SATELLITE MODULE FOR ATTITUDE DETERMINATION

Abstract

A satellite module for attitude determination includes a containment body comprising at least one data acquisition board and a connection interface, at least one first-type sensor selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, in communication with the data acquisition board and at least one second-type sensor, different from the first type, selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, and in communication with the data acquisition board. The connection interface may be mounted on a first face of the containment body, the first-type sensor may be mounted on a second face of the containment body, and the second-type sensor may be mounted on a third face of the containment body.

Claims

1. A satellite module for attitude determination comprising: a polyhedral containment body configured to be mounted on a satellite and comprising at least one data acquisition board and a connection interface for allowing the data acquisition board to be in signal communication and electrical communication with the satellite; at least one first-type sensor selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, in communication with the data acquisition board and at least one second-type sensor, different from the first type, selected from a sun sensor, an earth sensor, a stellar sensor, a horizon sensor, and in communication with the data acquisition board, wherein the connection interface is mounted on a first face the polyhedral containment body, the first-type sensor mounted on a second face of the polyhedral containment body, and the second-type sensor is mounted on a third face of the polyhedral containment body.

2. The satellite module according to claim 1, wherein the first-type sensor comprises an earth sensor or a sun sensor.

3. The satellite module according claim 1, wherein the first-type sensor has a field of view having an angular diameter of at least 60 degrees.

4. The satellite module according to claim 1, wherein the second-type sensor is a stellar sensor.

5. The satellite module according to claim 4, wherein the stellar sensor has a field of view having an angular diameter higher than 15 degrees.

6. The satellite module according to claim 1, comprising a third-type sensor, different from the first and second types, mounted on a fourth face of the polyhedral containment body.

7. The satellite module according to claim 6, wherein the third-type sensor is selected from a sun sensor, an earth sensor, a horizon sensor.

8. The satellite module according to claim 7, wherein both a first-type sensor and a third-type sensor are mounted on the second face and on the fourth face.

9. The satellite module according to claim 6, wherein said polyhedral containment body comprises a top face and a front face, between which a plurality of connecting faces extend; said second face or said fourth face being identified by two of said pluralities of connecting faces.

10. The satellite module according to claim 9, wherein at least one first-type sensor is mounted on each connecting face.

11. The satellite module according to claim 10, wherein both a first-type sensor and third-type sensor are mounted on each connecting face.

12. The satellite module according to claim 9, wherein the plurality of connecting faces comprises four connecting faces, each of which forms an angle between 10 degrees and 25 degrees with an adjacent connecting face.

13. The satellite module according to claim 9, wherein said polyhedral containment body comprises a lateral face and at least one intermediate face which extends between the lateral face and the front face; at least one first-type or third-type sensor being mounted on said intermediate face.

14. The satellite module according to claim 13, comprising two intermediate faces, wherein a first intermediate face forms an angle between 20 degrees and 40 degrees with a second intermediate face.

15. The satellite comprising at least two satellite modules according to claim 1.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0067] Further characteristics and advantages of the present invention will become clearer from the following detailed description of preferred embodiments thereof, with reference to the appended drawings and provided by way of an indicative and non-limiting example. In such drawings:

[0068] FIG. 1 is a schematic perspective view of a satellite module according to the present invention;

[0069] FIG. 2 is a functional schematic view of some parts of the satellite module of FIG. 1;

[0070] FIG. 3 is a schematic perspective view of a satellite comprising eight satellite modules according to the present invention; and

[0071] FIGS. 4 to 9 are bottom, top and side views of the satellite of FIG. 3.

DETAILED DESCRIPTION

[0072] With reference to the accompanying figures, a satellite module according to the present invention is generally indicated with the number 10.

[0073] The satellite module comprises a containment body 11 made of metallic material, for example of aluminium, aluminium alloy or titanium. In the preferred embodiment of the invention, the containment body can be inscribed in a cube having a smaller side of 10 centimetres, preferably of about 5 centimetres.

[0074] At least one first-type sensor 12, at least one second-type sensor 13 and at least one connection interface 14 are constrained to the containment body 11.

[0075] In the preferred embodiment of the invention, at least one third-type sensor 15 is also constrained to the containment body 11.

[0076] The first, second and third types of sensors (when the third type of sensors are present) belong to sensors that need to be directly exposed to, i.e. directly in view of, the quantity to be measured. The first-type sensor 12, the second-type sensor 13 and the third-type sensor 15 (when present) are mounted on external surfaces of the containment body 11 so that the respective devices sensitive to the quantity to be measured face the environment outside the containment body 11.

