Technology is disclosed herein for a payload dispensing hinge assembly. The payload dispensing hinge assembly has a first hinge-half and a second hinge-half that are joined by a hinge pin. The first hinge-half may be connected to a payload that is to be dispensed at a target angle. The second hinge-half may be connected to a payload base. The first hinge-half may have a first mounting bracket and a rotatable arm that are shaped to form interlocks that serve to dis-engageably link these two components. A biasing mechanism rotates the rotatable arm and hence the first mounting bracket and payload about a hinge line. The second hinge-half has a hinge stop that stops the rotation of the rotatable arm at a target angle, whereby the first bracket dis-engages from the rotatable arm to dispense the payload at the target angle.

A spacecraft system may include a storage portion (e.g., a first portion and a second portion) and a solar array apparatus that may be configurable in at least a stowed configuration and a deployed configuration. The solar array apparatus may include at least one solar array to collect incident radiation when the solar array apparatus is in the deployed configuration. In one or more embodiments, the at least one solar array may extend away from the storage portion. In one or more embodiments, the at least one solar array may extend between the first portion and the second portion. The solar array apparatus may also include an extendable boom operable to extend the at least one solar array apparatus from the stowed configuration to the deployed configuration.

Technology is disclosed herein for a payload dispensing hinge assembly. The payload dispensing hinge assembly has a first hinge-half and a second hinge-half that are joined by a hinge pin. The first hinge-half may be connected to a payload that is to be dispensed at a target angle. The second hinge-half may be connected to a payload base. The first hinge-half may have a first mounting bracket and a rotatable arm that are shaped to form interlocks that serve to dis-engageably link these two components. A biasing mechanism rotates the rotatable arm and hence the first mounting bracket and payload about a hinge line. The second hinge-half has a hinge stop that stops the rotation of the rotatable arm at a target angle, whereby the first bracket dis-engages from the rotatable arm to dispense the payload at the target angle.

A launch system comprises a nose section comprising a nose coupling surface, a tail section comprising a tail coupling surface facing the nose coupling surface and a mast coupling the nose and tail sections. The mast is configured to expand and retract to displace the nose and tail sections within a range of distances from one another. In a retracted state, the nose and tail sections are either structurally coupled to one another at the nose and tail coupling surfaces or structurally coupled to at least one integrated module located between the nose and the tail sections.

An exemplary starshade comprises a tensegrity truss structure having a central hub with radially extending, telescoping booms. Telescoping tension struts connected to the central hub and booms provide a compressive force on the booms during final truss deployment. Opaque petals, not supported by the tensegrity truss structure prior to its final deployment, are each sequentially placed on and attached to the tensegrity truss structure in side by side position to form a concentric ring of petals spaced apart from the central hub. A fan fold covering, not supported by the tensegrity truss structure prior to its final deployment, is placed on and attached to the tensegrity truss structure to form an opaque, concentric inner ring about the central hub. An outer edge of the inner ring is adjacent an interior edge of the concentric ring of petals to block light from the petals to the central hub.

A radiator for a satellite intended for being placed in geostationary orbit around the earth in a tilted plane relative to the plane of the ecliptic, includes at least one panel having at least one radiative surface, and including: a mounting foot supporting the panel; and control and motor elements for pivoting the mounting foot about an axis of rotation tilted relative to the radiative surface which is perpendicular to a radiation axis, the radiation axis and the axis of rotation being tilted relative to one another by a non-zero working angle, corresponding to the tilt angle of the plane of the orbit of the satellite relative to the plane of the ecliptic, the working angle being fixed, such that for any rotation of the mounting foot about the axis of rotation owing to the control and motor elements, the radiative surface remains parallel to the plane of the ecliptic.

A method to control a sunlight acquisition phase of a spacecraft with a nonzero angular momentum of an axis D.sub.H. The spacecraft includes a solar generator configured to rotate about an axis Y. The spacecraft actuators are controlled to place the spacecraft in an intermediate orientation in which the axis Y is substantially orthogonal to the axis D.sub.H. The solar generator is controlled to orientate the solar generator towards the sun. The spacecraft actuators are controlled to reduce the angular momentum of the spacecraft. The actuators of the spacecraft engine are controlled to place the spacecraft in an acquisition orientation in which the axis Y is substantially orthogonal to the direction of the sun with respect to the spacecraft.

A hinge assembly comprises first and second tape spring elements, wherein each of the spring elements is configured to connect a first element of a space structure to a second element of the space structure. Each of the first and the second tape spring elements is movable from a folded state into an unfolded state by releasing stored strain energy, to deploy the first and the second element of the space structure. The first tape spring element is connected to a first direct current source and configured to conduct direct current of a first polarity supplied to the first tape spring element from the first direct current source. The second tape spring element is connected to a second direct current source and configured to conduct direct current of a second polarity, opposite to the first polarity, supplied to the second tape spring element from the second direct current source.

An active tie rod for retaining and releasing appendages without shock is provided. The device comprises: a fixed base, a tie rod extending along an axis between two ends, a mechanism fastened to the base and temporarily retaining a first end of the tie rod, operating the mechanism to release the tie rod from the base, an end piece fastened to a second end of the tie rod, at least one appendage temporarily retained between the base and the end piece, a component disposed between the mechanism and the end piece, and an actuator of the component to modify a characteristic dimension of the component along the axis between two values, the component producing a tension in the tie rod for the first of the two values and the tension in the tie rod being reduced for a second of the two values.

A side-by-side dual-launch spacecraft arrangement is provided. The arrangement may include a dual-launch adaptor, a first spacecraft, and a second spacecraft. The first spacecraft and the second spacecraft may be mounted on the dual-launch adaptor and may be arranged side by side on the dual-launch adaptor. An aspect ratio of each of the first and second spacecraft may be within a range of 0.55 and 0.8.