Satellite autonomous attitude determination method

文档序号：83770 发布日期：2021-10-08 浏览：30次中文

阅读说明：本技术 一种卫星自主姿态确定方法 (Satellite autonomous attitude determination method ) 是由翟峻仪程春晓辛星霍宏伟黄丽雅金震王鹏飞于 2021-05-24 设计创作，主要内容包括：本发明一个实施例公开了一种卫星自主姿态确定方法,包括：S100、通过各敏感器测量信息进行敏感器姿态解算得到姿态信息,其中,各敏感器姿态解算包括陀螺递推姿态解算、双星敏姿态解算、单星敏姿态解算、太敏和地敏双矢量姿态解算、太敏和磁双矢量姿态解算以及地敏和磁双矢量解算；S102、将步骤S100中经陀螺和其他敏感器姿态解算得到的姿态信息进行有效组合完成滤波姿态解算,进一步得到连续的姿态信息,其中,滤波姿态解算包括互补滤波姿态解算和扩展卡尔曼滤波姿态解算即EKF姿态解算；S104、制定数据判断选用规则,考虑各敏感器及滤波定姿的精度,设计数据判断流程,在步骤S102的互补滤波姿态解算和EKF姿态解算的结果中选取最优化姿态信息。(One embodiment of the invention discloses a satellite autonomous attitude determination method, which comprises the following steps: s100, carrying out sensor attitude calculation through the measurement information of each sensor to obtain attitude information, wherein the sensor attitude calculation comprises gyro recursion attitude calculation, double-star-sensitive attitude calculation, single-star-sensitive attitude calculation, space-sensitive and ground-sensitive double-vector attitude calculation, space-sensitive and magnetic double-vector attitude calculation and ground-sensitive and magnetic double-vector calculation; s102, effectively combining attitude information obtained by resolving the attitude of the gyroscope and other sensors in the step S100 to complete filtering attitude resolving, and further obtaining continuous attitude information, wherein the filtering attitude resolving comprises complementary filtering attitude resolving and extended Kalman filtering attitude resolving, namely EKF attitude resolving; and S104, making a data judgment selection rule, designing a data judgment process by considering the accuracy of each sensor and the filtering attitude determination, and selecting the optimized attitude information from the complementary filtering attitude calculation and EKF attitude calculation results in the step S102.)

1. A method for autonomous attitude determination for a satellite, comprising:

s100, carrying out sensor attitude calculation through the measurement information of each sensor to obtain attitude information, wherein the sensor attitude calculation comprises gyro recursion attitude calculation, double-star-sensitive attitude calculation, single-star-sensitive attitude calculation, space-sensitive and ground-sensitive double-vector attitude calculation, space-sensitive and magnetic double-vector attitude calculation and ground-sensitive and magnetic double-vector calculation;

s102, effectively combining attitude information obtained by resolving the attitude of the gyroscope and other sensors in the step S100 to complete filtering attitude resolving, and further obtaining continuous attitude information, wherein the filtering attitude resolving comprises complementary filtering attitude resolving and extended Kalman filtering attitude resolving, namely EKF attitude resolving;

and S104, making a data judgment selection rule, designing a data judgment process by considering the accuracy of each sensor and the filtering attitude determination, and selecting the optimized attitude information from the complementary filtering attitude calculation and EKF attitude calculation results in the step S102.

2. The method of claim 1,

when the sensor attitude is calculated, besides gyro recursion attitude calculation, only one attitude calculation method is adopted according to the priority, and the sensor priority is double-star sensitivity, single-star sensitivity, space sensitivity, ground sensitivity, space sensitivity and magnetism or ground sensitivity and magnetism in turn.

3. The method of claim 1,

the complementary filtering algorithm has the following formula:

q_err＝q^-1×q_t

q＝q×q_tmp

in the formula, q is an attitude quaternion obtained by gyro recursion calculation; q. q.s_tAttitude quaternion obtained by resolving satellite sensitivity, satellite sensitivity and earth sensitivity, satellite sensitivity and magnetism or earth sensitivity and magnetism; q. q.s_errThe difference value of the posture quaternion of the two is obtained; k is a complementary filter gain coefficient; q _ tmp is q_errThe normalized value of (a).

4. The method of claim 1,

during the EKF attitude calculation, four EKF modes are adopted, including: and selecting an optimal filtering scheme to carry out attitude calculation to attitude information according to a filtering effectiveness judgment basis.

5. The method of claim 4, wherein the step of performing EKF attitude solution comprises:

s1020, determining measurement models of all sensors;

s1021, determining various EKF mode calculation methods;

and S1022, EKF validity judgment is carried out, validity of the four EKF modes is judged according to a preset rule, and an optimal item is selected.

6. The method of claim 5,

the predetermined rule is as follows: judging whether the failure time of the sensor exceeds the limit; if the sensor failure time length is not over the limit, judging whether the EKF is converged, and if the norm of the system estimation mean square error matrix is stable and is smaller than a set threshold value within a certain time, converging the EKF; if the EKF is converged and the star sensor attitude determination result is effective, comparing the EKF result with the star sensor attitude determination result, and if the error is not over limit, considering the EKF to be effective, preferentially using the filtering result.

