Satellite with mechanical structure

Abstract

The present invention concerns a rideshare satellite (1) comprising a satellite's body (10) including a nadir face (101) and a zenith face (102) opposed one to another. The satellite comprises a mechanical structure within the satellite's body (10). The mechanical structure comprises a structural hollow core element (20) extending from the zenith face (102) within the satellite's body. An extremity of the hollow core element (20) forms a fixation ring (201) on the zenith face (102) configured to be fixed to a payload adapter (A) of a spacecraft (S) when the rideshare satellite (1) is launched. The invention also concerns a launching method and a payload adapter.

Description

Satellite with mechanical structureTechnical domainThe present invention concerns a rideshare satellite provided with a mechanical structure, a method for launching a rideshare satellite and an adapter ring for rideshare satellites.Related art

[0001] Rideshare satellites, also known as secondary payload satellites, are small-sized satellites transported to orbit on a launch spacecraft which carries a primary payload. As a secondary payload, the rideshare satellite has to comply with the constraints imposed by the primary payload, in particular with the constraints of available space in the spacecraft, of launching schedule and of trajectory and orbit is typically determined by the entity launching the primary payload.

[0002] Therefore the entity launching the rideshare satellite needs to tailor the satellite and launch to those constraints which usually means specifically dimensioning the satellite and operating adaptation manoeuvres for the satellite to reach its orbit of operation.

[0003] Rideshare satellites are typically disposed around a central ring disposed around the longitudinal axis of the launch vehicle (i.e. spacecraft) which allows multiple secondary payloads to be secured radially around that axis. An industry standard is the family of EELV Secondary Payload Adapters (shortened as ESPA) consisting in a ring designed to support a primary payload on its top together with six secondary payloads laterally secured to the central ring through fixation ports.

[0004] This family of adapters allows securing secondary payloads weighting up to 700kg to fixation ports of a maximal diameter of 24STO12-158-PCTinches. Heavier payloads usually need to be secured to the spacecraft using the top port so as to be aligned with the axis of the spacecraft.

[0005] Hence, the available space for rideshare satellite corresponds to an annular region determined on the inside by the diameter of the central ring and on the outside by the inner diameter of the spacecraft. In order to maximize the efficiency of secondary payload satellites, manufacturers have to develop compact satellite configurations to match the geometry of available space and including a maximal number of features while complying with the weight requirements.

[0006] However, new classes of telecommunication satellites that are far smaller and cheaper than traditional telecom satellites are arising. It is desirable for such satellites to beneficiate of the ridesharing cost efficiency, however they usually exceed the maximal weight authorized by existing secondary payload adapters.

[0007] The new classes of satellites mentioned above are typically configured to operate in Geostationary orbit (GEO), Medium Earth orbit (MEO) and / or High Earth orbit (HEO).

[0008] Hence, an important feature of those satellites is a thruster architecture allowing them to be sent as rideshare satellite that is allowing them to switch from an elliptical or circular transfer orbit to which it is sent by the spacecraft to their orbit of operation. Additionally the thruster architecture also need to include station keeping thrusters allowing the satellites to maintain their orbit and orientation.

[0009] This architecture is taken into account in the seek for optimization of both the available space in the spacecraft and the weight efficiency. Indeed, powerful propulsion systems, such as those needed to switch from orbits, often require fuel tanks that are both bulky and heavy.STO12-158-PCT

[0010] In the state of the art, satellite's belonging to these new classes are attached to the launching spacecraft in a longitudinal orientation, i.e. the zenith / nadir direction of the satellite is aligned with the longitudinal axis of the spacecraft. Indeed, the existing separation rings and / or mass and / or mass distribution of the satellites prevent cantilevered orientation within the spacecraft. This prevents such satellites from being sent as secondary payloads, as there is only one such available spot in the spacecraft, which increases the mission's costs.

[0011] For telecommunication satellites, another important feature is the reflector system that allows signals to be transmitted from a satellite's antenna toward the Earth and vice-versa. Since the available space is limited, the size (and even the curvature) of the reflector(s) of rideshare satellites is limited as well. As a consequence, the distance between the emitting / receiving source and the corresponding reflector needs to be increased to achieve sufficient transmission quality. There is therefore a need for ways of increasing this distance while complying with the compactness, available space and weight requirements.Short disclosure of the invention

[0012] An aim of the present invention is the provision of a rideshare satellite that overcomes the shortcomings and limitations of the state of the art.