[0077] In the preferred embodiment of the invention, the first-type sensor 12 is a sun sensor, i.e. a sensor capable of providing the direction which connects the sun to the sensor itself. By way of example, the sun sensor 12 can be a photodiode (or an array of photodiodes) or a photodetector (or an array of photodetectors) capable of transforming an incident light into an electrical signal which is subsequently processed and compared to determine the direction of the incident light. The angular diameter of the field of view of the sun sensor 12 is between 60 degrees and 90 degrees, preferably of about 80 degrees.

[0078] The second-type sensor 13 of sensors is a stellar sensor, that is to say a sensor that detects the position of the stars through photocells or photographic equipment. The stellar sensor 13 comprises, in the preferred embodiment of the invention, an optical head 13a equipped with a baffle 13b (see FIG. 1) and electronics 13c (FIG. 2) configured to process the detected images and compare them with a preloaded stellar catalogue containing less than 1000 stars. The angular diameter of the field of view of the stellar sensor 13 is greater than 15 degrees, preferably between 20 degrees and 60 degrees, even more preferably of about 40 degrees.

[0079] The third-type sensor 15 is an earth sensor, that is to say a sensor capable of detecting the position of the earth with respect to the satellite. In the preferred embodiment of the invention, the earth sensor 15 is an infrared thermopile. The angular diameter of the field of view of the earth sensor 15 is between 60 degrees and 90 degrees, preferably of about 80 degrees.

[0080] Inside the containment body 11 there is also provided at least one further sensor 16 of a fourth type different from the first, second and third type of sensors.

[0081] The fourth-type sensor 16 is selected from the group comprising a magnetometer and a gyroscope. In the preferred embodiment of the invention, two fourth-type sensors 16 are provided, in particular a magnetometer and a gyroscope. The magnetometer is configured to measure the module of the earth's magnetic field or to measure the component of the earth's magnetic field along one or more directions. The gyroscope is an electronic gyroscope, for example a Mems gyroscope and is able to detect changes in orientation of the satellite.

[0082] As shown schematically in FIG. 2, a data acquisition board 17 is provided inside the containment body 11 placed in signal communication with the sensors of the first 12, second 13, third 15 and fourth type 16 in order to receive electrical signals from the sensors themselves.

[0083] In the preferred embodiment of the invention, the data acquisition board 17 integrates (or can be in signal communication with) a microcontroller 18 configured to process the electrical signals acquired by the sensors and generate output signals configured to be received by the satellite attitude control system and processed to change the satellite attitude.

[0084] The data acquisition board 17 is in signal connection and in electrical connection with the connection interface 14 to receive from it an electrical power supply, for example at 5V DC, taken from on-board systems of the satellite and to receive the output signals generated by the microcontroller 18. The data acquisition board 17 is also in electrical connection with one or more of the sensors to power them.

[0085] The connection interface 14 is configured to connect to a satellite connector to integrate the satellite module 11 into the ADCS system of the satellite.

[0086] As represented in FIG. 1, the containment body 11 is externally polyhedral in shape, so as to comprise a plurality of faces oriented differently from each other.

[0087] In particular, the containment body 11 comprises a first face 19 on which the connection interface 14 is mounted, a second face 20 on which the first-type sensor 12 is mounted, a third face 21 on which the second-type sensor 13 is mounted and a fourth face 22 on which the third-type sensor 15 is mounted.

[0088] In the embodiment shown in FIG. 1, the sun sensor 12 is mounted on the second face 20, the stellar sensor 13 is mounted on the third face 21 and the earth sensor 15 is mounted on the fourth face 22.

[0089] The second 20 and the fourth face 22 are faces of the containment body 11 which belong to a plurality of connecting faces 23 which extend between a top face 24 and a front face 25 of the containment body 11.

[0090] The front face 25 and the top face 24 are substantially perpendicular to each other.

[0091] In the embodiment of FIG. 1, the connecting faces 23 are four, they are consecutive to each other and uniformly connect the top face 24 and the front face 25. Each connecting face forms an angle of about 18 degrees with the previous and/or next connecting face. The connecting face 23 immediately adjacent to the front face 25 forms an angle of about 18 degrees with the front face 25 and the connecting face 23 immediately adjacent to the top face 24 forms an angle of about 18 degrees with the top face 24.

[0092] The connecting faces 23 are preferably plane and the top 24 and front 25 faces also are preferably planet.

[0093] The first face 19 defines a rear face of the containment body 11. The rear face 19 is opposite to the front face 25 and is preferably parallel to the front face 25.

[0094] The third face 21 defines a lateral face for the containment body 11. The lateral face 21 is substantially perpendicular to the top face 24 and the front face 25. As illustrated in FIG. 1, the lateral face 21 is also substantially perpendicular to the rear face 19.

