CA2111602A1 - Method for controlling the attitude of a satellite aimed towards a celestial object and a satellite suitable for implementing it - Google Patents
Method for controlling the attitude of a satellite aimed towards a celestial object and a satellite suitable for implementing itInfo
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- CA2111602A1 CA2111602A1 CA002111602A CA2111602A CA2111602A1 CA 2111602 A1 CA2111602 A1 CA 2111602A1 CA 002111602 A CA002111602 A CA 002111602A CA 2111602 A CA2111602 A CA 2111602A CA 2111602 A1 CA2111602 A1 CA 2111602A1
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- 238000012545 processing Methods 0.000 claims abstract description 13
- 238000005259 measurement Methods 0.000 claims description 30
- 235000019892 Stellar Nutrition 0.000 claims description 14
- 238000005286 illumination Methods 0.000 claims description 3
- 238000001514 detection method Methods 0.000 description 7
- 239000011159 matrix material Substances 0.000 description 7
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- 238000004364 calculation method Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 2
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- 241000283984 Rodentia Species 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 238000012937 correction Methods 0.000 description 1
- 125000004122 cyclic group Chemical group 0.000 description 1
- 238000013016 damping Methods 0.000 description 1
- 230000009977 dual effect Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000013642 negative control Substances 0.000 description 1
- 239000013641 positive control Substances 0.000 description 1
- 239000003380 propellant Substances 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
- 238000012163 sequencing technique Methods 0.000 description 1
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
- B64G1/245—Attitude control algorithms for spacecraft attitude control
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/363—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using sun sensors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/365—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using horizon or Earth sensors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/369—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using gyroscopes as attitude sensors
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- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Automation & Control Theory (AREA)
- Life Sciences & Earth Sciences (AREA)
- Environmental & Geological Engineering (AREA)
- General Life Sciences & Earth Sciences (AREA)
- Geochemistry & Mineralogy (AREA)
- Geology (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
- Navigation (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
METHOD FOR CONTROLLING THE ATTITUDE OF A SATELLITE
AIMED TOWARDS A CELESTIAL OBJECT AND A SATELLITE
SUITABLE FOR IMPLEMENTING IT
A method of controlling the attitude of a satellite (11) according to which the direction (S) of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector (.omega.) of the satellite is detected and, by means of an actuating assembly (16), torques are applied to the satellite which are defined by a control law as as to rotate the satellite about this direction whilst orienting an aiming axis (SR) related to the satellite in this direction, involves defining this direction of the predetermined celestial object in a frame of reference related to the satellite by a first quantity representing a first angle (.alpha.,.alpha.') measured between an axis of sight (Zs, Zs') and the projection of this direction onto a first reference plane containing this axis of sight and by a second quantity representing a second angle (.beta.') defined by this axis of sight and the projection of this direction onto a second reference plane containing this axis sight, this second angle being calculated from this first angle and from the instantaneous angular velocity vector of the satellite, whilst a satellite for implementing the method includes a body (11); a sensor (13a, 13b) having the axis of sight (Zs); the actuating assembly (16); an attitude control unit (20) for generating signals for applying the said torques to the satellite; a unit (15) for measuring the rotation velocity; and a processing unit (21) calculating the said second quantity.
METHOD FOR CONTROLLING THE ATTITUDE OF A SATELLITE
AIMED TOWARDS A CELESTIAL OBJECT AND A SATELLITE
SUITABLE FOR IMPLEMENTING IT
A method of controlling the attitude of a satellite (11) according to which the direction (S) of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector (.omega.) of the satellite is detected and, by means of an actuating assembly (16), torques are applied to the satellite which are defined by a control law as as to rotate the satellite about this direction whilst orienting an aiming axis (SR) related to the satellite in this direction, involves defining this direction of the predetermined celestial object in a frame of reference related to the satellite by a first quantity representing a first angle (.alpha.,.alpha.') measured between an axis of sight (Zs, Zs') and the projection of this direction onto a first reference plane containing this axis of sight and by a second quantity representing a second angle (.beta.') defined by this axis of sight and the projection of this direction onto a second reference plane containing this axis sight, this second angle being calculated from this first angle and from the instantaneous angular velocity vector of the satellite, whilst a satellite for implementing the method includes a body (11); a sensor (13a, 13b) having the axis of sight (Zs); the actuating assembly (16); an attitude control unit (20) for generating signals for applying the said torques to the satellite; a unit (15) for measuring the rotation velocity; and a processing unit (21) calculating the said second quantity.
Description
^"."~j ~l ~
.~.~j M~THOD FOR CONTROLLING THE ATTITUDE OF ~ E AIMED
TOWARDS A CELESTIAL OBJECT AND A SATELLI~'E SUITABLE FOR
IMPLEMENTING IT
BACKGROUND OF THE INVENTION
5 The present invention concerns a method for aiming a satellite with respect to a celestial or heavenly object (the sun or a star for example) and a device suitable for implementing it.
Satellite in this case means any artificial object moving in the solar system. This object may in particular be:
:
10 - in an orbit around the earth or any other planet in the solar system, - in an orbit around a satellite of any-planet in the solar system, - in a solar orbit, possibly a transfer orbit between two 15 planets.
As is known, satellites have, for the purpose of controlling their orbit and attitude, sensors and actuators connected to logic processor units within a system normally referred to as the attitude and orbit control system. The logic processor 20 units are often integrated within an on~board compu~er, the ackuators including ~or example thrusters (and/or magnetic coils, and/or gears3 whilst the energy consumed on board is provided by ~atteries, solar cells subjected to solar ~-~
radiation and/or propellants used for thrust propulsion in `
25 particular ~or orbit control.
As for the sensors, these are in practice of several types, depending on the nature, size, brightness, etc of the heavenly body (earth, sun or star) which they serve to detect. There are currently single-axis or dual-axis sensors.
30 Single-axis sensor means in general an appliance providing an angular coordinate of the directio~ of the heavenly body aimed at in a frame of re~fere.nce peculiar to the sensor. In practice an axis of sight and a sensing axis are defined ~or .
~....
- ..
` 2111602 this sensor, and this angular coordinate is for example the angle made by the axis of sight of the sensor to the projection of the direction of the heavenly body aimed at onto the plane containing the axis of sight perpendicular to the ~; 5 sensing axis~
A dual-axis sensor means here an appliance supplying two angular coordinates of the direction of the heavenly body aimed at in a frame of re~erence related to the sensor (which determines this direction completely). This appliance therefore has the function of two single-axis sensors with identical axes of sight and separate sensing axes, generally orthogonal. For detecting the sun, such a dual-axis sensor is in fact normally formed by associating two such single-axis sensors.
In fact a satellite may be caused to adopt several attitudes during its life after being released ~rom its launch vehicle.
Thus, fox example, the nominal attitude of a satellite of the triple-axis stabilised type moving on a circular terrestrial orbit consists of having a Z axis t known as the yaw axis, pointing towards the earth, a Y axis known as the pitch axis, perpendicular to the plane of the orbit, and an X axis known as the roll axis, perpendicular to the Z and Y axes and having -the same direction as the instantaneous linear velocity of the satellite on its orbit, the direction of the Y axis being such ~`~
that the frame of xeference (X, Y, Z) is positive. Such a nominal attitude is controlled by means of a terrestrial sensor, either alone or combined with solar sensors or a stellar sensor.
~ .
Other types of attitude may be envisaged, notably just before a satellite is put into its operational orbit, or after a serious failure of any type af~ecting in particular the attitude and orbit control system.
.
...
" 3 2111602 It is normal in such cases to attempt to put the satellite into a so-called sun-aimed attitude in which it is in slow rotation about an axis pointing towards the sun, chosen so as to be close to a principal inertia axis and such that the 5 satellite solar panels are illuminated. This enables the satellite to await subsequent commands, whilst ensuring its safety, that is to say:
- ensuring the recharging of the batteries with which the satellite is equipped, lO - powering the various items of equipment required by the satelliteo sensor, computer, heaters, remote control and telemetry, in particular, - ensuring illumination of the satellita by the sun so that ` the thermal configuration of the satellite is homogeneous and 15 maintains the equipment within the permitted temperature range.
.
Conventionally, this attitude control mode is referred to as the sun acquisition mode and is based on a sequencing of the ~ollowing type given by way o~ example:
~:
20 - seeking the sun by rotating the satellite about one axis ``~
~for example the roll axis), so that the field of view of at least one solar sensor encounters the sun, "..
- rotating the satellite about one axis (for example the pitch axis) so as to bring the direction o~ the sun towards the 25 desired direction, - rotating the satellite about the direction of the sun, and controlling this direction so that it becomes identical with the desired direction t~or example the roll axis).
The satellite thus being in a sun-aimed attitude, the ~ , r ` 4 ~ 6 0 2 requirements of the mission generally require it to come (or return) to its nominal attitude, that is in practice for it then to be aimed (or re-aimed) towards the earth (or a star).
The method of seeking the earth or star which is 5 conventionally used, in such an sarth acquisition mode or star acquisition mode, consists, during the seeking stage, of putting the satellite into slow rotation about an axis aimed towards the sun, this axis being chosen so that the field of view of the earth (or star) sensor necessarily encounters the 10 éarth (or star).
:
In some cases, when it is not possible to seek the earth directly, a reference star is initially sought, the Pole Star ~or example: once the star has been found as described above, the satellite is, from the measurements of the star sensor, 15 controlled with respect to rotation so as to bring the field of view of the earth sensor facing the earth: this is then referred to as acquisition of the earth via the star.
The rotation axis pointed towards the sun during the seeking of the earth or star forms, with the axis of sight of the 20 earth or star sensor, an angle at least approximately equal to the sun-satellite-earth angl~ or sun-satellite-star angle.
