Abstract: Dynamic resource management systems that improve communication over a satellite communication link. The system comprises one or more personal computers coupled by way of a network to a local area network edge device. The local area network edge device communicates by way of a non-processing satellite with a gateway that interfaces to an internet service provider or corporate network. The non-processing satellite implements a bent pipe communications link between the local area network edge device and the gateway. The non-processing (bent pipe) satellite provides a fixed connectivity to an aggregation point on the ground (the gateway). The networking protocol supported by the local area network edge device is an Internet protocol (IP), such as 802.3 Ethernet, universal system bus (USB), or 802.11 Ethernet protocols, for example. Signaling in the system is based on non-asynchronous transfer mode (ATM) protocols.

Abstract: A satellite communication system includes a plurality of spacecraft in geosynchronous orbit, where each spacecraft provides a plurality of beams on the surface of the earth, and a plurality of ground stations individual ones of which are located in one of the beams for transmitting uplink signals to one of the spacecraft. Each spacecraft has a plurality of receivers for receiving a plurality of the uplinked signals from ground stations, a frequency translator for translating the received uplink signals to a transmission frequency of a plurality of downlink signals, and a plurality of transmitters for transmitting the plurality of downlink signals within the same beams as the corresponding uplink signals. Each transmitter includes a combiner for combining together a plurality of frequency translated signals and a power amplifier, such as a TWTA, for amplifying the combined plurality of frequency translated signals.

Abstract: Array fed multiple beam antenna systems and methods for use on an orbiting spacecraft carrying a communication system. The antenna systems include a reflector and a relatively small array feed compared to the reflector. The array feed has feeds (radiators) that illuminate the reflector and that are disposed in a focal plane of the reflector. A power division network excites the radiators of the feeds. The antenna system is capable of very wide scan angle operation and may be used to provide multiple spot beam coverage over the surface of the Earth viewed from a synchronous orbit spacecraft. Phase aberration normally associated with scanning is corrected by adjusting excitation coefficients of each array feed. The spot beams are scanned across a field of regard by controlling the position of each feed in the focal plane and using the appropriate amplitude and phase distribution associated with a particular spot beam (array feed). The amplitude and phase distributions may be fixed or varied during operation.

Abstract: A method of charging a nickel-hydrogen battery which includes a positive electrode, a negative electrode and an electrolyte, after the battery has been subject to discharge is disclosed the method comprising the steps of: (a) subjecting the battery to a high charge rate for a period until the battery temperature increases at least &Dgr;T1 over the local minimum battery temperature; (b) switching the battery charge level to a lower rate of charge than in step (a) for a period until the battery temperature increases at least &Dgr;T2 over the local minimum battery temperature; (c) switching the battery charge level to a much lower rate of charge than step (b) for a period up to about 30 to 60 minutes before the next scheduled eclipse; and thereafter (d) switching the battery to a pulse charge equivalent to the high charge rate of step (a).

Abstract: After discharge, a rechargeable battery having a nickel hydroxide positive electrode is charged at a temperature T1 between a proximately −10° C. and −30° C., lower than a temperature T2, in the range of approximately −15° C. to +5° C., at which discharge begins, automatically determining the total charge to be returned to the battery for establishing the ideal charge for the battery leading into the next discharge cycle, and applying charge energy to the battery having the magnitude as automatically determined. After discharge, the battery is cooled into the T1 range, then heated to stabilize the temperature to that in the T1 range. Then the battery is charged according to a nominal profile of charge power as a function of time, the accumulated charge imparted to the battery sensed as cooling proceeds, the nominal power profile adjusted according to the accumulated charge obtained, and the remainder of the nominal charge profile adjusted accordingly.

Abstract: A control system and method that keeps solar arrays on a satellite pointing toward the sun. Exemplary control systems and methods are implemented as follows. An orbit propagator computes satellite orbital location data and computes a sun vector referenced to an inertial reference frame. A coordinate transformation processor processes the sun vector in the inertial reference frame and a sensed satellite attitude signal to generate a sun vector referenced to a satellite body reference frame. A solar array position processor processes a step count derived from a solar array pointing control system and sensed solar array position to generate an estimate of the position of the solar array. A summing device sums the solar array position estimate output, the body frame gun vector, and a bias signal derived from a ground command, to produce a solar array position error signal. A filter filters the solar array position error signal.

