CN108069050B - Spacecraft initial attitude capture control method and system - Google Patents
Spacecraft initial attitude capture control method and system Download PDFInfo
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Abstract
本发明提供一种航天器的初始姿态捕获控制方法及系统，所述航天器的初始姿态捕获控制方法包括：利用三轴磁力矩器根据地磁在星体的变化率对所述航天器施加控制磁矩，实现速率阻尼阶段控制；利用反作用轮组起旋，根据所述航天器的姿态信息对所述航天器施加三轴轮控，同时卸载所述三轴磁力矩器对所述航天器施加的控制磁矩，实现太阳捕获阶段控制。本发明不仅在姿态控制部件上只采用了反作用轮和磁力矩器，考虑了星箭分离偏差过大的应对措施，首先施加了速率阻尼，以反作用轮不致于饱和为前提条件，同时在太阳捕获阶段，仅依靠磁强计信息和太阳敏感器信息，施加三轴轮控，即可完成初始捕获控制。
The present invention provides an initial attitude capture control method and system of a spacecraft. The initial attitude capture control method of the spacecraft includes: applying a control magnetic moment to the spacecraft according to the change rate of geomagnetism in a star by using a three-axis magnetic torque device , to realize the rate damping stage control; use the reaction wheel set to spin, apply three-axis wheel control to the spacecraft according to the attitude information of the spacecraft, and unload the control imposed by the three-axis magnetic torque device on the spacecraft at the same time Magnetic moment to achieve solar capture phase control. The present invention not only adopts the reaction wheel and the magnetic torque device on the attitude control part, but also considers the countermeasures for the excessive separation deviation of the star and arrow, and firstly applies the rate damping. In the first stage, only relying on the information of the magnetometer and the solar sensor, and applying the three-axis wheel control, the initial capture control can be completed.
Description
Technical Field
The invention relates to the technical field of satellite control, in particular to an initial attitude acquisition control method and system for a spacecraft.
Background
Due to the particularity of the satellite, once the satellite breaks down during in-orbit flight, the satellite is difficult to make up by maintenance like other ground products, so that how to ensure the satellite to safely and reliably complete the task in the in-orbit is the first task in satellite development. The safety of the satellite in orbit is divided into two stages: the first stage is the normal safe capture of the initial of the track entry; the second phase is the completion of the normal security assurance tasks. The initial normal safe capture of the first stage of the orbit, namely the initial stage of the satellite orbit, how to eliminate the attitude deviation caused by the separation of the satellite and the arrow, safely and reliably realizes the initial capture of the sun, and meets the requirement of the satellite energy supply. This phase is the first phase to ensure that the second phase of the satellite is safe to complete the mission by orbiting.
At present, most of the initial capturing methods of the satellite attitude control system adopt a propulsion system for damping and then apply control of a reaction wheel and a magnetic torquer, namely the propulsion system, the reaction wheel and the magnetic torquer. The propulsion system is a main control means for initial acquisition of the satellite due to high control efficiency, but the requirements on the installation structure and related operations of the propulsion system are complex due to the fact that the propellant of the propulsion system is flammable, explosive and toxic, and meanwhile the propulsion system works at the cost of propellant consumption. Thus, once the propulsion system is installed, the satellite not only increases in weight and resources, but the system becomes more complex. Once the satellite system becomes complex, its reliability and safety are greatly reduced. According to past flight experience statistics, seventy-eight percent of satellite in-orbit faults are related to a propulsion system.
In view of this, how to find a safer and more reliable solution for acquiring the initial attitude of the satellite becomes a problem to be solved by those skilled in the art.
Disclosure of Invention
In view of the above drawbacks of the prior art, an object of the present invention is to provide a method and a system for controlling initial attitude capture of a spacecraft, so as to solve the problems of complex mounting structure and low safety caused by a propulsion system required during initial capture of the spacecraft during orbit entering in the prior art.
