CN112009723A - Method for unfolding two-dimensional plane extensible mechanism by satellite autorotation in-orbit - Google Patents
Method for unfolding two-dimensional plane extensible mechanism by satellite autorotation in-orbit Download PDFInfo
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
- B64G1/1007—Communications satellites
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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
- B64G1/222—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles for deploying structures between a stowed and deployed state
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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
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
Abstract
The invention discloses a method for unfolding a two-dimensional plane extensible structure by utilizing satellite autorotation in orbit, which comprises the following steps: 1) establishing a satellite in-orbit unfolding flexible multi-body dynamic model with a two-dimensional unfoldable structure; 2) acquiring mass inertia characteristics of a satellite with a two-dimensional deployable structure, hinge driving torque, friction torque and damper temperature damping relation data of the two-dimensional deployable structure, maximum bearable force and torque of each hinge and a weak structure, and a maximum angular velocity measurement range of a satellite gyroscope; 3) calculating to obtain a satellite three-axis angular velocity data list; 4) determining a satellite coordinate axis capable of assisting the expansion of the two-dimensional plane expandable structure; 5) determining satellite angular velocity for assisting the satellite-borne deployable structure to deploy through flexible multi-body dynamics analysis; 6) and (5) adjusting the attitude of the satellite, enabling the satellite to obtain the satellite spin angular velocity in the fifth step, and releasing the satellite deployable structure unfolding restraint device.
Description
Technical Field
The invention relates to a method for unfolding a two-dimensional plane extensible structure by utilizing satellite spin angular velocity, which can be used as a low-impact and high-reliability unfolding method for spacecrafts with two-dimensional plane extensible structures, such as large-scale high-power communication satellites, space-based radars and the like.
Background
Large-scale and ultra-large-scale communication satellites and space-based radar systems generally adopt large-scale solar wings as energy devices meeting high-power requirements of the large-scale and ultra-large-scale communication satellites and space-based radar systems, and provide required electric energy for a whole satellite platform and loads; the space-based radar system employs a large area planar deployable structure as a phased array antenna support structure to provide the required gain. As a typical large-scale planar deployable structure, with the continuous improvement of the output power of the solar wing by the satellite, the conventional linear deployable solar wing is difficult to meet the power requirement of the satellite, and a two-dimensional secondary deployable solar wing (as shown in fig. 1) is becoming equipment for a large-scale high-flux communication satellite power and an astrological SAR satellite.
The large and ultra-large solar wings have large on-orbit expansion impact, great damage to the solar wing self structure and great influence on satellite SADA safety, and if no measures are taken, the solar wing self damage is caused, and the satellite safety is endangered[1,2]. The conventional method for automatically triggering the unfolding of the solar wing side plates by unfolding each middle plate in place has the advantages of high SADA reversal angular velocity in unfolding in place, high impact force and high moment of flexible accessories on the satellite and difficulty in adapting to the condition that the solar wing with one side exceeding four side plates is difficult to adapt. In fig. 1, (a) and (b) are respectively a conventional one-sided six-plate and one-sided eight-fast-plate solar wing configuration. FIG. 2 is a schematic diagram showing the on-orbit unfolding process of the solar wing with six solar panels on one side[3]. The unfolding process of the solar wing is divided into two times, the first unfolding occurs after the separation of the satellite and the arrow, the satellite firstly unfolds the outer plates as shown in fig. 2(a) and (b), the second unfolding occurs after the fixed point of the satellite, the pressing devices for restraining the middle plates explode, the middle plates unfold to the folded position, then the side plates automatically release and unfold in place as shown in fig. 2(c) to (f). According to the conservation of angular momentum, the self vibration of the solar wing and the SADA counter-rotating angular velocity caused by the side plates during the second unfolding of the solar wing are high, the SADA root bears large impact torque, certain adverse effect is brought to the safety of a satellite solar wing hinge and other flexible structures, and the method faces the characteristics of multiple impacts caused by asynchronous unfolding and high SADA impact torque peak value along with the increase of the number of the side plates of the solar wing.