Abstract: A magnetic assembly is constructed comprising a first core portion, a second core portion, and a winding assembly. The first and second core portions each include a base portion and a plurality of projections extending from the base portion. The winding includes a plurality of stacked layers having conductive paths applied to their surfaces. The winding is constructed with a bore, wherein each conductive path encircles the bore. A number of the stacked conductive layers form a primary winding, and a number of the stacked conductive layers form a secondary winding. The winding is disposed over the first core portion in a manner so that a projection engages the opening of the winding to provide magnetic coupling of winding and core. The second core portion is disposed over the first core portion to form a closed magnetic circuit through and around the winding.

Abstract: In a presently preferred embodiment of this invention a spacecraft includes at least four thermal radiator panels, with a first pair of said panels being initially stowed along a North facing surface of said spacecraft and a second pair of said panels being initially stowed along a South facing surface of said spacecraft. Each of said panels is coupled to a hinge mechanism at an anti-Earth facing end of said spacecraft and is rotatable by said hinge mechanism from a stowed position to a deployed position. The deployed position is a position that is off of the anti-Earth facing end of the spacecraft. In the presently preferred embodiment of this invention, when in the deployed position a longitudinal axis of each of said panels forms a non-zero angle with an Earth-pointing axis of said spacecraft.

Abstract: An improved thermal harness or thermal heat strap that comprises carbon based (graphite) fiber heat conducting elements coupled to graphite, composite, polymer, or metal end attachment brackets. An outer casing comprising a tubular braided sleeve surrounds the heat conducting elements and eliminates possible contamination caused by the graphite fibers. The thermal harness may be advantageously used in electronic applications to dissipate heat from high temperature components.

Abstract: Deployable antenna systems for use on a spacecraft that is moveable from a stowed position to a deployed position. Two embodiments are disclosed that include one or two sets of main reflector assembly (and optional adjustment mechanism) and subreflectors. The antenna systems include a feed horn assembly fixedly attached to the spacecraft and a rotatable hinge attached to the spacecraft. A substantially rigid reflector support structure is attached to the hinge and rotates about a hinge axis. The main reflector assembly (and optional adjustment mechanism) is attached to a lower portion of the support structure and the subreflector is attached to an upper portion of the support structure. The subreflector has a fixed relation relative to the main reflector assembly and has a fixed relation relative to the feed horn assembly when the antenna system is in the deployed position so that the antenna systems generate predetermined beam coverage patterns.

Abstract: A system for supplying a space station or satellite comprises a low cost expendable single-stage to orbit launch vehicle having a moderate reliability in the range of about 0.5 to 0.8 for launching a payload of consumable items to low earth orbit. For propulsion, the launch vehicle includes a pressure-fed rocket engine, preferably with a nozzle for aiming products of combustion in a direction away from the launch vehicle and having a fixed orientation relative to the launch vehicle. A pressurant tank contains an inert pressurant gas under pressure. A conduit system serves to introduce the inert pressurant gas from the pressurant tank to a fuel tank and thence to the rocket engine and also to introduce the inert pressurant gas from the pressurant tank to an oxidizer tank and thence to the rocket engine for combustion with the fuel.

Abstract: An electromagnetic cavity filter is formed by at least two cavities having electrically conductive walls. Each cavity is the equivalent of two filter poles because two orthogonal modes of electromagnetic radiation can resonate within each cavity. Characterizing vector tuning elements are coupled to each of the cavities that are each aligned along respective axes. The tuning elements are used to provoke derivative orthogonal modes and determine the degree of coupling between orthogonal modes. One or more intercavity couplers interconnect the cavities and are rotated at arbitrary angles that are different from the axes of the characterizing vector tuning elements. Electrically adjacent and nonadjacent modes of proximate cavities can be coupled, permitting elliptic filter functions. Electrically nonadjacent modes are coupled by means of an iris interconnecting the two cavities. Electrically adjacent modes are coupled by means of an electrically conductive probe penetrating each of the cavities.