This axis may be khe direction which the sun should have in ~`
the frame of reference related to the satellite once the `
latter is in its attitude pointed towards the earth or star. ;~
. ~
25 These different attitude control modes therefore require the direction of the sun or star in question to be determined for the purpose of taking it into account in the control loops using the control law corresponding to the current stage of the attitude control mode: in particular in the last phase of 30 the sun acquisition mode, in the first phase of the earth or star acquisition modes and in the last phase of the acquisition of the earth via a star.
Amongst the documents which relate to such a change in nominal :: 5 2 1 ~ 16 ~ 2 attitude, a~ter an intermediate change into a sun-aimed attitude, can be cited the patents FR-2.407.860 (MESSERSCHMITT-BOLKOW-BLOHM), EP-0.338.687 (BRITISH
AEROSPACE), US-5.080.307 (HUGHES AIRCRAFT) and FR-2.649.809 (MESSERSCHMITT-BOLKOW-BLOHM).
In the first three patents, the direction of the sun is determined from the measurement of two single-axis solar sensors whose sensing axes are perpendicular, or from a twin-axis solar sensor.
:: :
Likewise, in the patent EP-0~338.687, the direction o~ the star is computed from the measurement of two single-axis stellar sensors whose sensing axes are perpendicular, or from a twin-axis stellar sensor.
The computation is carried out by determining the three coordinates o~ the unit vector o~ the instantaneous direction of the sun or star in the frame of reference of the sensor ~ ;
from two angular measurements defining the orientation of this direction in the frame of re~erence. `
.
Control is effected by applyin~ demands of the type~
.
U=-Kd*~-C*S~c]-Kp*~S~c A SR]
where:
U = demand signal to be applied as determined by the ~
control law ::
S~c = unit vector of the instantaneous direction of the ~:
heavenly body in question (sun or star) calculated from the angular measurements oE the sensors SR = unit vector of the reference direction of the heavenly body in question ~ A ~ V ~ :
`~:
~`` 6 21~1~02 ~ ~ = measured velocity vector : :
C = velocity of rotation demanded about the direction SR
Kd - velocity regulation gain 5 Xp = position regulation gain :~
`~ A = vector product (or, in English, cross product).
; During the sun, earth or star acquisition modes, the velocity o~ the satellite about its three axes is measured (by means o~
gyrometers, for example) and used in the attitude control ~
laws, solely ~or the purpose of damping the positional control ~:
or carrying out the ~elocity control.
~ :
An earth-seeking mode using a single solar sensor is described in the patent FR-2.649.809. This patent proposes a method in which the direction of the sun is not completely determined.
The control law is based on the ~act that the rotation of the satellite causes coupling between the errors in attitude according to the measurement axis and according to the non-measurement axis. Thus, when the sin~le-axis solar sensor detects an error, the control law generates a command suitable ~or cancelling out this error according to the measurement axis, whilst the non-measurable error on the other axis is eli~inated by coupling.
: ' ':
The control law used is o~ the type:
, U--Kd*[~-C*S~]~Kp*[SRAeM+(S~*sRT)eM-eM]*L(N,r-Nby) : . .
where:
U = demand to be applied SR = unit vector o~ the referenc~ direction of the sun ~-"`; . 7 2 ~ 0 2 SRT = transposed unit vector of the re~erence direction of the sun ~ = measured velocity vector ¦ C = rotation velocity demanded 5 Kd -- velocity regulation gain ~ .
: Kp - position regulation gain eM = measurement axis vector N~ = solar sensor measurement ~: Nby = correction of the solar sensor measurement :
10 ~ = identity matrix : :
;~ L = limitation ~actor A = vector product tor, in English, cross product). :-:~
;~It will be noted that this command includes a term relating to velocity control according to three axes and a term relating to positional control according to two axes, namely the measurement axis a~d an axis perpendicular to this measurement axis and to the reference direction of the sun, referred to as ~
the non-measurable component axis. :~:
This method has the following drawbacks:
- complexity o~ the control laws, which are very differen~
from the conventional laws of the earth and sun acquisition .
modes, . ~
. - poor acauracy of aiming about the non-measurable component .. :
axis since this axis is controlled passi~ely by coupling and the error in attitude about this axis is never measured or . determined, .
~ difficulty in using this control law to generate thruster commands on the aforesaid two axes since that requires:
. either a complicated logic for generating impulses of different durations on the different thrusters: the impulse modulator conventionally used on known satellites !. ,.
8 21~ 02 is then not applicable, :~
-~ . or using thrusters ha~ing specific orientations suitable for producing torques on both axes, which de-optimises the system of attitude control of the satellite by thrusters since the directions of these thrusters are then imposed by the datum of these two axes.
An earth-seeking mode using a single solar sensor based on coupling between axes due to the rotation is also described in the document: "The attitude and orbit control of the EUTELSAT
II spacècraft" 6.11 page 95, Symposium on Automatic Control in Space - IFAC - 17-21/7/1989. This document proposes an earth-seeking mode in which the reference direction of the rotation vector of the satellite during the earth seeking is chosen so that it forms, to the direction of the sun, an angle equal to the earth-satellite-sun angle and so that its component which cannot be measured is zero. This article deals with the pitch component, which amounts to saying that the sun is maintained in the XZ plane of the satellite throughout the earth-seeking phase and that the ~inal attitude at th~ time o~ sensing the earth has a yaw angular di~ference.
During the earth seeking, only veloci~y control about the direction SR is suggested, so that the control law used by this method is probably of the type:
U = -Kd*[~-C*SR]
25 where:
.~
~, U - demand to be applied SR = unit vector o~ the reference direction of the sun = velocity vector measured ~ - velocity o~ rotation about SR
; 30 Kd = velocity regulation gain , .
This method has the following drawbacks:
~, ~`
:: `::`
9 2 ~ 0 2 - the necessity for two additional stages: pitch rotation to bring the sun facing the sensor used, and then yaw rotation to cancel out the yaw error due to the particular choice of the reference direction of the sun, - the poor accuracy of the aiming, which deteriorates over time because of drift in the gyrometers and because the ~.
control law does not includa the positional control term.
This poor aiming accuracy: :
, : . prevents itsapplication to the sun acquisition mode, the long duration of which (typically several hours) would, with this type of law, result in latent aiming errors, ::
. prevents its application to the star acquisition method, which requires great aiming accuracy because of the need .
to recognise the star, I5 . the risk, in some cases, of resulting in ~ailures of the earth acquisition.
on the other hand, the present invention relates to a method for aiming the satelli~e..-towards a heavenly body such as the sun or a star:
- which is applicable to any triple-axis stabilised satellite whatever the arrangement of its thrusters (unlike the patent FR-2.649.809), - which does not call into question the logics conventionally used for the sun, earth or star acquisition modes (unlike the patent FR-2.649.809), - which uses the measurement of a single solar or stellar sensor with a single measurement axis (unlike the patents FR-2.407.860, EP-0.338.687 and US-5.080.307) but in combination with other available measurements, ~ ' .
` lo 2111~2 - which can easily be programmed in the on-board computer, - the aiming accuracy of which is optimum because of a positional control carried out aotively and continuously, by comparing the calculated direction of the heavenly body with the direction aimed at or reference direction.
SUMMARY OF THE INVENTION
To this end the invention proposes a method of controlling the attitude of a satellite according to which the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector of ~he satellite is detected and, by means of an actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate ~he satellite about this direction whilst orienting an aiming axis related to the satellite in this direction, characterised in that this direction of the predetermined celestial object is defined in the frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of this direction onto a first reference plane containing this axis o~ sight and by a second quantity representing a second angle defined by this axis of sight and the projection o~ this direction onto a second reference plane containing this axis of sight, this second angle being calculated from this first angle and ~rom the instantaneous angular velocity vector of the satellite.
Compared with the conventional methods known at the present time, this method maXes it possible to save on one axis for measuring the heavenly body aimed at ~namely the use of a single-axis sensor instead of a twin-axis sensor or saving on one single-axis sensor), its redundancy and, possibly, sensor control electronics, and therefore to reduce the weight and cost of the satellite attitude control system; all this without changing the structure of the sun, earth or star ~:~
aoquisition modes.
:
.. ... ..
``` 11 21~02 This method can al~o be appli~d to a satellite for dealing with a failure of some of its sensors. In such case, the satellite may already be in orbit.
The calculation method is based on the relationship between 5 the direction of the heavenly body aimed at and its change in a frame of reference related to the satellite, the single ; measurement of the sensor and its change and the angular velocity measurements.
According to preferred characteristics of this method:
~: :
- this axis of sight belongs to a sensor with at least one sensing axis, the first reference plane being defined as being perpendicular to the sensing axis, - the second reference plane is perpendicular to the first reference plane and is defined as containing the axis of sight and sensinq axi5, - the sensor is a single-axis sensor, - the sensor is a twin-axis sensor, a single output of which is used, - the first and second quantities representing the first and second angles are the tangents of these angles, - the control law is of the type:
U=-Kd*~-c*sslc]-Kp*~sslc ~ SR]
' where: ~
_ '.:
U = demand to be applied to the torque generator Ss~c = unit vector of the instantaneous direction of the celestial object SR = unit vector of the aiming axis forming the rotation reference axis . '' ~'~
;' `;``` 12 ~1116~2 = measured velocity vector C = rotation velocity demanded about SR
Kd = velocity regulation gain Kp = position regulation gain 5 A = vector product (or, in English, cross product) - the celestial object is the sun, the aiming axis is an axis which is at least approximately close to an inertia axis of the satellite, chosen so as to obtain continuous illumination of a solar generator installed on the satellite, by virtue of which the satellite is in a sun-aimed mode, ~-- the satellite has another sensor with a second axis of sight, suitable for detecting another predetermined celestial object, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the (sun) - satellite - ~or other celestial object) an~le, this other celestial object is a star, by ~irtue of which the satellite is in star acquisition mode, ` - this other celestial o~ject is the earth, by virtue of which the satellite is in earth acquisition mod~, - the celestial object i5 a star, ~
:~. ,,' - the satellite also has a terrestrial sensor with a second axis of sight, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the star-satellite-earth angle, hy virtue of which the satellite is in the mode for the acquisition of the earth from a star.