Fig. 3 is a schematic view of the connection state of the solar wing and the SADA. When the single side of the solar wing is increased from six plates to eight plates, the traditional mode that the middle plates are unfolded in place to trigger the unfolding of the solar wing side plates is still adopted, so that the SADA (solar array panel inversion angular velocity) reversal speed of the solar wing is too high, and the problems of mechanism damage or reverse voltage breakdown of electronic devices can occur.
Disclosure of Invention
The technical problem solved by the invention is as follows: in order to avoid the damage to the solar wing hinge and the line safety caused by the on-orbit expansion impact of the two-dimensional secondary space expandable structure, the method for expanding the two-dimensional plane expandable mechanism by utilizing the satellite autorotation on-orbit is provided.
The purpose of the invention is realized by the following technical scheme: a method for unfolding a two-dimensional plane deployable mechanism by utilizing satellite autorotation in orbit comprises the following steps:
firstly, establishing a satellite in-orbit unfolding flexible multi-body dynamic model with a two-dimensional unfoldable structure;
acquiring mass inertia characteristics of a satellite with a two-dimensional deployable structure, hinge driving torque, friction torque, damper temperature damping relation data of the two-dimensional deployable structure, maximum bearable force and torque of each hinge and a weak structure, and a maximum angular velocity measurement range of a satellite gyroscope;
step three, from the minimum satellite triaxial angular velocity omegaminAt the beginning, multiplying by 2 each time until the maximum satellite three-axis angular velocity omegamaxCalculating to obtain a satellite three-axis angular velocity data list;
determining a satellite coordinate axis K capable of assisting the expansion of the two-dimensional plane expandable structure;
step five, determining satellite angular velocity omega for assisting the satellite-borne deployable structure to deploy through flexible multi-body dynamics analysisobj；
And step six, the satellite adjusts the attitude, enables the satellite to obtain the satellite spin angular velocity in the step five, and releases the satellite deployable structure unfolding restraint device.
Firstly, establishing an in-orbit second-unfolding flexible multi-body dynamic equation of a satellite with two-dimensional second-unfolding solar wings as follows:
wherein Z is a system quality matrix,
the system generalized coordinate matrix is adopted, z is the generalized inertial force matrix, and phi is the Jacobian matrix of the closed-loop cut-off hinge constraint and the compaction release constraint equation during the unfolding of the solar wing side panel.
In step three, the minimum satellite three-axis angular velocity omegaminTaking 0.25deg/sec, the maximum satellite three-axis angular velocity omegamaxTake 64 deg/sec.
The concrete process of the step four is as follows:
41) setting the driving force or moment of the two-dimensional deployable structure to zero;
42) setting the X, Y, Z triaxial angular velocities of the satellite to be Mdeg/sec, wherein M is more than or equal to 10 and is more than or equal to 1;
43) respectively carrying out satellite multi-body dynamic simulation with a two-dimensional deployable structure, wherein the satellite multi-body dynamic simulation is that the satellite rotates around an X, Y, Z axis at Mdeg/sec, by using the satellite on-orbit unfolding flexible multi-body dynamic model obtained in the step one;
44) comparing the sum of the hinge unfolding angles of the satellite two-dimensional expandable structure obtained in the step 43), and taking the satellite coordinate axis K with the largest sum of the hinge unfolding angles of the two-dimensional planar expandable structure at the simulation termination time as a satellite coordinate axis for assisting the expansion of the two-dimensional planar expandable structure.
