CN115140318A - Magnetic damping control method suitable for racemization of micro-nano satellite at large angular rate - Google Patents
Magnetic damping control method suitable for racemization of micro-nano satellite at large angular rate Download PDFInfo
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- CN115140318A CN115140318A CN202210851877.6A CN202210851877A CN115140318A CN 115140318 A CN115140318 A CN 115140318A CN 202210851877 A CN202210851877 A CN 202210851877A CN 115140318 A CN115140318 A CN 115140318A
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- 238000013016 damping Methods 0.000 title claims abstract description 34
- 230000006340 racemization Effects 0.000 title claims abstract description 28
- 238000000034 method Methods 0.000 title claims abstract description 21
- 239000013598 vector Substances 0.000 claims abstract description 21
- 238000005259 measurement Methods 0.000 claims description 6
- 238000010586 diagram Methods 0.000 description 5
- 230000000694 effects Effects 0.000 description 5
- 238000004088 simulation Methods 0.000 description 4
- 108091092878 Microsatellite Proteins 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 230000005358 geomagnetic field Effects 0.000 description 1
- 239000011159 matrix material Substances 0.000 description 1
- 230000000630 rising effect Effects 0.000 description 1
- 238000005096 rolling process Methods 0.000 description 1
- 238000009987 spinning Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
- B64G1/245—Attitude control algorithms for spacecraft attitude control
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/369—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using gyroscopes as attitude sensors
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- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention discloses a magnetic damping control method suitable for large-angle-rate racemization of a micro-nano satellite, which adopts a magnetometer, a gyroscope and a magnetic torquer to perform angular rate damping and comprises the following steps: a magnetometer collects a geomagnetic vector in a current state, and a gyroscope collects an angular rate at the current moment; determining the output magnetic moment of the magnetic torquer according to the magnetic damping control law; on the basis of the traditional B-dot control law, discussing the constraint condition of racemization control to obtain the racemization period of the magnetic torquer; and sending a control command to the magnetic torquer to perform magnetic damping racemization, so as to realize racemization control of the micro-nano satellite. The invention is based on the B-dot control law, obtains a novel control method by adding the discussion of constraint conditions, and has the advantage that the traditional B-dot control law cannot despun under the condition of a micro-nano satellite large angular rate.
Description
Technical Field
The invention belongs to a micro-nano satellite technology, and particularly relates to a magnetic damping control method suitable for micro-nano satellite large-angle rate racemization.
Background
In general, in order to ensure that the attitude of the initial orbiting micro/nano satellite can be rapidly oriented to the ground or to the sun, firstly, the triaxial angular velocity of the micro/nano satellite is stabilized, that is, the attitude of the micro/nano satellite is despuned. In practical engineering, because the rotational inertia of the micro-nano satellite is small, compared with devices such as an air jet thruster and a momentum wheel, the magnetic torquer is more suitable to be used as an actuating mechanism.
The magnetic torquer can generate magnetic moment after being electrified, and the magnetic moment and the geomagnetic field at the position of the micro-nano satellite interact to generate magnetic control moment for controlling the attitude of the micro-nano satellite. At present, the research of the magnetic control theory is very deep, and in the micro-nano satellite with active control of the emitted attitude in recent years, magnetic damping control is mostly adopted for attitude racemization. The method is always true under the condition that the attitude angular velocity of the micro-nano satellite is small. However, when the attitude angular velocity of the micro-nano satellite is high, the magnetic torquer outputs a constant magnetic moment vector in the satellite body coordinate system in a control period, but the geomagnetic vector B rotates along with the attitude of the micro-nano satellite, and the expression under the satellite body coordinate system always changes, so that the direction of the output moment of the magnetic torquer changes along with the change of the satellite body coordinate system, and a moment with completely opposite direction may be output in a control period, thereby causing magnetic damping failure.
Disclosure of Invention
The invention provides a magnetic damping control method suitable for racemization of a micro-nano satellite at a large angular velocity (at least 2 rad/s). On the basis of a traditional B-dot control method, by discussing an upper limit value and a lower limit value of a control period, the problems that the desired racemization effect cannot be achieved due to too short control period and the direction of an expression of a geomagnetic vector B is overturned under a satellite body coordinate system due to too long control period are avoided, so that the desired racemization effect is achieved under the condition of the large angular velocity of the micro-nano satellite.