Abstract: A novel method for producing improved RF transmission lines for satellite beamforming networks and printed circuit antenna(s) comprising the steps of (1) bonding together (a) a central conductor strip or trace comprising a dielectric layer or circuit board having on one or both surfaces thereof conductive strip circuitry; (b) upper and lower core layers of lightweight closed-cell plastic foam bonded to (a), and (c) upper and lower surface layers or faceskins of conductive metal foil or of dielectric material bonded to metal foil layers inside and/or outside, to form ground planes, bonded to the surfaces of the adjacent foam core layers.

Abstract: The invention resides in a method and apparatus for conditioning an inner surface of a tube. A pressurized mixture of air and grit and is supplied to a nozzle having an opening with a given diameter through which diameter the pressurized grit and air pass. An elongated nozzle adapter is connected to said nozzle. The nozzle adapter has a passage having a diameter substantially equal to the diameter of said nozzle and said nozzle adapter is connected to said nozzle such that the opening in said nozzle and said passage are aligned with one another. The nozzle adapter is inserted into a tube and is axially moved and rotated in order to condition the inner surface thereof with a pressurized mixture of air and grit.

Abstract: A method for estimating motion of a dual-spin spacecraft having a gimballed momentum wheel. The method disengages the gimbal from its drive train in anticipation of a short disturbance. Then, the gimbal slip resulting from the disturbance is measured. The method may be used to stabilize any dual-spin spacecraft that uses a gimballed momentum wheel. To stabilize the dual-spin spacecraft, the gimbal angle and gimbal rate are measured during and after the disturbance to provide an indication of the inertial spacecraft motion along gimbal axes. The magnitude and direction of the disturbance are determined by comparing motion of the gimbal before and after the disturbance. Then, torques are applied to the gimbal to counteract the spacecraft motion resulting from the disturbance.

Abstract: A charging sensor that uses photoemission to directly measure the electrical potential of a spacecraft in orbit relative to the space environment. The photoemission based charging sensor uses photoemission from a photoemitting conductive plate of a charge probe as a reference point for voltage readings and provides direct measurements of spacecraft charging.

Abstract: A control method and improved equatorial-normal body-stabilized spacecraft that operates in an inclined orbit. The control method controls pointing of payloads (communication antennas and instruments) and solar arrays disposed on the spacecraft. The method controls the spacecraft to provide Earth-pointing payloads and sun-pointing solar arrays so that the spacecraft body and solar arrays each rotate only around a single axis, a pitch axis of the spacecraft.

Abstract: An integrated transformer and inductor assembly for use in soft switching or resonant power converters, and the like. The assembly has a planar structure and includes a planar transformer and a parallel inductor. The assembly has a transformer core with a central gap. Planar interleaved primary and secondary winding are separated by insulating layers and are disposed within the transformer core. The parallel inductor is provided by a concentric inductor (reactive) winding disposed adjacent the center of the transformer core, which may be wound around a bobbin. The concentric inductor (reactive) winding carries inductor current, while load current flows mostly in the planar windings. Loss due to magnetizing current is substantially reduced in the present invention.

Abstract: A method for operating a pulse width modulated switching power converter (1), and a pulse width modulated switching power converter that is constructed to operate in accordance with the method. The power converter has an input coupled to an output of a solar array (10) and an output providing an output current, the output being coupled to battery (14) for applying a charging current ICHARGE to the battery while also supplying a current ILOAD to a load. The method includes the steps of: (a) incrementally increasing (18, C1) a duty cycle of the pulse width modulated switching power converter so as to incrementally increase a magnitude of the output current of the switching power converter; (b) sensing (14a) a first magnitude of both ICHARGE and ILOAD; (c) storing (20) the sensed magnitude; (d) sensing a second magnitude of both ICHARGE and ILOAD; and (e) comparing (22) the stored first magnitude to the sensed second magnitude.

Abstract: A method of controlling the orbit of an inclined, eccentric geosynchronous satellite by maintaining its perigee location while simultaneously preventing a reference point on the orbit from moving too far relative to a normal position over the Earth. Doing this changes the semimajor axis of the orbit, which when perturbed away from the geosynchronous value, creates a drift rate relative to the Earth. To correct the argument of perigee and maintain the prescribed tolerance on ground track position, multiple maneuvers must be performed to control the drift rate. The relationship between maximum allowable velocity change and time between maneuvers is defined by the required correction to the argument of perigee and the allowable motion of the reference point.