The concrete process of the step five is as follows:
51) giving the data obtained in the step two to the satellite in-orbit unfolding flexible multi-body dynamic model with the two-dimensional unfoldable structure obtained in the step one;
52) according to the satellite rotation axis K determined in the step four, enabling the axis angular speed omegaiIs equal to omegamin，i＝1；
53) Using angular velocity omega about the K-axis of the satelliteiFlexible to satellite in-orbit expansionCarrying out simulation on the body dynamic model, wherein i is a positive integer;
54) post-processing the simulation result obtained in the step 52) to obtain the force borne by the weak structure of the satellite;
55) comparing the stress of the weak structure of the satellite obtained in the step 53) with 0.7 times of the maximum force which can be borne by the weak structure, and if the stress of the structure obtained in the step 53) is less than 0.7 times of the stress of the weak structure of the satellite and the maximum force which can be borne by the weak structure, multiplying the angular speed of the current shaft by 2 to serve as the angular speed omega of the next stepi+1And returns to step 53); if the stress of the structure obtained in the step 53) is more than 0.7 times of the stress of the weak structure of the satellite and the maximum force which can be borne by the weak structure, executing a step 56);
56) judge | ωi-ωi-1Whether | is less than or equal to Δ ωallowedIf so, will ωi-1As allowed satellite spin angular velocity ωobjIf not, the average value of the angular velocity at the current time and the angular velocity at the previous time is used as the angular velocity, omega, for calculationi+1＝(ωi+ωi-1) /2, return to step 53); the Δ ωallowedIndicating the allowed value of the angular velocity error.
The concrete process of the step six is as follows:
61) adjusting the posture of the satellite to ensure that the sun vector is vertical to the surface of the unfolded two-dimensional plane deployable structure, opening a heater of a damper of the deployable structure, and waiting for the temperature of the damper of the two-dimensional deployable structure to reach a preset value;
62) starting the corresponding thruster group pair of the satellite to make the satellite spin angular velocity along the K axis equal to omegaobj；
63) The firer cutter of the solar wing pressing mechanism is detonated, and the solar wing is unfolded in place and locked under the action of the self-rotating angular velocity and the self-unfolding driving action.
Compared with the prior art, the invention has the advantages that:
1) compared with the existing method for unfolding the two-dimensional deployable structure under the condition that the spin angular velocity of the satellite is zero, the auxiliary unfolding coordinate axis and the angular velocity of the two-dimensional deployable structure of the satellite are calculated through flexible multi-body dynamics analysis, a basis is provided for unfolding the two-dimensional plane deployable structure under the assistance of the spin of the satellite, the reliability and the efficiency of driving the two-dimensional structure to be unfolded by replacing elastic potential energy with the conservation of angular momentum are improved, the rigidity and the unfolding in-place impact of an elastic potential energy driving element are further reduced, the influence of the unfolding in-place impact of the two-dimensional deployable structure on other active driving mechanisms and weak links on the satellite is greatly relieved, and the influence of the unfolding impact of the satellite-borne two-dimensional deployable structure;
2) compared with the method that the satellite adopts the two-dimensional mechanism to unfold the fault along the three-axis translation direction of the satellite, the method for determining the most sensitive rotation coordinate axis and the angular speed of the auxiliary two-dimensional deployable structure through the flexible multi-body dynamics simulation analysis method is provided, the scientificity of strategy determination of on-orbit unfolding fault handling of the two-dimensional plane deployable structure is improved, unfolding fault removing measures of the two-dimensional deployable structure can be determined in an auxiliary mode, and the safety of the satellite is saved.
Drawings
Fig. 1 is a schematic diagram of an in-orbit unfolding state of a solar wing of a large communication satellite, wherein 1(a) is in a one-side six-plate configuration, and 1(b) is in a one-side eight-plate configuration.
Fig. 2 is an on-orbit unfolding configuration of six solar wings on one side in two-dimensional secondary, wherein 2(a) is that the solar wings are in a furled and compressed state, 2(b) is that the outer plates are in a 90-degree locked state, 2(c) is that the solar wings are in a second unfolding process state, 2(d) is that the middle plates of the two-dimensional unfolded solar wings are unfolded and locked, 2(e) is that the side plates of the solar wings are unfolded and locked, and 2(f) is that the side plates of the two-dimensional unfolded solar wings are unfolded and locked.
Fig. 3 is a schematic view of the connection state of the solar wing and the SADA.
