High-precision all-zero Doppler attitude guidance control method under satellite three-axis arbitrary attitude

Abstract

The invention discloses a high-precision all-zero Doppler attitude guidance control method under any three-axis attitude of a satellite. The method comprises the following steps: calculating a two-dimensional guidance attitude under an orbit system; calculating a bias target attitude quaternion according to a ground pointing demand; calculating a guidance attitude quaternion expressed under the satellite body system; calculating attitude quaternion deviation and angular velocity deviation for control; and calculating a control instruction torque to realize attitude guidance control. According to the invention, high-precision all-zero Doppler center frequency two-dimensional guide attitude control can be realized.

Description

Technical Field

[0001] This invention relates to satellite attitude control technology, primarily a high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes. It belongs to the field of spacecraft attitude control technology. Background Technology

[0002] A synthetic aperture radar satellite is an Earth observation satellite that uses a spaceborne synthetic aperture radar (SAR) as its payload.

[0003] Doppler center frequency f D This is one of the most critical parameters in SAR satellite data processing. It is related to the relative motion between the satellite and the ground target and is greatly affected by the Earth's rotation. For X-band SAR satellites, the variation globally is approximately ±23 kHz, which poses difficulties for imaging processing. Therefore, motion compensation is required, and applying yaw guidance to the satellite's attitude motion has proven to be a feasible solution. If the effect of orbital eccentricity is considered, pitch guidance is added, and together with yaw guidance, they are referred to as "two-dimensional guidance."

[0004] Many SAR satellites are in a fixed side-looking state, and their two-dimensional guidance and attitude control is relatively simple, making them unsuitable for zero-Doppler attitude guidance and control under arbitrary attitudes. Summary of the Invention

[0005] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes. For microwave imaging satellites with arbitrary attitude offsets, the method calculates the two-dimensional guidance law of arbitrary offset attitude based on imaging target information and orbit information, and realizes high-precision two-dimensional guidance attitude control with all-zero Doppler center frequency.

[0006] The technical solution of this invention is:

[0007] A high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes includes:

[0008] Step 1: Calculate the two-dimensional guidance attitude in the orbital system and obtain the yaw guidance angle ψ. c Pitch guidance angle θ c ;

[0009] Step 2: Calculate the offset target attitude quaternion q based on the ground pointing requirements. m ;

[0010] Step 3: Based on the obtained yaw guidance angle ψ c Pitch guidance angle θ c and the biased target pose quaternion q m Calculate the quaternion q of the guidance attitude in the satellite's intrinsic system. d ;

[0011] Step 4: Based on the yaw guidance angle ψ c Pitch guidance angle θ c And the guidance attitude quaternion q represented in the satellite system d Calculate the attitude quaternion deviation q used for control. e and angular velocity deviation ω e ;

[0012] Step 5: Based on the attitude quaternion deviation q used for control e and angular velocity deviation ω e Calculate the control command torque T c This enables attitude guidance and control.

[0013] Preferably, step one obtains the yaw guidance angle ψ. c The method is as follows:

[0014]

[0015] Where i is the orbital inclination, u is the latitudinal argument, ω0 is the satellite's real-time orbital angular velocity in orbit, and ω e It is the Earth's angular velocity of rotation, ω e =7.2921158×10 -5 rad / s; N is the ratio of the satellite's real-time orbital angular velocity to the Earth's rotational angular velocity.

[0016] Preferably, step one obtains the pitch guidance angle θ. c The method is as follows:

[0017]

[0018] Where e is the orbital eccentricity and f is the true anomaly angle.

[0019] Preferably, step two involves calculating the biased target attitude quaternion q. m The method is as follows:

[0020]

[0021] in, From a downward perspective, θ L It is an oblique perspective. The perspective is from the left and right sides. θ L and Obtained from the input of the superior; θ bm This refers to the elevation offset angle of the satellite system.

[0022] Preferably, step three involves calculating the quaternion q representing the satellite's attitude in its own system. d The method is as follows:

[0023]

[0024] in, q represents quaternion multiplication; c Let be the quaternion representing the guidance attitude in the orbital system.

