Then the calibration value S of the structural parameter * as follows

[0110]

[0111] Using the method of the present invention, various index parameters, ideal structural parameters and structural parameter deviations of the star sensor and the inertial unit are set, and the calibration accuracy of the method is verified by simulation. The three-axis structural parameter deviations are set to -30.60", 19.76" and 40.85", respectively. The estimated results of the structural parameter deviations are as follows: Figure 2 As shown in the figure, the estimated structural parameter deviation quickly converges to the true value of the structural parameter deviation after experiencing the initial fluctuation. The root mean square error of the estimated result after convergence is less than 1", which meets the high-precision structural parameter calibration requirements for the combined application of star sensors and inertial units.

[0112] A second aspect of the present invention provides a flight calibration system for star sensor and inertial unit structural parameters, comprising:

[0113] A setting module is used to set the sensor accuracy, seeing, attitude increment integration time and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data manual and the weather conditions during calibration, and to set the initial state of the structural parameter deviation estimator according to the above parameters;

[0114] A time propagation module is used to perform time propagation of the estimated values ​​and covariance of the gyroscope's constant bias, bias drift, and structural parameter deviation according to the parameter model after the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information;

[0115] A construction module is used to construct the attitude increments of the inertial unit and the star sensor using the obtained angular velocity measurements and star observation information, and then construct the measurement equations of the structural parameter deviations and the measurement noise covariance;

[0116] An iterative estimation module is used to solve the corresponding regularized least squares problem based on the measurement equation using a regularized robust filtering framework, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation;

[0117] The compensation module is used to compensate the structural parameters after obtaining the estimated value of the stable structural parameter deviation.

[0118] In summary, the present invention relates to a method and system for flight calibration of structural parameters of a star sensor and an inertial unit. The calibration method comprises the following steps: setting parameters and environmental parameters of the star sensor and the inertial unit, and setting the initial state of the structural parameter deviation estimator; after the inertial unit and the star sensor output data, time propagation of the estimated values ​​and estimated covariance of the constant zero bias, zero bias drift and structural parameter deviation; calculating the attitude increment of the inertial unit and the star sensor using angular velocity and stargazing information, and then constructing the measurement equation of the structural parameter deviation; solving the corresponding regularized least squares problem using a regularized robust filtering framework, completing the update of the estimated values ​​and estimated covariance of the constant zero bias, zero bias drift and structural parameter deviation; after obtaining the stable estimated value of the structural parameter deviation, compensating the structural parameters. The method of the present invention can achieve high-precision calibration of the structural parameters of the star sensor and the inertial unit because it avoids the introduction of attitude accumulation error and starlight atmospheric refraction. In addition, each of the above-mentioned attitude increments is based on their own measurement coordinate system, so the method of the present invention does not need to use timing positioning information to perform complex astronomical-ground coordinate conversion.

[0119] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principle of the present invention. These improvements and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A method for flight calibration of star sensor and inertial unit structural parameters, characterized in that: The following steps are involved: Step 1: Set the sensor accuracy, seeing, attitude increment integration time, and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data manual and the weather conditions during calibration, and set the initial state of the structural parameter deviation estimator according to the above parameters; Step 2: after the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information, the estimated values ​​and estimated covariances of the gyroscope's constant bias, bias drift, and structural parameter deviations are time-propagated according to the parameter model; Step 3: Use the angular velocity measurement values ​​and star observation information obtained in step 2 to construct the attitude increments of the inertial unit and the star sensor, and then construct the measurement equation of the structural parameter deviation and the measurement noise covariance; Step 4: Based on the measurement equation described in step 3, the regularized robust filtering framework is used to solve the corresponding regularized least squares problem, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation; Step 5: After obtaining the estimated value of the stable structural parameter deviation, the structural parameters are compensated.

2. The method for flight calibration of a star sensor and inertial unit structural parameters as claimed in claim 1, characterized in that: Among the parameters that need to be set in step 1, the sensor accuracy includes the zero bias repeatability σ of the three-axis gyroscope in the inertial unit bx , σ by , σ bz , angle random walk N Gx , N Gy , N Gz , rate random walk K Gx , K Gy , K Gz , the star measurement accuracy of the star sensor σ w ; Seeing degree σ η Determined by the half-width of the seeing disk, which is 2.355σ under Gaussian distribution. η ; The attitude increment integration time is represented by the symbol Δk; The structural parameters refer to the installation structural parameters of the star sensor and the inertial unit, which are represented by an orthogonal matrix S, which describes the coordinate transformation from the star sensor coordinate system to the inertial unit coordinate system. The pre-calibrated value of this matrix is ​​S o , the deviation between the pre-calibrated value and the actual value is the structural parameter deviation matrix M, and their relationship is S = MS o , the structural parameter deviation M represents a small angle rotation, which is approximated by the corresponding rotation vector θ: in,[·] × Represents the antisymmetric matrix of the vector, the uncertainty of the structural parameters is the standard deviation of θ1, θ2, θ3, expressed as σ θx , σ θy , σ θz .