Fig. 4 is a flow chart of the technical solution of the present invention.
Detailed Description
Some details of the technical solution of the present invention are further explained with reference to fig. 4, and the implementation includes the following steps:
firstly, establishing a satellite in-orbit unfolding flexible multi-body dynamic model with a two-dimensional unfoldable structure;
acquiring mass inertia characteristics of a satellite with a two-dimensional deployable structure, hinge driving torque, friction torque, damper temperature damping relation data of the two-dimensional deployable structure, maximum bearable force and torque of each hinge and a weak structure, and a maximum angular velocity measurement range of a satellite gyroscope;
step three, satellite three-axis angular velocity is from omegaminStarting with each multiplication by 2 until ωmaxAnd calculating to obtain a satellite three-axis angular velocity data list, generally omegaminTake 0.25deg/sec, omegamaxTaking 64 deg/sec;
and step four, determining a satellite coordinate axis K capable of assisting the satellite-borne two-dimensional plane expandable structure to expand.
(1) Setting the driving force or moment of the two-dimensional deployable structure to zero;
(2) setting the X, Y, Z triaxial angular velocities of the satellite to be Mdeg/sec, generally taking M to be more than or equal to 1 and M to be less than or equal to 10;
(3) respectively carrying out satellite multi-body dynamic simulation with a two-dimensional deployable structure, wherein the satellite multi-body dynamic simulation is that the satellite rotates around an X, Y, Z axis at Mdeg/sec, by using the satellite on-orbit unfolding flexible multi-body dynamic model obtained in the step one;
(4) and (4) comparing the sum of the hinge unfolding angles of the satellite two-dimensional expandable structure obtained in the step (3), and taking the satellite coordinate axis K with the largest sum of the hinge unfolding angles of the two-dimensional planar expandable structure at the simulation termination time as a satellite coordinate axis for assisting the expansion of the two-dimensional planar expandable structure.
Step five, determining satellite angular velocity omega for assisting satellite-borne deployable structure to deployobj。
(1) Giving the data obtained in the step two to the satellite in-orbit unfolding flexible multi-body dynamic model with the two-dimensional unfoldable structure obtained in the step one;
(2) according to the satellite rotation axis K determined in the step four, enabling the axis angular speed omegaiEqual to 0.25deg/sec, i ═ 1;
(3) using angular velocity omega about the K-axis of the satelliteiDeveloping simulation of the satellite on-orbit expansion flexible multi-body dynamic model;
(4) carrying out post-processing on the simulation result of the step (2) in the fifth step to obtain the force borne by the weak structure of the satellite;
(5) step (3) is carried outComparing the stress of the weak structure of the satellite with 0.7 times of the maximum force which can be borne by the weak structure, if the structural stress obtained in the step (3) is less than 0.7 times of the stress of the weak structure of the satellite and the maximum force which can be borne by the weak structure of the satellite, multiplying the angular speed of the current shaft by 2 to serve as the angular speed omega of the next stepi+1Executing the step (3) of the fifth step; if the structural stress obtained in the step (3) is greater than 0.7 times of the stress of the weak structure of the satellite and the maximum force which can be borne by the weak structure, executing the next step;
(6) judging the current angular velocity omegaiAngular velocity omega of previous stepi-1Whether or not Δ ω is less than or equal toallowedIf so, will ωi-1As allowed satellite spin angular velocity ωobjIf not, the average value of the angular velocity at the current time and the angular velocity at the previous time is used as the angular velocity, omega, for calculationi+1＝(ωi+ωi-1) Step 2, executing the step 3 of the step five;
sixthly, adjusting the satellite attitude, enabling the satellite to obtain the satellite spin angular velocity in the fifth step, and releasing the satellite deployable structure unfolding restraint device;
(1) adjusting the posture of the satellite to ensure that the sun vector is vertical to the surface of the unfolded two-dimensional deployable structure, opening a heater of a damper of the deployable structure, and waiting for the temperature of the damper of the two-dimensional deployable structure to reach a preset value;
(2) starting the corresponding thruster group pair of the satellite to make the satellite spin angular velocity along the K axis equal to omegaobj；
(3) Initiating the initiating explosive cutter of the solar wing pressing mechanism, and unfolding the solar wing in place and locking the solar wing under the action of the self-rotating angular velocity and the self-unfolding driving action;
those skilled in the art will appreciate that the invention may be practiced without these specific details.