[0013] Another aim of the invention is to provide a rideshare satellite suitable for a cantilevered attachment within its launching spacecraft.

[0014] Another aim of the invention is to provide a rideshare satellite with increased inner surface for accommodating payload components.

[0015] According to the invention these aims are attained by the object of the appended claims.STO12-158-PCTBrief description of the figures

[0016] Examples of implementations of the invention are indicated in the description illustrated by the appended figures in which:• Figures 1 A and 1 B illustrate perspective views of a satellite's body according to the invention with a structural hollow core element.• Figures 2A and 2B illustrate perspective views of a satellite's body according to the invention with a structural hollow core element and four shear walls.• Figures 3A and 3B illustrate perspective views of a satellite's body according to the invention with a structural hollow core element, four shear walls and one mid-deck.• Figures 4A and 4B illustrate perspective views of a satellite's body according to the invention with a structural hollow core element, and eight shear walls.• Figures 5A and 5B illustrate perspective views of a satellite's body according to the invention with a structural hollow core element, eight shear walls and one mid-deck.• Figure 6 illustrates a tubular payload adapter.• Figures 7A-7C illustrate respectively a perspective, a lateral and a top view of a tubular payload adapter with one rideshare satellite according to the invention attached to it.STO12-158-PCTFigures 8A and 8B illustrate respectively a perspective and a top view of a tubular payload adapter with three rideshare satellite according to the invention attached to it.• Figures 9A-9C illustrate views of a satellite's body according to the invention with a structural hollow core element, and respectively four, six and eight shear walls.• Figure 10 illustrates a view of a satellite's body according to the invention with a structural hollow core element and six shear walls.• Figures 11A-11C illustrate various embodiments related to the geometry of the structural hollow core element.Examples of embodiments of the present invention

[0017] Throughout the present text, the expression "zenith" and "nadir" are used to designate opposite directions respectively to the satellite when it is in operation mode, i.e. when the satellite is in its final transmission mode. The "zenith" designates the direction opposed to the Earth while the "nadir" designates the direction of Earth. Derivative expressions such as "zenith face" or "nadir face" are therefore to be understood as "face facing the zenith direction", i.e. "face opposed to the Earth" and "face facing the nadir direction", i.e. "face directed toward the Earth.

[0018] According to an aspect illustrated in Fig. 1, the present invention relates to a rideshare satellite comprising a satellite's body 10 including a nadir face 101 and a zenith face 102 opposed one to another. The satellite comprises a mechanical structure for increasing the structural integrity of the satellite's body 10. In particular, the mechanical structure includes a structural hollow core element 20 extending from the zenith face 102STO12-158-PCTwithin the satellite's body 10. This structural hollow core element 20 forms on the outside of the zenith face 102, a fixation ring 201 configured to be fixed to a payload adapted A of a spacecraft S to which the rideshare satellite is fixed at launching.

[0019] Although always comprised within the satellite's body 10, the nadir face 101 and one of the lateral faces of the satellite's body are not illustrated in Fig. 1 A-5B in order to see the inside of the satellite's body.

[0020] Advantageously, the structural hollow core element 20 strengthen the integrity of the satellite's body 10, which allows the satellite to be fixed as secondary payload on the spacecraft S, i.e. fixed in a cantilevered way with respect to the longitudinal axis of the spacecraft as illustrated in Figs 8A and 8B. This greatly reduces mission's costs as secondary payloads are much cheaper than primary payload. It therefore allows to launch heavier satellites as rideshare satellites because of the reinforced structure provided by the structural hollow core element 20.

[0021] The hollow core element 20 may extend from the zenith face 102 within the satellite's body along a portion of the total distance between the zenith face 102 and nadir face 101 (i.e. the height of the satellite's body) or along the entirety of this total distance so as to connect the zenith and nadir faces. In one embodiment, the hollow core element 20 extends along at least 30% of this total distance.

[0022] As illustrated in Figures 9A-9C, the hollow core element 20 may extend only partially between the zenith face and the nadir face. For example, it can extend only form the zenith face to a mid-deck disposed within the satellite's body 10.