In order to achieve the above and other related objects, the present invention provides a method for controlling initial attitude acquisition of a spacecraft, including: applying a control magnetic moment to the spacecraft by using a triaxial magnetic torquer according to the change rate of the geomagnetism on a star body, so as to realize rate damping stage control; and starting rotation by utilizing the reaction wheel set, exerting three-axis wheel control on the spacecraft according to the attitude information of the spacecraft, and unloading the control magnetic moment exerted by the three-axis magnetic torquer on the spacecraft to realize control in the solar capture stage.
In an embodiment of the present invention, a control magnetic moment applied by the three-axis magnetic torquer to the spacecraft according to a change rate of geomagnetism in a star is:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
In an embodiment of the present invention, another control magnetic moment applied by the three-axis magnetic torquer to the spacecraft according to a change rate of geomagnetism in a star is:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z.
In an embodiment of the present invention, the rotation of the reaction wheel set applies a three-axis wheel-controlled control torque to the spacecraft by:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcA control torque for the reaction wheel set; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; and I is a spacecraft inertia matrix.
In an embodiment of the present invention, an implementation process of the three-axis magnetic torquer for unloading the control magnetic moment applied to the spacecraft includes:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient. The control mode of the reaction wheel set for exerting three-axis wheel control on the spacecraft comprises the following steps: implementing PD control on the pitch of the spacecraft; and carrying out nutation and precession composite control on the rolling and yawing of the spacecraft.
The invention also provides an initial attitude capture control system of a spacecraft, which comprises: the attitude sensor is used for acquiring attitude information of the spacecraft; the attitude controller is in communication connection with the attitude sensor and is used for judging the current state of the spacecraft according to the attitude information of the spacecraft and sending a rate damping control instruction when the current state is a rate damping stage or sending a sun capture control instruction when the current state is a sun capture stage; the attitude control component is in communication connection with the attitude controller and comprises a three-axis magnetic torquer and a reaction wheel set; the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism on the star body under the control of the rate damping control instruction; the reaction wheel set rotates under the control of the sun capturing control instruction, three-axis wheel control is applied to the spacecraft according to the attitude information of the spacecraft, and meanwhile, the control magnetic moment applied to the spacecraft by the three-axis magnetic torquer is unloaded.
In an embodiment of the present invention, a control magnetic moment applied by the three-axis magnetic torquer to the spacecraft according to a change rate of geomagnetism in a star is:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
In an embodiment of the present invention, another control magnetic moment applied by the three-axis magnetic torquer to the spacecraft according to a change rate of geomagnetism in a star is:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z.
In an embodiment of the present invention, the rotation of the reaction wheel set applies a control moment of three-axis wheel control to the spacecraft, which is:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcA control torque for the reaction wheel set; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; and I is a spacecraft inertia matrix.
In an embodiment of the present invention, the unloading model for unloading the control magnetic moment applied to the spacecraft by the three-axis magnetic torquer is:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient.
In an embodiment of the present invention, an implementation structure of the three-axis magnetic torquer includes: three independent magnetic rods with the same performance; the three magnetic rods are respectively arranged along the X, Y, Z three axes of the spacecraft; the reaction wheel set comprises an X-direction reaction wheel, a Y-direction reaction wheel, a Z-direction reaction wheel and an inclined reaction wheel; the obliquely-installed reaction wheel forms a preset included angle with the X, Y, Z triaxial; the three-axis direction of the reaction wheel set is the same as the three-axis direction of the spacecraft.
As described above, the method and system for controlling initial attitude capture of a spacecraft of the present invention have the following beneficial effects:
1) the simplest configuration is achieved: all control from the orbit entering to the task completion of the spacecraft can be realized by only one set of reaction wheel and one set of magnetic torquer.
2) The configuration of the control system has the characteristics of safety, reliability and low power consumption, and as long as a power supply is provided, the magnetic torquer which can be used infinitely is combined with the reaction wheel to replace a propulsion system, the reaction wheel and the magnetic torquer, so that the resources are saved, the cost is reduced, the reliability is enhanced, the risk is reduced, and the service life of the spacecraft can be prolonged.