7. The method of claim 1,

according to the data judgment and selection rule, the optimal precision of double-star-sensitive attitude determination is obtained through comprehensive calculation, then the star-sensitive EKF attitude determination is carried out, and then the satellite-sensitive EKF attitude determination and the ground-sensitive EKF attitude determination are carried out sequentially, the single-star-sensitive attitude determination is carried out sequentially, the satellite-sensitive EKF attitude determination and the ground-sensitive EKF attitude determination are carried out sequentially, and the satellite-sensitive EKF attitude determination and the ground-sensitive EKF attitude determination are carried out sequentially, or the satellite-sensitive EKF attitude determination and the magnetic EKF attitude determination are carried out sequentially, or the ground-sensitive EKF attitude determination and the magnetic EKF attitude determination are carried out sequentially.

8. The method of claim 7,

if the EKF attitude determination is selected, assigning angular velocity information and attitude information obtained by the EKF calculation to corresponding system information, and if the sensor attitude determination is adopted, the data obtained by the EKF calculation is not needed; and when the sensor attitude determination is invalid and the EKF is not converged, attitude information resolved by the EKF is still adopted.

9. A computer-readable storage medium, on which a computer program is stored which, when being executed by a processor, carries out the method according to any one of claims 1-8.

10. A computer device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, characterized in that the processor implements the method according to any of claims 1-8 when executing the program.

Technical Field

The invention relates to the technical field of spaceflight. And more particularly, to a satellite autonomous attitude determination method, a computer device, and a computer-readable storage medium.

Background

With the advance of aerospace industry, aerospace missions present a flexible and diversified development situation, and small satellites have the characteristics of high function density, low development and emission cost, small size, light weight, short development period, flexible emission means, easiness in networking and the like, the space application demand is gradually increased, and the marketization and miniaturization become the mainstream trend of future development of satellites. The control system is used as a core part for reliable operation of the satellite, and more accurate attitude input is very important to obtain.

To ensure reliable performance of satellite tasks, satellites are typically equipped with a variety of measurement sensors, such as star sensors, sun sensors, and infrared horizon. At present, a star sensor is usually adopted as a main sensor for determining the attitude, a gyroscope is adopted as a basic angular velocity measuring device, however, the star sensor can not provide continuous attitude information, and the problems of unstable measurement and the like caused by loss of star maps exist; the high-precision gyroscope has the problems of complex structure, heavy weight and large volume. The sensors are low-cost medium-low precision devices under the limit of the conditions of the cost, the weight and the like of the small satellite. Therefore, the high-performance and high-reliability attitude determination target is realized by designing with the aid of medium-low precision sensors, and the high-performance and high-reliability attitude determination target is an urgent need for rapid development of small satellites.

Disclosure of Invention

The invention aims to provide a satellite autonomous attitude determination method. To solve at least one of the problems of the prior art.

In order to achieve the purpose, the invention adopts the following technical scheme:

in a first aspect, the present invention provides a method for determining an autonomous attitude of a satellite, including:

In one embodiment of the present invention, the substrate is,

the complementary filtering algorithm has the following formula:

q_err＝q^-1×q_t

q＝q×q_tmp

in the formula, q is an attitude quaternion obtained by gyro recursion calculation; q. q.s_tAttitude quaternion obtained by resolving satellite sensitivity, satellite sensitivity and earth sensitivity, satellite sensitivity and magnetism or earth sensitivity and magnetism; q. q.s_errThe difference value of the posture quaternion of the two is obtained; k is complementary filteringA gain factor; q _ tmp is q_errThe normalized value of (a).

In one embodiment of the present invention, the substrate is,

In one embodiment, the step of performing EKF attitude solution comprises:

s1020, determining measurement models of all sensors;

s1021, determining various EKF mode calculation methods;

and S1022, EKF validity judgment is carried out, validity of the four EKF modes is judged according to a preset rule, and an optimal item is selected.

In one embodiment of the present invention, the substrate is,

In a second aspect, the present invention also provides a computer readable storage medium having stored thereon a computer program which, when executed by a processor, performs the method as provided in the first aspect of the application.

In a third aspect, the present invention also provides a computer device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, wherein the processor executes the program to implement the method as provided in the first aspect of the present application.

The invention has the following beneficial effects:

the method for determining the autonomous attitude of the satellite can effectively avoid attitude information disorder caused by the fault of a single sensor and ensure the stability and reliability of the attitude. According to the satellite autonomous attitude determination method, high-precision attitude information is obtained through data fusion processing and resolving by means of the medium-low precision sensor, and the requirement of satellite service operation is met.

Drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings needed to be used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts.

FIG. 1 shows a flow diagram of a method for satellite autonomous attitude determination, according to one embodiment of the invention.

Fig. 2 shows a schematic diagram of a satellite autonomous attitude determination method according to an embodiment of the invention.

Figure 3 shows a satellite orbital system and a body system schematic according to one embodiment of the invention.

FIG. 4 shows a complementary filter attitude solution diagram according to one embodiment of the invention.

FIG. 5 illustrates an EKF attitude resolution flow diagram in accordance with one embodiment of the invention.

FIG. 6 illustrates an EKF attitude solution schematic in accordance with one embodiment of the invention.

FIG. 7 illustrates an EKF validity determination flow diagram in accordance with one embodiment of the present invention.

FIG. 8 illustrates a schematic view of a star sensor 1 attitude solution attitude angle error in accordance with one embodiment of the present invention.

