CN1025995C - Attitude pointing error correction system and method for geosynchronous satellites - Google Patents
Attitude pointing error correction system and method for geosynchronous satellites Download PDFInfo
- Publication number
- CN1025995C CN1025995C CN89109644A CN89109644A CN1025995C CN 1025995 C CN1025995 C CN 1025995C CN 89109644 A CN89109644 A CN 89109644A CN 89109644 A CN89109644 A CN 89109644A CN 1025995 C CN1025995 C CN 1025995C
- Authority
- CN
- China
- Prior art keywords
- satellite
- momentum vector
- around
- yaw axis
- wobble shaft
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
Images
Classifications
-
- H—ELECTRICITY
- H04—ELECTRIC COMMUNICATION TECHNIQUE
- H04B—TRANSMISSION
- H04B7/00—Radio transmission systems, i.e. using radiation field
- H04B7/14—Relay systems
- H04B7/15—Active relay systems
- H04B7/185—Space-based or airborne stations; Stations for satellite systems
- H04B7/19—Earth-synchronous stations
Abstract
The present invention relates to an attitude control system and a method for saving fuel and prolonging the service life of a synchronous satellite, which can compensate the error of the rolling rotation and the yawing direction caused by the orbit deviating from the plane of a standard equatorial orbit. A universal rack device combines the momentum vector determined by inertia with a planet to form a single free degree relationship, preferably a double free degree relationship, with the momentum vector. In a double free degree embodiment, the momentum vector is established by the self rotation of a satellite or a momentum wheel provided with an axis and a roller and coincided with a yawing axis. A rotating torque generator of the universal rack applies the rotating torque of a winding amount vector with the time change so as to correct the error. In addition, a common propelling device system correcting the error of a pitching axle and an orbit so as to complete a position keeping function is arranged.
Description
The present invention relates to the attitude control of synchronous satellite, more particularly, relate to track is departed from nominal equatorial orbit plane and attitude control system and method that the lift-over that causes and driftage error in pointing compensate.
Communication and maritime satellite one boat are arranged on the circuit orbit that is called synchronous orbit or geostationary orbit, have the rotation period identical with the earth so that synchronous speed to be provided.On the orbit plane that this satellite should be arranged on theoretically with earth equatorial plane overlaps, therefore, satellite antenna can point to the position of the earth of expectation.In general, synchronous satellite by around self rotation or install a momentum wheel and realize that momentum is stable, make spin axis keep vertical, and calibrate the earth wave beam optical axis and make it perpendicular to spin axis with the equatorial orbit plane.Under this ideal state, the earth wave beam optical axis points to the sub-satellite point zone all the time when satellite and the earth rotate synchronously.
Some factors can cause the track drift, make satellite orbit favour nominal equatorial orbit plane.This track tilts and can produce lift-over and driftage error in pointing along with time integral.Specifically, the Sun and the Moon can cause the orbit perturbation effect to the variation of the earth gravitational field that non-sphere produced of the graviational interaction of satellite and the earth, and this makes the satellite orbit plane inclination in desirable equatorial plane.The clean effect of these orbit perturbations influence is the inclination that causes satellite orbit, thereby with annual 0.75 ° to 0.95 ° the slow drift of speed.
Because the increase of orbit inclination, lift-over and driftage error in pointing cause that the earth illuminance model of satellite antenna is from desirable target area drift.For example, as illustrated in fig. 1 and 2, satellite ' S ' is along the orbital motion around the earth, this track and equatorial orbit plane inclination i angle, with the equatorial orbit Plane intersects in an ascending node Na and a descending node Nd, at ascending node Na place, satellite runs to the Northern Hemisphere from the southern hemisphere, and at descending node Nd, satellite runs to the southern hemisphere from the Northern Hemisphere.At satellite when it rises a some Na and moves to its north latitude and spend, through its northern antipoints Nn, and when its descending node Nd moves to its south latitude and spends, through its southern antipoints Ns.
Because the inclination angle i between satellite actual track and nominal equator, satellite invest the harmful effect that the antenna illuminance model of earth surface will be subjected to south-northern sinusoidal variations and rotate sinusoidal variations, respectively corresponding to the roll error and the yaw error of outer space vehicle.For example, as shown in Figure 2, the spin axis of satellite put down perpendicular to the bevelled track situation under, when satellite during through its ascending node Na, the roll error of earth illuminance model (Fig. 3 A) is zero, and yaw error (Fig. 3 B maximum.When satellite during to its northern antipoints Nn operation, roll error increases until reaching maxim at northern antipoints Nn, and yaw error is kept to zero.As shown in Figure 2, when satellite during at its northern antipoints Nn, the S of the directed earth surface of the earth wave beam optical axis
1And when satellite left to descending node Nd from its northern antipoints Nn, roll error was decreased to zero, and yaw error is increased to its maxim once more.When satellite walks to its southern antipoints Ns, as shown in Figure 2, the earth wave beam optical axis with directed earth table some S
2
Rolling that orbit inclination causes and yaw error depend on the direction of outer space vehicle spin axis.Generally speaking, when spin axis from perpendicular to the axle tilt alpha angle of equatorial plane the time, roll error will be (1,178i-α) sin nt, yaw error is-α cos nt, and i is an orbit inclination in the formula, and n is an orbit angular velocity, t is the time, t=0 when at ascending node.Obviously, lift-over and yaw error have functional relation, and one of them can be used as another function and obtains.