3) The speed damping method of the magnetic torquer is added before the solar capture stage, the separation speed deviation is reduced to the capacity that the reaction wheel can absorb, and the risk brought by the separation deviation of the star and the arrow in case of overlarge separation deviation is avoided.
Drawings
Fig. 1 is a schematic flow chart illustrating an implementation of a spacecraft initial attitude capture control method according to an embodiment of the present invention.
Fig. 2 is a schematic structural diagram illustrating an implementation of an initial attitude capture control system of a spacecraft according to an embodiment of the present invention.
Fig. 3 is a schematic structural diagram illustrating an implementation of an attitude control component of an initial attitude capture control system of a spacecraft according to an embodiment of the present invention.
Description of the element reference numerals
Initial attitude capture control system of 200 spacecraft
210 attitude sensor
220 posture controller
230 attitude control part
231 triaxial magnetic torquer
2311X magnetic rod
2312Y magnetic bar
2313Z magnetic rod
232 reaction wheel set
2321X-direction reaction wheel
2322Y-direction reaction wheel
2323Z-direction reaction wheel
2324 reaction wheel is obliquely installed
S101 to S102
Detailed Description
The embodiments of the present invention are described below with reference to specific embodiments, and other advantages and effects of the present invention will be easily understood by those skilled in the art from the disclosure of the present specification. The invention is capable of other and different embodiments and of being practiced or of being carried out in various ways, and its several details are capable of modification in various respects, all without departing from the spirit and scope of the present invention.
It should be noted that the drawings provided in the present embodiment are only for illustrating the basic idea of the present invention, and the components related to the present invention are only shown in the drawings rather than drawn according to the number, shape and size of the components in actual implementation, and the type, quantity and proportion of the components in actual implementation may be changed freely, and the layout of the components may be more complicated.
For satellites with low control precision requirements and less maneuvering, a bias momentum stabilization mode can be adopted to avoid a propulsion system, and a control method of a bias momentum wheel and a magnetic torquer is adopted to complete a task. However, it is difficult to meet the performance requirement of the method of "offset momentum wheel + magnetic torquer" for the zero momentum satellite with high precision and large inertia. The invention aims to realize the initial acquisition of the zero momentum satellite safely and reliably by only combining the reaction wheel and the magnetic torquer without configuring the propelling jet.
The invention not only adopts the reaction wheel and the magnetic torquer on the control part, but also considers the corresponding measure of the overlarge star-arrow separation deviation on the control method, firstly applies the speed damping, and takes the precondition that the reaction wheel is not saturated, and at the same time, in the sun capturing stage, only depending on the information of the magnetometer and the sun sensor, the three-axis wheel control is applied, and the initial capturing control can be completed, and the specific realization mode is as follows.
Referring to fig. 1, the present embodiment provides an initial attitude capture control method for a spacecraft, where the initial attitude capture control method for a spacecraft includes:
s101, applying a control magnetic moment to the spacecraft by using a three-axis magnetic torquer according to the change rate of geomagnetism on a star body so as to control the angular speed of the spacecraft and realize rate damping stage control. Specifically, in a speed damping stage, the three-axis magnetic torquer is controlled to implement B-dot magnetic control on X, Y, Z three channels of the spacecraft by using geomagnetic variation, so that the angular speed of the spacecraft is reduced, and the reaction wheel saturation in a solar capture stage is avoided.
Further, the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism in the star, wherein the control magnetic moment is as follows:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
The other control magnetic moment applied to the spacecraft by the triaxial magnetic torquer according to the change rate of the geomagnetism in the star body is as follows:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z.
S102, utilizing the reaction wheel set to rotate, applying three-axis wheel control (namely that all reaction wheels in the reaction wheel set act and control at the same time) to the spacecraft according to the attitude information of the spacecraft, and unloading the control magnetic moment applied to the spacecraft by the three-axis magnetic torquer to realize control in a solar capture stage; the reaction wheel set comprises an X-direction reaction wheel, a Y-direction reaction wheel, a Z-direction reaction wheel and an inclined reaction wheel; the inclined reaction wheel and the X, Y, Z three shafts form a preset included angle. Specifically, the reaction wheel set is controlled to implement three-axis wheeling on the spacecraft during the solar capture phase, and magnetic torquer unloading (i.e., the attitude controller 220 unloads the control magnetic moment applied to the spacecraft before the reaction wheel set is controlled during the solar capture phase). The function of the angled reaction wheels is to maintain the full star at zero momentum and to act as a backup control in the event of a failure of one of the reaction wheels.