[0025] Preferably, step four involves calculating the attitude quaternion deviation q for control. e The method is as follows:

[0026]

[0027] Where, q bo Let be the attitude quaternion of the satellite body relative to the orbital system.

[0028] Preferably, the angular velocity deviation ω in step four e The method is as follows:

[0029]

[0030] ω bo ω represents the angular velocity of the satellite relative to its orbital system in the satellite's own system. c The guiding angular velocity is expressed in the orbital frame;

[0031] remember The corresponding rotation matrix:

[0032]

[0033] in, Represent θ respectively c ψ c Take the derivative with respect to time; q4 is a scalar.

[0034] Preferably, step five involves calculating the control command torque T. c The method is as follows:

[0035]

[0036] Among them, K, D, K int For the controller parameters related to the satellite's moment of inertia, K = (0.01-1)I, D = (0.1-1.5)I, K int = (0.01-0.1)K, where I is the satellite's principal inertia matrix;

[0037] Indicates K int ∫ qe vd t The amplitude is limited according to Tmax_int, where Tmax_int ranges from 0.01 to 0.1 Nm;

[0038] q ev_vsat For q ev =[q e1 q e2 q e3 ] T According to the quaternion limit value q emax The value for vector limiting output is calculated as follows:

[0039]

[0040] Where max(q) e1 | / q emax ,|q e2 | / q emax ,|q e3 | / q emax ,1) means taking |q e1 | / q emax ,|q e2 | / q emax ,|q e3 | / q emax The largest value in 1;

[0041] q ev The attitude quaternion deviation q e The quaternion vector part, q e =[q ev q e4 ] T =[q e1 q e2 q e3 q e4 ] T ;

[0042] q emax This is the quaternion limit value, ranging from 0.01 to 0.2.

[0043] The advantages of this invention compared to the prior art are:

[0044] This invention converts the two-dimensional guidance attitude angles and offset attitude angles represented in the orbital frame into attitude quaternions. The guidance attitude quaternions represented in the satellite's own frame are calculated by attitude quaternion multiplication. The quaternion deviation is calculated and a quaternion vector limiting control algorithm is used for satellite attitude control. The attitude quaternion deviation is limited to produce attitude angular velocity limiting effect, thus realizing high-precision all-zero Doppler two-dimensional guidance control of the satellite in any three-axis attitude. Attached Figure Description

[0045] Figure 1 This diagram illustrates the comparison between traditional fixed-side-looking SAR satellites and arbitrary-attitude pointing SAR satellites.

[0046] Figure 2 A graph showing the guidance angle of a SAR satellite with a fixed side-looking attitude.

[0047] Figure 3 A graph showing the guidance angle curves for SAR satellites pointing in any orientation. Detailed Implementation

[0048] This invention provides a high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes, comprising:

[0049] Step 1: Obtain the yaw guidance angle ψ c Pitch guidance angle θ c ;

[0050] Yaw guidance angle ψ c :

[0051]

[0052] Where i is the orbital inclination, u is the latitudinal argument, ω0 is the satellite's real-time orbital angular velocity in orbit, and ω e It is the Earth's angular velocity of rotation, ω e =7.2921158×10 -5 rad / s. N is the ratio of the satellite's real-time orbital angular velocity to the Earth's rotational angular velocity.

[0053] Pitch guidance angle θ c :

[0054]

[0055] Where e is the orbital eccentricity and f is the true anomaly angle.

[0056] For SAR satellites with a fixed side-looking attitude, the two-dimensional guidance in the orbital frame and the two-dimensional guidance in the fixed side-looking coordinate system are consistent, such as... Figure 2 As shown.

[0057] Step 2: Calculate the offset target attitude quaternion based on the ground pointing requirements;

[0058] Pointing a SAR satellite in any orientation requires specifying the following viewpoint. Oblique angle θ L Left and right side views like Figure 1 As shown.

[0059] Based on the known input from the lower perspective Oblique angle θ L Left and right side views Calculate the satellite's intrinsic elevation offset angle θ bm :

[0060]

[0061] Calculate the biased target attitude quaternion:

[0062]

[0063] Step 3: Calculate the quaternion q representing the satellite's attitude in its own system. d ;

[0064]

[0065] in, This represents quaternion multiplication. q c Let be the quaternion representing the guidance attitude in the orbital system.

