CN103921959B - Two-dimentional pointing system configuration designing method on star - Google Patents
Two-dimentional pointing system configuration designing method on star Download PDFInfo
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- CN103921959B CN103921959B CN201410161957.4A CN201410161957A CN103921959B CN 103921959 B CN103921959 B CN 103921959B CN 201410161957 A CN201410161957 A CN 201410161957A CN 103921959 B CN103921959 B CN 103921959B
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Abstract
The invention provides two-dimentional pointing system configuration designing method on a kind of star, belong to AEROSAT technical field.In Configuration Design of the present invention, the intersection point of the barycenter and turning cylinder that arrange load overlaps, and removes load is connected joint drive motor with satellite body, load is connected by free joint with satellite body, load is installed control moment gyroscope group, to provide the moment pointed to and control to need.Configuration designing method of the present invention can reduce the disturbance of load to satellite body, improves degree of stability and the stable state accuracy of satellite body attitude, completes the energy that satellite body attitude stabilization consumes when being greatly reduced in the flyback of load high speed.
Description
Technical field
The present invention relates to AEROSAT technical field, be specifically related to the Configuration Design of the two-dimentional pointing system that a kind of satellite uses.
Background technology
Along with the high speed development of aerospace industry, AEROSPACE APPLICATION is also further extensive.Complicated practical application not only requires satellite platform attitude stabilization, also requires platform to be installed the load (as camera, antenna etc.) with direction-pointing function, to complete actual task.As tracking and data relay satellite, actual task requires the accurate point and track towards user satellite of single access antenna that it is installed, and requires that satellite platform keeps attitude stabilization simultaneously; Missile defence forewarn system then requires that star pointing to tracking load (as infrared camera) has revolution at a high speed, flyback and following function, requires satellite platform attitude stabilization simultaneously.
Sensing load on star and satellite body constitute complicated multi-body system.The top priority of satellite needs to maintain the stable of attitude, and load need point to target simultaneously.Load, completing in sensing task process, can produce the moment of reaction to satellite body, and then reduces the degree of stability of platform.How to reduce the hot issue that the disturbance of load to satellite body is current research.
On current star, pointing system adopts frame system mostly, and by motor output torque, the sensing completing load controls.The subject matter of such design is, the moment that motor acts in load directly creates the moment of reaction acted on body, and this is very unfavorable to the attitude stabilization of body.Current research emphasis is how to design composite controller, and the moment of momentum of load can be compensated in time, to improve the stable state accuracy of body attitude.But, although the attitude control accuracy of body improves, control energy problem and be not resolved all the time.Its basic reason is, the applied moment that motor provides creates the moment of reaction on satellite body, and the directing mechanism on body also will compensate this moment, to maintain the attitude stabilization of body.Especially load have over a long time fast retrace task time, the overall energy consumption problem of system is even more serious.
Summary of the invention
The present invention is directed to the spacecraft multi-body system having and point to mission payload, the problem that the satellite body integral energy loss existed is serious, provide two-dimentional pointing system configuration designing method on a kind of star, be intended to for domestic now provide technical support with pointing system design on star in the future, to the disturbance to satellite body during reduction loading movement, reduce the energy consumption of whole system simultaneously.
On star of the present invention, two-dimentional pointing system configuration designing method comprises the following aspects:
The first step: the intersection point of the barycenter and turning cylinder that arrange load overlaps, and is arranged on satellite body by load by barycenter;
Second step: remove the drive motor be arranged on satellite body and intermediate connector, load is connected by free joint with satellite body;
3rd step: install one group of control moment gyroscope group (CMGs) in load, to provide the moment pointed to and control to need.
The new configuration of the inventive method design drives by cancelling traditional motor, have employed the concept that low-disturbance moment drives, by stagger angle momentum exchange device (CMGs) in load, reduce the Dynamics Coupling degree between multi-body system well.New configuration makes the disturbance amplitude of load to body significantly reduce, and improves body and maintains the stable stable state accuracy of platform stance and degree of stability, reduced the energy consumption pointing to tracking and platform attitude stabilization task simultaneously.
Accompanying drawing explanation
Fig. 1 is the overall configuration figure that under conventional configuration, satellite platform and load are formed;
Fig. 2 is the overall configuration figure adopting the satellite platform of the inventive method and load to form;
Fig. 3 be that load completes identical sensing task and satellite body is not controlled time attitude misalignment Comparative result figure;
Fig. 4 is that load completes identical sensing task and satellite body completes attitude stabilization task, the attitude misalignment comparison diagram of body under identical controller;
Fig. 5 is that load completes identical sensing task and satellite body completes attitude stabilization task, the attitude stability comparison diagram of body under identical controller;
Fig. 6 is that load completes identical sensing task and satellite body completes attitude stabilization task, the control torque amplitude comparison diagram of body under identical controller.