[0023] As illustrated in Fig. 11 A, the cross-section of the structural hollow core element 20 measured in a plane which is parallel to the zenith face can typically be circular, triangular, quadrilateral, polygonal with more than four sides, elliptical, star-shaped, etc. It may also consist of aSTO12-158-PCTcombination of rectilinear portions and curved portions. The cross-section may also differ from the cross-section of the fixation ring 201.

[0024] The diameter of the structural hollow core element 20 may vary across a portion or the whole height of the satellite's body 10, that is along a longitudinal of the hollow core element. As illustrated in Fig. 11 B, the hollow core element may be for example conical, pyramidal, a combination of a truncated cone and a cylinder, etc. In particular and as illustrated in Fig. 11C, the diameter of a portion of the hollow core element 20 may be different from the diameter of the fixation ring 201.

[0025] Fig. 8A illustrates a perspective view of an arrangement of a plurality of rideshare satellites 1 provided with a mechanical structure according to the present invention allowing their fixation onto the payload adapter A in the depicted cantilevered orientation. The payload adapter A is part of the spacecraft S onto which the satellites are attached. The longitudinal axis of the payload adapter A coincides with the longitudinal axis of the spacecraft so that the rideshare satellites are attached in a cantilevered configuration through a fixation ring 201. The satellites may be attached to the payload adapter A so that a longitudinal axis of the payload adapter A forms an angle with a longitudinal axis of the separation ring of each satellite ranging from 45° to 135°.

[0026] In an embodiment and in order to take further advantage of the available space within the spacecraft, the nadir face 101 is larger than the zenith face 102 to better match the geometry of the available volume within the spacecraft S. This allows to obtain a nadir face (or deck) with a larger area than satellites of the state of the art so that more components can be accommodated withing the rideshare satellite 1.

[0027] Advantageously, the diameter of the structural hollow core element 20 is strictly greater than 24 inches, preferentially greater than 28 inches. This enables heavier satellites to be fixed as secondary payload in a spacecraft. Indeed, larger diameter of the structural hollow core elementSTO12-158-PCTincrease the integrity of the satellite's body, so that a heavier satellite can be fixed in a cantilevered way on a payload adapter A.

[0028] The structural hollow core element 20 may surprisingly and advantageously accommodate a fuel tank of the rideshare satellite 1. Fuel are usually bulky and take a large amount of usable space within the satellite (or outside as the available space within spacecraft is limited). Therefore, one can take advantage of the inner volume of the structural element to accommodate the fuel tank in order to free space in the rest of the satellite, so that more components can be brought on board.

[0029] In an embodiment, the fuel tank partially protrude from the fixation ring 201 of structural hollow core element 20. This portion of the fuel tank therefore protrudes within the aperture of the corresponding fixation ring of the payload adapter A when the satellite is fixed to the spacecraft. As a consequence, the center of gravity of the rideshare satellite 1 is shifted toward the zenith face and therefore toward the payload adapter. This further increase the maximal weight of the rideshare satellite that can be fixed to the payload adapter without compromising its structural integrity.

[0030] In an embodiment, the structural hollow core element 20 further accommodates at least one satellite thruster (orbit raising or station keeping thruster). Thrusters are usually bulky and heavy, and can advantageously be disposed within the hollow core element 20 so spare usable space within or outside the satellite's body.

[0031] In an embodiment, the satellite thruster accommodated within the hollow core element 20 protrudes from the fixation ring 201 and therefore protrudes within the aperture of the corresponding fixation ring of the payload adapter A when the satellite is fixed to the spacecraft. As a consequence, the center of gravity of the rideshare satellite 1 is shifted toward the zenith face102 and therefore toward the payload adapter. ThisSTO12-158-PCTfurther increase the maximal weight of the rideshare satellite that can be fixed to the payload adapter without compromising its structural integrity.

[0032] In an embodiment, the satellite is entirely at electric solar propulsion.

[0033] The weight of the rideshare satellite 1 does not exceed 1800kg.

[0034] In an embodiment, the external dimensions of the rideshare satellite 1 in stowed configuration as shown in Fig. 7A (reflectors and solar arrays stowed) does not exceed 4m in width, 2.5m in height and 3m in depth. The width typically corresponds to the dimension measured in the East-West direction, the height to the dimension measured in the zenithnadir direction and the depth to the dimension measured in the North- South direction, when the satellite is in regular transmission configuration.