[0066] For a SAR satellite pointing in any attitude, the guidance attitude quaternion q will be... d The conversion to attitude angles involves components along all three axes, rather than just the pitch and yaw axes. Figure 3 As shown.

[0067] Step 4: Calculate the attitude quaternion deviation q for control. e and angular velocity deviation ω e ;

[0068] Attitude quaternion deviation q e for:

[0069]

[0070] Where, q bo The attitude quaternion of the satellite body relative to the orbital system;

[0071] Angular velocity deviation ω e for:

[0072]

[0073] ω bo ω represents the angular velocity of the satellite relative to its orbital system in the satellite's own system. c The guiding angular velocity is expressed in the orbital frame;

[0074] remember Its corresponding rotation matrix:

[0075]

[0076] in, Represent θ respectively c ψ c Take the derivative with respect to time. q4 is a scalar.

[0077] Step 5: Calculate the control command torque to achieve attitude guidance control;

[0078]

[0079] Among them, K, D, K int For the controller parameters related to the satellite's moment of inertia, K = (0.01-1)I, D = (0.1-1.5)I, K int = (0.01-0.1)K, where I is the satellite's principal inertia matrix; Indicates K int ∫q ev d t The amplitude is limited according to Tmax_int, which is typically set to 0.01-0.1 Nm; q ev_vsat For q ev =[q e1 q e2 q e3 ] T According to q emax The value for vector limiting output is calculated as follows:

[0080]

[0081] Where max(q) e1 | / q emax ,|q e2 | / q emax ,|q e3 | / q emax ,1) means taking |q e1 | / q emax ,|q e2 | / q emax ,|q e3 | / q emax The largest value in 1.

[0082] q ev The attitude quaternion deviation q e The quaternion vector part, q e =[q ev q e4 ] T =[q e1 q e2 q e3 q e4 ] T ;

[0083] q emax This is the quaternion limit value, typically taken as 0.01-0.2.

[0084] Example

[0085] The satellite's three-axis rotational inertia is [400 600 800] kgm 2 The orbital inclination is 35°, the eccentricity is 0.001, the perigee argument is 90°, the side angle is 30°, the downward angle is 30°, and the oblique angle is 25°.

[0086] Step 1: Calculate the two-dimensional guidance attitude in the orbital system

[0087] Yaw guidance angle:

[0088]

[0089] Where i = 35 * pi / 180, and u is the latitude argument; ω0 = 0.0011; ω e It is the Earth's angular velocity of rotation, ω e =7.2921158×10 -5 rad / s.

[0090] Pitch guidance angle:

[0091]

[0092] Where e = 0.001, and f is the true anterior angle.

[0093] Step 2: Calculate the offset target attitude quaternion based on the ground requirements.

[0094] According to the lower perspective θ L Calculate the elevation offset angle θ of the satellite's intrinsic system. bm

[0095]

[0096] Calculate the biased target attitude quaternion

[0097]

[0098] in, θ L =25*pi / 180, θ=25*pi / 180.

[0099] Step 3: Calculate the quaternion q of the satellite's guidance attitude in its own system. d

[0100]

[0101] in, This represents quaternion multiplication.

[0102] Step 4: Calculate the attitude quaternion deviation and angular velocity deviation for control.

[0103] Attitude quaternion deviation is

[0104]

[0105] Angular velocity deviation is

[0106] ω e =ω bo -A bc ω c

[0107] in, A(q) represents the quaternion q = [q1q2q3q4] T The rotation matrix corresponding to (q4 is a scalar)

[0108]

[0109]

[0110] Step 5: Calculate the control command torque to achieve attitude guidance control.

[0111]

[0112] Where K = [293 450 574] T D = [391 600 766] T ;K int =[14 22 28] T , This indicates that the amplitude is limited according to Tmax_int, where Tmax_int = 0.01 Nm; q ev_vsat For q ev =[q e1 q e2 q e3 ] T According to q emax =0.07 is the value used for vector limiting output, and the calculation method is as follows:

[0113]

[0114] Here, max(a,b,c,d) means taking the largest value among a,b,c,d.

