Abstract: A radiative fin for a spacecraft is disclosed having an end fitting of heat conductive material, configured to be mounted on the spacecraft, a flexible radiative laminate, connected to the end fitting at one end and having an opposite free end, at least one pyrolytic graphite sheet, and at least one heat emission layer in contact with the pyrolythic graphite sheet on at least part of the surface of the pyrolythic graphite sheet, and a flexible rod, extending from the end fitting along at least part of a side of the flexible radiative laminate and being affixed to the latter. The flexible rod is adapted to occupy a folded position and a deployed position and to exert, while in the folded position, a deployment torque adapted to bring the flexible rod back to the deployed position.

Abstract: A multi-angle imager (10) comprises an imaging array (Mij) configured to receive light beams (Li) via one or more entrance pupils (A1) according to distinct fields of view (Vi) of an object (P0) along each of multiple entry angles (?i). The imaging array (Mij) comprises multiple imaging branches (M1j, M2j) configured to form respective optical paths for the light beams (L1, L2) through the imager (10) for imaging respective subsections (S1, S2) of the object (P0). Each imaging branch (M1j) comprises a distinct set of optical elements (M11, M21) configured to receive the respective light beam (L1) along the respective entry angle (?1) and redirect the respective light beam (L1) towards the imaging plane (P1). The light beams (L1, L2) from each of the multiple imaging branches (M1j, M2j) are redirected to travel in a common direction (y) between the imaging array (Mij) and the imaging plane (P1).

Abstract: A multi-angle imager (10) comprises an imaging array (Mij) configured to receive light beams (Li) via one or more entrance pupils (A1) according to distinct fields of view (Vi) of an object (P0) along each of multiple entry angles (?i). The imaging array (Mij) comprises multiple imaging branches (M1j, M2j) configured to form respective optical paths for the light beams (L1,L2) through the imager (10) for imaging respective subsections (S1,S2) of the object (P0). Each imaging branch (M1j) comprises a distinct set of optical elements (M11,M21) configured to receive the respective light beam (L1) along the respective entry angle (?1) and redirect the respective light beam (L1) towards the imaging plane (P1). The light beams (L 1,L2) from each of the multiple imaging branches (M1j ,M2j) are redirected to travel in a common direction (y) between the imaging array (Mij) and the imaging plane (P1).

Abstract: A spacecraft (10), comprising a body (12), a solar array (30) with a support panel (32) which is connected to the body, and a thermal radiator (50) that is connected to the body and which includes a radiator substrate (52) that is thermally coupled to the body via at least one heat link (64). The solar array and thermal radiator are configured to be transitioned from a stowed state wherein the support panel and the radiator substrate are held fixed in an overlapping arrangement along and near the body, to a deployed state wherein the solar array is unfolded with the support panel positioned at a distance from the body and the radiator substrate is folded away from the body and the solar array. Preferably, the solar array and thermal radiator are flexible, to allow them to be kept in an overlapping and temporarily bent shape in the stowed state.

Abstract: A solar array (50) for a spacecraft (10), comprising a solar concentrator that is provided with photovoltaic cells and reflective areas configured for reflecting solar radiation towards the photovoltaic cells, wherein the reflective areas and the photovoltaic cells are provided on opposite surfaces of concentrator reflector sheet members (56) that are repositionable from a retracted state wherein the concentrator reflector sheet members are in a substantially flat arrangement, to a extended state wherein the concentrator reflector sheet members are raised to allow the reflective areas to reflect solar radiation towards the exposed photovoltaic cells. Alternatively or in addition, the solar array may comprise a support panel, which may be at least partially flexible for retaining the support panel in a bent panel shape when the solar array is in the stowed state fixed at a position near a body of the spacecraft.

Abstract: A radiator comprising at least one heat conductive layer having an in-plane heat conductivity of at least 500 W/mK and at least one heat emission layer in contact with the heat conductive layer, wherein the emission layer has an exposed surface with an emissivity of at least 0.7. Preferably, the heat conductive layer comprises a carbon-based material. The radiator is to be used in combination with a space vehicle structure, and due to its flexible character may be conformed to the particular shapes of such structure.

Abstract: A radiator, comprising at least one heat conductive layer with pyrolytic graphite material and an in-plane heat conductivity of at least 500 W/m·K. The radiator further comprises at least one heat emission layer that is in contact with the heat conductive layer, wherein the emission layer has an exposed surface with an emissivity of at least 0.7. The radiator is to be used in combination with a space vehicle structure, and due to its flexible character may be conformed to the particular shapes of such structure.

Abstract: A spacecraft (10), comprising a body (12), a solar array (30) with a support panel (32) which is connected to the body, and a thermal radiator (50) that is connected to the body and which includes a radiator substrate (52) that is thermally coupled to the body via at least one heat link (64). The solar array and thermal radiator are configured to be transitioned from a stowed state wherein the support panel and the radiator substrate are held fixed in an overlapping arrangement along and near the body, to a deployed state wherein the solar array is unfolded with the support panel positioned at a distance from the body and the radiator substrate is folded away from the body and the solar array. Preferably, the solar array and thermal radiator are flexible, to allow them to be kept in an overlapping and temporarily bent shape in the stowed state.

Abstract: A solar array (50) for a spacecraft (10), comprising a solar concentrator that is provided with photovoltaic cells and reflective areas configured for reflecting solar radiation towards the photovoltaic cells, wherein the reflective areas and the photovoltaic cells are provided on opposite surfaces of concentrator reflector sheet members (56) that are repositionable from a retracted state wherein the concentrator reflector sheet members are in a substantially flat arrangement, to a extended state wherein the concentrator reflector sheet members are raised to allow the reflective areas to reflect solar radiation towards the exposed photovoltaic cells. Alternatively or in addition, the solar array may comprise a support panel, which may be at least partially flexible for retaining the support panel in a bent panel shape when the solar array is in the stowed state fixed at a position near a body of the spacecraft.