A kind of is to have a mind to make the aircraft spin axis to favour the equatorial orbit normal for reducing the technology that the lift-over error in pointing proposes.As shown in Figure 2, satellite spin axle (dotted line is represented) is tilted an angle θ so that make the earth wave beam visual angle of satellite be positioned to be in the region S O that satellite obtained in the equatorial orbit once more effectively.Though roll error will be reduced to zero effectively, yaw error is owing to the spin axis inclination angle [theta] increases, and its value is-(i+ θ) cos nt.When utilizing communication of figure polarization or narrow spot beam (narrow spot beam), unacceptable between the increase of this yaw error.
In common satellite system, utilize the propelling unit consume fuel regularly to revise the inclination of track.This method is called the locus, north and south and keeps method (north-south station-keep-ing).Specifically, during 10 years, the maintenance effect of this position may need 20% of the initial gross weight of satellite, and wherein propellant accounts for main portion, and about 90% is used for the track correction, and all the other are used for handling in other track, comprise pitch error correction (Fig. 3 c).In general, the work life of satellite is subjected to the restriction that the locus keeps required fuel, thereby locus, inactive north and south keeps system can prolong its work life.But the locus, north and south of stopping using keeps system can cause the attitude error that must revise.
For tilt correction is handled the understanding that needs satellite to carry fuel in a large number, various attitude control systems had once been proposed to revise the attitude error that track tilts to cause.For example, US Patent 4,084, a kind of aircraft lift-over/yaw direction Adjustment System is disclosed for No. 772, wherein, aircraft comes stable by a momentum wheel, the angular rate of momentum wheel changes with the aircraft momentum that changes association with sinusoidal manner in the orbit rotation process and makes the aircraft roll attitude make sinusoidal variations with orbit rotation each time.US Patent 4,062 provides the means of a kind of lift-over and yaw attitude control thereby disclose a kind of torque system in No. 509 to set up aircraft magnetic field and earth magnetic field mutual action.
The invention provides a kind of system and method for synchronous satellite attitude control, be used for track is departed from nominal equatorial orbit plane and the error in pointing that causes compensates.For satellite is set up a momentum vector, by a gimbal table system be fixed on inertia fixed between and be attached on the satellite, this gimbal table system provides the relation of at least one degree of freedom between aircraft and this momentum vector.The gimbal table axle is provided with each satellite of gimbal table torque generator along at least one main shaft of satellite such as wobble shaft and/or yaw axis provides the torque of the momentum vector of determining around inertia to revise attitude error.Depend on the moment of momentum direction and can obtain according to the function of orbit inclination and satellite position in orbit by the tilt lift-over cause and yaw error of track with analysis method.Depend on concrete structure, the gimbal table torque generator rotates correction with the error in indication of satellite in the earth rotates according to appropriate time relationship around the axis of rolling and/or yaw axis.Remove beyond the lift-over and yaw error that causes by orbit inclination, the additional error in indication that is caused by other external disturbance torque such as sun torque is by common attitude control system correction, and this common posture control system is made up of an earth sensor and appearance control torque generator.
In the first embodiment of the present invention, the satellite of one spin stabilization has a spin component, the momentum vector that it can provide an inertia to determine, also has a despun antenna assembly, this assembly is attached to spin component by the gimbal table device of a double freedom, this gimbal table device has one first gimbal table, and it can rotate around wobble shaft, and another gimbal table then can rotate around yaw axis.Be provided with the gimbal table torquer and provide torque to the gimbal table relevant respectively, thereby apply the lift-over and the error in pointing of going off course that tilts to cause with the correction track around the torque of the definite momentum vector of the inertia that freely revolves part foundation to antenna assembly with wobble shaft correction or yaw axis correction.Lift-over and driftage gimbal table torquer are driven with a kind of sinusoidal manner in 24 hours periods.Realize that as the moment of momentum device and by double freedom gimbal table device the method that selected controlled connection is arranged with antenna assembly provides a kind of important means of error in pointing correction to need not consume fuel and carried out the track tilt correction by utilizing satelloid.Though preferably adopt double freedom relation,, the single degree of freedom relation that also can adopt in wobble shaft or the yaw axis at least one is to realize the correction along this.
In another embodiment of the present invention, a momentum wheel is connected on the aircraft by the gimbal table device of a double freedom, is provided with torquer to realize the rotation of aircraft around the definite momentum vector of inertia in rolling of gimbal table device upper edge and yaw axis.The same with the situation of first embodiment, lift-over and driftage gimbal table torquer are driven with a kind of sinusoidal manner in 24 hours periods.
In another embodiment of the present invention, when selecting the moment of momentum direction, make one of two kinds of errors, promptly roll error or the yaw error that is tilted to cause by track is zero.Another error then is provided with a single degree of freedom gimbal table and revises, this gimbal table is provided with and aircraft is rotated to compensate that error along this, this is to be finished together by many torques that act on gimbal table, and the aircraft momentum vector definite with respect to inertia rotated.