The invention also proposes a satellite having a body, a sensor with an axis of sight suitable for detecting a ;
,~
f~ .
13 2 i l~ 6 ~ 2 predetermined celestial object and supplying a first quantity representing a first angle measured between the axis of sight and the projection of the instantaneous direction o~ the celestial object onto a ~irst reference plane containing this axis of sight, an actuating unit, an at~itude control unit suitable for generating, from this Lirst quantity and a second quantity representing an angle defined by the axis of sight and the projection of this instantaneous direction of the zelestial object onto a second reference plane containing this axis of sight, separate from the first reference plane, signals suitable for applying to the satellite, through the .
actuating unit, torques suitable for rotating the satellite about this direction and orienting an aiming axis in this direction, and a unit for measuring the instantaneous rotation velocity of the satellite, characterised in that this second representative quantity is applied to the attitude control unit by a preliminary processing unit suitable ~or calculating this second quantity from the first quantity and from the :
output signal of the instantaneous rotation velocity measuremen~ unit.
According to pre~erred characteristics of this satellite:
- this sensor is a sensor with a single sensing axis, the ~irst reference plane being perpendicular to this sensing axis, and the second réferenc~ plane containing this sensing axis, - the sensor is a solar sensor, - the satellite also has a stellar sensor, - the satellite also has a terrestrial sensor, - the sensor is a stellar sensor, :~
- the actuating unit inclucles thrusters, ~ `
~`" ` 21~602 - the uni~ for measuring the instantaneous rotation velocity of the satellite includes gyrometers.
BRIEF DESCRIPTION OF THE DRAWINGS
Objects, characteristics and advantages of the invention will emerge from the following description, given by way of non-limitative example, with reference to the accompanying drawings in which:
:~
- Figure 1 is a diagram of a single-axis sensor, - Figure 2 is a diagrammatic view in perspective of a satellite according to the in~ention, :
- Figure 3 is a functional diagram of an aiming device according to the invention, - Figure 4 is a perspective view of the satellite of Figure 2 in the earth seeking phase, and - Figure 5 is a diagrammatic view of this satellite after sensinq of the earth.
. DESCRIPTION OF THE PREFERRED EMBODIMENT :.
A single-axis sensor as shown diagrammatically under the reference numeral 1 in Figure 1 is in practice a slot extending along an axis Ys referred to as the sensing axis, perpendicular to an axis Xs referred to as the non-measurement axis, whilst the axis Zs perpendicular to the plane XsYs of ~-the sensor is referred to as the optical axis or axis of -sight, intended to be oriented sufficiently close to the celestial object to be detected for the latter to come within the field of view of the sensor.
..
What has just been stated applies particularly where the sensor is a solar sensor. A stellar sensor generally consists of a charge transfer detector matrix (CCD matrix) onto which an image of the star is projected. The measurement of the sensor in fact consists of two coordinates of the pixel or ::
~., ; 15 2111~02 pixels illuminated by the star. This stellar sensor, generally with dual axes, is able to supply only a single-axis indication if a measurement processing fault results in the ¦ provision of a single coordinate instead o~ two. In such 5 case, the invention can be applied.
If S is the direction of this celestial object, the sensor 1 provides a measurement signal representing the angle a between the axis of sight and the projection of this direction in a reference plane passing through the axis of sight and lO perpendicular to the sensing axis.
It is important to note that the angle B between the axis of sight and the projection of S in a second reference plane defined by the axis of sight and the sensing axis is on the other hand not measured.
The atti~ude drifts which the satellite may undergo about the axes Xs, Ys and Zs which, for the sensor, are roll, pitch and yaw axes, are designated by ~, ~ and ~.
Figure 2 shows diagrammatically the body 11 of a triple-axis stabilised satellite according to the invention, including in 2b addition, in a con~entional manner, a solar generator with one or more panels (not shown).
In the case considered here of a given earth orbit, preferably ~ -geos~nchronous, of low inclination (typically less than 10) the three axes along which it is sought to stabilise this ;~
satellite ~that is to say its body) in nominal attitude are, respectively, an axis directed towards the earth, usually termed Z and refexred to as the yaw axis, an axis perpendicular to the plane of the orbit and directed towards the south, termed Y and referred to as the pitch axis, and an X axis forming, with the pitch and yaw axes, a positive orthonormal axis system (X, Y, Z), and referred to as the roll axis. In practice, when the orbit is circular, this roll axis . `
` é ~ :
.,.,.",...............
:`` 16 2 ~ 0 2 is tangent to the orbit and preferably has the same direction as the velocity at which the satellite moves on the orbit.
Generally the solar generator extends parallel to the pitch axis.
5 The body 11 of the satellite has an attitude and orbit control system comprising:
::~
: - a terrestrial deteation system 12, here formed by a single twin-axis detector, having an optical axis Zst parallel or ; close to the Z axis and two measurement axes (or sensing axes) Xst and Yst (in practice orthogonal) transverse to the optical axis, advantageously close or even parallel to the X and Y
axes of the satellite respectively (the field of view is for example + 14 around the axis Ys and + 5 around the axis Xs), - the solar detection system including a plurality of singla-axis solar.sensors having axes of sight of different orientations: two sensors 13a and 13b are shown here, with axes o~ sight Zs and Zs' respectively oriented in the plane XZ
at least approximately a~ 45 from the axes -X and -Z on the .:
one hand and X and -~ on the other hand, and having sensing :
axes at least approximately parallel to Y, . ' '.' .
- an optional stellar detection system 1~, here formed by a twin-axis detector, having an optical axis Zsp pointed towards the north so as to be able to detect the Pole Star (as a :
reminder, the Pole Star is very close to the north) and two ~ .
measurement axes Xsp and Ysp parallel to X and Z; this stellar :~
' detection system 14 is in practice offset from the Y axis by a .
sufficient distance to prevent the solar generator encroaching -~
appreciably into its field o view;
- an ang~lar velocity detection unit 15, for example gyrometers, for measuring the angular velocities of the satellite about three axis preferably parallel to the X, Y and . : .
::
~ ~ ~,. ~.,., ,.~`,''`,', . .~::,~,~: ~ ,j ::,' -~
~: 17 ~ 6 ~ 2 Z axes, - a~ actuating unit, in this case formed by thrusters 16 at least four in number (in this case six), for generating positive or negative control torques about the X, Y and Z
: 5 axes, and - an analogue or digital processing unit 17, for processing the measurements supplied by the detection systems and, by means of control laws which are per se conventional, in the :; nominal attitude control reg:ime or in reacquisition mode, ~: 10 co~mands intendad ~or the actuating unit (through filters, limiters and modulators).
: ':
The process of the invention i9 described below with regard to the functicnal diagram in Figure 3, involving any one of the ~
singl~-axis solar sensors (chosen according to the desired .. :
direction o~ rotation) and the angular velocity detection unit 15. In the example in question, aiming is effected by means of the solar sensor 13a.
The various quantities which will be used in the iterative calculation defined below, have the followin~ definitions: .
. '.~
.20 sx~cl S~c Sz~c : components of the unit vector of the .~.
direction of the heavenly body in satellite axes, SX, SY, SZ : components of the unit vector of the :;-direction of the heavenly body in sensor :~
' axes at the previous moment, SXI, SY', SZ' : components of the unit vector o~ the direction of the heavenly body in sensor axes at the present moment, 30 ~ : angle between the axis Zs and the projection of the direction of the heavenly body in the plane Xs, Zs in the ' :
.
: .`` 18 2111602 .sensor frame of re~erence, B : angle between the axis Zs and the projection of the direction of the heavenly body in the plane Yszs in the ~: 5 s~nsor ~rame of reference : tan~' = SX'/SZ 7 current measurement of ~ by the sensor : tan~ = SX/S~ : previous measurement of ~ by the sensor used tanB' = SY'/SZ' : current estimation of B
10 (tan*')f current filtered value of tanB
tanB) f : previous filtered value of tanB ~::
: roll, pitch and yaw micro-rotations : between the sensor frames of reference at the previous and current moments ::.
15 Dt : duration of the calculation cycle ~xs= Dt ' ~Yg Dt ' ~5 D~
roll, pitch and yaw velocity in sensor ~
axes ~;
~c' ~y~o' .... ~, 20 ~z~c : roll, pitch and yaw velocities measured by ~
satellite axis :.
M : matrix for converting between the ~.
satellite axes and sensor axes .
: inverse matrix of the matrix M
It should first of all be noted that there is a change from the heavenly body according to sensor axes, between the -~
previous and current moments, by means of the matrix equation~
~sxl f ~ 1 fs~
sY = ~ sY' .:
sz ~ ~ 1 s~', .~, .