FIG. 9 shows a schematic diagram of a gyro + star sensor 1 complementary filtered attitude angle error, according to one embodiment of the present invention.

FIG. 10 is a schematic diagram of attitude angle errors in the form of four EKFs in accordance with one embodiment of the present invention

FIG. 11 illustrates a schematic diagram of attitude angle errors for the four forms of EKF after convergence, in accordance with one embodiment of the present invention.

FIG. 12 illustrates a flow diagram for pose data fusion selection in accordance with one embodiment of the present invention.

FIG. 13 shows a schematic block diagram of a computer device suitable for use in implementing embodiments of the present application.

FIG. 14 shows a schematic view of attitude angle errors in the form of a satellite sensitive EKF, according to one embodiment of the invention.

Detailed Description

In order to make the technical solutions and advantages of the present invention clearer, the following will describe embodiments of the present invention in further detail with reference to the accompanying drawings. Similar parts in the figures are denoted by the same reference numerals. It is to be understood by persons skilled in the art that the following detailed description is illustrative and not restrictive, and is not to be taken as limiting the scope of the invention.

First embodiment

As shown in fig. 1 and 2, an embodiment of the present invention discloses a method for determining an autonomous attitude of a satellite, including:

Specifically, the measurement sensors of the satellite equipment comprise 1 set of three-axis gyroscope (hereinafter referred to as gyroscope), 2 sets of star sensor (hereinafter referred to as star sensor), 1 set of three-axis sun sensor (hereinafter referred to as satellite sensor), 1 set of earth sensor (hereinafter referred to as earth sensor) and 1 set of three-axis magnetometer (hereinafter referred to as magnetism), and all the sensors belong to medium-low precision grades. The measurement accuracy of the gyroscope is set to be 0.5 degree/h, the measurement accuracies of the star sensor 1 and the star sensor 2 are respectively 5 angular seconds and 10 angular seconds, the measurement accuracies of the satellite sensor and the earth sensor are respectively 0.05 degree and 0.15 degree, and the magnetic measurement accuracy is 500 nT.

In this embodiment, the pose information is defined as the pose of the main system relative to the orbital system; the inertia system is an equatorial inertia coordinate system; orbital system O_cXoYoZoTaking the center of mass of the satellite as the origin O of the coordinate system_c，O_cZoThe axis pointing in the opposite direction of the vector of the earth's center of the satellite, O_cXoShaft passing through O_cPoint of being in contact with O_cZoIn a plane perpendicular to the axis and pointing in the direction of speed, O_cYoThe axes spiral with them according to the right hand rule. Origin of coordinates O of body system_c，O_cZbThe axis pointing in a direction normal to the floor, O_cYbPer O_cPoint of being in contact with O_cZbPointing in the vertical plane to the direction of the negative normal of the orbital plane in the standard flight attitude of the satellite, O_cXbThe axes spiral with them according to the right hand rule; the installation is determined by the coordinates of the installation position of each sensor on the satellite body. The satellite orbits and the system are schematically shown in FIG. 3.

In the present embodiment, first, the principle of dual vector orientation will be explained. It is known that there are two reference vectors V in the reference coordinate system that are not parallel to each other₁，V₂Their observed vector in the present system is U₁，U₂. The current attitude transformation matrix A satisfies U₁＝AV₁，U₂＝AV₂By utilizing the non-parallelism of the reference vector, a new orthogonal coordinate system R, S is created in the reference coordinate system and the main system, respectively, and the unit vector of each coordinate axis is

Thus, a 3-dimensional matrix M_R＝[R₁ R₂ R₃]，M_S＝[S₁ S₂ S₃]Is a direction cosine array of R and S in two coordinate systems. Meanwhile, the transformation matrix between the two coordinate systems of R and S is also A, so that the attitude transformation matrix can be obtained

In this embodiment, a description is given to a gyro recursive attitude solution, in which a gyroscope outputs an angular velocity of a satellite relative to an inertial space, and a quaternion q from an inertial system to a body system at a current time is obtained through one-step recursion by using a quaternion from the inertial system to the body system at a previous time_iAnd the quaternion q from the inertial system to the orbital system can be obtained by combining the satellite orbit information_oiTherefore, the attitude quaternion q of the satellite body under the orbital system can be obtained, and attitude information can be obtained through calculation.

In this embodiment, the double-star-sensitive attitude solution and the single-star-sensitive attitude solution are explained.

And outputting attitude quaternion of the installation coordinate system relative to an inertial space by the star sensors, and resolving by adopting double star sensors when the two star sensors work effectively, or resolving by adopting single star sensor attitude.

In the double-star-sensitive attitude calculation method, a coordinate conversion matrix of a star-sensitive installation system relative to an inertial system can be obtained according to attitude quaternion output by the star-sensitive system, namely

In the formula (I), the compound is shown in the specification,andthe matrix is a three-axis measurement vector under two satellite-sensitive mounting systems and can be simultaneously represented by coordinate axis unit vectors of the satellite-sensitive mounting systems. Assuming that the optical axes of the two star sensors are Z axes, the matrix M is installed through the star sensors_sThe coordinates of the two star sensors Z axis in the inertial system and the main system can be obtained to form dual vectors,

V₁＝F_si1·[0 0 1]^T，V₂＝F_si2·[0 0 1]^T

U₁＝M_s1·[0 0 1]^T，U₂＝M_s2·[0 0 1]^T

and then, obtaining an attitude transformation matrix from the inertial system to the main system by adopting a double-vector attitude determination principle, further obtaining an attitude quaternion of the star sensor in the inertial system, and finally resolving to obtain attitude information by combining with orbit data.

