Space Systems/loral, LLC

Macdonald, Dettwiler and Associates Corporation

Other

John Douglas Lymer

Robert Edward Helmer

Tej Sachdev

Robert Erik Schwarz

Peter Krimbalis

Rangaswamy Ravindran

Richard Andrew Bryan

Branko Francetic

Paul Robert Peacock

Andrew Allen

Kerry Spring

Andrew E Turner

Roderick C Tennyson

Alfred Heikal Tadros

Raymond M Measures

W Dayle Hogg

Paul Anthony Briggs

Neil D Glossop

Peter Panagiotis Krimbalis

Techniques for performing spacecraft rendezvous and/or docking include operating a first orbiting spacecraft, the first spacecraft including a sensor arrangement, a first processor and a first inter-satellite link (ISL) arrangement and performing one or both of a rendezvous operation and a docking operation with the first spacecraft and a second orbiting spacecraft, the second spacecraft including one or more actuators. The performing one or both of the rendezvous operation and the docking operation includes determining a pose and pose rate of the second spacecraft relative to the first spacecraft using observations made by the sensor arrangement and determining a desired approach trajectory for the second spacecraft.

Spacecraft rendezvous and docking techniques

A spacecraft includes a plurality of deployable module elements, at least one of the deployable module elements including a robotic manipulator, the spacecraft being reconfigurable from a launch configuration to an on-orbit configuration. In the launch configuration, the deployable module elements are disposed in a launch vehicle in a first arrangement. In the on-orbit configuration, the deployable module elements are disposed in a second configuration. The spacecraft is self-assembled by the robotic manipulator reconfiguring the spacecraft from the launch configuration, through a transition configuration, to the on-orbit configuration. The deployable module elements may be in a stacked arrangement in the launch configuration and may be in a side-by-side arrangement in the on-orbit configuration.

Self-assembling persistent space platform

A spacecraft including a main body structure and at least a first deployable element is reconfigured from a launch configuration to an on-orbit configuration. In the launch configuration, the first deployable element is mechanically attached with the spacecraft main body structure by way of a first arrangement. In the on-orbit configuration, the first deployable element is mechanically attached with the spacecraft main body structure by way of a second arrangement. Reconfiguring the spacecraft includes detaching the first deployable element from the first arrangement, moving the first deployable element with respect to the spacecraft main body structure; and attaching the first deployable element to the second arrangement.

On-orbit assembly of communication satellites

There is disclosed a Payload Ejection System (PES) able to store any set of payloads for launch and eject that set of payloads at a controlled speed with a low tumble rate while accommodating any offset centre of mass within a restricted volume. The need for ballasting or balancing is eliminated thus freeing up the design space for these payloads. Insensitivity to centre of mass location is enabled by the use of a deployment hinge assembly arrangement which uses two or more non-parallel folding hinge arrangements that allow for linear motion of the output link in one direction while restricting the motion all other directions. One embodiment of the current PES concept uses four (4) hinge panel assemblies, selected to provide optimal stiffness around the entire mechanism. The stiffness of the PES is integral to managing offset centre-of-mass locations by allowing the mechanism to translate the effective force vector to the center of mass location.

Spacecraft payload ejection system

There is disclosed a Payload Ejection System (PES) able to store any set of payloads for launch and eject that set of payloads at a controlled speed with a low tumble rate while accommodating any offset center of mass within a restricted volume. The need for ballasting or balancing is eliminated thus freeing up the design space for these payloads. Insensitivity to center of mass location is enabled by the use of a deployment hinge assembly arrangement which uses two or more non-parallel folding hinge arrangements that allow for linear motion of the output link in one direction while restricting the motion all other directions. One embodiment of the current PES concept uses four (4) hinge panel assemblies, selected to provide optimal stiffness around the entire mechanism. The stiffness of the PES is integral to managing offset center-of-mass locations by allowing the mechanism to translate the effective force vector to the center of mass location.

Herein is disclosed a propellant transfer system and method for refueling on-orbit spacecraft. The system and method are configured to allow for resupply of spacecraft configured to be fueled by either a bipropellant (oxidizer and fuel) or a monopropellant (typically hydrazine). The system and method are particularly suited for resupply of satellites not originally prepared for refueling as well but the system may also be used for as satellites specifically designed for refueling.

Propellant transfer system and method for resupply of fluid propellant to on-orbit spacecraft

Propellant transfer system and method for resupply of propellant to on-orbit spacecraft

A damage detection and evaluation system utilizes optical fibers to detect and assess damage to a structure. The optical fibers are intermittently etched in a controlled manner to weaken them so they fracture when a desired load is applied to the structure. Where visual inspection of the optical fibers is used to check for damage, a reference set of optical fibers is used. Where photodetection is used and minimal connections to the structure are required, one end of the optical fiber is made reflecting so an optical connection is only made to the other end. To further simplify the connection, a single input/output optical connection to a demultiplexing/multiplexing chip mounted on or within the structure is provided. Where orthogonal grids of fibers are not possible, the invention provides interdigitated optical fibers and volume backscattering to locate the position of the first fracture along the length of the optical fiber. The entire periphery of a damaged area can then be determined with a set of such interdigitated optical fibers. Again, a single optical interface to the structure can be provided using a demultiplexing/multiplexing chip. In the case of composite materials, the optical fibers are located between specified plys and at a specified orientation for optimum sensitivity.

Damage evaluation system and method using optical fibers

Growing community of inventors

Toronto, Canada

John Douglas Lymer