Claims (6)
1. A method for unfolding a two-dimensional plane deployable mechanism by utilizing satellite autorotation in orbit is characterized by comprising the following steps:
firstly, establishing a satellite in-orbit unfolding flexible multi-body dynamic model with a two-dimensional unfoldable structure;
acquiring mass inertia characteristics of a satellite with a two-dimensional deployable structure, hinge driving torque, friction torque, damper temperature damping relation data of the two-dimensional deployable structure, maximum bearable force and torque of each hinge and a weak structure, and a maximum angular velocity measurement range of a satellite gyroscope;
step three, from the minimum satellite triaxial angular velocity omegaminAt the beginning, multiplying by 2 each time until the maximum satellite three-axis angular velocity omegamaxCalculating to obtain a satellite three-axis angular velocity data list;
determining a satellite coordinate axis K capable of assisting the expansion of the two-dimensional plane expandable structure;
step five, determining satellite angular velocity omega for assisting the satellite-borne deployable structure to deploy through flexible multi-body dynamics analysisobj；
And step six, the satellite adjusts the attitude, enables the satellite to obtain the satellite spin angular velocity in the step five, and releases the satellite deployable structure unfolding restraint device.
2. The method for automatically deploying a two-dimensional planar deployable mechanism in orbit by a satellite according to claim 1, wherein the method comprises the following steps: firstly, establishing an in-orbit second-unfolding flexible multi-body dynamic equation of a satellite with two-dimensional second-unfolding solar wings as follows:
wherein Z is a system quality matrix,
the system generalized coordinate matrix is adopted, z is the generalized inertial force matrix, and phi is the Jacobian matrix of the closed-loop cut-off hinge constraint and the compaction release constraint equation during the unfolding of the solar wing side panel.
3. The method for automatically deploying a two-dimensional planar deployable mechanism in orbit by a satellite according to claim 1, wherein the method comprises the following steps: in step three, the minimum defenseAngular velocity omega of three axes of starminTaking 0.25deg/sec, the maximum satellite three-axis angular velocity omegamaxTake 64 deg/sec.
4. The method for automatically deploying a two-dimensional planar deployable mechanism in orbit by a satellite according to claim 1, wherein the method comprises the following steps: the concrete process of the step four is as follows:
41) setting the driving force or moment of the two-dimensional deployable structure to zero;
42) setting the X, Y, Z triaxial angular velocities of the satellite to be Mdeg/sec, wherein M is more than or equal to 10 and is more than or equal to 1;
43) respectively carrying out satellite multi-body dynamic simulation with a two-dimensional deployable structure, wherein the satellite multi-body dynamic simulation is that the satellite rotates around an X, Y, Z axis at Mdeg/sec, by using the satellite on-orbit unfolding flexible multi-body dynamic model obtained in the step one;
44) comparing the sum of the hinge unfolding angles of the satellite two-dimensional expandable structure obtained in the step 43), and taking the satellite coordinate axis K with the largest sum of the hinge unfolding angles of the two-dimensional planar expandable structure at the simulation termination time as a satellite coordinate axis for assisting the expansion of the two-dimensional planar expandable structure.