In the single-satellite-sensitive attitude calculation method, attitude information can be calculated by combining orbit data with the aid of attitude quaternion of an inertial system output by satellite sensors.

In the present embodiment, the calculation of the attitude of the space-sensitive and ground-sensitive double vectors, the space-sensitive and magnetic double vectors, and the space-sensitive and magnetic double vectors is explained.

The solar vector Sb under the relative installation coordinate system of the space sensitive output; the earth sensitivity outputs an earth center vector Eb relative to the installation coordinate system, and the magnetometer outputs a geomagnetic vector Bb relative to the installation coordinate system. Combining the components So, Eo, Bo of the sun vector, geocentric vector and magnetic vector in the orbital data under the orbital system,

U₁₁＝Sb，V₁₂＝So

U₂₁＝Eb，V₂₂＝Eo

U₃₁＝Bb，V₃₂＝Bo

and (3) respectively obtaining the attitude transformation matrix of the orbit system big-disk book system by adopting a double-vector attitude determination principle, further obtaining corresponding attitude quaternions, and resolving attitude information.

In one embodiment, in the sensor attitude solution, besides the gyro recursive attitude solution, only one attitude solution method is adopted according to the priority, and the sensor priority is sequentially double star sensitivity, single star sensitivity, both earth sensitivity and earth sensitivity, both earth sensitivity and magnetism, or both earth sensitivity and magnetism.

S102, effectively combining attitude information obtained by resolving the attitude of the gyroscope and other sensors in the step S100 to complete filtering attitude resolving, and further obtaining high-precision continuous attitude information, wherein the filtering attitude resolving comprises complementary filtering attitude resolving and extended Kalman filtering attitude resolving, namely EKF attitude resolving;

specifically, the complementary filtering is a filter technology for obtaining optimized output by utilizing two or more than two inputs, and in the attitude calculation of each sensor of the satellite, the dynamic response characteristic of a gyroscope is good, but an accumulated error is generated; the star-sensitive, the earth-sensitive and the magnetic calculation postures have no accumulated error, but the dynamic response characteristic is relatively poor. Therefore, the characteristics of the two are complementary in the frequency domain, the attitude data obtained by the two can be fused by adopting a complementary filtering algorithm, and the accuracy and the dynamic reliability of the resolving attitude are improved.

In one embodiment, shown in fig. 4, complementary filtering, a filter technique that uses two or more inputs to obtain an optimized output, the inputs typically having different frequency characteristics, i.e., complementary relationships in the frequency domain, is performed for pose resolution.

The input signal x (t) comprises two complementary observed quantities in frequency domains, wherein one observed quantity has high reliability in a high-frequency region, and the other observed quantity has high reliability in a low-frequency region. Complementary filtering by high-pass filtering a first observed quantity, i.e. 1-G (t), and low-pass filtering another observed quantity, i.e. G (t), combining the two inputs to obtain an output value y (t) superior to any input,

y(t)＝[x(t)+δ(t)][1-G(t)]+[x(t)+η(t)]G(t)

＝x(t)+δ(t)[1-G(t)]+η(t)G(t)

in the formula, δ (t), η (t) are signal noise.

In the attitude calculation of each sensor of the satellite, the dynamic response characteristic of the gyroscope is good, but an accumulated error is generated; the star-sensitive, the earth-sensitive and the magnetic or the earth-sensitive and the magnetic calculation attitude have no accumulated error, but the dynamic response characteristic is relatively poor. Therefore, the characteristics of the two are complementary in the frequency domain, the attitude data obtained by the two can be fused by adopting a complementary filtering algorithm, and the accuracy and the dynamic reliability of the resolving attitude are improved.

In this embodiment, the complementary filtering algorithm has the following formula:

q_err＝q^-1×q_t

q＝q×q_tmp

Specifically, the EKF attitude solution is a state estimation method, and the state quantity is estimated according to observation information by establishing a state equation and an observation equation of state quantity change by adopting an estimation algorithm and becomes the optimal estimation under a certain criterion. The method comprises the steps of considering nonlinear factors in practical application, determining the satellite attitude by adopting an extended Kalman filtering algorithm, wherein the core of the algorithm is to carry out local linearization on a nonlinear state equation and a nonlinear measurement equation by using Taylor expansion, and then calculating by using a Kalman filtering principle.

As shown in fig. 5, the EKF algorithm can be generally divided into two phases, a prediction phase and a correction phase. The core idea is that the system state at the current moment is predicted according to the system state at the previous moment, and then the real state at the current moment is obtained by combining the observed value of the system state at the current moment and correcting. The estimation process establishes prior estimation of the current state, the correction process is responsible for feedback, the estimation state is corrected by using the current measurement variable, and the posterior estimation of state improvement is established.

In one embodiment, four EKF modes are designed in the EKF attitude calculation by considering the use characteristics of each sensor, and the four EKF modes comprise: and selecting an optimal filtering scheme to carry out attitude calculation to attitude information according to a filtering effectiveness judgment basis. FIG. 6 is an EKF attitude solution schematic.