30 The variation ~ in the tangent of the measured anqle ~ is : ;~
equal to: ~ - tana' - tan~, that is to say:
I
~5}~
: `
19 2 ~ 0 2 ~=tan~ tan~=~SX/ SX= SX~ S~-~*Sy/+~*
SZI SZ SZI SZ~*S~-~*SX~
or again:
*tan~*tan@/_- tana~*tan~*~ - tan~
1-tan~*~-tana/*~
5 The est.imation of the non-measured angular component B' of the direction of the heavenly body in the sensor frame of reference is effected:
by cal~ulating the instantaneous velocities of rotation in ;~
the sensor frame of reference according to the change of frame 10 of reference formula:
~xsl ~)x~c~
~i)yS = M I ~ys/C : ~ ~
,~)Z5 l~ZS/C, . ` . '^
- whence tanB' is calculated by the equation:
~yS~tan*tanal*~yS*
~z9+~xg*~an~
' ~`~'`' ~ `' ' Advantageously a filtering of tanB' is carried out in order to eliminate the effect of the cyclic determination of this tangent. This filterinq is for example a first order 20 filtering of the type~
~tanB')r - a(tanB)f ~ (l-a)tanB' ~`
, where the subscript f corresponds to the filtered values and where a is a constant. ~-Finally, there is a progression, in a ~nown manner, from these 25 tangents to the components of the unit vector of the direction of the heavenly body in the sensor frame of reference according to the e~uations:
~ "~ ' .ul t~
~ . 20 2111~2 I_ tan~l SX-- - -~l+tan2a,~ 2 p j .
_ tan~ :
~/l+tan2~tan2~/
~ ~ 5Z/~
tan2a/+tall,2~1 ~:; and then by change o~ ~rame of reference these components are 5 determined in the satellite frame of re~erence: :~
~ = M ISYI
~ sz . ~
The attitude control proper is then obtained by means of a ::
lO conventional law, *or example the one already presented above : in the prea~ble:
U=-~d*~-c*s~c~-Kp*~s~c A SR]
where SR has the same meàn~ng as before.
In Figure 3 can thus be seen the sensor whose field of view ~
15 contains the satellite-heavenly body direction (in this case ~ :
the single-axis sensor 13a), the angular velocity detection.
unit 15 Pormied by gyromieters, the processing unit 17 and the actuating means 16.
The processing unit 17 includes a control unit 20, ~:~
20 conventional in itself, suitable for supplying attitude control signals U fromi two quantities characteristic of the angular orientation of the satellite-heavenly body direction : :
(in this case the tangents of the instantaneous angles ~' and B').
25 According to the invention one of these quantities is derived, not directly and solely from the output of an angular position ,~,,",,.,.,, .. ,.,.,.,,.,.. ,.,.,,.. .. , .,, ,. ~
;~ 21 21 11~2 sensor, but from the instantaneous angular velocity of the satellite in the frame of reference (X, Y, Z) related to it.
The processing unit 17 includes for this purpose a preliminary processing unit 21 suitable for calculating as indicated abova, from the single quantity tan~ supplied by the sensor 13a and the instantaneous angular velocity ~, the other quantity tan~ required by the control unit.
The satellite 11 in earth seeking phase can be seen in Figure 4. The satellite 11 is rotating about the axis SR pointed ;~
towards the sun. This rotation brings the axis of sight of ; the terrestrial sensor (which is identical with the yaw axis of the satellite~ facing the earth, thus enabling the latter to be detected. This situation is shown diagrammatically in ~-Figure 5.
It goes without saying that the above description has been put forward only by way of non-limitative example and that many variants could be proposed by a person skilled in the art without departing from the scope of the invention. Thus, for example, the stellar sansor could be omitted: a single-axis solar sensor with an axis of sight approximately directed along Z could then be added in order to supplement the field in which it is possible, at any time, to search for the sun. ~`
Likewise the process of the invention can be applied without any problem:
- to the case of a satellite provided with at least one twin-axis solar or stellar sensor formed by two coupled single-axis sensors, one o~ which happened to fail, - to the case of a satellite provided with at least one twin-axis stellar sensor, the measurements of which along one of the axes is unavailable because of a failure.
.~.~j M~THOD FOR CONTROLLING THE ATTITUDE OF ~ E AIMED
TOWARDS A CELESTIAL OBJECT AND A SATELLI~'E SUITABLE FOR
IMPLEMENTING IT
BACKGROUND OF THE INVENTION
5 The present invention concerns a method for aiming a satellite with respect to a celestial or heavenly object (the sun or a star for example) and a device suitable for implementing it.
Satellite in this case means any artificial object moving in the solar system. This object may in particular be:
:
10 - in an orbit around the earth or any other planet in the solar system, - in an orbit around a satellite of any-planet in the solar system, - in a solar orbit, possibly a transfer orbit between two 15 planets.
As is known, satellites have, for the purpose of controlling their orbit and attitude, sensors and actuators connected to logic processor units within a system normally referred to as the attitude and orbit control system. The logic processor 20 units are often integrated within an on~board compu~er, the ackuators including ~or example thrusters (and/or magnetic coils, and/or gears3 whilst the energy consumed on board is provided by ~atteries, solar cells subjected to solar ~-~
radiation and/or propellants used for thrust propulsion in `
25 particular ~or orbit control.
As for the sensors, these are in practice of several types, depending on the nature, size, brightness, etc of the heavenly body (earth, sun or star) which they serve to detect. There are currently single-axis or dual-axis sensors.
30 Single-axis sensor means in general an appliance providing an angular coordinate of the directio~ of the heavenly body aimed at in a frame of re~fere.nce peculiar to the sensor. In practice an axis of sight and a sensing axis are defined ~or .
~....
- ..
` 2111602 this sensor, and this angular coordinate is for example the angle made by the axis of sight of the sensor to the projection of the direction of the heavenly body aimed at onto the plane containing the axis of sight perpendicular to the ~; 5 sensing axis~
A dual-axis sensor means here an appliance supplying two angular coordinates of the direction of the heavenly body aimed at in a frame of re~erence related to the sensor (which determines this direction completely). This appliance therefore has the function of two single-axis sensors with identical axes of sight and separate sensing axes, generally orthogonal. For detecting the sun, such a dual-axis sensor is in fact normally formed by associating two such single-axis sensors.
In fact a satellite may be caused to adopt several attitudes during its life after being released ~rom its launch vehicle.
Thus, fox example, the nominal attitude of a satellite of the triple-axis stabilised type moving on a circular terrestrial orbit consists of having a Z axis t known as the yaw axis, pointing towards the earth, a Y axis known as the pitch axis, perpendicular to the plane of the orbit, and an X axis known as the roll axis, perpendicular to the Z and Y axes and having -the same direction as the instantaneous linear velocity of the satellite on its orbit, the direction of the Y axis being such ~`~
that the frame of xeference (X, Y, Z) is positive. Such a nominal attitude is controlled by means of a terrestrial sensor, either alone or combined with solar sensors or a stellar sensor.
~ .
Other types of attitude may be envisaged, notably just before a satellite is put into its operational orbit, or after a serious failure of any type af~ecting in particular the attitude and orbit control system.
.
...
" 3 2111602 It is normal in such cases to attempt to put the satellite into a so-called sun-aimed attitude in which it is in slow rotation about an axis pointing towards the sun, chosen so as to be close to a principal inertia axis and such that the 5 satellite solar panels are illuminated. This enables the satellite to await subsequent commands, whilst ensuring its safety, that is to say:
- ensuring the recharging of the batteries with which the satellite is equipped, lO - powering the various items of equipment required by the satelliteo sensor, computer, heaters, remote control and telemetry, in particular, - ensuring illumination of the satellita by the sun so that ` the thermal configuration of the satellite is homogeneous and 15 maintains the equipment within the permitted temperature range.
.
Conventionally, this attitude control mode is referred to as the sun acquisition mode and is based on a sequencing of the ~ollowing type given by way o~ example:
~:
20 - seeking the sun by rotating the satellite about one axis ``~
~for example the roll axis), so that the field of view of at least one solar sensor encounters the sun, "..
- rotating the satellite about one axis (for example the pitch axis) so as to bring the direction o~ the sun towards the 25 desired direction, - rotating the satellite about the direction of the sun, and controlling this direction so that it becomes identical with the desired direction t~or example the roll axis).
The satellite thus being in a sun-aimed attitude, the ~ , r ` 4 ~ 6 0 2 requirements of the mission generally require it to come (or return) to its nominal attitude, that is in practice for it then to be aimed (or re-aimed) towards the earth (or a star).
The method of seeking the earth or star which is 5 conventionally used, in such an sarth acquisition mode or star acquisition mode, consists, during the seeking stage, of putting the satellite into slow rotation about an axis aimed towards the sun, this axis being chosen so that the field of view of the earth (or star) sensor necessarily encounters the 10 éarth (or star).
:
In some cases, when it is not possible to seek the earth directly, a reference star is initially sought, the Pole Star ~or example: once the star has been found as described above, the satellite is, from the measurements of the star sensor, 15 controlled with respect to rotation so as to bring the field of view of the earth sensor facing the earth: this is then referred to as acquisition of the earth via the star.
The rotation axis pointed towards the sun during the seeking of the earth or star forms, with the axis of sight of the 20 earth or star sensor, an angle at least approximately equal to the sun-satellite-earth angl~ or sun-satellite-star angle.
This axis may be khe direction which the sun should have in ~`
the frame of reference related to the satellite once the `
latter is in its attitude pointed towards the earth or star. ;~
. ~
25 These different attitude control modes therefore require the direction of the sun or star in question to be determined for the purpose of taking it into account in the control loops using the control law corresponding to the current stage of the attitude control mode: in particular in the last phase of 30 the sun acquisition mode, in the first phase of the earth or star acquisition modes and in the last phase of the acquisition of the earth via a star.
Amongst the documents which relate to such a change in nominal :: 5 2 1 ~ 16 ~ 2 attitude, a~ter an intermediate change into a sun-aimed attitude, can be cited the patents FR-2.407.860 (MESSERSCHMITT-BOLKOW-BLOHM), EP-0.338.687 (BRITISH
AEROSPACE), US-5.080.307 (HUGHES AIRCRAFT) and FR-2.649.809 (MESSERSCHMITT-BOLKOW-BLOHM).
In the first three patents, the direction of the sun is determined from the measurement of two single-axis solar sensors whose sensing axes are perpendicular, or from a twin-axis solar sensor.