The invention has the advantages that attitude control system and method that a kind of synchronous satellite is provided, can compensate the drift lift-over that causes and the error in pointing of going off course that depart from the nominal equatorial orbit by track easily, come system and method fuel saving of the present invention and prolonged the work life of satellite greatly compared with the prior art system and the device that need consume fuel to revise orbit error.
Contrast the following drawings now, by way of example, describe the present invention in detail:
Fig. 1 be equatorial orbit plane and bevelled around the schematic perspective view of Earth Orbit Plane, show each node of orbit and antipoints;
Fig. 2 is the two-dimensional representation of inclination shown in Figure 1 and equatorial orbit;
Fig. 3 A is earth surface and the wobble shaft error in pointing view to the effect of earth antenna illuminance model;
Fig. 3 B is earth surface and the yaw axis error in pointing view to the effect of earth antenna illuminance model;
Fig. 3 C is earth surface and faces upward the view of an error in pointing to the effect of earth antenna illuminance model of bowing;
Fig. 4 is the schematic perspective view of first embodiment of the invention;
Fig. 5 be an employed double freedom gimbal table device embodiment illustrated in fig. 4 solid;
Fig. 6 realizes the schematic block diagram to the typical control loop of gimbal table control;
Fig. 7 is the schematic perspective view of second embodiment of the invention;
Fig. 8 is the schematic perspective view of third embodiment of the invention;
Fig. 9 is the schematic perspective view of fourth embodiment of the invention.
Shown in Figure 4 is the schematic perspective view that adopts a satellite 10 of the present invention.Satellite 10 is the spin stabilization types that adopt in synchronous orbit, comprises a spin component 12 and a despun antenna tower 14.Spin component 12 is used for around aircraft principal axis A X rotation, is a kind of common design, comprises a cylindrical basically housing, for example, is the longeron formula, also comprises a racemization motor and bearing assembly 16.Depend on its flight mission, satellite 10 is being equipped suitable tracking, remote measurement, and instruction repertoire; Main power system, thermal control system, and propulsion system.As shown in Figure 4, satellite 10 is equipped with a propeller control system, and it comprises the first, the second and triple screw T
1, T
2And T
3These propelling units Tn is common design, with ejection propellant (generally being hydrazine hydrazine), is used for changing the moment of momentum of satellite 10 in response to signal control valve work.Propelling unit Tn shown in Figure 4 depends on the structure of satellite just for example, also can adopt other pusher mechanism.
Tilt the additional error that also may exist other external disturbance torque such as sun torque to cause the lift-over and direction point error cause by track except above-mentioned.These additional errors can adopt common attitude control system to revise, and common attitude control system is made of earth sensor and appearance control torquer.The output valve of earth sensor to attitude controller for example propelling unit provide, then propelling unit to external world the satellite attitude that causes of perturbing torque revise.The preceding also U.S. discloses a kind of earth sensor 4,084, No. 722.
The gimbal table device 22 that schematically draws in Fig. 4 and draw in detail in Fig. 5 is connected between racemization motor and bearing assembly 16 and the antenna tower 14.It can be that lift-over and yaw axis are done selectively to tilt along two axis with respect to spin component 12 that gimbal table device 22 makes antenna tower 14.Therefore, antenna tower spindle axis Aant can be controlled, and it is overlapped with the spindle axis Ax of spin component 12, perhaps makes itself and principal axis A x be an oblique angle.As shown in Figure 5, gimbal table device 22 comprises the Inner Gimbal 28 that 24, one outer gimbal tables 26 of a carrier ring that structurally links to each other with antenna tower 14 and link to each other with bearing assembly 16 by suitable member such as hollow pillar (drawing) and racemization motor.Outer gimbal table 26 is connected to carrier ring 24 by an outer gimbal table torquer 30 and an outer gimbal pick-up 32, and for example pivot center separately can be aimed at a lift-over line jointly.Link to each other with gimbal table 28 with an Inner Gimbal position transduser 36 by an Inner Gimbal torquer 34 with the outer gimbal table 26 of a kind of similar mode, its pivot center separately is common and yaw axes ground is accurate.Gimbal table torquer 30 and 34 is a boat design, as an electro-motor train of gears so that relevant gimbal table is relatively rotated.Gimbal pick- up 32 and 36 provides and originates in gimbal table relative angle position and concern output signal, and gimbal pick- up 32 and 34 for example can comprise that resolver or optical encoder are to provide position, the angle information of needs.Be provided with the gimbal table hill holder so that the angular transposition of gimbal table is limited in the acceptable limit.
