Space Systems/loral, LLC

Maxar Space LLC

Lenny Low

Gordon C Wu

Randy Pon

Susan Sebata

David Marlow

David Younghee Oh

Joel Boccio

Harry A Yates

Jorge Delgado

Maria Eugenia Torres

Chris Goodman

A spacecraft includes a semi-rigid solar array and a main body structure, the main body structure configured as a convex polyhedron and including an aft and a forward face disposed opposite to the aft face and at least four side faces disposed between and approximately orthogonal to the aft face and the forward face. The solar array includes a number of panels linked together with flexible couplings. In an undeployed configuration, panels of the solar array cover at least two adjacent side faces, the flexible couplings providing an articulable joint approximately aligned with a line along which the two adjacent side faces are joined and connecting a first panel of the solar array and a second panel of the solar array, the first panel being proximal to a first side face and the second panel being proximal to a second side face.

Spacecraft design with semi-rigid solar array

A spacecraft includes a propulsion subsystem including at least two electric thrusters, an electrical interface assembly that couples electrical conductors from the thrusters to a spacecraft harness, a pneumatic interface assembly that controls flow rate of propellant to the thrusters and a thruster support module (TSM) including a pointing arrangement and a mounting arrangement. A proximal portion of the mounting arrangement is coupled with a distal portion of the pointing arrangement; the at least two electric thrusters are disposed on a distal portion of the mounting arrangement; the electrical interface assembly and the pneumatic interface assembly are disposed on the proximal portion of the mounting arrangement. The mounting arrangement is configured to limit heat transfer between the thrusters and (b) one or more of the proximal portion of the mounting arrangement, the electrical interface assembly and the pneumatic interface assembly.

Multiple thruster firing on a single articulable module

Heat transfer systems and methods 30 for use on a spacecraft that use a loop heat pipe to transfer heat from a remotely located heat source to a thermal radiator or other heat dissipating apparatus. The loop heat pipe is a two phase heat transfer device that has an evaporator coupled to the heat source and a condenser coupled to the thermal radiator or other heat dissipating apparatus. The loop heat pipe comprises thin walled tubing to connect the evaporator and condenser. The thin walled tubing allows the loop heat pipe to be flexible. This flexibility offers significant advantages in terms of routing and accommodating design

Heat transfer of a remote heat source using a loop heat pipe

A spacecraft heat dissipation method, and a spacecraft having an improved thermal radiator system that uses two-sided deployable thermal radiators that dissipates heat from both front and back surfaces thereof. The use of the two-sided deployable thermal radiators enables the thermal radiator system to have approximately 50% more heat dissipating capability than a system with just one surface exposed to dissipate heat. The spacecraft includes a body, one or more solar arrays, and the present radiator system which comprises opposite facing fixed payload radiators that are thermally coupled to selected ones of the deployable radiators by way of heat pipes. In an exemplary method a spacecraft is configured to have a body, one or more solar arrays opposite facing fixed payload radiators, and one or more two-sided deployable radiators selectively coupled to the fixed payload radiators. The spacecraft is launched into orbit. When in orbit, heat coupled to the opposite facing fixed payload radiators is transferred to the two-sided deployable radiators for radiation into space from both sides thereof.

Two-sided deployable thermal radiator system and method

Heat dissipating apparatus or system for dissipating heat generated by a payload on a spacecraft having a plurality of surfaces. The system includes at least one radiator-condenser coupled to one or more of the surfaces of the spacecraft and a heat pump including an evaporator, a compressor, and an expansion valve coupled in a closed-loop manner to the radiator-condenser. The system components are coupled together using small diameter, thin- and smooth-walled tubing. The system enables the radiator-condenser to be elevated above the source or payload temperature, and, along with the use of thin walled tubing, reduces the mass of the spacecraft. Because the radiator-condenser temperatures are elevated, multiple surfaces of the spacecraft (west, east, earth and aft) can be effectively used as radiating surfaces.

Spacecraft radiator system using a heat pump

A spacecraft, along with a spacecraft radiator system and spacecraft heat dissipation method are disclosed. The spacecraft comprises a body, a plurality of solar arrays, and the spacecraft radiator system. The spacecraft radiator system comprises first and second opposite facing payload radiators, first and second opposite facing deployable radiators, and one or more coupling or loop heat pipes cross coupling opposite facing payload and deployable radiators so that they function in tandem. By cross-coupling the opposite facing payload and deployable radiators, one of the two radiators acting in tandem is always in the shade during solstice seasons. Consequently, the solar load processed by the radiator system is minimized, thereby, increasing the thermal dissipation capability of the radiator system by approximately 15%.

Spacecraft radiator system and method using cross-coupled deployable thermal radiators

Heat dissipating apparatus for use in dissipating heat produced by electronic components of a spacecraft. The heat dissipating apparatus comprises first and second radiator panels each having an outer panel faceskin 18 and inner panel faceskin. A heat pipe matrix is disposed between outer and inner panel faceskins of each radiator panel, and a honeycomb core is disposed between the inner and outer panel faceskins that surrounds the heat pipe matrix. A transverse panel interconnects the first and second radiator panels. A plurality of crossing heat pipes extend between and thermally couple to the heat pipe matrices of the first and second radiator panels. The plurality of crossing heat pipes extend outside the plane of the transverse panel.

Spacecraft radiator system using crossing heat pipes

Systems and methods that employ a phase change material to provide thermal control of electric propulsion devices (thrusters). A spacecraft is configured to have an electric propulsion thruster. The electric propulsion thruster is surrounded with a phase change material. Suitable phase change materials include high-density polyethylene (HDPE), waxes, paraffin materials, and eutectic salts. The spacecraft is launched into orbit. The electric propulsion thruster is fired for a predetermined period of time. Heat generated by the electric propulsion thruster is absorbed and stored in the phase change material while the thruster is firing. The stored heat is dissipated into space after the thruster has stopped firing.

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Hillsborough, CA, United States of America

Lenny Low