:: :
Likewise, in the patent EP-0~338.687, the direction o~ the star is computed from the measurement of two single-axis stellar sensors whose sensing axes are perpendicular, or from a twin-axis stellar sensor.
The computation is carried out by determining the three coordinates o~ the unit vector o~ the instantaneous direction of the sun or star in the frame of reference of the sensor ~ ;
from two angular measurements defining the orientation of this direction in the frame of re~erence. `
.
Control is effected by applyin~ demands of the type~
.
U=-Kd*~-C*S~c]-Kp*~S~c A SR]
where:
U = demand signal to be applied as determined by the ~
control law ::
S~c = unit vector of the instantaneous direction of the ~:
heavenly body in question (sun or star) calculated from the angular measurements oE the sensors SR = unit vector of the reference direction of the heavenly body in question ~ A ~ V ~ :
`~:
~`` 6 21~1~02 ~ ~ = measured velocity vector : :
C = velocity of rotation demanded about the direction SR
Kd - velocity regulation gain 5 Xp = position regulation gain :~
`~ A = vector product (or, in English, cross product).
; During the sun, earth or star acquisition modes, the velocity o~ the satellite about its three axes is measured (by means o~
gyrometers, for example) and used in the attitude control ~
laws, solely ~or the purpose of damping the positional control ~:
or carrying out the ~elocity control.
~ :
An earth-seeking mode using a single solar sensor is described in the patent FR-2.649.809. This patent proposes a method in which the direction of the sun is not completely determined.
The control law is based on the ~act that the rotation of the satellite causes coupling between the errors in attitude according to the measurement axis and according to the non-measurement axis. Thus, when the sin~le-axis solar sensor detects an error, the control law generates a command suitable ~or cancelling out this error according to the measurement axis, whilst the non-measurable error on the other axis is eli~inated by coupling.
: ' ':
The control law used is o~ the type:
, U--Kd*[~-C*S~]~Kp*[SRAeM+(S~*sRT)eM-eM]*L(N,r-Nby) : . .
where:
U = demand to be applied SR = unit vector o~ the referenc~ direction of the sun ~-"`; . 7 2 ~ 0 2 SRT = transposed unit vector of the re~erence direction of the sun ~ = measured velocity vector ¦ C = rotation velocity demanded 5 Kd -- velocity regulation gain ~ .
: Kp - position regulation gain eM = measurement axis vector N~ = solar sensor measurement ~: Nby = correction of the solar sensor measurement :
10 ~ = identity matrix : :
;~ L = limitation ~actor A = vector product tor, in English, cross product). :-:~
;~It will be noted that this command includes a term relating to velocity control according to three axes and a term relating to positional control according to two axes, namely the measurement axis a~d an axis perpendicular to this measurement axis and to the reference direction of the sun, referred to as ~
the non-measurable component axis. :~:
This method has the following drawbacks:
- complexity o~ the control laws, which are very differen~
from the conventional laws of the earth and sun acquisition .
modes, . ~
. - poor acauracy of aiming about the non-measurable component .. :
axis since this axis is controlled passi~ely by coupling and the error in attitude about this axis is never measured or . determined, .
~ difficulty in using this control law to generate thruster commands on the aforesaid two axes since that requires:
. either a complicated logic for generating impulses of different durations on the different thrusters: the impulse modulator conventionally used on known satellites !. ,.
8 21~ 02 is then not applicable, :~
-~ . or using thrusters ha~ing specific orientations suitable for producing torques on both axes, which de-optimises the system of attitude control of the satellite by thrusters since the directions of these thrusters are then imposed by the datum of these two axes.
An earth-seeking mode using a single solar sensor based on coupling between axes due to the rotation is also described in the document: "The attitude and orbit control of the EUTELSAT
II spacècraft" 6.11 page 95, Symposium on Automatic Control in Space - IFAC - 17-21/7/1989. This document proposes an earth-seeking mode in which the reference direction of the rotation vector of the satellite during the earth seeking is chosen so that it forms, to the direction of the sun, an angle equal to the earth-satellite-sun angle and so that its component which cannot be measured is zero. This article deals with the pitch component, which amounts to saying that the sun is maintained in the XZ plane of the satellite throughout the earth-seeking phase and that the ~inal attitude at th~ time o~ sensing the earth has a yaw angular di~ference.
During the earth seeking, only veloci~y control about the direction SR is suggested, so that the control law used by this method is probably of the type:
U = -Kd*[~-C*SR]
25 where:
.~
~, U - demand to be applied SR = unit vector o~ the reference direction of the sun = velocity vector measured ~ - velocity o~ rotation about SR
; 30 Kd = velocity regulation gain , .
This method has the following drawbacks:
~, ~`
:: `::`
9 2 ~ 0 2 - the necessity for two additional stages: pitch rotation to bring the sun facing the sensor used, and then yaw rotation to cancel out the yaw error due to the particular choice of the reference direction of the sun, - the poor accuracy of the aiming, which deteriorates over time because of drift in the gyrometers and because the ~.
control law does not includa the positional control term.
This poor aiming accuracy: :
, : . prevents itsapplication to the sun acquisition mode, the long duration of which (typically several hours) would, with this type of law, result in latent aiming errors, ::
. prevents its application to the star acquisition method, which requires great aiming accuracy because of the need .
to recognise the star, I5 . the risk, in some cases, of resulting in ~ailures of the earth acquisition.
on the other hand, the present invention relates to a method for aiming the satelli~e..-towards a heavenly body such as the sun or a star:
- which is applicable to any triple-axis stabilised satellite whatever the arrangement of its thrusters (unlike the patent FR-2.649.809), - which does not call into question the logics conventionally used for the sun, earth or star acquisition modes (unlike the patent FR-2.649.809), - which uses the measurement of a single solar or stellar sensor with a single measurement axis (unlike the patents FR-2.407.860, EP-0.338.687 and US-5.080.307) but in combination with other available measurements, ~ ' .
` lo 2111~2 - which can easily be programmed in the on-board computer, - the aiming accuracy of which is optimum because of a positional control carried out aotively and continuously, by comparing the calculated direction of the heavenly body with the direction aimed at or reference direction.
SUMMARY OF THE INVENTION
To this end the invention proposes a method of controlling the attitude of a satellite according to which the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector of ~he satellite is detected and, by means of an actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate ~he satellite about this direction whilst orienting an aiming axis related to the satellite in this direction, characterised in that this direction of the predetermined celestial object is defined in the frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of this direction onto a first reference plane containing this axis o~ sight and by a second quantity representing a second angle defined by this axis of sight and the projection o~ this direction onto a second reference plane containing this axis of sight, this second angle being calculated from this first angle and ~rom the instantaneous angular velocity vector of the satellite.
Compared with the conventional methods known at the present time, this method maXes it possible to save on one axis for measuring the heavenly body aimed at ~namely the use of a single-axis sensor instead of a twin-axis sensor or saving on one single-axis sensor), its redundancy and, possibly, sensor control electronics, and therefore to reduce the weight and cost of the satellite attitude control system; all this without changing the structure of the sun, earth or star ~:~
aoquisition modes.
:
.. ... ..
``` 11 21~02 This method can al~o be appli~d to a satellite for dealing with a failure of some of its sensors. In such case, the satellite may already be in orbit.
The calculation method is based on the relationship between 5 the direction of the heavenly body aimed at and its change in a frame of reference related to the satellite, the single ; measurement of the sensor and its change and the angular velocity measurements.
According to preferred characteristics of this method:
~: :
- this axis of sight belongs to a sensor with at least one sensing axis, the first reference plane being defined as being perpendicular to the sensing axis, - the second reference plane is perpendicular to the first reference plane and is defined as containing the axis of sight and sensinq axi5, - the sensor is a single-axis sensor, - the sensor is a twin-axis sensor, a single output of which is used, - the first and second quantities representing the first and second angles are the tangents of these angles, - the control law is of the type:
U=-Kd*~-c*sslc]-Kp*~sslc ~ SR]
' where: ~
_ '.:
U = demand to be applied to the torque generator Ss~c = unit vector of the instantaneous direction of the celestial object SR = unit vector of the aiming axis forming the rotation reference axis . '' ~'~
;' `;``` 12 ~1116~2 = measured velocity vector C = rotation velocity demanded about SR
Kd = velocity regulation gain Kp = position regulation gain 5 A = vector product (or, in English, cross product) - the celestial object is the sun, the aiming axis is an axis which is at least approximately close to an inertia axis of the satellite, chosen so as to obtain continuous illumination of a solar generator installed on the satellite, by virtue of which the satellite is in a sun-aimed mode, ~-- the satellite has another sensor with a second axis of sight, suitable for detecting another predetermined celestial object, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the (sun) - satellite - ~or other celestial object) an~le, this other celestial object is a star, by ~irtue of which the satellite is in star acquisition mode, ` - this other celestial o~ject is the earth, by virtue of which the satellite is in earth acquisition mod~, - the celestial object i5 a star, ~
:~. ,,' - the satellite also has a terrestrial sensor with a second axis of sight, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the star-satellite-earth angle, hy virtue of which the satellite is in the mode for the acquisition of the earth from a star.
The invention also proposes a satellite having a body, a sensor with an axis of sight suitable for detecting a ;
,~
f~ .
13 2 i l~ 6 ~ 2 predetermined celestial object and supplying a first quantity representing a first angle measured between the axis of sight and the projection of the instantaneous direction o~ the celestial object onto a ~irst reference plane containing this axis of sight, an actuating unit, an at~itude control unit suitable for generating, from this Lirst quantity and a second quantity representing an angle defined by the axis of sight and the projection of this instantaneous direction of the zelestial object onto a second reference plane containing this axis of sight, separate from the first reference plane, signals suitable for applying to the satellite, through the .
actuating unit, torques suitable for rotating the satellite about this direction and orienting an aiming axis in this direction, and a unit for measuring the instantaneous rotation velocity of the satellite, characterised in that this second representative quantity is applied to the attitude control unit by a preliminary processing unit suitable ~or calculating this second quantity from the first quantity and from the :
output signal of the instantaneous rotation velocity measuremen~ unit.