Gimbal table 26 and 28 motion and last location are by a universal control loop control, and Fig. 6 shows an example of this loop block diagram.As shown in Figure 6, torquer actuating device 38 receives one from a command source 40 and specifies incoming signal ' CMD ' generation of ideal position and provide a suitable electrical output signal to torquer, and torquer then will drive new location with the gimbal table (dotting among Fig. 5) of its mechanical connection.Command source 40 can be partly by from GCI ground control installation or from the star the instruction of disposal system one incoming signal ' CMD ' is provided.One time meter CLK provides a timing wave t of 24 hours, is being in ascending node Na(Fig. 1) time, t=0.Therefore, when satellite when the earth rotates, the gimbal table control signal can sinusoidal manner change in time.Specifically, hereinafter will tell about more comprehensively, command signal ' CMD ' comprises that one is used for the function sin nt of wobble shaft correction and the function cos nt that is used for the yaw axis correction.Gimbal pick-up provides a feedback electric signal to torquer actuating device 38, and indication is loaded with the position of the gimbal table of the torquer that actuating device 38 controlled, so that realize towards the motion of ideal position and remain on this position.
For lift-over and the yaw axis correction that embodies satellite shown in Figure 4, satellite spin axle Ax be preferably in be registered at first with perpendicular position, bevelled satellite orbit plane on (see figure 2), this inclination also makes the normal on spin rotating shaft rail plane be angle i.30 of wobble shaft gimbal table torquers are by time dependent sinusoidal cmd signal-θ sin nt control, the maxim of θ is (0.178i) in the formula, and yaw axis gimbal table torquer is simultaneously by a time dependent sinusoidal cmd signal i cos nt control.Value 0.178i promptly determines that by earth radius and track profile n represents the angular rate of track by the geometric configuration of synchronous orbit.Biasing factor to the one-θ of earth sensor control loop introducing of satellite sin nt makes roll error biasing-θ sin nt, setovers with realization-θ sin nt.Obviously, antenna tower 14 will be aimed at again, make axis Ax that its axis Aant aims at spin component 12 with the lift-over that realizes satellite 10 each commentaries on classics and driftage to error continually varying correction in time.Though, initial to punctual satellite spin axle preferably perpendicular to the bevelled orbit plane of satellite 10, this aligning is not necessary, can carry out the aligning of alternate manner yet, and is definite by inertia as long as the momentum vector of spin component 12 remains.
The embodiment that shows of Fig. 4 is described by two degree of freedom.If desired, the momentum vector that is provided by spin component 12 can be to connect by the single degree of freedom of gimbal table axle along wobble shaft or yaw axis aligning with being connected of antenna tower 14.Therefore, under the situation that this single degree of freedom connects with the yaw axis axle is aimed at, in the earth sensor control loop of satellite, introduce of the correction of a biasing factor, and in the gimbal table actuating spindle rotates, revise the yaw axis error termly with implementation wobble shaft error in pointing.
Fig. 7 has drawn and has adopted second embodiment of satellite 50 of the present invention.As shown in Figure 7, satellite 50 is a parallelepiped, and wherein a part is cut open to represent its inner structure.For clarity sake, the sun wing has been shown in Fig. 7 part omitted, antenna and propelling unit.
As embodiment illustrated in fig. 5, embodiment illustrated in fig. 7 by aircraft spin or provide a momentum vector by the spin that other establishes object, and be connected with momentum vector by a double freedom device.So that provide torque to aircraft, thus make its can around its lift-over and yaw axis rotate to since satellite orbit revise with respect to the lift-over and the error in pointing of going off course that the inclination of nominal equatorial plane causes.Remove use a foregoing two degrees of freedom gimbal table device to one of satellite effect beyond the torque of the momentum vector H that inertia is determined, it will also be appreciated that, also can use a single gimbal table momentum wheel and a propeller control system to carry out the correction of error in pointing, this point will be below to talking about in Fig. 8 and the description embodiment illustrated in fig. 9.
As shown in Figure 8, a satellite 70 is provided with an earth sensor 72 and a gimbal table momentum wheel 74, and this is taken turns 74 and can rotate to form a momentum vector H around axis AmW.Momentum wheel 74 is connected in Flight Vehicle Structure by a single degree of freedom gimbal table 76.An one gimbal table torquer 78 and a gimbal pick-up and yaw detector 80 are collinear relationship with yaw axis when mounted.Gimbal table 76 is controlled by the sort of control loop shown in Figure 6 along moving in the whole process of track operation.Programmable gimbal table actuating device comprises a time meter of on-cycle when the gimbal table precession is gone by 24.Wobble shaft is biased a θ sin nt, and θ is the inclination angle in the formula, and its maxim is 0.178i, and i is an orbit inclination.Use one or more propelling units implement lift-over control in the ordinary way in case the momentum wheel spin axis from the normal slope θ angle of inclined plane.Then, the aircraft momentum wheel axis Anw definite with respect to inertia rotates (i+ θ) cos nt angle to revise the error that track tilts to cause around yaw axis.Roll error biasing and gimbal table can be provided by disposal system on GCI ground control installation or the star along the instruction of yaw axis precession.
