Abstract: The present invention provides a system (200) for determining the ephemeris and attitude for a spacecraft based on optical payload pointing directions and the ephemeris of other spacecraft. An ephemeris determining subsystem (300) obtains ephemeris data (308) for one or more reference spacecraft. Optical payload pointing information (306) is obtained for the optical payload pointed to the reference spacecraft. The spacecraft ephemeris (312) is then calculated with ephemeris determination algorithms (302) based on the ephemeris data (308) and optical payload pointing information (306). Other available information, such as stored results of previous measurements and calculations, time (310), spacecraft attitude (304), information from non-payload sensors and ranging information may be used to enhance the accuracy or decrease the complexity of the ephemeris calculations. Spacecraft attitude can also be determined using reference spacecraft ephemeris data and optical payload pointing information.

Abstract: A relatively high average power optical laser utilizes a diode pumped multi-mode dual-clad fiber amplifier configured for double pass amplification to provide a relatively higher average power than single mode fiber amplifiers. In order to compensate for mode scrambling and depolarization from the multi-mode fiber amplifier, the output beam from the multi-mode fiber amplifier is applied to a vector phase conjugator after a single pass through the amplifier and reflected back into the amplifier. The vector phase conjugator conjugates the electric field amplitude, phase and polarization and reflects a conjugate wave back into the multi-mode fiber amplifier in a reverse direction. The propagation of the conjugate wave through the amplifier unscrambles mode mixing and restores the beam back to its original polarization state. A Faraday rotator and polarization beam splitter are used to outcouple an output beam after the second pass through the multi-mode fiber amplifier.

Abstract: A motion control apparatus is provided for controlling oscillations of an appendage 10 coupled to a spacecraft 12. A plurality of guidelines 20 extend between remote portions 22 of the appendage 10 and the appendage base 14. A plurality of motion control elements 28 are interposed between the appendage base 14 and the spacecraft 12 and/or along the guidelines 20 between the appendage base 14 and the remote portions 22 of the appendage 10 or both. The motion control elements 28 include visco-elastic fluid, electro-rheological fluid, magnetic-rheological fluid, or a piezo-electric stack so that a motion control property of the motion control elements 28 may be varied according to the magnitude of an electric or magnetic field source 30 coupled thereto.

Abstract: A compact delay stage for phase delaying a light signal to provide increased amounts of phase delay to a light signal. A plurality of reflective structures are configured to provide a delay path to the light signal. One of the reflective structures is configured to accept the light signal and direct the light signal between the reflective structures a predetermined selectable number of times until the light signal has completed the delay path thereby translating the light signal into a delay signal. One of the reflective structures is configured to provide egress to the delayed signal to exit the compact delay stage.

Abstract: A thruster rocket engine throat insert (12) has a thin walled shell (54) made from a high strength, oxidation resistant material. The shell (54) having a throat (48) of reduced cross-section and a radially extending annular stiffening ring (60) located at the throat (48). A casing (56) made from a material having a thermal conductivity at least 10 times greater than that of shell (54) is molded around an outer surface (58) of shell (54) and has a generally cylindrical exterior surface (59). Shell (54) resists yielding and oxidation caused by the extreme temperatures of rocket fuel combustion products passing through the throat insert (12), while the casing (56) acts to efficiently transfer heat from the shell (54).

Abstract: A rocket propulsion system for spacecraft achieves greater economy, reliability and efficiency rocket by incorporating monopropellant RCS thrusters (1a-1f) for attitude control and bipropellant SCAT thrusters (5a-5c) for velocity control. Both sets of thrusters are designed to use the same liquid fuel, supplied by a pressurized non-pressure regulated tank, and operate in the blow down mode. In the propulsion system such station keeping and attitude control thrusters may function in conjunction with a large thrust apogee kick engine, which may also be of the SCAT thruster construction, that uses the same propellent fuel. Hydrazine and Binitrogen tetroxide are preferred as the fuel and oxidizer, respectively. The new system offers a simple conversion of existing monopropellant systems to a high performance bipropellant dual mode system without the extreme complexity and cost attendant to a binitrogen tetroxide--hydrazine bipropellant system.

Abstract: The present invention provides a method for custom coding uplink signals and downlink beams in a satellite communications system. The method includes the step of applying, at a ground station (100), an outer code (102) to an uplink signal to produce an outer coded uplink signal. Next, the ground station (100) applies a selected reduced complexity inner code (104) to the outer coded uplink signal to produce a concatenated coded uplink signal. The ground station (100) transmits the concatenated coded uplink signal to a satellite (110). At the satellite (110), the inner code of the concatenated coded uplink signal is decoded (116) to produce an outer coded satellite data stream. Next, the satellite (110) applies a selected inner code (120) to a portion of the outer coded satellite data stream to produce a concatenated coded satellite data stream. The satellite (110) then transmits the concatenated coded satellite data stream in a downlink beam to a destination ground station (126).