Further, the reaction wheel set rotates to apply three-axis wheel-controlled control torque to the spacecraft, and the three-axis wheel-controlled control torque comprises the following steps:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcA control torque for the reaction wheel set; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; i is a spacecraft inertia matrix,
one implementation of the three-axis magnetic torquer to unload the control magnetic moment applied to the spacecraft comprises:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient.
The control mode of the reaction wheel set for exerting three-axis wheel control on the spacecraft comprises the following steps: implementing PD control (i.e. proportional-derivative control, which is a classical control method) on the pitch of the spacecraft; and carrying out nutation and precession composite control on the rolling and yawing of the spacecraft.
The protection scope of the method for controlling initial attitude capture of a spacecraft according to the present invention is not limited to the execution sequence of the steps listed in this embodiment, and all the solutions implemented by adding, subtracting, and replacing the steps in the prior art according to the principles of the present invention are included in the protection scope of the present invention.
The invention also provides an initial attitude capture control system of a spacecraft, which can realize the initial attitude capture control method of the spacecraft, but an implementation device of the initial attitude capture control method of the spacecraft provided by the invention comprises but is not limited to the structure of the initial attitude capture control system of the spacecraft listed in the embodiment, and all structural deformation and replacement in the prior art made according to the principle of the invention are included in the protection scope of the invention.
Referring to fig. 2, the initial attitude acquisition control system 200 of the spacecraft includes: attitude sensor 210, attitude controller 220, attitude control component 230.
The attitude sensor 210 is used for acquiring attitude information of the spacecraft. The spacecraft includes a satellite. The attitude sensor 210 includes a sun sensor or/and a magnetometer.
The attitude controller 220 is in communication connection with the attitude sensor 210, and is configured to determine a current state of the spacecraft according to the attitude information of the spacecraft, and issue a rate damping control instruction when the current state is a rate damping stage, or issue a sun capture control instruction when the current state is a sun capture stage.
The attitude control component 230 is in communication connection with the attitude controller 220 and comprises a three-axis magnetic torquer 231 and a reaction wheel set 232; the three-axis magnetic torquer 231 applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism in the star body under the control of the rate damping control instruction so as to control the angular speed of the spacecraft; the reaction wheel set 232 rotates under the control of the sun capturing control instruction, exerts three-axis wheel control on the spacecraft according to the attitude information of the spacecraft, and unloads the control magnetic moment exerted by the three-axis magnetic torquer on the spacecraft.
The attitude controller 220 controls the three-axis magnetic torquer to implement B-dot magnetic control on X, Y, Z three channels of the spacecraft by using geomagnetic variation in a speed damping stage so as to reduce the angular speed of the spacecraft and avoid saturation of a reaction wheel in a solar capture stage. The attitude controller 220 controls the reaction wheel set to perform three-axis steering on the spacecraft during the solar capture phase, and magnetic torquer unloading (i.e., the attitude controller 220 unloads the control magnetic moment applied to the spacecraft before the reaction wheel set is controlled by the sun capture phase on the magnetic torquer).
Further, referring to FIG. 3, one implementation of the reaction wheel set 232 (not shown) includes an X-direction reaction wheel 2321, a Y-direction reaction wheel 2322, a Z-direction reaction wheel 2323 and a skewed reaction wheel 2324; the inclined reaction wheel and the X, Y, Z three shafts form a preset included angle. In this embodiment, the predetermined angle between the three axes of the obliquely-installed reaction wheels 234 and X, Y, Z may preferably be set to 57.3 degrees, but is not limited to this angle. The angular momentum of each reaction wheel may preferably be 15Nms (newton meters per second), the mass may preferably be 10kg (kilogram), and the average power consumption may preferably be 12W (watts), although the scope of the present invention is not limited to this preferred condition.