Another embodiment of the present invention as shown in Figure 9, with embodiment illustrated in fig. 8 the same, it comprises a momentum wheel that links to each other with Flight Vehicle Structure by a single degree of freedom gimbal table.As shown in the figure, satellite 90 comprises an earth sensor 92 and a momentum wheel 94, and momentum wheel 94 rotates to form a momentum vector H around axis Amw.Momentum wheel 94 is connected in Flight Vehicle Structure by a single degree of freedom gimbal table 96.When one a gimbal table torquer 98 and a gimbal table sensor 100 are installed and wobble shaft be collinear relationship.The motion of gimbal table 96 in the whole process of track operation is subjected to the control of the sort of control loop shown in Figure 6, and programmable gimbal table actuating device comprises can make the gimbal table precession by 24 hours on-cycle time meters.Momentum wheel 94 from spinning roller Anw nominal ground along axle holding position, north and south, perpendicular to the equatorial orbit plane.Around wobble shaft pivot angle-(i+ θ) sin nt, θ is the song angle to satellite 90 in the formula with respect to the definite momentum wheel axis Amw of inertia, and its maxim is 0.178i, roll error is biased-and the bias of θ sin nt.Instruction about roll error biasing and wobble shaft gimbal table angle is provided by disposal system on GCI ground control installation and the star.GCI ground control installation is considered the size of variation regular update roll error biasing of orbit inclination and the angle of wobble shaft gimbal table.
The invention has the advantages that provides a kind of system and method for controlling synchronous satellite, the error in pointing that can the drift of the relative nominal equatorial plane of a kind of fuel-saving mode serial update orbit inclination angle causes. Therefore the present invention service life that can greatly improve synchronous satellite, this is owing to greatly reduced the result that the position of satellite keeps needed fuel.
Claims (47)
1, the method for the error in pointing that causes of a kind of correction synchronous satellite (10) orbit inclination, described synchronous satellite has a momentum vector (H), with the connecting device of the single degree of freedom of at least one coincidence selected in wobble shaft and the yaw axis described momentum vector (H) and satellite (10) are connect mutually, with realize satellite (10) with respect to momentum vector (H) around the above-mentioned selected wobble shaft or the rotation of yaw axis, it is characterized in that:
Set up described momentum vector (H) along one that is not orthogonal to the equatorial orbit plane; And
With with the synchronous time dependent mode of the track of satellite (10) and according to the function of orbit inclination i, apply torque around described selected axle to satellite (10).
2, in accordance with the method for claim 1, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device (22) that overlaps with wobble shaft connects described momentum vector and satellite (10) mutually, with realization satellite (10) with respect to of the rotation of described momentum vector around wobble shaft.
3, in accordance with the method for claim 2, it is characterized in that:
In the described step that applies torque, apply commentaries on classics distance around wobble shaft according to function-θ sin nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
4, in accordance with the method for claim 1, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device (22) that overlaps with yaw axis connects described momentum vector and satellite (10) mutually, with realization satellite (10) with respect to of the rotation of described momentum vector around yaw axis.
5, in accordance with the method for claim 4, it is characterized in that:
In the described step that applies torque, apply torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate.
6, in accordance with the method for claim 5, it is characterized in that:
In the described step that applies torque, along setting up described momentum vector with the direction at the normal slope θ angle of inclined plane;
Apply torque around described yaw axis according to function (i+ θ) cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
7, in accordance with the method for claim 1, it is characterized in that:
In the described connection step, by a double freedom, with the connecting device (22) that lift-over and yaw axis overlap described momentum vector and satellite (10) are connect mutually respectively, to realize that satellite (10) is with respect to the rotation of described momentum vector around wobble shaft and yaw axis.
8, in accordance with the method for claim 7, it is characterized in that:
In the described step that applies torque, apply torque according to function-θ sin nt to satellite (10) around wobble shaft, apply torque according to function i cos nt to satellite (10) around yaw axis, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
9, in accordance with the method for claim 8, it is characterized in that:
The described step that applies torque is set up momentum vector along a normal on inclined plane plane;
Apply torque around wobble shaft according to function-θ sin nt to satellite (10), apply torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
10, the method for the error in pointing that causes of a kind of correction synchronous satellite (10) orbit inclination, satellite (10) is for having one from the sort of type of executing a part (12) and racemization part (14), described spin component (12) and disappearing is executed part (14) by at least one single degree of freedom, be tied with at least one connecting device that overlaps (22) selected in wobble shaft and the yaw axis, to realize that racemization part (14) with respect to the rotation of described momentum vector around described selected axle, is characterized in that having following steps:
Set up described momentum vector along one that is not orthogonal to the equatorial orbit plane; And
With the synchronous time dependent mode of a kind of and satellite (10) track and according to the function of orbit inclination i, execute part (14) and apply torque around described selected axle to described disappearing.
11, in accordance with the method for claim 10, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device (22) that overlaps with wobble shaft connects the momentum vector of described spin component (12) mutually with racemization part (14), with the realization racemization partly (10) with respect to of the rotation of described momentum vector around wobble shaft.
12, in accordance with the method for claim 10, it is characterized in that:
Described revolving in the step that adds torque applies torque around lift-over according to function-θ sin nt to racemization part (14), and t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
13, in accordance with the method for claim 10, it is characterized in that: by at least one single degree of freedom, the connecting device (22) that overlaps with yaw axis connects the momentum vector of described spin component (12) mutually with racemization part (14), so as the realization racemization partly (14) with respect to the rotation of momentum vector around yaw axis.