In one embodiment, the step of performing EKF attitude solution comprises:

s1020, determining measurement models of all sensors;

gyro measurement model

ω＝ω_b+b+v_g

In the formula, omega is the output angular velocity of the gyroscope; omega_bIs the true angular velocity; b is the constant drift of the gyroscope, and b is set to satisfyv_gFor the measurement noise of the gyro, let v_gIs white noise, satisfies

Star sensitivity measuring model

The output model of the star sensor measurement is

q_bm＝q_b+v_m

In the formula, q_bmAs a star sensitivity measurement, q_bIs the true value, v_mNoise was measured for star sensitivity. And recording the satellite sensitive measurement residual error as a difference value between the satellite sensitive measurement output quaternion and an estimated value.

Δq_b＝H_mΔq+v_m

H_m＝I_3×3

In the formula, H_mIs a star sensitivity measurement matrix.

Space-sensitive measuring model

Sun vector S under the system_b＝[S_bx S_by S_bz]The sensitive measuring angle eta, xi is set

D_η＝tanη＝-S_bx/S_by

D_ξ＝tanξ＝S_bz/S_by

The output model of the space-sensitive measurement is

In the formula, v_sRecording the error of the space-sensitive measurement as the difference between the output of the space-sensitive measurement and the estimated value,

will D_η、D_ξIn thatIs developed according to Taylor formula, first order approximation is taken,

in the formula, the sun vector estimates the residual errorΔ q is the error quaternion Δ q_boThe vector portion of (2). It is possible to obtain,

in the formula, H_sIs an space sensitive measurement matrix.

Ground sensitivity measurement model

In the formula, gamma_mAnd theta_mFor earth-sensitive measurements, γ and θ are the true roll and pitch angles of the satellite, γ_biasAnd theta_biasFor ground-sensitive measurement of deviation, v_eTo measure the noise, set to white noise, satisfy

The attitude angle estimation error is considered to be a small angle and is expressed as delta gamma approximately equal to 2 delta q₁，δθ≈2Δq₂. Is provided withAndfor an estimate of the earth-sensitive measurement deviation obtained by filtering, the earth-sensitive measurement residual can then be defined as

In the formula, H_eIs a ground sensitive measurement matrix.

Magnetic measurement model

In order to facilitate the calculation and simplify the magnetic measurement model,

B_bm＝B_b+v_c

in the formula, B_bmAs a measure of the magnetic properties of the system, B_bIs a true magnetic vector, v_cIs the magnetic measurement noise. The magnetic measurement residual is the difference between the magnetic measurement output value and the estimated value.

ΔB_b＝2[B_bm×]Δq+v_c

＝H_cΔq+v_c

In the formula, H_cIs a magnetic measurement matrix.

S1021, determining various EKF mode calculation methods;

star sensor + gyroscopic EKF

In the attitude calculation of the satellite-sensitive and gyroscope EKF, satellite attitude kinematics model is combined, satellite-sensitive measurement data is used as a measurement vector, attitude quaternion is used as a system state quantity, gyroscope measurement data is introduced into a system state equation, and filtering observation of attitude can be established. The filtered output value includes an optimal estimate of the gyro's constant drift error in addition to an optimal estimate of the attitude.

i. System equation of state

The system state quantity is selected as attitude error quaternion and gyro drift, i.e.

X^T＝[Δq₁ Δq₂ Δq₃ Δb_x Δb_y Δb_z]^T

According to the quaternion motion equation and the output model of the gyroscope, an equation expressed by error quaternion can be obtained,

Δq₀＝0

in the formula, Δ q is an error quaternion Δ q_boThe vector portion of (a) is,is an estimated value of the angular velocity of the star relative to the Chi-inertia system, and delta b is a measurement residual error of the gyro constant drift, so as to satisfy the requirements

The system state equation is

Wherein F (t), G (t) are respectively a state variance matrix and a system noise variance matrix, W_kA noise sequence is excited for the system. The subsequent definitions herein are consistent therewith.

Discretizing system state equations

X(k+1)＝Φ(k+1，k)·X(k)+Γ(k+1，k)·W

Wherein Φ (k) is I₆+ f (k) · Δ t, Γ (k) ═ g (t) · Δ t, Δ t is the sampling step, Φ (k +1, k) is t_kTime to t_k+1A one-step transfer matrix of time; Γ (k +1, k) is the system noise driving array. The subsequent definitions herein are consistent therewith.

Systematic measurement equation

The system measurement is expressed by quaternion, the residual vector part of the star sensor measurement is the difference value between quaternion measured by two star sensors and integral quaternion of the gyroscope, and is recorded as [ delta q [ ]₁ Δq₂]^T。

The system measurement equation is

Z(t)＝H·X(t)+V

In the formula, measuring arrayMeasuring noiseOutput noise of star sensor, satisfy

Taimin and Gemin + Gyroscope EKF

In the attitude calculation of the satellite attitude and the ground sensitivity plus the gyroscope EKF, the satellite attitude kinematics model is combined, the ether sensitivity measurement data and the ground sensitivity measurement data are used as measurement vectors, the attitude quaternion is used as a system state quantity, the gyroscope measurement data and the ground sensitivity measurement data are introduced into a system state equation, and the filtering observation of the attitude can be established. Besides the optimal estimation of the attitude, the filtering output value also comprises the optimal estimation of the measurement error of the gyroscope and the ground sensitive constant value.

i. System equation of state

The system state quantity is selected as attitude error quaternion, gyro drift and ground sensitive measurement error, i.e.