[0035] In an embodiment, the end-of-life electric power of the rideshare satellite is inferior to 4kW.

[0036] The restrictions on the weight, the dimensions and / or the end-of- life power typically correspond to classes of satellites that can greatly beneficiate of being launched as secondary payload. Indeed, they are low- cost satellites for which the mission costs are critical and must be kept minimal to preserve attractiveness.

[0037] Figs 2A to 5B illustrates various embodiments in which shear walls 21 and mid-decks 22 are provided to increase both structural integrity and available space for fixing radiofrequency, electronic and / or electric components.

[0038] Shear walls are walls protruding within the satellite's body 10 from any of the lateral face. They may extend along the whole width / length of the satellite's body 10, that is joining two oppositionSTO12-158-PCTlateral faces, or they may extend only partially within the satellite's body, for example until meeting the structural hollow core element 20. These shear walls can be vertical, i.e. in a plane parallel to the zenith-nadir direction, but can also be inclined according to the mission payload's needs. They may extend partially or entirely between the zenith and the nadir face.

[0039] In an embodiment illustrated in Figs 2A and 2B, the mechanical structure comprises at least one shear wall 21 extends radially from the structural hollow core element 20 towards the outer lateral faces 103, forming four inner volumes within the satellite's body.

[0040] The four shear walls 21 can be angularly spaced at 90° around the hollow core structural element 20 so as to delimitate four equal inner volumes within the satellite's body. Any other suitable angular configuration of the four shear walls can be considered by the skilled person.

[0041] Alternatively, or complementarily, the mechanical structure can comprise at least one mid-deck 22 extending radially from the hollow core element 20 and perpendicularly to the nadir-zenith direction. This mid-deck also allows to increase the available surface within the satellite's body.

[0042] Such mid-deck may be disposed at any height along the direction perpendicular to the zenith face.

[0043] Figs 3A and 3B illustrate such a mid-deck 22 in combination with four shear walls 21 that together delimitate eight inner volumes.

[0044] In some embodiments, the structural hollow core element 20 only extend partially within the satellite's body 10, e.g. 50% of the height of the satellite's body. As illustrated in Figs. 9A-9C, a mid-deck 22 may extend radially from such a partially extending hollow core element 20. Therefore,STO12-158-PCTthe inner available space for components is increased in the upper part of the satellite's body as is it free from the hollow core element as compared to the embodiments described above.

[0045] Fig. 9A illustrates an embodiment in which the lower part, that is the part between the zenith face 102 and the mid-deck 22 is provided with four radially extending shear walls 21. The upper part, that is the part between the mid-deck 22 and the nadir face (not illustrated) is devoid of any shear wall.

[0046] Fig. 9B illustrates an embodiment in which the lower part is provided with four radially extending shear walls 21. The upper part is provided with two shear walls 21 extending between two opposite lateral faces 103 (one of which is not illustrated).

[0047] The lower part, that is the part between the zenith face 102 and the mid-deck 22 is provided with four radially extending shear walls 21. The upper part is provided with four shear walls 21 extending from the four lateral face 103 (one of which is not illustrated) and meeting in the center of the satellite's body.

[0048] According to another embodiment illustrated in Fig. 10, the mechanical structure comprises six shear walls extending from the hollow core element 20. As a non-limitative example, these six walls may be arranged as two pairs of parallel walls and two other walls at 90° respectively to the pairs of parallel walls. The mechanical structure may additionally also include one or more mid-decks.

[0049] In another embodiment illustrated in Figs 4A and 4B, the mechanical structure comprises eight shear walls 21 delimitating eight inner volumes within the satellite's body. The eight shear walls may for example be arranged in four pairs of parallel walls extending toward the outer lateral faces 103. Alternatively, the eight shear walls may also be radially distributed around the hollow core element 20.STO12-158-PCT

[0050] As previously, a mid-deck 22 can also be combined with this shear wall configuration as illustrated in Figs. 5A and 5B so as to delimitate at least twelve inner volumes.

[0051] According to another aspect, the rideshare satellite 1 of the present invention may comprise at least one power battery 30 disposed on an outer side of the zenith face 102. This configuration allows to further free inner space for components by putting the power batteries outside of the satellite's body. Moreover, they are supported by the zenith face 102 so as to once again shift the center of gravity of the rideshare satellite toward the payload adapter A.