14, in accordance with the method for claim 13, it is characterized in that:
Described revolving in the step that adds torque applies torque around yaw axis according to function i cos nt to racemization part (14), and t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate.
15, in accordance with the method for claim 14, it is characterized in that:
Described revolving in the step that adds torque is along setting up described momentum vector with the direction at inclined plane normal slope θ angle;
Apply torque around described yaw axis according to (i+ θ) cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and 0 maxim is 0.178i.
16, in accordance with the method for claim 10, it is characterized in that:
In the described connection step, by a double freedom, the connecting device (22) that overlaps with wobble shaft and yaw axis connects the momentum vector of spin component (12) mutually with racemization part (14) respectively, with the realization racemization partly with respect to of the rotation of described momentum vector around wobble shaft and yaw axis.
17, in accordance with the method for claim 16, it is characterized in that:
In the described step that applies torque, apply torque according to function-θ sin nt to satellite (10) around wobble shaft, apply torque according to function i cos nt to satellite (10) around yaw axis, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
18, in accordance with the method for claim 17, it is characterized in that:
In the described step that applies torque, set up described momentum vector along a normal on inclined plane plane;
Apply torque around wobble shaft according to function-θ sin nt to satellite (10), apply torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
19, the method for the error in pointing that causes of a kind of correction synchronous satellite (10) orbit inclination, synchronous satellite (10) has at least one quality (54,74,94) around a rotation to set up a momentum vector along this, by at least one single degree of freedom, with at least one selected in wobble shaft and the yaw axis connecting device that overlaps (56,58,76,96) described momentum vector is connect mutually with satellite (10), to realize that satellite (10) with respect to the rotation of described momentum vector around described selected axle, is characterized in that having following steps:
Set up described momentum vector along one that is not orthogonal to the equatorial orbit plane; And
Apply torque with the synchronous time dependent mode of a kind of and satellite (10) track and according to the function of orbit inclination i to satellite (10) around described selected axle.
20, in accordance with the method for claim 19, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device that overlaps with wobble shaft (58,96) connects described momentum vector mutually with satellite (10), so that realize that satellite (10) is with respect to the rotation of described momentum vector around wobble shaft.
21, in accordance with the method for claim 20, it is characterized in that:
In the described step that applies torque, apply torque around wobble shaft according to function-θ sin nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
22, in accordance with the method for claim 19, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device that overlaps with wobble shaft (56,72) connects described momentum vector mutually with satellite (10), so that realize that satellite (10) is with respect to the rotation of described momentum vector around wobble shaft.
23, in accordance with the method for claim 22, it is characterized in that:
In the described step that applies torque, apply torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate.
24, in accordance with the method for claim 23, it is characterized in that:
In the described step that applies torque, along setting up described momentum vector with an axis at inclined plane normal slope θ angle;
Apply torque around yaw axis according to (i+ θ) cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
25, in accordance with the method for claim 19, it is characterized in that:
In the described connection step, by at least one single degree of freedom, the connecting device (56,58) that overlaps with wobble shaft and driftage connects described momentum vector mutually with satellite (10) respectively, so that realize that satellite (10) is with respect to the rotation of described momentum vector around wobble shaft and yaw axis.
26, in accordance with the method for claim 25, it is characterized in that:
The described step that applies torque applies torque around wobble shaft according to function # sin nt to satellite (10), applies torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
27, in accordance with the method for claim 26, it is characterized in that:
In the described step that applies torque, set up described momentum vector along a normal on inclined plane plane;
Apply torque around wobble shaft according to function-sin nt to satellite (10), apply torque around yaw axis according to function i cos nt to satellite (10), t is the time in the formula, and at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
28, in accordance with the method for claim 27, wherein synchronous satellite (10) is to have earth sensor (52) control loop to revise the sort of type of described attitude to determine satellite (10) with respect to the attitude of the earth and in response to the output of earth sensor (52), the method is characterized in that further comprising the steps of:
Send into a bias so that setover the described effect that applies the torque step to earth sensor (52) control loop.
29, the attitude control system of a kind of synchronous satellite (10), this satellite (10) is by a quality (12,54,74,94) set up a momentum vector around the rotation of spin axis, and a gimbal table (22,56,58,76,96), be used at least a portion of the synchronous satellite (10) of momentum vector is connect mutually, thereby control satellite (10) is at least one selected relatively rotating in wobble shaft or yaw axis, it is characterized in that:
Described momentum vector is to set up along an axle that is not orthogonal to the equatorial orbit plane; And
One control setup realizes that satellite (10) is with the synchronous time dependent mode of a kind of and satellite (10) track and according to the function of orbit inclination i, around described selected controlled rotation thereby be used to drive described gimbal table.
30, according to the attitude control system of the described synchronous satellite of claim 29 (10), it is characterized in that: the described device of setting up momentum vector is one and can rotates to form along the momentum wheel (54 of a momentum vector of this spin axis around a spin axis, 74,94).
31, according to the attitude control system of the described synchronous satellite of claim 29 (10), it is characterized in that: the spin component (12) that the described device of setting up momentum vector is satellite (10), it rotates around a spin axis with respect to the racemization part (14) of satellite (10).
