The Dual Spacecraft Deployment System (DSDS) is a concept developed for deploying Dual Satellites into independent orbits from a single launch vehicle. The Deployment Concept, Deployment System and Spacecraft Design with Representative Instruments are presented herein. The design employs an innovative half-hex design for the spacecraft buses which allows the spacecraft to be accommodated in a parallel mounted configuration on the launch vehicle. The Spacecraft are arranged on the launch vehicle with a central deployment structure with a deployment mechanism which allows independent lateral deployment of two spacecraft. The Spacecraft Design can have a large volume, e.g., sized to accommodate Spacecraft Sub-systems and Instrument Systems. This concept allows the launch vehicle to accommodate two large Spacecraft deployed to different orbits with a single launch doubling the efficiency of the launch vehicle.

According to the invention, a link is established between the fixed part (4A) and the removable part (4B) of an enclosure (4) in which the second satellite is previously enclosed, said link being such that said removable part (4B) remains attached to said fixed part (4A) when said removable part (4B), initially in a first position for which said enclosure (4) is closed and encloses said second satellite, lies in a second position for which said enclosure (4) is open and able to release said satellite.

A multichannel actuation system for a launch vehicle. The multichannel actuation system includes a control unit operably connected to a single interface unit, and a plurality of individual hold-and-release mechanisms (HRMs) configured for releasably holding a plurality of payloads to the launch vehicle. Each of the plurality of HRMs is operably connected to the interface unit. The interface unit is configured for controlling actuation of the plurality of HRMs based on release operation instructions received from the control unit.

A satellite apparatus is disclosed, including a housing having first and second opposing walls, and a support structure inside the housing spanning the first and second walls. The support structure is structurally connected to the housing only at the first and second walls, and an end portion of the support structure is configured for connection to a launch vehicle by a separation system.

A nanosatellite configured for launching from a cube satellite launcher is disclosed. The nanosatellite has a lower portion sized to the fit within the launcher housing, and has top, bottom, left, right, front and rear sides forming a payload housing. The top side has a top surface configured to engage with a front end of the spring of the launcher. The nanosatellite has an upper portion extending from the top surface of the lower portion, and is sized to fit within the spring of the launcher. A control subsystem contained within the upper portion. The nanosatellite may have one or more a solar panel arrays connected to the lower portion. In some embodiments, the solar arrays are movable between a stowed position against the lower portion and a deployed position extending outwardly from the lower portion. An additional 1U-3U of payload may be connected to the lower portion.