According to pre~erred characteristics of this satellite:
- this sensor is a sensor with a single sensing axis, the ~irst reference plane being perpendicular to this sensing axis, and the second réferenc~ plane containing this sensing axis, - the sensor is a solar sensor, - the satellite also has a stellar sensor, - the satellite also has a terrestrial sensor, - the sensor is a stellar sensor, :~
- the actuating unit inclucles thrusters, ~ `
~`" ` 21~602 - the uni~ for measuring the instantaneous rotation velocity of the satellite includes gyrometers.
BRIEF DESCRIPTION OF THE DRAWINGS
Objects, characteristics and advantages of the invention will emerge from the following description, given by way of non-limitative example, with reference to the accompanying drawings in which:
:~
- Figure 1 is a diagram of a single-axis sensor, - Figure 2 is a diagrammatic view in perspective of a satellite according to the in~ention, :
- Figure 3 is a functional diagram of an aiming device according to the invention, - Figure 4 is a perspective view of the satellite of Figure 2 in the earth seeking phase, and - Figure 5 is a diagrammatic view of this satellite after sensinq of the earth.
. DESCRIPTION OF THE PREFERRED EMBODIMENT :.
A single-axis sensor as shown diagrammatically under the reference numeral 1 in Figure 1 is in practice a slot extending along an axis Ys referred to as the sensing axis, perpendicular to an axis Xs referred to as the non-measurement axis, whilst the axis Zs perpendicular to the plane XsYs of ~-the sensor is referred to as the optical axis or axis of -sight, intended to be oriented sufficiently close to the celestial object to be detected for the latter to come within the field of view of the sensor.
..
What has just been stated applies particularly where the sensor is a solar sensor. A stellar sensor generally consists of a charge transfer detector matrix (CCD matrix) onto which an image of the star is projected. The measurement of the sensor in fact consists of two coordinates of the pixel or ::
~., ; 15 2111~02 pixels illuminated by the star. This stellar sensor, generally with dual axes, is able to supply only a single-axis indication if a measurement processing fault results in the ¦ provision of a single coordinate instead o~ two. In such 5 case, the invention can be applied.
If S is the direction of this celestial object, the sensor 1 provides a measurement signal representing the angle a between the axis of sight and the projection of this direction in a reference plane passing through the axis of sight and lO perpendicular to the sensing axis.
It is important to note that the angle B between the axis of sight and the projection of S in a second reference plane defined by the axis of sight and the sensing axis is on the other hand not measured.
The atti~ude drifts which the satellite may undergo about the axes Xs, Ys and Zs which, for the sensor, are roll, pitch and yaw axes, are designated by ~, ~ and ~.
Figure 2 shows diagrammatically the body 11 of a triple-axis stabilised satellite according to the invention, including in 2b addition, in a con~entional manner, a solar generator with one or more panels (not shown).
In the case considered here of a given earth orbit, preferably ~ -geos~nchronous, of low inclination (typically less than 10) the three axes along which it is sought to stabilise this ;~
satellite ~that is to say its body) in nominal attitude are, respectively, an axis directed towards the earth, usually termed Z and refexred to as the yaw axis, an axis perpendicular to the plane of the orbit and directed towards the south, termed Y and referred to as the pitch axis, and an X axis forming, with the pitch and yaw axes, a positive orthonormal axis system (X, Y, Z), and referred to as the roll axis. In practice, when the orbit is circular, this roll axis . `
` é ~ :
.,.,.",...............
:`` 16 2 ~ 0 2 is tangent to the orbit and preferably has the same direction as the velocity at which the satellite moves on the orbit.
Generally the solar generator extends parallel to the pitch axis.
5 The body 11 of the satellite has an attitude and orbit control system comprising:
::~
: - a terrestrial deteation system 12, here formed by a single twin-axis detector, having an optical axis Zst parallel or ; close to the Z axis and two measurement axes (or sensing axes) Xst and Yst (in practice orthogonal) transverse to the optical axis, advantageously close or even parallel to the X and Y
axes of the satellite respectively (the field of view is for example + 14 around the axis Ys and + 5 around the axis Xs), - the solar detection system including a plurality of singla-axis solar.sensors having axes of sight of different orientations: two sensors 13a and 13b are shown here, with axes o~ sight Zs and Zs' respectively oriented in the plane XZ
at least approximately a~ 45 from the axes -X and -Z on the .:
one hand and X and -~ on the other hand, and having sensing :
axes at least approximately parallel to Y, . ' '.' .
- an optional stellar detection system 1~, here formed by a twin-axis detector, having an optical axis Zsp pointed towards the north so as to be able to detect the Pole Star (as a :
reminder, the Pole Star is very close to the north) and two ~ .
measurement axes Xsp and Ysp parallel to X and Z; this stellar :~
' detection system 14 is in practice offset from the Y axis by a .
sufficient distance to prevent the solar generator encroaching -~
appreciably into its field o view;
- an ang~lar velocity detection unit 15, for example gyrometers, for measuring the angular velocities of the satellite about three axis preferably parallel to the X, Y and . : .
::
~ ~ ~,. ~.,., ,.~`,''`,', . .~::,~,~: ~ ,j ::,' -~
~: 17 ~ 6 ~ 2 Z axes, - a~ actuating unit, in this case formed by thrusters 16 at least four in number (in this case six), for generating positive or negative control torques about the X, Y and Z
: 5 axes, and - an analogue or digital processing unit 17, for processing the measurements supplied by the detection systems and, by means of control laws which are per se conventional, in the :; nominal attitude control reg:ime or in reacquisition mode, ~: 10 co~mands intendad ~or the actuating unit (through filters, limiters and modulators).
: ':
The process of the invention i9 described below with regard to the functicnal diagram in Figure 3, involving any one of the ~
singl~-axis solar sensors (chosen according to the desired .. :
direction o~ rotation) and the angular velocity detection unit 15. In the example in question, aiming is effected by means of the solar sensor 13a.
The various quantities which will be used in the iterative calculation defined below, have the followin~ definitions: .
. '.~
.20 sx~cl S~c Sz~c : components of the unit vector of the .~.
direction of the heavenly body in satellite axes, SX, SY, SZ : components of the unit vector of the :;-direction of the heavenly body in sensor :~
' axes at the previous moment, SXI, SY', SZ' : components of the unit vector o~ the direction of the heavenly body in sensor axes at the present moment, 30 ~ : angle between the axis Zs and the projection of the direction of the heavenly body in the plane Xs, Zs in the ' :
.
: .`` 18 2111602 .sensor frame of re~erence, B : angle between the axis Zs and the projection of the direction of the heavenly body in the plane Yszs in the ~: 5 s~nsor ~rame of reference : tan~' = SX'/SZ 7 current measurement of ~ by the sensor : tan~ = SX/S~ : previous measurement of ~ by the sensor used tanB' = SY'/SZ' : current estimation of B
10 (tan*')f current filtered value of tanB
tanB) f : previous filtered value of tanB ~::
: roll, pitch and yaw micro-rotations : between the sensor frames of reference at the previous and current moments ::.
15 Dt : duration of the calculation cycle ~xs= Dt ' ~Yg Dt ' ~5 D~
roll, pitch and yaw velocity in sensor ~
axes ~;
~c' ~y~o' .... ~, 20 ~z~c : roll, pitch and yaw velocities measured by ~
satellite axis :.
M : matrix for converting between the ~.
satellite axes and sensor axes .
: inverse matrix of the matrix M
It should first of all be noted that there is a change from the heavenly body according to sensor axes, between the -~
previous and current moments, by means of the matrix equation~
~sxl f ~ 1 fs~
sY = ~ sY' .:
sz ~ ~ 1 s~', .~, .
30 The variation ~ in the tangent of the measured anqle ~ is : ;~
equal to: ~ - tana' - tan~, that is to say:
I
~5}~
: `
19 2 ~ 0 2 ~=tan~ tan~=~SX/ SX= SX~ S~-~*Sy/+~*
SZI SZ SZI SZ~*S~-~*SX~
or again:
*tan~*tan@/_- tana~*tan~*~ - tan~
1-tan~*~-tana/*~
5 The est.imation of the non-measured angular component B' of the direction of the heavenly body in the sensor frame of reference is effected:
by cal~ulating the instantaneous velocities of rotation in ;~
the sensor frame of reference according to the change of frame 10 of reference formula:
~xsl ~)x~c~
~i)yS = M I ~ys/C : ~ ~
,~)Z5 l~ZS/C, . ` . '^
- whence tanB' is calculated by the equation:
~yS~tan*tanal*~yS*
~z9+~xg*~an~
' ~`~'`' ~ `' ' Advantageously a filtering of tanB' is carried out in order to eliminate the effect of the cyclic determination of this tangent. This filterinq is for example a first order 20 filtering of the type~
~tanB')r - a(tanB)f ~ (l-a)tanB' ~`
, where the subscript f corresponds to the filtered values and where a is a constant. ~-Finally, there is a progression, in a ~nown manner, from these 25 tangents to the components of the unit vector of the direction of the heavenly body in the sensor frame of reference according to the e~uations:
~ "~ ' .ul t~
~ . 20 2111~2 I_ tan~l SX-- - -~l+tan2a,~ 2 p j .
_ tan~ :
~/l+tan2~tan2~/
~ ~ 5Z/~
tan2a/+tall,2~1 ~:; and then by change o~ ~rame of reference these components are 5 determined in the satellite frame of re~erence: :~
~ = M ISYI
~ sz . ~
The attitude control proper is then obtained by means of a ::
lO conventional law, *or example the one already presented above : in the prea~ble:
U=-~d*~-c*s~c~-Kp*~s~c A SR]
where SR has the same meàn~ng as before.