X^T＝[Δq₁ Δq₂ Δq₃ Δb_x Δb_y Δb_z Δγ_bias Δθ_bias]^T

The system filter state equation is

The filter state equation is discretized and,

X(k+1)＝Φ(k+1，k)·X(k)+Γ(k+1，k)·W

wherein Φ (k) is I₈+ f (k) · Δ t, Γ (k) ═ g (t) · Δ t, Δ t is the sampling step.

Systematic measurement equation

If the satellite is in the illumination area, the earth sensitivity and the earth sensitivity are normally output, and the system quantity measurement adopts the measurement residual error of the earth sensitivity and is recorded as Z^T＝[ΔD_η ΔD_ξ Δγ Δθ]^T。

The system measurement equation is

Z(t)＝H·X(t)+V

In the formula (I), the compound is shown in the specification,v_s、v_erespectively, too sensitive and ground sensitive output noise.

If the satellite is in a shadow area and is subjected to the space sensitivity failure, the system quantity measurement adopts the measurement residual error of the space sensitivity, and is recorded as Z^T＝[0 0 Δγ Δθ]^TH in the measurement matrix_s＝0，v_s0; if the earth sensitivity fails, the system quantity measurement adopts the hypersensitive measurement residual error, Z^T＝[ΔD_η ΔD_ξ 0 0]^TH in the measurement matrix_e＝0，v_e＝0。

Taimin and magnet + gyrekf

In the process of resolving the attitude of the satellite attitude kinematics model, the satellite attitude kinematics model is combined, the ether-sensitive and magnetic measurement data are used as measurement vectors, the attitude quaternion is used as a system state quantity, the gyro measurement data are introduced into a system state equation, and the filtering observation of the attitude can be established. Besides the optimal estimation of the attitude, the filtering output value also comprises the optimal estimation of the gyro constant value measurement error.

i. System equation of state

The system state quantity is selected as an attitude error quaternion and a gyro error, namely

X^T＝[Δq₁ Δq₂ Δq₃ Δb_x Δb_y Δb_z]^T

The system filter state equation is

The filter state equation is discretized and,

X(k+1)＝Φ(k+1，k)·X(k)+Γ(k+1，k)·W

wherein Φ (k) is I₆+ f (k) · Δ t, Γ (k) ═ g (t) · Δ t, Δ t is the sampling step.

Systematic measurement equation

If the satellite is in the illumination area, both the satellite sensitivity and the satellite magnetism are normally output, and the system quantity measurement adoptsThe measurement residual of the sensitivity and magnetism, denoted as Z^T＝[ΔD_η ΔD_ξ ΔB_bx ΔB_by ΔB_bz]^T。

The system measurement equation is Z (t) H.X (t) V

In the formula (I), the compound is shown in the specification,v_s、v_crespectively, the output noise of the sensitivity and the magnetism.

If the satellite is in a shadow area and is subjected to the space-sensitive failure, the system quantity measurement adopts magnetic measurement residual error which is recorded as Z^T＝[0 0 ΔB_bxΔB_by ΔB_bz]^TH in the measurement matrix_s＝0，v_s0; if the magnetic failure occurs, the system quantity measurement adopts the hypersensitive measurement residual error, Z^T＝[ΔD_η ΔD_ξ 0 0 0]^TH in the measurement matrix_c＝0，v_c＝0。

Gemini and magnetic + gyrekf

In the attitude calculation of the ground sensitivity and the magnetism + gyroscope EKF, a satellite attitude kinematics model is combined, the ground sensitivity and magnetism measurement data are used as measurement vectors, the attitude quaternion is used as a system state quantity, the gyroscope and ground sensitivity measurement data are introduced into a system state equation, and the filtering observation of the attitude can be established. Besides the optimal estimation of the attitude, the filtering output value also comprises the optimal estimation of the measurement error of the gyroscope and the ground sensitive constant value.

i. System equation of state

The system state quantity is selected as attitude error quaternion, gyro drift and ground sensitive measurement error, i.e.

X^T＝[Δq₁ Δq₂ Δq₃ Δb_x Δb_y Δb_z Δγ_bias Δθ_bias]^T

The system filter state equation is

The filter state equation is discretized and,

X(k+1)＝Φ(k+1，k)·X(k)+Γ(k+1，k)·W

wherein Φ (k) is I₈+ f (k) · Δ t, Γ (k) ═ g (t) · Δ t, Δ t is the sampling step.

Systematic measurement equation

If the earth sensitivity and the magnetism are effective, the measurement residual error of the earth sensitivity and the magnetism is adopted for the system quantity measurement and is recorded as Z^T＝[Δγ Δθ ΔB_bx ΔB_by ΔB_bz]^T。

The system measurement equation is Z (t) H.X (t) V

In the formula (I), the compound is shown in the specification,v_e、v_crespectively ground-sensitive and magnetic output noise.