The technical solution for realizing the invention is as follows: a magnetic damping control method suitable for micro-nano satellite large angular rate racemization is used for resolving measurement data of a gyroscope and a magnetometer on a micro-nano satellite to obtain the actually-measured angular velocity and geomagnetic vector of the gyroscope of the micro-nano satellite, and after a control period is calculated, the large angular rate racemization of the micro-nano satellite is realized by outputting torque through a magnetic torquer, and the method comprises the following specific steps:
Wherein I is the rotational inertia of the micro/nano satellite; m is 0 Is the output magnetic moment of the magnetic torquer; the geomagnetic vector B can be regarded as unchanged in the magnitude direction in an inertia system in a short time; k is a control coefficient, and k is a control coefficient,
a component representing a derivative of the geomagnetic vector B with respect to time; omega B Representing the actually measured angular speed of the gyroscope;
a component representing a derivative of a gyroscope measured angular velocity; t is 0 Representing the kinetic energy of the micro-nano satellite rotating around the mass center of the micro-nano satellite; t represents time.
And (5) turning to the step 3.
Step 5, setting a proper control period T to enable T max ＜T＜T min (ii) a And (6) sending a power-on command to the magnetic torquer, outputting the magnetic moment M = c by the magnetic torquer, and switching to the step 6.
And 6, sending a power-off instruction to the magnetic torquer after the timing time is reached according to the set control period T, and turning to the step 7.
Step 7, collecting the actually measured angular velocity of the gyroscope after the magnetic torquer is powered off
And (3) judging whether the attitude of the current micro-nano satellite meets a triaxial stable state, if so, ending the magnetic damping control, and if not, returning to the step (2) until the micro-nano satellite is in the triaxial stable state.
Compared with the prior art, the invention has the remarkable advantages that: the method can be suitable for the scene of the spinning of the micro-nano satellite at the large angular rate, and achieves the purpose of despinning the micro-nano satellite at the large angular rate by continuously updating the control period and the control law.
Drawings
Fig. 1 is a schematic diagram of a satellite body coordinate system and an inertial coordinate system.
Fig. 2 is a schematic diagram of a magnetic damping control cycle.
Fig. 3 is a comparison diagram of effect example data, wherein (a) in fig. 3 is a simulation diagram of an embodiment of the control method of the present invention, and (B) in fig. 3 is a simulation diagram of an embodiment of a conventional B-dot control.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without inventive step based on the embodiments of the present invention, are within the scope of protection of the present invention.
The following further introduces specific embodiments, technical difficulties and inventions of the present invention with reference to the design examples.
With reference to fig. 2, the magnetic damping control method for large-angle rate racemization of a micro-nano satellite according to the invention is used for resolving measurement data of a gyroscope and a magnetometer on the micro-nano satellite to obtain an actually-measured angular velocity and a geomagnetic vector of the gyroscope of the micro-nano satellite, and outputting torque through a magnetic torquer after a control period is calculated to realize large-angle rate racemization of the micro-nano satellite, and comprises the following steps:
Wherein I is the rotational inertia of the micro/nano satellite; m is 0 Is the output magnetic moment of the magnetic torquer; the geomagnetic vector B can be regarded as unchanged in the magnitude direction in the inertial system in a short time; k is a control coefficient, and k is a control coefficient,
a component representing a derivative of the geomagnetic vector B with respect to time; omega B Representing the actually measured angular speed of the gyroscope;
a component representing a derivative of a gyroscope measured angular velocity; t is 0 Kinetic energy for representing rotation of micro-nano satellite around self mass center(ii) a t represents time.
And (5) turning to the step 3.
s3-1: the coordinate system is established as in FIG. 1, where Sx i y i z i Is an inertial coordinate system with its origin S at the center of the earth, sx i The axis pointing in the equatorial plane to the vernal equinox, sz i Sy coinciding with the direction of rotation of the earth and pointing to the north pole as positive i Forming a right-hand rectangular coordinate system with the other two shafts; oxyz is t 1 A satellite body coordinate system of time, ox ' y ' z ' being t 2 A satellite body coordinate system of a moment; the origin O of the satellite body coordinate system is positioned at the centroid of the micro-nano satellite, and Ox' point to the flight direction of the micro-nano satellite and are called rolling axes; the Oz and Oz' axes are referred to as yaw axes; the Oy and Oy' axes are referred to as pitch axes.