The three-axis direction of the reaction wheel set is the same as the three-axis direction of the spacecraft. One implementation of the three-axis magnetic torquer 231 (not shown) includes: three independent magnetic bars (i.e., X bar 2311, Y bar 2312, Z bar 2313) of the same performance; the three magnetic bars (namely X magnetic bar 2311, Y magnetic bar 2312 and Z magnetic bar 2313) are respectively arranged along the same lineThe X, Y, Z triaxial mounting of the spacecraft. In this embodiment, the magnetic moment of the three-axis magnetic torquer 231 may preferably be (-120 to +120) Am2(amperes/meter), but is not limited to this range of magnetic moments. Each magnetic rod may preferably have a mass of 3.8kg (kilogram) and a power consumption of 6.5W (watt), but the scope of the present invention is not limited to this preferred condition.
Further, the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism in the star, wherein the control magnetic moment is as follows:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
When k is1,k2,k31, and BdotOnly taking a sign, when the magnetic moment applies the maximum magnetic moment, the three-axis magnetic torquer applies another control magnetic moment to the spacecraft according to the change rate of the geomagnetism in the star body:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z. The dependence on the performance of the magnetometer can be reduced by controlling the magnetic moment, and the control method is simplified.
The reaction wheel set rotates to apply three-axis wheel control to the spacecraft, and one control moment is as follows:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcA control torque for the reaction wheel set; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; and I is a spacecraft inertia matrix. If the inertial system of the spacecraft is the attitude angular velocity omegabiIs omegabi＝[ωx,ωy,ωz]T，ωx,ωy,ωzThe angular velocities are x, y and z under an inertial system, namely the angular velocities of three axes under an inertial system coordinate system; then matrix omegabi[×]Can be expressed as:
the unloading model of the control magnetic moment applied to the spacecraft by the triaxial magnetic torquer in unloading is as follows:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient.
The invention can be directly suitable for the whole-process control of other spacecrafts except for high-inertia zero-momentum satellites, and various key technologies and innovative design ideas can be popularized to the design of attitude control systems of various space aircrafts.
The technical problem solved by the invention is as follows: the zero-momentum high-inertia spacecraft in-orbit initial capture can be realized only by means of a set of reaction wheel set, a set of magnetic torquer and an attitude controller storing a control algorithm, and the complexity and the insecurity of a satellite caused by using propulsion jet control in the traditional capture method are avoided, so that the on-board resources are saved, and the safety and the reliability of the system are enhanced; meanwhile, before the reaction wheel set controls capture, the speed damping of the magnetic torquer is increased, and the fault that the star-arrow separation deviation is too large can be dealt with. When the separation deviation of the star and the arrow is overlarge, the reaction wheel is only used for absorbing redundant momentum brought to the spacecraft by the deviation, and the saturation failure of the reaction wheel is easily caused. The invention not only adopts a reaction wheel and a magnetic torquer on the attitude control part, but also considers the corresponding measure of overlarge satellite-rocket separation deviation on the control method, firstly applies velocity damping on the premise that the reaction wheel is not saturated, and at the same time, in the sun capturing stage, only depending on the information of a magnetometer and the sun sensor, three-axis wheel control is applied, and the initial capturing control can be completed.
According to the invention, through the optimization design, the initial capture of the spacecraft is realized by the simplest system configuration, and meanwhile, the spacecraft has the capability of coping with the overlarge satellite-rocket separation deviation.
The invention has the following advantages:
1) the simplest configuration is achieved: all control from the orbit entering to the task completion of the spacecraft can be realized by only one set of reaction wheel and one set of magnetic torquer.
2) The configuration of the control system has the characteristics of safety, reliability and low power consumption, and as long as a power supply is provided, the magnetic torquer which can be used infinitely is combined with the reaction wheel to replace a propulsion system, the reaction wheel and the magnetic torquer, so that the resources are saved, the cost is reduced, the reliability is enhanced, the risk is reduced, and the service life of the spacecraft can be prolonged.