32, according to the attitude control system of the described synchronous satellite of claim 29 (10), it is characterized in that: described hookup is a single degree of freedom gimbal table device (96), its axis overlaps with wobble shaft, to realize that satellite (10) is with respect to the rotation of described momentum vector around wobble shaft.
33, according to the attitude control system of the described synchronous satellite of claim 32 (10), it is characterized in that: described gimbal table device is controlled according to function-θ sin nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
34, according to the attitude control system of the described synchronous satellite of claim 29 (10), it is characterized in that: described hookup is a single degree of freedom gimbal table device, and its axis overlaps with yaw axis to realize that satellite (10) is with respect to the rotation of described momentum vector around yaw axis.
35, according to the attitude control system of the described synchronous satellite of claim 34 (10), it is characterized in that: described gimbal table device is controlled according to function icos nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
36, according to the attitude control system of the described synchronous satellite of claim 35 (10), it is characterized in that: the described device of setting up momentum vector is along setting up described momentum vector with an axis at the normal slope θ angle of bevelled orbit plane, and described gimbal table device is controlled according to (i+ θ) cos nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and i is an orbit inclination.
37, according to the attitude control system of the described synchronous satellite of claim 29 (10), it is characterized in that: described hookup is a double freedom gimbal table (56, a 58) device, and its axis overlaps with the wobble shaft and the yaw axis of satellite (10).
38, according to the attitude control system of the described synchronous satellite of claim 37 (10), it is characterized in that: wobble shaft is controlled according to function-θ sin nt by described actuating device, and t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
39, according to the attitude control system of the described synchronous satellite of claim 38 (10), it is characterized in that: the described device of setting up momentum vector is along setting up described momentum vector with a normal of bevelled orbit plane, and wobble shaft is controlled according to-(i+ θ) sin nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and the maxim of θ is 0.178i.
40, according to the attitude control system of the described synchronous satellite of claim 37 (10), it is characterized in that: yaw axis is controlled according to function icos nt by described actuating device, and t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and i is an orbit inclination.
41, according to the attitude control system of the described synchronous satellite of claim 40 (10), it is characterized in that: the described device of setting up momentum vector is along setting up described momentum vector with a normal of bevelled orbit plane, and yaw axis is controlled according to i cos nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and i is an orbit inclination.
42, according to the attitude control system of the described synchronous satellite of claim 37 (10), it is characterized in that: wobble shaft is controlled according to-θ sin nt by described actuating device, yaw axis is controlled according to function i cos nt by described actuating device, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and i is an orbit inclination, θ has a maxim, is equivalent to earth radius multiply by orbit inclination to synchronous altitude ratio.
43, the attitude control system of a kind of synchronous satellite (10), satellite (10) is to have a spin component (12) of setting up a momentum vector and one to disappear and execute the sort of type of part (14), and be included in double freedom hookup (22) between spin component (12) and the racemization part (14), be used for realizing the controlled rotation of the spin component (12) of satellite and racemization part (14) around the wobble shaft and the yaw axis of quadrature, described wobble shaft and yaw axis are limited by described double freedom hookup (22), wobble shaft is tangential on orbital path, yaw axis perpendicular to wobble shaft, is characterized in that in orbit plane:
One control setup is used to control described hookup, realizing controlled relatively rotate of described spin component (12) and racemization part between (14), described controlled relatively rotating is synchronously to change in time and carry out according to the function of orbit inclination i around the track of the earth with satellite (10).
44, according to the described attitude control system of claim 43, it is characterized in that: described momentum vector is to set up along the axis that is not orthogonal to the equatorial orbit plane.
45, according to the attitude control system of the described synchronous satellite of claim 44 (10), it is characterized in that: described control setup control described hookup (22) with provide around wobble shaft according to relatively rotating of carrying out of function-θ sin nt and around yaw axis according to relatively rotating that function i cos nt carries out, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, i is an orbit inclination, θ has a maxim, is equivalent to earth radius multiply by orbit inclination to synchronous altitude ratio.
46, according to the attitude control system of the described synchronous satellite of claim 44 (10), it is characterized in that: described control setup is controlled described hookup (22) to provide around wobble shaft according to relatively rotating that-θ sin nt carries out, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, i is an orbit inclination, and the θ maxim is 0.178i.