In Figure 3 can thus be seen the sensor whose field of view ~
15 contains the satellite-heavenly body direction (in this case ~ :
the single-axis sensor 13a), the angular velocity detection.
unit 15 Pormied by gyromieters, the processing unit 17 and the actuating means 16.
The processing unit 17 includes a control unit 20, ~:~
20 conventional in itself, suitable for supplying attitude control signals U fromi two quantities characteristic of the angular orientation of the satellite-heavenly body direction : :
(in this case the tangents of the instantaneous angles ~' and B').
25 According to the invention one of these quantities is derived, not directly and solely from the output of an angular position ,~,,",,.,.,, .. ,.,.,.,,.,.. ,.,.,,.. .. , .,, ,. ~
;~ 21 21 11~2 sensor, but from the instantaneous angular velocity of the satellite in the frame of reference (X, Y, Z) related to it.
The processing unit 17 includes for this purpose a preliminary processing unit 21 suitable for calculating as indicated abova, from the single quantity tan~ supplied by the sensor 13a and the instantaneous angular velocity ~, the other quantity tan~ required by the control unit.
The satellite 11 in earth seeking phase can be seen in Figure 4. The satellite 11 is rotating about the axis SR pointed ;~
towards the sun. This rotation brings the axis of sight of ; the terrestrial sensor (which is identical with the yaw axis of the satellite~ facing the earth, thus enabling the latter to be detected. This situation is shown diagrammatically in ~-Figure 5.
It goes without saying that the above description has been put forward only by way of non-limitative example and that many variants could be proposed by a person skilled in the art without departing from the scope of the invention. Thus, for example, the stellar sansor could be omitted: a single-axis solar sensor with an axis of sight approximately directed along Z could then be added in order to supplement the field in which it is possible, at any time, to search for the sun. ~`
Likewise the process of the invention can be applied without any problem:
- to the case of a satellite provided with at least one twin-axis solar or stellar sensor formed by two coupled single-axis sensors, one o~ which happened to fail, - to the case of a satellite provided with at least one twin-axis stellar sensor, the measurements of which along one of the axes is unavailable because of a failure.
Claims (29)
1. A method of controlling the attitude of a satellite according to which the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector of the satellite is detected and, by means of an actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate the satellite about this direction whilst orienting an aiming axis related to the satellite in this direction, characterised in that this direction of the predetermined celestial object is defined in the frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of this direction onto a first reference plane containing this axis of sight and by a second quantity representing a second angle defined by this axis of sight and the projection of this direction onto a second reference plane containing this axis of sight, this second angle being calculated from this first angle and from the instantaneous angular velocity vector of the satellite.
2. The method according to Claim 1, characterised in that this axis of sight belongs to a sensor with at least one sensing axis, the first reference plane being defined as being perpendicular to the sensing axis.
3. The method according to Claim 2, characterised in that the second reference plane is perpendicular to the first reference plane and defined as containing the axis of sight and the sensing axis.
4. The method according to Claim 2, characterised in that the sensor is a single-axis sensor.
5. The method according to Claim 3, characterised in that the sensor is a single-axis sensor.
6. The method according to Claim 2, characterised in that the sensor is a twin-axis sensor, a single output of which is used.
7. The method according to Claim 3, characterised in that the sensor is a twin-axis sensor, a single output of which is used.
8. The method according to Claim 1, characterised in that the first and second quantities representing the first and second angles are the tangents of these angles.
9. The method according to Claim 1, in which the control law is of the type:
where:
? = demand to he applied to the torque generator = unit vector of the instantaneous direction of the celestial object = unit vector of the aiming axis forming the rotation reference axis ? = measured velocity vector C = rotation velocity demanded about Kd = velocity regulation gain Kp = position regulation gain ? = vector product.
where:
? = demand to he applied to the torque generator = unit vector of the instantaneous direction of the celestial object = unit vector of the aiming axis forming the rotation reference axis ? = measured velocity vector C = rotation velocity demanded about Kd = velocity regulation gain Kp = position regulation gain ? = vector product.
10. The method according to Claim 1, in which the celestial object is the sun.
11. The method according to Claim 10, characterised in that the aiming axis is an axis which is at least approximately close to an inertia axis of the satellite, chosen so as to obtain continous illumination of a solar generator installed on the satellite, by virtue of which the satellite is in a sun-aimed mode.
12. The method according to Claim 10, characterised in that the satellite has another sensor with a second axis of sight, suitable for detecting another predetermined celestial object, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the (sun) - satellite - (or other celestial object) angle.
13. The method according to Claim 12, characterised in that this other celestial object is a star, by virtue of which the satellite is in star acquisition mode.
14. The method according to Claim 12, characterised in that this other celestial object is the earth, by virtue of which the satellite is in earth acquisition mode.
15. The method according to Claim 1, characterised in that the celestial object is a star.
16. The method according to Claim 15, characterised in that the satellite also has a terrestrial sensor with a second axis of sight, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least aproximately equal to the star-satellite-earth angle, by virtue of which the satellite is in the mode for the acquisition of the earth from a star.
17. The method according to Claim 1, characterised in that the satellite is triple-axis stabilised on a geosynchronous orbit with an inclination of less than 10 degrees.
18. Satellite having a body, a sensor with an axis of sight suitable for detecting a predetermined celestial object and supplying a first quantity representing a first angle measured between the axis of sight and the projection of the instantaneous direction of the celestial object onto a first reference plane containing this axis of sight, an actuating unit, an attitude control unit suitable for generating, from this first quantity and a second quantity representing an angle defined by the axis of sight and the projection of this instantaneous direction of the celestial object onto a second reference plane containing this axis of sight, separate from the first reference plane, signals suitable for applying to the satellite, through the actuating unit, torques suitable for rotating the satellite about this direction and orienting an aiming axis in this direction and a unit for measuring the instantaneous rotation velocity of the satellite, characterised in that this second representative quantity is applied to the attitude control unit by a preliminary processing unit suitable for calculating the second quantity from the first quantity and from the output signal of the instantaneous rotation velocity measurement unit.
19. The satellite according to Claim 18, characterised in that this sensor is a sensor with a single sensing axis, the first reference plane being perpendicular to this sensing axis, and the second reference plane containing this sensing axis.
20. The satellite according to Claim 18, characterised in that the sensor is a solar sensor.
21. The satellite according to Claim 19, characterised in that the sensor is a solar sensor.
22. The satellite according to Claim 18, characterised in that the satellite also has a stellar sensor.
23. The satellite according to Claim 18, characterised in that the actuating unit includes thrusters.
24. The satellite according to Claim 18, characterised in that the unit for measuring the instantaneous rotation velocity of the satellite includes gyrometers.
25. A method of controlling the attitude of a satellite having an actuating assembly, and defining an aiming axis related to the satellite in which:
a) the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, b) the instantaneous angular velocity vector of the satellite is detected and, c) by means of the actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate the satellite about this said direction whilst orienting said aiming axis in this said direction, wherein:
d) said direction of the predetermined celestial object is defined in the frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of this direction onto a first reference plane containing this axis of sight and by a second quantity representing a second angle defined by this axis of sight and the projection of this direction onto a second reference plane containing this axis of sight, this second angle being calculated from this first angle and from the instantaneous angular velocity vector of the satellite.
a) the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, b) the instantaneous angular velocity vector of the satellite is detected and, c) by means of the actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate the satellite about this said direction whilst orienting said aiming axis in this said direction, wherein:
d) said direction of the predetermined celestial object is defined in the frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of this direction onto a first reference plane containing this axis of sight and by a second quantity representing a second angle defined by this axis of sight and the projection of this direction onto a second reference plane containing this axis of sight, this second angle being calculated from this first angle and from the instantaneous angular velocity vector of the satellite.
26. A satellite having:
a body which has an aiming axis;
a sensor having an axis of sight suitable for detecting a predetermined celestial object and supplying a first quantity representing a first angle measured between the axis of sight and the projection of the instantaneous direction of the celestial object onto a first reference plane containing this axis of sight;
an actuating unit;
an attitude control unit suitable for generating, from this first quantity and a second quantity representing an angle defined by the axis of sight and the projection of this instantaneous direction of the celestial object onto a second reference plane containing this axis of sight, separate from the first reference plane, signals suitable for applying to the satellite, through the actuating unit, torques suitable for rotating the satellite about this direction and orienting the aiming axis in this said direction;
and a unit for measuring the instantaneous rotation velocity of the satellite, wherein a preliminary processing unit suitable for calculating the second quantity from the first quantity and from the output signal of the instantaneous rotation velocity measurement unit is arranged to apply the second representative quantity to the attitude control unit.
a body which has an aiming axis;
a sensor having an axis of sight suitable for detecting a predetermined celestial object and supplying a first quantity representing a first angle measured between the axis of sight and the projection of the instantaneous direction of the celestial object onto a first reference plane containing this axis of sight;
an actuating unit;
an attitude control unit suitable for generating, from this first quantity and a second quantity representing an angle defined by the axis of sight and the projection of this instantaneous direction of the celestial object onto a second reference plane containing this axis of sight, separate from the first reference plane, signals suitable for applying to the satellite, through the actuating unit, torques suitable for rotating the satellite about this direction and orienting the aiming axis in this said direction;
and a unit for measuring the instantaneous rotation velocity of the satellite, wherein a preliminary processing unit suitable for calculating the second quantity from the first quantity and from the output signal of the instantaneous rotation velocity measurement unit is arranged to apply the second representative quantity to the attitude control unit.