[0115] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any person skilled in the art can make possible variations and modifications to the technical solutions of the present invention using the disclosed methods and techniques without departing from the spirit and scope of the invention. Therefore, any simple modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the content of the technical solutions of the present invention, shall fall within the protection scope of the present invention. Where there is no conflict, the embodiments of this application and the technical features thereof can be combined with each other.

[0116] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes, characterized in that, include: Step 1: Calculate the two-dimensional guidance attitude in the orbital system and obtain the yaw guidance angle ψ. c Pitch guidance angle θ c ; Step 2: Calculate the offset target attitude quaternion q based on the ground pointing requirements. m ; Step 3: Based on the obtained yaw guidance angle ψ c Pitch guidance angle θ c and the biased target pose quaternion q m Calculate the quaternion q of the guidance attitude in the satellite's intrinsic system. d ; Step 4: Based on the yaw guidance angle ψ c Pitch guidance angle θ c And the guidance attitude quaternion q represented in the satellite system d Calculate the attitude quaternion deviation q used for control. e and angular velocity deviation ω e ; Step 5: Based on the attitude quaternion deviation q used for control e and angular velocity deviation ω e Calculate the control command torque T c This enables attitude guidance and control.

2. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 1, characterized in that, Step 1: Obtain the yaw guidance angle ψ c The method is as follows: Where i is the orbital inclination, u is the latitudinal argument, ω0 is the satellite's real-time orbital angular velocity in orbit, and ω e It is the Earth's angular velocity of rotation, ω e =7.2921158×10 -5 rad / s; N is the ratio of the satellite's real-time orbital angular velocity to the Earth's rotational angular velocity.

3. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 2, characterized in that, Step 1: Obtain the pitch guidance angle θ c The method is as follows: Where e is the orbital eccentricity and f is the true anomaly angle.

4. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 3, characterized in that, Step 2: Calculate the biased target attitude quaternion q m The method is as follows: in, From a downward perspective, θ L It is an oblique perspective. The perspective is from the left and right sides. θ L and Obtained from the input of the superior; θ bm This refers to the elevation offset angle of the satellite system.

5. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 4, characterized in that, Step 3: Calculate the quaternion q representing the satellite's attitude in its own system. d The method is as follows: in, q represents quaternion multiplication; c Let be the quaternion representing the guidance attitude in the orbital system.

6. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to any one of claims 1-5, characterized in that, Step 4: Calculate the attitude quaternion deviation q for control. e The method is as follows: Where, q bo Let be the attitude quaternion of the satellite body relative to the orbital system.

7. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 6, characterized in that, Step 4 angular velocity deviation ω e The method is as follows: ω bo ω represents the angular velocity of the satellite relative to its orbital system in the satellite's own system. c The guiding angular velocity is expressed in the orbital frame; remember The corresponding rotation matrix: in, Represent θ respectively c ψ c Take the derivative with respect to time; q4 is a scalar.

8. The high-precision all-zero Doppler attitude guidance and control method for satellites in arbitrary three-axis attitudes according to claim 7, characterized in that, Step 5: Calculate the control command torque T c The method is as follows: Among them, K, D, K int For the controller parameters related to the satellite's moment of inertia, K = (0.01-1)I, D = (0.1-1.5)I, K int = (0.01-0.1)K, where I is the satellite's principal inertia matrix; Indicates K int ∫q ev d t The amplitude is limited according to Tmax_int, where Tmax_int ranges from 0.01 to 0.1 Nm; q ev_vsat For q ev =[q e1 q e2 q e3 ] T According to the quaternion limit value q emax The value for vector limiting output is calculated as follows: Where max(q) e1 | / q emax q e2 | / q emax q e3 / q emax ,1) means taking |q e1 | / q emax ,|q e2 | / q emax ,|q e3 | / q emax The largest value in 1; q ev The attitude quaternion deviation q e The quaternion vector part, q e =[q ev q e4 ] T =[q e1 q e2 q e3 q e4 ] T ; q emax This is the quaternion limit value, ranging from 0.01 to 0.2.

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