[0095] Two intermediate faces 26 extend between the lateral face 21 and the front face 25. The intermediate faces 26 are consecutive to each other and uniformly connect the lateral face 21 and the front face 25. The intermediate faces 26 form an angle of about 33 degrees between them. The intermediate face 26 immediately adjacent to the front face 25 forms an angle of about 33 degrees with the front face 25 and the intermediate face 26 immediately adjacent to the lateral face 21 forms an angle of about 33 degrees with the lateral face 21. The intermediate faces 26 are preferably plane. The intermediate faces 26 also form an edge with two connecting faces 23, in particular they form an edge with the two consecutive connecting faces 23 that are the closest to the front face 25.

[0096] In the preferred embodiment of the invention, an earth sensor 15 and a sun sensor 12 are mounted on each connecting face 23, as illustrated in FIG. 1. Preferably, an earth sensor 15 and a sun sensor 12 are also mounted on both intermediate faces 26.

[0097] The satellite module 10 creates a stand alone unit capable of operating autonomously and independently of other satellite modules 10.

[0098] FIG. 3 shows a possible and preferred application of a plurality of satellite modules 10 on a satellite 100, for example a cubic-shaped satellite or in any case a satellite having eight edges.

[0099] For display convenience, the satellite modules 10 with which the satellite 100 is equipped have been represented with Roman numerals I to VII to distinguish them from each other despite being identical.

[0100] Eight satellite modules 100 which are identical to each other but differently oriented between them are mounted on the satellite 100. The satellite modules 10 are mounted at the eight edges of the satellite 100.

[0101] FIG. 4 shows a bottom view of the satellite 100 of FIG. 3 with the orientations of the various sensors mounted on the containment body 11 highlighted. For display convenience, a single stellar sensor 13, a single sun sensor 12 and a connection interface 14 have been represented. Furthermore, the containment body 11 has been represented cubic in shape. In any case, in accordance with the satellite module described above, the containment body 11 comprises a first face on which the connection interface 14 is mounted, a second face on which a sun sensor 12 is mounted and a third face on which a stellar sensor 13 is mounted. A person skilled in the art will easily understand that the following also applies to a satellite module 10 whose containment body 11 and whose sensor arrangement is identical to the satellite module of FIG. 1.

[0102] FIG. 5 is a top view of the satellite 100 of FIG. 3 and FIGS. 6 to 9 are side views of the satellite 100 of FIG. 3.

[0103] The satellite module 10 indicated by the Roman numeral V in FIG. 4 has been represented with the stellar sensor 13 which is directed towards the viewer of the figure and the sun sensor 12 has been represented directed downwards for the viewer of the figure. The satellite module 10 indicated by the Roman numeral VI in FIG. 4 has been represented with the stellar sensor 13 which is directed downwards for the viewer of the figure and the sun sensor 12 has been represented directed towards the viewer of the figure. The satellite module 10 indicated by the Roman numeral V in FIG. 4 has been represented with the stellar sensor 13 which is directed to the right of the viewer of the figure and the sun sensor 12 has been represented directed towards the viewer of the figure. The satellite module 10 with no Roman numeral indication in FIG. 4 has been represented with the stellar sensor 13 which is directed towards the viewer of the figure and the sun sensor 12 has been represented directed to the left of the viewer of the figure.

[0104] Identical symbols to represent the orientation of the sun sensor 12 and the stellar sensor 13 have been adopted in figures from 5 to 9.

[0105] As can be appreciated from FIGS. 4 to 9, for each side of the satellite 100 there is at least one stellar sensor 13 which has its own field of view directed moving away from this side.

[0106] Likewise, for each side of the satellite 100 there is at least one sun sensor 12 which has its own field of view directed moving away from this side.

[0107] In this way, whatever the attitude of the satellite 100 during a mission, there is always at least one stellar sensor 13 whose field of view frames only stars and at least one sun sensor whose field of view only frames the sun.

[0108] In the case of a satellite module whose containment body 11 is like that represented in FIG. 1, regardless of the attitude of the satellite 100 during a mission, there will always be at least one stellar sensor 13 whose field of view frames only stars and at least four sun sensors and four earth sensors whose respective field of views only frame the sun and earth.

[0109] This obtained minimum redundancy of the sensors that simultaneously measure the same quantity (for some attitudes the numbers of sensors of the same type that are measuring the same quantity can be even greater) allows the use of sensors with reduced sensitivity and also allows to correct or in any case to filter incorrect readings of other sensors.

[0110] Furthermore, the obtained redundancy of sensors that simultaneously measure the same quantity allows to have a slow degradation in the attitude determination of the satellite 100, since a possible failure of a sensor is compensated by the redundant sensors obtaining, for each sensor failure, a small degradation in the accuracy of the attitude determination (at least as long as at least one of the sensors continues to work).