27. A method for controlling the attitude of a satellite with respect to a predetermined celestial body, said satellite comprising a geometrical frame of reference, an aiming axis and an axis of sight both fixed with respect to said geometrical frame of reference, a first reference plane and a second reference plane both containing said axis of sight, said method comprising the steps of sensing a first value representative of a first angle defined by said axis of sight with the projection of the instantaneous direction of said predetermined celestial body onto said first reference plane, sensing the instantaneous angular velocity vector of the satellite within said geometrical frame of reference, calculating from said first value and from said instantaneous angular velocity vector a second value representative of a second angle defined by said axis of sight with the projection of said instantaneous direction of said predetermined celestial body onto said second reference plane, applying torques to said satellite, said torques being defined by a predetermined control law from said first and second values so as to aim said aiming axis at said predetermined celestial body as well as to rotate said satellite about said direction of said predetermined celestial body.
28. A satellite having a body with a geometrical frame of reference, an aiming axis fixed with respect to said geometrical frame of reference, a viewing means adapted to sense a predetermined celestial body and having an axis of sight and a first reference plane and a second reference plane both including said axis of sight, said sensing means being adapted to provide a first value representative of a first angle between said axis of sight and the projection of the instantaneous direction of
29 said predetermined celestial body onto said first reference plane, a rotation velocity sensing means for sensing the instantaneous rotation velocity of the satellite and providing rotation velocity values, a preliminary processing means connected to said viewing means and to said rotation velocity sensing means for calculating from said first value and said rotation velocity values a second value representative of a second angle between said axis of sight and the projection of the instantaneous direction of said predetermined celestial body onto said second reference plane, attitude control means connected to said viewing means and to said preliminary processing means for generating from said first and second values, torque signals adapted to aim said aiming axis at said predetermined celestial body and to rotate said satellite about said direction of said predetermined celestial body, and actuating means for applying torques to the satellite as a response to said torque signals.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| FR9215231 | 1992-12-17 | ||
| FR9215231A FR2699701B1 (en) | 1992-12-17 | 1992-12-17 | Method of controlling the attitude of a satellite pointed towards a celestial object and satellite adapted to its implementation. |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2111602A1 true CA2111602A1 (en) | 1994-06-18 |
Family
ID=9436717
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002111602A Abandoned CA2111602A1 (en) | 1992-12-17 | 1993-12-16 | Method for controlling the attitude of a satellite aimed towards a celestial object and a satellite suitable for implementing it |
Country Status (6)
| Country | Link |
|---|---|
| US (1) | US5458300A (en) |
| EP (1) | EP0603058B1 (en) |
| CA (1) | CA2111602A1 (en) |
| DE (1) | DE69309015T2 (en) |
| ES (1) | ES2099928T3 (en) |
| FR (1) | FR2699701B1 (en) |
Families Citing this family (22)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE19519732A1 (en) * | 1995-06-02 | 1996-12-05 | Daimler Benz Aerospace Ag | Nozzle arrangement and method for controlling the nozzles |
| FR2748721A1 (en) * | 1996-05-17 | 1997-11-21 | Matra Marconi Space France | APPARATUS FOR ADJUSTING THE ROTATION OF A SPATIAL VESSEL AROUND A AXIS |
| US6076774A (en) * | 1998-08-12 | 2000-06-20 | Hughes Electronics Corporation | Fuel and thermal optimal spiral earth acquisition |
| US6435457B1 (en) * | 2001-01-25 | 2002-08-20 | The Boeing Company | Thruster systems for spacecraft station changing, station keeping and momentum dumping |
| US6732977B1 (en) | 2002-02-11 | 2004-05-11 | Lockheed Martin Corporation | System for on-orbit correction of spacecraft payload pointing errors |
| US6695263B1 (en) | 2002-02-12 | 2004-02-24 | Lockheed Martin Corporation | System for geosynchronous spacecraft rapid earth reacquisition |
| US6561462B1 (en) * | 2002-02-21 | 2003-05-13 | The Boeing Company | Spacecraft power/sun aquistion algorithm using slit sun sensors |
| US7051980B2 (en) * | 2002-02-26 | 2006-05-30 | Lockheed Martin Corporation | Efficient orbit sparing system for space vehicle constellations |
| US6702234B1 (en) | 2002-03-29 | 2004-03-09 | Lockheed Martin Corporation | Fault tolerant attitude control system for zero momentum spacecraft |
| US6637701B1 (en) * | 2002-04-03 | 2003-10-28 | Lockheed Martin Corporation | Gimbaled ion thruster arrangement for high efficiency stationkeeping |
| DE10342866A1 (en) | 2003-09-15 | 2005-04-21 | Eads Astrium Gmbh | Method for determining the position of a spacecraft with the aid of a direction vector and a total twist measurement |
| US7630798B2 (en) * | 2005-08-05 | 2009-12-08 | The Boeing Company | Heading reference command and control algorithm systems and methods for aircraft turn-to-target maneuvers |
| US7835826B1 (en) | 2005-12-13 | 2010-11-16 | Lockheed Martin Corporation | Attitude determination system for yaw-steering spacecraft |
| CN102267575B (en) * | 2011-04-29 | 2013-09-11 | 航天东方红卫星有限公司 | Posture compensation method for dealing with mismatching of push-scanning speed for imaging along random direction of quick satellite |
| CN102267574B (en) * | 2011-04-29 | 2013-05-01 | 航天东方红卫星有限公司 | Posture compensation method in case of mismatched side-swing push-broom velocity in imaging of agile satellite |
| CN102410842B (en) * | 2011-07-26 | 2013-05-08 | 西安费斯达自动化工程有限公司 | Visual attitude measuring method based on vertical spinning top and charge coupled device (CCD) linear array |
| FR2997386B1 (en) * | 2012-10-31 | 2015-05-29 | Thales Sa | OPTIMIZED PROPULSION DEVICE FOR ORBIT CONTROL AND SATELLITE ATTITUDE CONTROL |
| US9334068B2 (en) * | 2014-04-04 | 2016-05-10 | NOA Inc. | Unified orbit and attitude control for nanosatellites using pulsed ablative thrusters |
| CN105539884B (en) * | 2016-02-05 | 2017-10-27 | 上海微小卫星工程中心 | A kind of satellite driftage control guidance method |
| CN110228605B (en) * | 2019-06-18 | 2021-03-23 | 北京电子工程总体研究所 | Sun sensor-based satellite safety sun control method |
| CN111717415B (en) * | 2020-06-22 | 2022-05-17 | 中国科学院微小卫星创新研究院 | Solar cell array sun tracking method for constellation satellite |
| CN112810840B (en) * | 2021-02-26 | 2022-07-01 | 中国人民解放军国防科技大学 | On-orbit operation method and device for satellite combining response activation and standby latency |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE3417661A1 (en) * | 1983-05-13 | 1984-11-15 | Mitsubishi Denki K.K., Tokio/Tokyo | System for controlling the orientation of an artificial satellite |
| FR2583873B1 (en) * | 1985-06-20 | 1987-09-11 | Matra | METHOD AND DEVICE FOR INJECTING SATELLITE ONTO GEOSTATIONARY ORBIT WITH STABILIZATION FOLLOWING THE THREE AXES |
| DE3885883D1 (en) * | 1987-09-16 | 1994-01-05 | Deutsche Aerospace | DEVICE FOR SETPOINT CONTROL AND / OR STABILIZATION OF MOVABLE BODIES WITH STORED TWIST. |
| GB8809247D0 (en) * | 1988-04-20 | 1988-05-25 | British Aerospace | Attitude recovery for spacecraft |
| FR2637565B1 (en) * | 1988-10-06 | 1991-01-11 | Aerospatiale | ACTIVE CONTROL SYSTEM BASED ON THREE AXES OF THE ATTITUDE OF A GEOSTATIONARY SATELLITE |
| US4997146A (en) * | 1989-05-30 | 1991-03-05 | Ford Aerospace Corporation | Electronic sun incidence avoidance device |
| DE3922761C2 (en) * | 1989-07-11 | 1994-08-25 | Deutsche Aerospace | Method and device for aligning a geostationary satellite in a reference direction |
| US5080307A (en) * | 1990-05-14 | 1992-01-14 | Hughes Aircraft Company | Spacecraft earth-pointing attitude acquisition method |
| US5255879A (en) * | 1991-11-27 | 1993-10-26 | Hughes Aircraft Company | Three axes stabilized spacecraft and method of sun acquisition |
| US5227385A (en) * | 1992-03-13 | 1993-07-13 | Wake Forest University | Method for treatment of neurodegenerative diseases |
| US5348255A (en) * | 1992-06-02 | 1994-09-20 | Hughes Aircraft Company | System and method for sensing attitude of a spacecraft with equilized star tracker errors along three orthogonal axes |
-
1992
- 1992-12-17 FR FR9215231A patent/FR2699701B1/en not_active Expired - Fee Related
-
1993
- 1993-12-14 DE DE69309015T patent/DE69309015T2/en not_active Expired - Fee Related
- 1993-12-14 ES ES93403018T patent/ES2099928T3/en not_active Expired - Lifetime
- 1993-12-14 EP EP93403018A patent/EP0603058B1/en not_active Expired - Lifetime
- 1993-12-16 CA CA002111602A patent/CA2111602A1/en not_active Abandoned
- 1993-12-16 US US08/168,931 patent/US5458300A/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| FR2699701A1 (en) | 1994-06-24 |
| DE69309015D1 (en) | 1997-04-24 |
| EP0603058A1 (en) | 1994-06-22 |
| DE69309015T2 (en) | 1997-07-03 |
| ES2099928T3 (en) | 1997-06-01 |
| EP0603058B1 (en) | 1997-03-19 |
| US5458300A (en) | 1995-10-17 |
| FR2699701B1 (en) | 1995-03-24 |
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