S3-2: separately calculate t 1 And t 2 Change rate of kinetic energy of moment micro-nano satellite rotating around centroid to time t
And
then t 1 Time: m = k · (ω) B1 ×B 1 )，
t 2 Time: m' = k · (ω) B1 ×B 1 )，
Wherein, t 1 The time is the starting time of a certain control period t 2 The moment is the ending moment of the control period; m is t 1 The magnetic moment vector of the moment magnetic torquer, m' is rotated to t along with the attitude of the micro-nano satellite 2 A magnetic moment vector of the moment magnetic torquer; omega B1 Is t 1 Angular velocity, omega, of a time-of-day microsatellite B2 Is t 2 Angular velocity of the micro-nano satellite at the moment; b is 1 Is t 1 Geomagnetic vector of position of time micro-nano satellite, B 2 Is t 2 Geomagnetic vectors of positions of the micro-nano satellites at the moment; omega B1 And B 1 Are all t 1 Expression, omega, in a global coordinate system of a time satellite B2 And B 2 Are all t 2 Expressed in a satellite body coordinate system of the time.
S3-3: when m' and ω B2 ×B 2 Less than 90, i.e.
Within a short time omega B And B varies little with respect to the inertial frame, and ω B2 ×B 2 All the time with omega B Perpendicular, then ω B2 ×B 2 The included angle between m' is equivalent to omega in the control period B Induced vector rotation by an angle approximately equal to ω B ·(t 2 -t 1 ). When omega B ·(t 2 -t 1 ) When the temperature is less than 90 degrees,
at the moment, the kinetic energy of the micro-nano satellite is gradually reduced, and racemization control can be realized, so that
May be used as an upper limit of the control period.
And (5) turning to step 4.
Step 5, setting a proper control period T to enable T max ＜T＜T min (ii) a And (6) sending a power-on command to the magnetic torquer, outputting the magnetic moment M = c by the magnetic torquer, and turning to the step 6.
Step 6, timing T-T according to the set control period T r -2T d And then, sending a power-off command to the magnetic torquer, and turning to the step 7.
Step 7, collecting the actually measured angular velocity omega of the gyroscope after the magnetic torquer is powered off B0 And judging whether the attitude of the current micro-nano satellite meets a triaxial stable state, if so, ending the magnetic damping control, and if not, returning to the step 2 until the micro-nano satellite is in the triaxial stable state.
Take a derotation control task of a certain six-unit cube star as an example. The inertia tensor matrix for the satellite is: [0.130.010.01;0.010.10.01;0.010.010.05]kgm 2 (ii) a Initial attitude was [40-4040]Angular velocity of initial attitude of [20-20120]](ii) DEG/s; the maximum magnetic moment output by the magnetic torquer is 0.2Am 2 The rising time of the magnetic torquer is 80ms, and the falling time is 70ms; the kinetic resolution period for the numerical simulation was 50ms. According to the magnetic damping racemization control method provided by the invention, the upper limit of the control period is 0.82s, the lower limit of the control period is 0.3s, and the control period of the magnetic damping control law selected by simulation is 0.8s.
With reference to fig. 3, compared with the prior art, the method provided by the invention can enable the micro-nano satellite to achieve the expected racemization effect through magnetic damping control on the premise that the initial attitude angular velocity is [20-20120] °/s, but the conventional Bdot control cannot achieve the expected racemization effect.
Claims (4)
1. A magnetic damping control method suitable for racemization of micro-nano satellites at a large angular rate is characterized by comprising the following steps: and resolving the measurement data of the gyroscope and the magnetometer on the micro-nano satellite to obtain the actually measured angular velocity and the geomagnetic vector of the gyroscope of the micro-nano satellite, and outputting torque through the magnetic torquer after calculating a control period so as to despin the micro-nano satellite at a large angular rate.