[0052] In an embodiment illustrated in 7C, two, three or four batteries are disposed on corner regions of the zenith face 102.

[0053] Being outside the satellite's body, the power batteries may be exposed to important solar radiation. Therefore, in one embodiment one or more of the power batteries may comprise an optical solar radiator element disposed on an outer surface of the battery exposed to radiation.

[0054] According to another aspect, the rideshare satellite 1 of the present invention may comprise an array of laterally deployable solar panels 50 attached to an outer lateral face 103 of the satellite's body. Advantageously, each solar panel of the array has a long side 51 and a short side 52 opposed to the long side. The individual solar panels are attached one to another either on both long sides or on both short sides, so that when the array is stowed against the lateral face 103, the shape of the stowed array matches the shape of the lateral face 103.

[0055] As illustrated in Fig. 7A the solar panel array 50 is stowed against an outer lateral face 103 of the satellite's body. Since the nadir face 101 is smaller than the zenith face 102, the outer lateral face 103 is trapezoid-like. The solar panel array is stowed configuration is therefore also trapezoidlike. The expression trapezoid-like refers to a shape having two parallelSTO12-158-PCTedges, one shorter than the other, those two sides being joined together by polygonal lateral side.

[0056] This matching between the geometry of the stowed solar array and of the lateral face 103 allows to maximize the size of the solar panels while not exceeding the surface of the lateral outer face.

[0057] The present invention also relates to a method for launching a rideshare satellite 1 as described above. The method comprises the steps of a. attaching the rideshare satellite 1 to a payload adapter A of a spacecraft S using fastening means between the fixation ring 201 and the payload adapter A, the tubular mechanical element 20 being cantilevered with respect to a longitudinal axis of the spacecraft S, b. bringing the spacecraft into a suitable orbit, for example the orbit of the primary payload, c. launching the rideshare satellite 1 from the spacecraft by releasing the fastening means.

[0058] According to a further aspect, the present invention relates to a payload adapter A for attaching a rideshare satellite 1 as described above. As illustrated in Fig. 6, the payload adapter A comprises a tubular body including a plurality of lateral openings, each lateral opening comprising a reinforced fixation ring 60 configured to be attached to the fixation ring 201 of the rideshare satellite using fastening means. The diameter D of each lateral opening is strictly greater than 24 inches, preferentially greater than 30 inches. This allows heavier rideshare satellites to be attached to the payload adapter A than in the state of the art.

[0059] Figs 7A to 7B illustrate various views of a rideshare satellite 1 attached to such a payload adapter A.STO12-158-PCT

[0060] Figs 8A and 8B illustrate a perspective and a top view of a payload adapter A supporting a plurality of rideshare satellites 1 according to the invention. The outer wall of the spacecraft S is depicted in Fig. 8B.STO12-158-PCT

Claims

Claims1. A rideshare satellite (1) comprising a satellite's body (10) including a nadir face (101) and a zenith face (102) opposed one to another, the satellite comprising a mechanical structure within the satellite's body (10) characterized in that the mechanical structure comprises a structural hollow core element (20) extending from the zenith face (102) within the satellite's body (10), an extremity of the hollow core element (20) forming a fixation ring (201) on the zenith face (102) configured to be fixed to a payload adapter (A) of a spacecraft (S) when the rideshare satellite (1) is launched.

2. Rideshare satellite (1) according to claim 1, wherein the structural hollow core element extends within the satellite's body (10) along a height of at least 30% of a total distance separating the zenith face (102) from the nadir face (101), preferably 50% of the total distance.

3. Rideshare satellite (1) according to any of the preceding claims, wherein the structural hollow core element (20) has a circular, triangular, quadrilateral, polygonal with more than five sides or elliptical cross-section in a plane parallel to the zenith face (102).

4. Rideshare satellite (1) according to any of the preceding claims, wherein a diameter of at least a portion of the structural hollow core element (20) is variable along a longitudinal direction of the structural hollow core element (20).

5. Rideshare satellite (1) according to any of the preceding claims, wherein the zenith face (102) is parallel and opposed to the nadir face (101), and wherein an area of the zenith face (102) is smaller than an area of the nadir face (101).STO12-158-PCT6. Rideshare satellite (1) according to any of the preceding claims, wherein a diameter of the structural hollow core element (20) is strictly greater than 24 inches, preferentially greater than 28 inches.