Detailed description of the invention
Below in conjunction with accompanying drawing, describe advantage of the present invention in detail.For by advantage displayization, the control imitation result below in conjunction with some spacecraft two systems explains the present invention program.
As shown in Figure 1, for the conventional configuration that satellite platform and load are formed, wherein load 2 is arranged on satellite platform 1 by turning cylinder 3, azimuth axis is connected joint and is provided with motor 4 with pitch axis, be used for respectively driving intermediate connector and load, the joint motor 4 of azimuth axis is arranged on satellite body 1, and the joint motor 4 of pitch axis is arranged on intermediate connector.Intermediate connector is the part connecting satellite body 1 and load 2.Load 2 is provided with tracker.Satellite platform 1 is provided with control moment gyroscope group (CMGs).Joint motor 4, for providing control torque, completes sensing task.In conventional configuration, export sensing control torque by joint motor 4, this makes in load 2 continuously and healthily flyback process, and satellite body 1 is subject to strong disturbance, and it is very unfavorable that this maintains attitude stabilization to satellite body 1; Actuating unit on satellite body 1 needs constantly to export compensating moment, and to maintain attitude stabilization, and attitude accuracy and degree of stability also reduce greatly.
Adopt configuration designing method provided by the invention, obtain configuration as shown in Figure 2.Wherein, the barycenter arranging load 2 overlaps with the intersection point of turning cylinder 3, is arranged on satellite body 1 by barycenter by load 2.Remove for exporting the drive motor pointing to control torque, namely the joint motor 4 being arranged on satellite body 1 and on intermediate connector is removed in Fig. 1, but retain original connection articulation structure, load 1 is connected by free joint 5 with satellite body 1, retain joint freedom degrees.Load 2 is installed one group of control moment gyroscope group (CMGs), to drive and the attitude of control load 2 relative to satellite body 1, and then points to tracking specific objective.As shown in Figure 2, body and load are equipped with a set of control moment gyroscope group (CMGs) respectively, complete sensing task to provide applied moment.Present invention employs the concept that low-disturbance moment drives, by stagger angle momentum exchange device (CMGs) in load 2, reduce the Dynamics Coupling degree between multi-body system well.
By the kinetic model of Kane establishing equation two body spacecraft, adopt the method that sliding formwork controls, respectively to satellite body and load design controller, and coupling is each other considered as to extraneous interference simultaneously.
Situation 1:
Simulate a low-orbit satellite, in this, as the sensing target of load optical axis.Control moment gyroscope group on satellite body not output torque completely, but the actuating unit in load still normally works, to ensure that load points to low-orbit satellite all the time.Because spacecraft body is subject to the coupling torque effect of load, its attitude inherently overturns.Because the sensing campaign of load in inertial system is identical, so the upset degree contrasting body attitude on this basis can reach the target judging Dynamics Coupling degree size.
In Fig. 3 to Fig. 5, the modulus value of quaternion arrow portion q characterizes satellite body attitude misalignment tolerance.
As shown in Figure 3, for load complete identical sensing task and satellite body is not controlled time attitude misalignment Comparative result figure, from figure, in 200 second time period, satellite body attitude misalignment tolerance is known, and conventional configuration is about 3 times of the new configuration of the present invention, and result is as shown in table 1.Obviously, under this illustrates the new configuration of the present invention, the attitude impact of load twirl on body is much less than conventional configuration.
Satellite body attitude misalignment tolerance under conventional configuration shown in table 1 and the new configuration of the present invention
| Conventional configuration | New configuration | |
| Attitude misalignment is measured | 0.042 | 0.015 |
Situation 2:
Assumed load is operated in the fast retrace stage, and be equivalent to make conical pendu0 at inertial space and point to motion, the scanning period is set as 5 seconds.Satellite body reaches attitude stabilization simultaneously.
Suppose that the actuating unit of body and the actuating unit of load normally work on star simultaneously, can the instruction moment of accurate o controller.
As shown in Figure 1, under conventional configuration, the sensing control torque of load is as the motor of joint, the applied moment that load is subject to such as will inevitably to produce at the moment of reaction reverse greatly, this moment of reaction acts on satellite body, and then produces the strong disturbance continued to the attitude stabilization of satellite body.The pose stabilization control device of satellite body, in order to maintain body attitude stabilization, will produce control command, and to offset or partial offset disturbs outward, this control torque is performed by the directing mechanism that satellite body is installed.
But as shown in Figure 2, under new configuration of the present invention, the sensing motion control moment of load is as the control moment gyroscope group be arranged in load, and turning cylinder place is free joint.The fast retrace of load will largely reduce the impact of satellite body, also will significantly reduce so satellite body is the energy maintaining attitude stabilization consumption.
