Patent Application
20150083865
This invention relates generally to a spacecraft launch system and method.
Many modern satellites are designed to be deployed in a Geostationary Earth Orbit (GEO), rather than a Lower Earth Orbit (LEO). A GEO is a higher-Earth orbit and the cost of launching a satellite into a GEO (or other higher-Earth orbits such as Medium Earth Orbit and Highly Elliptical Orbit) is significantly higher than launching into an LEO. To reduce the launch costs, a satellite may instead be launched into a much lower parking or transfer orbit and then moved to a higher-Earth orbit using a propulsion system incorporated into the satellite. A solar electric propulsion thruster system is now commonly used in such satellites, which typically includes solar arrays, at least one energy storage device, a propellant fuel storage tank, control electronics and a thruster engine. Examples of solar electric propulsion thruster systems include, for example, a Xenon ion propulsion thruster, a Hall Effect thruster, an ion thruster, a pulsed induction thruster, a FARAD, and a VASIMR. The traditional propulsion system required in a satellite (or other type of spacecraft) necessary for movement from a parking or transfer orbit to a higher-Earth orbit is significantly larger and consequently heavier and more expensive than the propulsion systems included in satellites launched directly into a higher-Earth orbit since such systems are used only for maintaining orbit and for orbit correction.
The present disclosure is addressed to a system and method for propelling spacecraft. The system includes a common base stage, an electrical propulsion system mounted on the base stage, and one or more spacecraft couplers mounted on the base stage. Each of the spacecraft couplers is configured to securedly attach a spacecraft to the base stage. Each spacecraft includes an internal power source. Each spacecraft coupler preferably includes an electrical connection for coupling the internal power source to the electrical propulsion system. Each electrical connection may also be configured to transfer control signals between a controller within the associated spacecraft and a controller coupled to the electrical propulsion system. The internal power source may comprise at least one solar collecting component and/or at least one battery. A power regulation circuit may be coupled between the electrical propulsion system and each internal power source. The power regulation circuit is preferably configured to draw an equal and proportional amount of power from each spacecraft. Each spacecraft may be a satellite and the electrical propulsion system may be configured to propel the base stage and attached satellites from a lower-Earth orbit to a higher-Earth orbit. Each satellite preferably includes an associated electrical propulsion system that is only capable of providing propulsion for orbit maintenance and maneuvering and is not capable of providing propulsion for orbit raising from a lower-Earth orbit to a higher-Earth orbit.
In a further embodiment, the system also includes a non-spacecraft coupler mounted on the base stage which is configured to securedly attach a non-spacecraft storage container to the base stage.
In a still further embodiment, the system also includes a spacecraft portion permanently affixed to the base stage. The spacecraft portion may be a satellite portion.
According to the method for propelling a spacecraft, a plurality of spacecraft are securedly attached to a base stage having an electrical propulsion system mounted thereon. An electrical power source in each of the plurality of spacecraft is coupled to the electrical propulsion system. The electrical propulsion system is operated to propel the base stage and attached spacecraft using electrical power from each electrical power source. Further, a controller within the associated spacecraft may be coupled to a controller coupled to the electrical propulsion system in the base stage, to transfer control signals between the controller within associated spacecraft and the controller coupled to the electrical propulsion system. Still further, a power regulation circuit may be coupled between the electrical propulsion system and each internal power source. The power regulation circuit may be configured to draw an equal and proportional amount of power from each spacecraft. In the method, each spacecraft may be a satellite, with the electrical propulsion system is configured to propel the base stage and attached satellite from a lower-Earth orbit to a higher-Earth orbit. Each satellite may include an associated electrical propulsion system that is only capable of providing propulsion for orbit maintenance and maneuvering and which is not capable of providing propulsion for orbit raising from a lower-Earth orbit to a higher-Earth orbit.
The following detailed description, given by way of example and not intended to limit the present invention solely thereto, will best be understood in conjunction with the accompanying drawings in which:
In the present disclosure, like reference numbers refer to like elements throughout the drawings, which illustrate various exemplary embodiments of the present invention. The embodiments disclosed herein provide a spacecraft launch system for moving a plurality of spacecraft (e.g., satellites) and/or non-spacecraft storage containers from a lower parking or transfer orbit to a higher-Earth orbit. The parking or transfer orbit may be an LEO or may simply be any desired orbit lower than a higher-Earth orbit. The system employs a common solar electric propulsion stage (base) that mates with a plurality of spacecraft and receives electrical power from the spacecraft (e.g., generated by solar collecting components mounted on such spacecraft). The solar collecting components may be solar panels. As explained in more detail below, the common propulsion stage includes a solar electric engine and associated propellant storage tank of the type required for the orbit-raising operation (i.e., the movement from the lower parking or transfer orbit to the higher-Earth orbit). Using a common propulsion stage eliminates the need for such costly parts on each spacecraft. Instead, each spacecraft will only require a smaller, lighter and much less expensive solar electric engine and associated propellant tank used only in orbit maintenance and maneuvering.
Referring now to the drawings and in particular to
After system 100 is launched into the lower parking or transfer orbit on a launching rocket, it separates from the launching rocket and may perform an orbit-raising operation (i.e., the transition from the lower parking or transfer orbit to a desired higher-Earth orbit). During orbit raising, each satellite 101, 102, 103 may provide electrical power to the common propulsion engine 141 on base stage 104 (via the solar panels 111, 112, 113 and internal batteries in each satellite). In this manner, the solar panels 111, 112, 113 for each satellite 101, 102, 103 (and an associated internal power regulation/control system and internal battery) may be deployed to provide one-third of the electrical power necessary for the common propulsion engine 141 (since each satellite provides a proportional portion of the power needed for base stage 104). As discussed below, the inclusion of three satellites 101, 102, 103 on base stage 104 is merely exemplary and one of ordinary skill in the art will readily recognize that the electrical power requirements supplied to base stage 104 from the satellite is a fractional proportion determined by the number of satellites mounted on base stage 104. The use of a common propulsion engine 141 eliminates the need for a larger propulsion engine and larger associated propellant storage tank for each satellite 101, 102, 103. The system 100, after deployment at the lower parking or transfer orbit, moves up to a position near to the desired higher-Earth orbit, and then each satellite 101, 102, 103 is detached and moved into the final desired orbit. By eliminating the propulsion engine and associated propellant storage tank (sized for an orbit raising operation) from each satellite, significant cost-savings and weight-savings can be achieved for each satellite.
As one of ordinary skill in the art will readily recognize, the number of satellites included on base stage 104 is an arbitrary design choice. The benefits provided by the embodiment of
Although the present invention has been particularly shown and described with reference to the preferred embodiments and various aspects thereof, it will be appreciated by those of ordinary skill in the art that various changes and modifications may be made without departing from the spirit and scope of the invention. It is intended that the appended claims be interpreted as including the embodiments described herein, the alternatives mentioned above, and all equivalents thereto.