47, according to the attitude control system of the described synchronous satellite of claim 44 (10), it is characterized in that: described control setup is controlled described hookup (22) to provide around yaw axis according to relatively rotating that function i cos nt carries out, t is the time in the formula, at the t=0 of ascending node place, n is the track angular rate, and i is an orbit inclination.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN89109644A CN1025995C (en) | 1989-12-30 | 1989-12-30 | Attitude pointing error correction system and method for geosynchronous satellites |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN89109644A CN1025995C (en) | 1989-12-30 | 1989-12-30 | Attitude pointing error correction system and method for geosynchronous satellites |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN1052828A CN1052828A (en) | 1991-07-10 |
| CN1025995C true CN1025995C (en) | 1994-09-28 |
Family
ID=4858102
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN89109644A Expired - Fee Related CN1025995C (en) | 1989-12-30 | 1989-12-30 | Attitude pointing error correction system and method for geosynchronous satellites |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN1025995C (en) |
Families Citing this family (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2955313B1 (en) * | 2010-01-19 | 2012-11-16 | Thales Sa | METHOD AND DEVICE FOR OPTIMIZING THE MASS OF A SATELLITE |
| FR2959490B1 (en) * | 2010-04-28 | 2012-07-13 | Astrium Sas | SATELLITE HAVING A SIMPLIFIED, ALLEGE AND ECONOMIC STRUCTURE AND ITS IMPLEMENTING METHOD |
| CN101907039B (en) * | 2010-07-23 | 2013-04-03 | 北京航空航天大学 | Nitrogen cold air micro-propeller adopting three cylindrical propelling agent storage boxes |
| CN101941527B (en) * | 2010-09-08 | 2012-11-14 | 北京航空航天大学 | Multi-sensor integrated attitude determining system structure |
| CN103274056B (en) * | 2013-04-24 | 2016-03-16 | 上海卫星工程研究所 | A kind of method utilizing data in-orbit to demarcate drift local time the in of satellite descending node |
| CN105197261B (en) * | 2015-08-24 | 2017-11-03 | 哈尔滨工业大学 | Towards the method for work of the quick Tum bling Target racemization cell sail of in-orbit service |
| CN109443813B (en) * | 2018-09-28 | 2020-06-09 | 中国空间技术研究院 | Rotation test method for satellite electric propulsion vector adjusting mechanism |
| FR3094479B1 (en) * | 2019-03-25 | 2021-08-27 | Airbus Defence & Space Sas | Device and method for determining the attitude of a satellite equipped with gyroscopic actuators, and a satellite carrying such a device |
| CN110466803B (en) * | 2019-07-03 | 2021-11-30 | 中国人民解放军63686部队 | Spinning stabilized satellite attitude prediction method based on equal-inclination-angle attitude control |
| CN111483618B (en) * | 2020-04-09 | 2021-10-01 | 上海航天控制技术研究所 | Multi-directional driving task autonomous management control method suitable for Mars detection |
| CN115196046B (en) * | 2022-09-19 | 2022-12-13 | 航天东方红卫星有限公司 | Method for determining orbit control strategy for super-life operation of sun-synchronous orbit satellite |
-
1989
- 1989-12-30 CN CN89109644A patent/CN1025995C/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| CN1052828A (en) | 1991-07-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US4911385A (en) | Attitude pointing error correction system and method for geosynchronous satellites | |
| AU639504B2 (en) | Satellite roll and yaw attitude control method | |
| RU2114770C1 (en) | Satellite with three-axis attitude stabilization and method of control of this satellite | |
| RU2212363C2 (en) | Method and system for injection of spacecraft into desired orbit by means of engines possessing high specific impulse | |
| US5697582A (en) | Method of adjusting the position of satellites by means of solar pressure torques | |
| EP0769736B1 (en) | Method for inclined orbit attitude control for momentum bias spacecraft | |
| EP0356602B1 (en) | Method of orienting a synchronous satellite | |
| US5692707A (en) | Universal spacecraft attitude steering control system | |
| JP2542094B2 (en) | Satellite control system | |
| US9027887B2 (en) | Method of controlling the attitude of a satellite and an attitude-controlled satellite | |
| US4230294A (en) | Closed loop roll control for momentum biased satellites | |
| CN1174982A (en) | Dynamic bais for controlling mavar orbital deviation | |
| US4262867A (en) | Apparatus for modifying position and attitude of a spacecraft | |
| EP0544198B1 (en) | Method and apparatus for controlling a solar wing of a satellite using a sun sensor | |
| CN1025995C (en) | Attitude pointing error correction system and method for geosynchronous satellites | |
| CA2094215A1 (en) | Spacecraft attitude control and momentum unloading using gimballed and throttled thrusters | |
| JPH04308906A (en) | Attitude controlling system for triaxially stabilized satellite in near equatorial orbit | |
| JPS5912000A (en) | Method of maintaining station northern and southern axis of artificial satellite in equator plane | |
| EP0668212A1 (en) | Optical satellite attitude control system | |
| US7343228B2 (en) | Transient cancellation technique for spacecraft solar wing stepping | |
| CN112650260B (en) | Solar sailboard variable-speed driving method under inclined orbit satellite yaw guidance | |
| CA2006199C (en) | Attitude pointing error correction system and method for geosynchronous satellites | |
| EP0544241A1 (en) | Method and apparatus for dynamic precompensation of solar wing stepping motions of a satellite | |
| JP2971494B2 (en) | Attitude control method and apparatus for geostationary satellite | |
| CA1305772C (en) | Method of orienting a synchronous satellite |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| C15 | Extension of patent right duration from 15 to 20 years for appl. with date before 31.12.1992 and still valid on 11.12.2001 (patent law change 1993) | ||
| OR01 | Other related matters | ||
| C19 | Lapse of patent right due to non-payment of the annual fee | ||
| CF01 | Termination of patent right due to non-payment of annual fee |