Load completes identical sensing task and satellite body completes attitude stabilization task, under identical controller, the attitude misalignment contrast of satellite body as shown in Figure 4, as can be seen from the figure, under new configuration of the present invention, the stable state accuracy of satellite body is about 0.1 degree, and under conventional configuration, being about 0.3 degree, stable state accuracy is much improved.
Load completes identical sensing task and satellite body completes attitude stabilization task, under identical controller, the attitude stability of satellite body as shown in Figure 5, as can be seen from the figure, under new configuration of the present invention the attitude stability of satellite body be about 0.02 degree per second, and be about under conventional configuration 0.4 degree per second, attitude stability significantly improves.
Load completes identical sensing task and satellite body completes attitude stabilization task, under identical controller, the control torque amplitude of satellite body as shown in Figure 6, as can be seen from the figure, under new configuration of the present invention, the control torque amplitude of satellite body is about 0.02 N of rice, and under conventional configuration, being about 5 Ns of rice, control torque amplitude significantly reduces.
Under identical system motion state, contrast the controlling quantity consumed, as shown in table 2.
The control energy consumption comparison of the new configuration of the present invention and conventional configuration under table 2 identical systems state of kinematic motion
| Conventional configuration | New configuration | |
| Satellite body controls energy consumption | 14054 | 52.4590 |
| Load controls energy consumption | 27218 | 27172 |
| Overhead control energy consumption | 41272 | 27224 |
As can be seen here, no matter be new configuration or conventional configuration, the control energy consumption of load is identical, but the control energy consumption of satellite body then gap is very far away, illustrate that the disturbance that satellite body is subject under new configuration of the present invention significantly reduces, demonstrate the preceence of new configuration further to a great extent, namely reduce Dynamics Coupling degree.The controlling quantity total amount completing identical motion control consumption under new configuration is only 65.96% of conventional configuration, saves energy consumption 34.04%.The advantage of this energy consumption will have larger meaning when researching and developing long-life satellite.
Claims (1)
1. two-dimentional pointing system configuration designing method on a star, it is characterized in that, based on the configuration that satellite platform and load are formed, wherein load is arranged on satellite platform by turning cylinder, azimuth axis is connected joint with pitch axis and is provided with motor, be used for respectively driving intermediate connector and load, the joint motor of azimuth axis is arranged on satellite body, and the joint motor of pitch axis is arranged on intermediate connector; Intermediate connector is the part connecting satellite body and load; Load is provided with tracker; Satellite platform is provided with control moment gyroscope group; Joint motor is used for providing control torque, completes sensing task;
On described star, two-dimentional pointing system configuration designing method comprises following aspect:
The first step: the intersection point of the barycenter and turning cylinder that arrange load overlaps;
Second step: remove the drive motor be arranged on satellite body and intermediate connector, load is connected by free joint with satellite body;
3rd step: install one group of control moment gyroscope group in load, to provide the moment pointed to and control to need.
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| CN201410161957.4A CN103921959B (en) | 2014-04-22 | 2014-04-22 | Two-dimentional pointing system configuration designing method on star |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN105867435A (en) * | 2016-05-11 | 2016-08-17 | 西北工业大学 | Smooth and steady pointing maneuvering control method for satellite optical load |
| CN106313031A (en) * | 2016-10-11 | 2017-01-11 | 北京航空航天大学 | Space mechanical arm driven by control torque spinning tops |
| CN110826251B (en) * | 2019-11-13 | 2020-10-20 | 北京理工大学 | Liquid-filled flexible spacecraft dynamics modeling method based on Kane equation |
| CN114408215B (en) * | 2021-12-27 | 2024-02-09 | 航天东方红卫星有限公司 | Satellite configuration suitable for rapid maneuvering ultra-stable imaging |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4375878A (en) * | 1980-10-28 | 1983-03-08 | Lockheed Missiles & Space Company, Inc. | Space satellite with agile payload orientation system |
| CN2204069Y (en) * | 1994-08-05 | 1995-07-26 | 电子工业部第五十四研究所 | Two axil aerial device with top stabilising platform |
| US6135389A (en) * | 1998-03-16 | 2000-10-24 | Hughes Electronics Corporation | Subterranean target steering strategy |
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Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4375878A (en) * | 1980-10-28 | 1983-03-08 | Lockheed Missiles & Space Company, Inc. | Space satellite with agile payload orientation system |
| CN2204069Y (en) * | 1994-08-05 | 1995-07-26 | 电子工业部第五十四研究所 | Two axil aerial device with top stabilising platform |
| US6135389A (en) * | 1998-03-16 | 2000-10-24 | Hughes Electronics Corporation | Subterranean target steering strategy |
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