3) The speed damping method of the magnetic torquer is added before the solar capture stage, the separation speed deviation is reduced to the capacity that the reaction wheel can absorb, and the risk brought by the separation deviation of the star and the arrow in case of overlarge separation deviation is avoided.
In conclusion, the invention has the significance of realizing the performance of a complex and high-cost control system by utilizing high-reliability, low-cost and simplified system configuration, and realizing the rapid capture and stability after the spacecraft is in orbit, good control precision and high reliability. Compared with international similar systems, the comprehensive performance reaches an advanced level.
In addition, the successful exploration of the control method plays a positive promoting role in the development of the aerospace control technology: the traditional propulsion system is abandoned; the satellite and arrow separation fault handling capacity is increased.
The foregoing embodiments are merely illustrative of the principles and utilities of the present invention and are not intended to limit the invention. Any person skilled in the art can modify or change the above-mentioned embodiments without departing from the spirit and scope of the present invention. Accordingly, it is intended that all equivalent modifications or changes which can be made by those skilled in the art without departing from the spirit and technical spirit of the present invention be covered by the claims of the present invention.
Claims (10)
1. An initial attitude capture control method of a spacecraft is characterized by comprising the following steps: the initial attitude capture control method of the spacecraft comprises the following steps:
applying a control magnetic moment to the spacecraft by using a triaxial magnetic torquer according to the change rate of the geomagnetism on a star body, so as to realize rate damping stage control;
starting rotation by utilizing a reaction wheel set, exerting three-axis wheel control on the spacecraft according to attitude information of the spacecraft, and unloading a control magnetic moment exerted by the three-axis magnetic torquer on the spacecraft at the same time to realize control in a solar capture stage;
the reaction wheel set rotates to apply three-axis wheel control torque to the spacecraft, and the control torque is as follows:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcFor reaction wheelsThe control moment of the group; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; i is a spacecraft inertia matrix,
ωx,ωy,ωzthe angular velocities x, y, and z are respectively in the inertial system.
2. A spacecraft initial attitude acquisition control method according to claim 1, wherein: the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism on a star body, wherein the control magnetic moment is as follows:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
3. A spacecraft initial attitude acquisition control method according to claim 1, wherein: the other control magnetic moment applied to the spacecraft by the triaxial magnetic torquer according to the change rate of the geomagnetism in the star body is as follows:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z.
4. A method as claimed in claim 1, wherein one implementation of the three-axis magnetic torquer to unload the control magnetic moment applied to the spacecraft comprises:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient.
5. A spacecraft initial attitude acquisition control method according to claim 1, wherein: the control mode of the reaction wheel set for exerting three-axis wheel control on the spacecraft comprises the following steps:
implementing PD control on the pitch of the spacecraft;
and carrying out nutation and precession composite control on the rolling and yawing of the spacecraft.
6. An initial attitude acquisition control system for a spacecraft, comprising:
the attitude sensor is used for acquiring attitude information of the spacecraft;
the attitude controller is in communication connection with the attitude sensor and is used for judging the current state of the spacecraft according to the attitude information of the spacecraft and sending a rate damping control instruction when the current state is a rate damping stage or sending a sun capture control instruction when the current state is a sun capture stage;
the attitude control component is in communication connection with the attitude controller and comprises a three-axis magnetic torquer and a reaction wheel set; the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism on the star body under the control of the rate damping control instruction; the reaction wheel set rotates under the control of the sun capture control instruction, three-axis wheel control is applied to the spacecraft according to the attitude information of the spacecraft, and meanwhile, the control magnetic moment applied to the spacecraft by the three-axis magnetic torquer is unloaded;
the reaction wheel set rotates to apply three-axis wheel control to the spacecraft, and one control moment is as follows:
Tc＝Kp(ΦT-Φbr)-KDωbr+T0，
T0＝ωbi[×]·(I·ωbi+hXYZ)(T0dimension 3 × 1);
wherein, TcA control torque for the reaction wheel set; phibr＝[φ,θ,ψ]TIs the Euler angle of the attitude of the orbital system; phiTEuler angles for the desired attitude; kpIs a diagonal matrix of scale coefficients, KDDiagonal matrix of differential coefficients, T0For coupling moment, omegabrIs the attitude angular velocity of the track system; omegabiIs the attitude angular velocity of the inertial system; h isXYZIs the three-axis angular momentum induced by the reaction wheel set; i is a spacecraft inertia matrix,
ωx,ωy,ωzthe angular velocities x, y, and z are respectively in the inertial system.