7. Rideshare satellite (1) according to any of the preceding claims, wherein a fuel tank is at least partially accommodated within the structural hollow core element (20).

8. Rideshare satellite (1) according to the preceding claim, wherein a portion of the fuel tank projects outside the fixation ring (201).

9. Rideshare satellite (1) according to any of the preceding claims, wherein a satellite thruster is at least partially accommodated within the structural hollow core element (20).

10. Rideshare satellite (1) according to the preceding claim, wherein a portion of the satellite thruster projects outside the fixation ring (201).

11. Rideshare satellite (1) according to any of the preceding claims, wherein the satellite is at solar electric propulsion.

12. Rideshare satellite (1) according to any of the preceding claims, wherein a maximal weight of the satellite is less than 1800kg.

13. Rideshare satellite (1) according to any of the preceding claims, wherein maximal dimensions of the satellite in a stowed configuration are less than 4m in width, 2.5m in height and 3m in depth.

14. Rideshare satellite (1) according to any of the preceding claims, wherein an end of life payload power of the satellite is less than 4kW.STO12-158-PCT15. Rideshare satellite (1) according to any of the preceding claims, wherein the mechanical structure further comprises at least one shear wall (21) extending from a lateral face (103) within the satellite's body (10).

16. Rideshare satellite (1) according to the preceding claim, wherein the least one shear wall (21) extends from lateral face (103) to the structural hollow core element (20).

17. Rideshare satellite (1) according to any of the claims 15 to 16, wherein the at least one shear wall (21) further extends from the zenith face (102) to the nadir face (101).

18. Rideshare satellite (1) according to any of the claims 16 to 17, wherein the mechanical structure comprise four shear walls (21) angularly spaced at 90° around the structural hollow core element (20) so as to delimitate four inner volumes within the satellite's body.

19. Rideshare satellite (1) according to any of the preceding claims, wherein the mechanical structure comprises at least one mid-deck (22) extending radially from the structural hollow core element (20) towards lateral outer faces (103) of the satellite's body (10).

20. Rideshare satellite (1) according to claim 15, wherein the mechanical structure comprises eight shear walls (21) so as to delimitate eight inner volumes within the satellite's body, the eight shear walls (21) being arranged in four pairs of parallel walls.

21. Rideshare satellite (1) according to claim 18, wherein the mechanical structure comprises at least one mid-deck (22) extending radially from the structural hollow core element (20) towards lateral outer faces of the satellite's body so as to delimitate, together with the eight shear walls (21), at least twelve inner volumes within the satellite's body.STO12-158-PCT22. Rideshare satellite (1) according to any of the preceding claims, comprising at least one power battery (30) disposed on an outer side of the zenith face (102).

23. Rideshare satellite (1) according to the preceding claim, comprising at least one optical solar radiator element disposed on an outer surface the at least one power battery (30).

24. Rideshare satellite (1) according to claim 5, comprising a set of laterally deployable solar panels (50) attached to an outer lateral face (103) of the satellite's body (10) each solar panel having a long side (51) and a short side (52) so that when the set of laterally deployable solar panel (50) is in a stowed position against the lateral face (103), the shape of each solar panel matches the shape of the lateral face.

25. Method for launching a rideshare satellite (1) according to any of the preceding claims, comprising the steps of attaching the rideshare satellite (1) to a payload adapter (A) of a spacecraft (S) using fastening means between the fixation ring (201) and the payload adapter (A), the structural hollow core element (20) being cantilevered with respect to a longitudinal axis of the spacecraft (S), bringing the spacecraft into a suitable orbit, launching the rideshare satellite (1) from the spacecraft by releasing fastening means.

26. Payload adapter (A) for attaching a rideshare satellite (1) according to any of the claims 1 to 24 to a spacecraft (S), the payload adapter (A) comprising a tubular body including a plurality of lateral openings, each lateral opening comprising a reinforced fixation ring (60) configured to be attached to the fixation ring (201) of the rideshare satellite using fastening means, characterized in that a diameter (D) of each lateral opening is strictly greater than 24 inches, preferentially greater than 30 inches.STO12-158-PCT

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