3. The method for flight calibration of a star sensor and inertial unit structural parameters as claimed in claim 1, characterized in that: In step 2, the parameter model includes the model of the constant bias, bias drift and structural parameter deviation of the gyroscope. The bias drift d k Modeled as a first-order Markov process d k =d k-1 +ε k-1 , where ε k-1 is the rate random walk noise, with a constant zero bias b k and structural parameter deviation θ k It can be modeled as a constant quantity to be estimated. Assuming that at time k, the inertial unit outputs a new angular velocity measurement and the star sensor outputs new star observation information, the estimated value and covariance of the state quantity composed of the gyroscope's constant zero bias, zero bias drift, and structural parameter deviation are propagated from time k-1 to time k as follows: P x,k|k-1 =P x,k-1|k-1 +Q k-1 in, and P x,k-1|k-1 They represent the posterior estimate and posterior covariance of the state quantity at time k-1 respectively, and P x,k|k-1 They represent the prior estimate and prior covariance of the state quantity at time k, Q k-1 The covariance matrix of the process noise is Where T s is the sampling interval.

4. The method for flight calibration of a star sensor and inertial unit structural parameters as claimed in claim 1, characterized in that: In step 3, the construction of the attitude increments of the inertial unit and the star sensor includes the following steps: (1) The starting time of the calculated attitude increment is set to l = k - Δk, and the ending time is set to k; (2) Let the angular velocity measurement value of the inertial unit at time k be ω gm,k , then in the interval from time i-1 to time i (l<i≤k), the rotation vector of the inertial unit coordinate system can be estimated as: in, is the a posteriori estimate of the zero bias drift at time i, and the matrix Λ contains the scale error λ x ,λ y ,λ z and the non-orthogonality error δ xy , δ xz , δ yz , has the following form: Rotation Vector Can be converted into rotation matrix form Therefore, the attitude increment of the inertial unit can be estimated as (3) Let the direction vector of the i-th star observed by the star sensor at time k in the star sensor coordinate system be w i,k At the adjacent time k and time l, the star sensor observes several common stars, and the collection of these stars is recorded as Star sensor attitude increment Use the QUEST algorithm or SVD decomposition method to solve the following Wahba problem: 。 5. The method for flight calibration of a star sensor and inertial unit structural parameters as claimed in claim 1, characterized in that: In step 4, the regularized least squares problem to be solved is: Among them, z l，k , H k and W k are the measured values, sensitivity matrix and measurement noise covariance of the measurement equation established in step 3, respectively, ||·|| W represents the weighted Euclidean norm of the vector, δH k Denotes the sensitivity matrix H k Possible deviations come from possible disturbances of the matrix Λ during operation. Λ Δ Λ E Λ , the matrix δH k It can be expressed as: ΔΛ=diag(Rλ x ,Rδ xy ,Rδ xz ,Rλ y ,Rδ yz ,Rλ z ) Among them, if λ′ is used x Represents λ x The disturbance amount, then Uλ x Represents λ′ x The limit of Rλ x =λ′ x / Uλ x , other variables are expressed in the same way.

6. A flight calibration system for star sensor and inertial unit structural parameters, characterized in that: include: A setting module is used to set the sensor accuracy, seeing, attitude increment integration time and pre-calibrated values ​​and uncertainties of the structural parameters in the calibration according to the data manual and the weather conditions during calibration, and to set the initial state of the structural parameter deviation estimator according to the above parameters; A time propagation module is used to perform time propagation of the estimated values ​​and covariance of the gyroscope's constant bias, bias drift, and structural parameter deviation according to the parameter model after the inertial unit outputs new angular velocity measurements and the star sensor outputs new star observation information; A construction module is used to construct the attitude increments of the inertial unit and the star sensor using the obtained angular velocity measurements and star observation information, and then construct the measurement equations of the structural parameter deviations and the measurement noise covariance; An iterative estimation module is used to solve the corresponding regularized least squares problem based on the measurement equation using a regularized robust filtering framework, complete the estimated value update and estimated covariance update of the constant zero bias, zero bias drift and structural parameter deviation, and realize iterative estimation; The compensation module is used to compensate the structural parameters after obtaining the estimated value of the stable structural parameter deviation.

Images