If the earth sensitivity fails, the system quantity measurement adopts magnetic measurement residual error, and is recorded as Z^T＝[0 0 ΔB_bx ΔB_by ΔB_bz]^TH in the measurement matrix_s＝0，v_s0; if the magnetic failure occurs, the system quantity measurement adopts the ground-sensitive measurement residual error, Z^T＝[Δγ Δθ 0 0 0]^TH in the measurement matrix_c＝0，v_c＝0。

And S1022, EKF validity judgment is carried out, validity of the four EKF modes is judged according to a preset rule, and an optimal item is selected.

In the filtering attitude calculation, four EKF modes can be calculated to obtain attitude information, corresponding rules need to be formulated to judge the effectiveness of the EKF modes, and the optimal items are selected. In this embodiment, as shown in fig. 7, the predetermined rule is: judging whether the failure time of the sensor exceeds the limit; if the sensor failure time length is not over the limit, judging whether the EKF is converged, and if the norm of the system estimation mean square error matrix is stable and is smaller than a set threshold value within a certain time, converging the EKF; if the EKF is converged and the star sensor attitude determination result is effective, comparing the EKF result with the star sensor attitude determination result, and if the error is not over limit, considering the EKF to be effective, preferentially using the filtering result.

And S104, making a data judgment selection rule, designing a data judgment process by considering the accuracy of each sensor and the filtering attitude determination, and selecting the optimized attitude information from the complementary filtering attitude calculation and EKF attitude calculation results in the step S102. The attitude data fusion selection flowchart is shown in fig. 12.

For star sensors, the optical axis error is relatively large, and a measure of optical axis error compensation is adopted in the double star sensors attitude determination of the sensors; in order to ensure that data in the satellite sensitive filtering is used maximally, the filtering observation quantity is satellite sensitive original data, and optical axis errors are not eliminated, in a specific embodiment, according to the data judgment and selection rule, the double satellite sensitive attitude determination precision is obtained through comprehensive calculation, then satellite sensitive EKF attitude determination is carried out, and then satellite sensitive EKF attitude determination and ground sensitive EKF attitude determination, single satellite sensitive attitude determination, satellite sensitive and ground sensitive attitude determination, satellite sensitive and magnetic EKF attitude determination, ground sensitive and magnetic EKF attitude determination, satellite sensitive and magnetic attitude determination or ground sensitive and magnetic attitude determination are carried out in sequence.

In one particular embodiment of the present invention,

The present invention will be further described with reference to specific examples.

In this example, the satellite is in a tilted circular orbit with an orbit height of 1125km, an inclination of 86.5 °, and a period of 108.8 min. In the satellite equipment, 1 set of gyroscope, 2 sets of star sensor, 1 set of satellite sensor, 1 set of earth sensor and 1 set of magnetometer with low precision are used as attitude measurement sensors. The simulation time length is 6528s of a track period, and the calculation period is 0.25 s. Setting an initial attitude angle as [5 degrees ] in a three-axis ground-to-ground stable state; 5 degrees; 5 DEG ] and the angular velocity is 0, and the initial settings of all sensors are effective except for the star sensor 2.

S100, carrying out sensor attitude calculation through the measurement information of each sensor to obtain attitude information.

In the present example, the star sensor attitude resolution portion employs star sensor 1 attitude resolution. The star sensor 1 outputs quaternion of the installation system relative to the inertial system, the quaternion of the system relative to the inertial system can be obtained through an installation matrix of the star sensor 1 under the system, and attitude information can be obtained through calculation by combining with orbit data.

It can be seen from fig. 8 that the error of the star sensor 1 attitude solution is within 0.03 °.

specifically, in the example, the sensor attitude determination accuracy is considered, and gyro + star sensor 1 complementary filtering attitude calculation and star sensor EKF filtering attitude calculation are adopted in the filtering attitude calculation part.

Carrying out complementary filtering attitude calculation, gyroscope + star sensor 1 complementary filtering, attitude quaternion q obtained by gyroscope recursion calculation, and attitude quaternion q obtained by star sensor 1 calculation_tThe gain factor k is set to 0.1,

q_err＝q^-1×q_t

q＝q×q_tmp

Therefore, as shown in fig. 9, the attitude quaternion after the complementary filtering data fusion processing can be obtained, and the attitude information is finally obtained by calculation in combination with the orbit data, and the error is within 0.005 °.

Attitude angle solutions of four EKF forms were performed, respectively, and the results are shown in fig. 10. As shown in FIG. 11, after convergence, the attitude error of the satellite sensitive EKF is within 0.001, and the attitude errors of the remaining three filtered terms are within 0.01.

In this example, according to the attitude data fusion selection flow chart shown in fig. 12, the satellite-sensitive EKF is finally selected as the attitude solution, and the high-precision attitude angle information is obtained, and the attitude angle error in the form of the satellite-sensitive EKF is shown in fig. 14.

The method for determining the multimode redundant autonomous attitude of the small satellite can effectively avoid attitude information disorder caused by single sensor faults and ensure the stability and reliability of the attitude. According to the multimode redundant autonomous attitude determination method for the small satellite, high-precision attitude information is obtained through data fusion processing and resolving by means of a medium-low precision sensor, and the requirement for satellite service operation is met.