7. An initial attitude acquisition control system of a spacecraft as claimed in claim 6, wherein: the three-axis magnetic torquer applies a control magnetic moment to the spacecraft according to the change rate of the geomagnetism on a star body, wherein the control magnetic moment is as follows:
M＝-k·Bdoti.e. by
Wherein Mx, My and Mz are respectively control magnetic moments exerted on the spacecraft by the triaxial magnetic torquer in the direction X, Y, Z; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotA rate of change component in the direction X, Y, Z; k is a radical of1,k2,k3Respectively X, Y, Z direction control coefficients.
8. An initial attitude acquisition control system of a spacecraft as claimed in claim 6, wherein: the other control magnetic moment applied to the spacecraft by the triaxial magnetic torquer according to the change rate of the geomagnetism in the star body is as follows:
wherein M ismaxSign (B) for maximum magnetic momentdot) Is shown in BdotReturns 1 when it is positive, at BdotReturns 0 when zero, at BdotIf the number is negative, returning to-1; bxdot，Bydot，BzdotRespectively the rate of change B of the earth magnetism in the stardotThe rate of change component in the direction X, Y, Z.
9. An initial attitude acquisition control system of a spacecraft as claimed in claim 6, wherein: the unloading model of the control magnetic moment applied to the spacecraft by the triaxial magnetic torquer in unloading is as follows:
wherein, PbFor unloaded control moment, Δ H is the angular momentum of the spacecraft, BbK is the control coefficient.
10. An initial attitude acquisition control system for a spacecraft as claimed in claim 6, wherein one implementation of said three-axis magnetic torquer comprises: three independent magnetic rods with the same performance; the three magnetic rods are respectively arranged along the X, Y, Z three axes of the spacecraft; the reaction wheel set comprises an X-direction reaction wheel, a Y-direction reaction wheel, a Z-direction reaction wheel and an inclined reaction wheel; the obliquely-installed reaction wheel forms a preset included angle with the X, Y, Z triaxial; the three-axis direction of the reaction wheel set is the same as the three-axis direction of the spacecraft.
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| CN109649691B (en) * | 2018-12-27 | 2021-07-13 | 上海航天控制技术研究所 | Single flywheel and magnetic combined control method and system for offset momentum satellite |
| CN109677638B (en) * | 2019-01-30 | 2020-09-01 | 上海微小卫星工程中心 | An Improved Pure Magnetron Spin-to-Sun Orientation Method Based on Geomagnetic Field Measurement Parameters |
| CN110481815B (en) * | 2019-08-21 | 2020-12-15 | 中国科学院力学研究所 | A transfer type contra-rotating geomagnetic energy storage-release delivery system and method |
| CN110949690B (en) * | 2019-10-10 | 2024-07-02 | 中国科学院力学研究所 | Spacecraft structure for low-orbit geomagnetic energy storage on-orbit delivery |
| CN111056045B (en) * | 2019-12-26 | 2021-03-23 | 航天行云科技有限公司 | Control method and device of three-axis magnetic torquer |
| CN111722637B (en) * | 2020-06-30 | 2022-09-13 | 中国科学院微小卫星创新研究院 | Spacecraft attitude determination method and system |
| CN112607064B (en) * | 2020-12-25 | 2021-08-06 | 上海交通大学 | Magnetic damping control method for micro-nano satellites to avoid entering spin state |
| CN112937920B (en) * | 2021-03-30 | 2024-07-19 | 湖南揽月机电科技有限公司 | Multi-redundancy satellite intelligent attitude control assembly and working method thereof |
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