2. The magnetic damping control method suitable for micro-nano satellite large angular rate racemization according to claim 1, characterized by comprising the following specific steps:
step 1, collecting actually-measured angular velocity omega of gyroscope at current moment B And magnetic strengthCalculating an output geomagnetic vector B, judging whether the current micro-nano satellite is in a three-axis stable state, if so, ending magnetic damping control, and if not, turning to the step 2;
step 2, a satellite body coordinate system is established according to the current attitude of the micro-nano satellite, and the change rate of the rotation kinetic energy of the micro-nano satellite around the mass center of the micro-nano satellite to time is obtained
Set damping control law c to
Make it
Wherein I is the rotational inertia of the micro/nano satellite; m is a unit of 0 Is the output magnetic moment of the magnetic torquer; the geomagnetic vector B can be regarded as unchanged in the magnitude direction in the inertial system in a short time; k is a control coefficient, and k is a control coefficient,
a component representing a derivative of the geomagnetic vector B with respect to time; omega B Representing the actually measured angular velocity of the gyroscope;
a component representing a derivative of a gyroscope measured angular velocity; t is 0 Representing the kinetic energy of the micro-nano satellite rotating around the mass center of the micro-nano satellite; t represents time;
turning to the step 3;
step 3, actually measuring the angular velocity omega according to the gyroscope B Calculating the upper limit T of the control period of the current magnetic damping racemization max (ii) a Turning to the step 4;
step 4, obtaining the rise time T of the magnetic torquer according to actual measurement r A fall time T d Further, a lower limit T of the control period is calculated min Turning to step 5;
step 5, setting a proper control period T to enable T max ＜T＜T min (ii) a Sending a power-on instruction to the magnetic torquer, outputting a magnetic moment M = c by the magnetic torquer, and turning to the step 6;
step 6, sending a power-off instruction to the magnetic torquer after the set control period T reaches the timing time, and turning to step 7;
step 7, collecting the actually measured angular velocity of the gyroscope after the magnetic torquer is powered off
And (3) judging whether the attitude of the current micro-nano satellite meets a triaxial stable state, if so, ending the magnetic damping control, and if not, returning to the step (2) until the micro-nano satellite is in the triaxial stable state.
3. The magnetic damping control method suitable for large-angle rate racemization of the micro-nano satellite according to claim 2, characterized in that: in step 3, the upper limit of the control period of the current magnetic damping racemization
4. The magnetic damping control method suitable for micro-nano satellite high-angular-rate racemization according to claim 3, characterized in that: in step 4, the lower limit T of the control period min ＝2(T r +T d )。
The magnetic damping control method suitable for large-angle-rate racemization of the micro-nano satellite according to claim 3, characterized in that: in step 6, the timing time is T-T r -2T d 。
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5123617A (en) * | 1990-03-05 | 1992-06-23 | General Electric Company | Spacecraft momentum unloading using controlled magnetic torques |
| CN105667838A (en) * | 2016-03-14 | 2016-06-15 | 西北工业大学 | Modular attitude determination and control device and method of Pico-satellite |
| CN107054697A (en) * | 2017-03-09 | 2017-08-18 | 西北工业大学 | A kind of Nano satellite magnetic torquer space temperature compensates attitude control method |
| CN109823571A (en) * | 2019-01-23 | 2019-05-31 | 清华大学 | A kind of multistage attitude control method of remote sensing micro-nano satellite |
| CN112607064A (en) * | 2020-12-25 | 2021-04-06 | 上海交通大学 | Micro-nano satellite magnetic damping control method for avoiding entering spinning state |
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- 2022-07-20 CN CN202210851877.6A patent/CN115140318B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5123617A (en) * | 1990-03-05 | 1992-06-23 | General Electric Company | Spacecraft momentum unloading using controlled magnetic torques |
| CN105667838A (en) * | 2016-03-14 | 2016-06-15 | 西北工业大学 | Modular attitude determination and control device and method of Pico-satellite |
| CN107054697A (en) * | 2017-03-09 | 2017-08-18 | 西北工业大学 | A kind of Nano satellite magnetic torquer space temperature compensates attitude control method |
| CN109823571A (en) * | 2019-01-23 | 2019-05-31 | 清华大学 | A kind of multistage attitude control method of remote sensing micro-nano satellite |
| CN112607064A (en) * | 2020-12-25 | 2021-04-06 | 上海交通大学 | Micro-nano satellite magnetic damping control method for avoiding entering spinning state |
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| Title |
|---|
| 刘海颖 等: "主动磁控微卫星姿态控制" * |
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