Second embodiment

Fig. 13 shows a schematic structural diagram of a computer device according to another embodiment of the present application. The computer device 50 shown in fig. 13 is only an example, and should not bring any limitation to the functions and the scope of use of the embodiments of the present application. As shown in fig. 13, the computer device 50 is in the form of a general purpose computing device. The components of computer device 50 may include, but are not limited to: one or more processors or processing units 500, a system memory 516, and a bus 501 that couples various system components including the system memory 516 and the processing unit 500.

Bus 501 represents one or more of any of several types of bus structures, including a memory bus or memory controller, a peripheral bus, an accelerated graphics port, and a processor or local bus using any of a variety of bus architectures. By way of example, such architectures include, but are not limited to, Industry Standard Architecture (ISA) bus, micro-channel architecture (MAC) bus, enhanced ISA bus, Video Electronics Standards Association (VESA) local bus, and Peripheral Component Interconnect (PCI) bus.

Computer device 50 typically includes a variety of computer system readable media. Such media may be any available media that is accessible by computer device 50 and includes both volatile and nonvolatile media, removable and non-removable media.

The system memory 516 may include computer system readable media in the form of volatile memory, such as Random Access Memory (RAM)504 and/or cache memory 506. The computer device 50 may further include other removable/non-removable, volatile/nonvolatile computer system storage media. By way of example only, storage system 508 may be used to read from and write to non-removable, nonvolatile magnetic media (not shown in FIG. 13, often referred to as a "hard disk drive"). Although not shown in FIG. 13, a magnetic disk drive for reading from and writing to a removable, nonvolatile magnetic disk (e.g., a "floppy disk") and an optical disk drive for reading from or writing to a removable, nonvolatile optical disk (e.g., a CD-ROM, DVD-ROM, or other optical media) may be provided. In these cases, each drive may be connected to the bus 501 by one or more data media interfaces. Memory 516 may include at least one program product having a set (e.g., at least one) of program modules that are configured to carry out the functions of embodiment one.

A program/utility 510 having a set (at least one) of program modules 512 may be stored, for example, in memory 516, such program modules 512 including, but not limited to, an operating system, one or more application programs, other program modules, and program data, each of which examples or some combination thereof may comprise an implementation of a network environment. Program modules 512 generally perform the functions and/or methodologies of the embodiments described herein.

Computer device 50 may also communicate with one or more external devices 70 (e.g., keyboard, pointing device, display 60, etc.), with one or more devices that enable a user to interact with the computer device 50, and/or with any devices (e.g., network card, modem, etc.) that enable the computer device 50 to communicate with one or more other computing devices. Such communication may occur via input/output (I/O) interfaces 502. Also, computer device 50 may communicate with one or more networks (e.g., a Local Area Network (LAN), a Wide Area Network (WAN) and/or a public network, such as the Internet) through network adapter 514. As shown in FIG. 13, network adapter 514 communicates with the other modules of computer device 50 via bus 501. It should be appreciated that although not shown in FIG. 13, other hardware and/or software modules may be used in conjunction with computer device 50, including but not limited to: microcode, device drivers, redundant processing units, external disk drive arrays, RAID systems, tape drives, and data backup storage systems, among others.

The processor unit 500 executes various functional applications and data processing by executing programs stored in the system memory 516, for example, to implement a method for determining an autonomous attitude of a satellite according to an embodiment of the present application.

Aiming at the existing problems, the method for determining the satellite autonomous attitude makes computer equipment suitable for a satellite autonomous attitude determination method, effectively avoids attitude information disorder caused by single sensor faults, and ensures stability and reliability of the attitude.

Third embodiment

Another embodiment of the present application provides a computer-readable storage medium, on which a computer program is stored, which when executed by a processor implements the method provided by the first embodiment. In practice, the computer-readable storage medium may take any combination of one or more computer-readable media. The computer readable medium may be a computer readable signal medium or a computer readable storage medium.

A computer readable storage medium may be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the foregoing. More specific examples (a non-exhaustive list) of the computer readable storage medium would include the following: an electrical connection having one or more wires, a portable computer diskette, a hard disk, a Random Access Memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing. In the present embodiment, a computer readable storage medium may be any tangible medium that can contain, or store a program for use by or in connection with an instruction execution system, apparatus, or device.

A computer readable signal medium may include a propagated data signal with computer readable program code embodied therein, for example, in baseband or as part of a carrier wave. Such a propagated data signal may take many forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. A computer readable signal medium may also be any computer readable medium that is not a computer readable storage medium and that can communicate, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device.

Program code embodied on a computer readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable, RF, etc., or any suitable combination of the foregoing. Computer program code for carrying out operations for aspects of the present application may be written in any combination of one or more programming languages, including an object oriented programming language such as Java, Smalltalk, C + +, and conventional procedural programming languages, such as the "C" programming language or similar programming languages. The program code may execute entirely on the user's computer, partly on the user's computer, as a stand-alone software package, partly on the user's computer and partly on a remote computer or entirely on the remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a Local Area Network (LAN) or a Wide Area Network (WAN), or the connection may be made to an external computer (for example, through the Internet using an Internet service provider).

It should be understood that the above-mentioned embodiments of the present invention are only examples for clearly illustrating the present invention, and are not intended to limit the embodiments of the present invention, and it will be obvious to those skilled in the art that other variations or modifications may be made on the basis of the above description, and all embodiments may not be exhaustive, and all obvious variations or modifications may be included within the scope of the present invention.

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Satellite autonomous attitude determination method

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