5. The method for automatically deploying a two-dimensional planar deployable mechanism in orbit by a satellite according to claim 1, wherein the method comprises the following steps: the concrete process of the step five is as follows:
51) giving the data obtained in the step two to the satellite in-orbit unfolding flexible multi-body dynamic model with the two-dimensional unfoldable structure obtained in the step one;
52) according to the satellite rotation axis K determined in the step four, enabling the axis angular speed omegaiIs equal to omegamin，i＝1；
53) Using angular velocity omega about the K-axis of the satelliteiCarrying out simulation on a satellite in-orbit expansion flexible multi-body dynamic model, wherein i is a positive integer;
54) post-processing the simulation result obtained in the step 52) to obtain the force borne by the weak structure of the satellite;
55) comparing the stress of the weak structure of the satellite obtained in the step 53) with 0.7 times of the maximum force which can be borne by the weak structure, and if the stress of the structure obtained in the step 53) is smaller than the stress of the weak structure of the satelliteThe angular velocity of the shaft is multiplied by 2 to be used as the angular velocity omega of the next step when the weak structure can bear 0.7 times of the maximum forcei+1And returns to step 53); if the stress of the structure obtained in the step 53) is more than 0.7 times of the stress of the weak structure of the satellite and the maximum force which can be borne by the weak structure, executing a step 56);
56) judge | ωi-ωi-1Whether | is less than or equal to Δ ωallowedIf so, will ωi-1As allowed satellite spin angular velocity ωobjIf not, the average value of the angular velocity at the current time and the angular velocity at the previous time is used as the angular velocity, omega, for calculationi+1＝(ωi+ωi-1) /2, return to step 53); the Δ ωallowedIndicating the allowed value of the angular velocity error.
6. The method for automatically deploying a two-dimensional planar deployable mechanism in orbit by a satellite according to claim 1, wherein the method comprises the following steps: the concrete process of the step six is as follows:
61) adjusting the posture of the satellite to ensure that the sun vector is vertical to the surface of the unfolded two-dimensional plane deployable structure, opening a heater of a damper of the deployable structure, and waiting for the temperature of the damper of the two-dimensional deployable structure to reach a preset value;
62) starting the corresponding thruster group pair of the satellite to make the satellite spin angular velocity along the K axis equal to omegaobj；
63) The firer cutter of the solar wing pressing mechanism is detonated, and the solar wing is unfolded in place and locked under the action of the self-rotating angular velocity and the self-unfolding driving action.
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| FR2996526A1 (en) * | 2012-10-05 | 2014-04-11 | Thales Sa | SATELLITE WITH DEPLOYABLE USEFUL LOAD MODULES |
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| CN104943876A (en) * | 2015-05-25 | 2015-09-30 | 沈阳航空航天大学 | Space solar power station solar wing two-dimensional unfolding device and unfolding methods thereof |
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| CN106184817A (en) * | 2016-07-08 | 2016-12-07 | 北京空间飞行器总体设计部 | Towards the spacecraft plane deployable supporting construction of load Two-Dimensional Quadratic and using method |
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2020
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CA1018140A (en) * | 1974-10-30 | 1977-09-27 | Robert E. Coltrin | Spacecraft with deployable solar cell panels |
| FR2996526A1 (en) * | 2012-10-05 | 2014-04-11 | Thales Sa | SATELLITE WITH DEPLOYABLE USEFUL LOAD MODULES |
| CN103072701A (en) * | 2013-01-30 | 2013-05-01 | 北京控制工程研究所 | Racemization control method for under-actuated satellite |
| CN103901894A (en) * | 2014-04-14 | 2014-07-02 | 西北工业大学 | Spinning unfolding and folding optimum control method of dual-body star space tethered formation system |
| CN104943876A (en) * | 2015-05-25 | 2015-09-30 | 沈阳航空航天大学 | Space solar power station solar wing two-dimensional unfolding device and unfolding methods thereof |
| CN105160051A (en) * | 2015-06-30 | 2015-12-16 | 中国空间技术研究院 | Truss antenna reflector deployment dynamics modeling method based on multi-body analysis test |
| CN106184817A (en) * | 2016-07-08 | 2016-12-07 | 北京空间飞行器总体设计部 | Towards the spacecraft plane deployable supporting construction of load Two-Dimensional Quadratic and using method |
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