Abstract: A circuit module construction and method for its fabrication, in which radio-frequency (RF) or millimeter-wave circuit components (10) are electromagnetically isolated in a circuit module (12), to allow for location of multiple circuit components in close proximity without concern for signal loss or interference between components. Multiple RF or millimeter-wave circuit components (10) are installed on a dielectric substrate (14) and are separated by at least one metal isolation wall (20), which extends in depth all the way through the dielectric substrate (14) to a metal layer (16) formed under the substrate. Each isolation wall (20) is formed by first cutting a channel through the dielectric substrate (14), preferably using a laser that selectively removes the dielectric material but not the metal layer (16). Then the channel is filled with metal by electroplating, to provide continuous electromagnetic isolation in a lateral direction, parallel to the plane of the substrate (14).

Abstract: A device, and a method for its operation, for verifying a person's identity by comparing in a correlator (18) selected features of a reference fingerprint image, generated in a sensor (10) from the fingerprint of a person of known identity, with a subject fingerprint image taken from a person whose identity is to be verified. In an enrollment procedure, a fingerprint from a person of known identity is analyzed in an enrollment processor (14) to locate multiple reference patches that together are distinctive to that person's fingerprint. The reference patch images are stored, together with their locations in the image, in a reference image storage unit (16). When a subject fingerprint image is later provided, every reference patch is compared with every possible patch of similar size in the subject image, to find a set of candidate match locations in the subject image.

Abstract: There is provided a technique for removing unwanted heat from a heat generating device 12 operating aboard a space vehicle orbiting in outer space by placing a heat conductive interface 10 between the device and a support platform 14. The interface 10 comprises an array of spaced apart strips 16 made of aluminum aligned in spaced apart relationship so that the array covers the heat dissipating surface 20 at the underside of the device, the strips 16 are configured by deformable protuberances, such corrugations 22 which flatten under compression forces forming a heat conductive interface between the device 12 and the platform 14.

Abstract: The present invention provides a system (200) for determining the ephemeris of a space vehicle based on distances to and ephemeris of other spacecraft. An ephemeris determining subsystem (300) obtains ephemeris data (308) for one or more reference spacecraft. Distance information (306) is obtained between the space vehicle and the reference spacecraft. The spacecraft ephemeris (312) is then calculated with ephemeris determination algorithms (302) based on the ephemeris data (308) and distance information (306). Other available information, such as stored results of previous measurements and calculations, time (310), information from other on-board sensors and attitude information may be used to enhance the accuracy or decrease the complexity of the ephemeris determination.

Abstract: A length and velocity measurement apparatus based upon an interference pattern created between the amplitude modulation component of a transmitted signal, generated by an RF/microwave modulated continuous wave (CW) laser, and the amplitude modulation component of a signal reflected back from the target. The distance between the system and the target is measured by counting the periods in the interference pattern created between the amplitude modulation component of the transmitted signal and the amplitude modulation component of the reflected signal while the modulation frequency is swept from a frequency F.sub.1 to a frequency F.sub.2. In order to provide a fine adjustment of the apparatus, the apparatus may be mounted on a movable carriage to adjust the periods in the interference pattern to an integral number. Velocity measurements are made at a single frequency based upon the Doppler effect induced by the moving target.

Abstract: A propulsion system and, in particular, a microthruster for a microsatellite. The microthruster may be formed as a resistojet type thruster formed with a chamber, closed by a diaphragm which acts as a blow-out disk. A fluid, such as an inert gas, is disposed within the chamber. Heating of the gas causes the gas pressure to increase until the diaphragm ruptures, which, in turn, causes the gas to flow out of the chamber, acting as a propellant, and providing a small unit force. The microthruster is adapted to be formed by known batch processing methods with 10.sup.4 -10.sup.6 microthrusters per wafer. The unit of force can be easily scaled by varying the number of microthruster 20 used for an application, the geometry of the chamber 22, as well as the type of fluid used within the chamber, to suit the microsatellite application.

Abstract: An antenna nulling system (28) for nulling a jamming signal having a multibeam antenna (48), a correlator (72), and antenna pattern calculator (94), a sequential updater (90) and a beamformer (70) is provided. The multibeam antenna (48) includes a plurality of antenna elements (50) and is operable to receive the plurality of signals. The correlator (72) is operable to receive at least one sample signal from one of the antenna elements (50) and a composite signal from the plurality of antenna elements (50). The correlator (72) determines a cross-correlation of the sample signal and the composite signal. The antenna pattern calculator (94) calculates a difference in pattern magnitude of an adapted antenna pattern and a quiescent antenna pattern of the multibeam antenna (48). The sequential updater (90) sequentially calculates a new weight for each of the antenna elements (50) based upon an existing weight of each antenna element (50), the cross-correlation and the difference in pattern magnitude.

Abstract: A wide band, single aperture antenna system (70) that provides simultaneous detection of both RHCP and LHCP signals. The antenna system (70) includes a multiple arm spiral antenna (10) that has spiral arm elements (12, 14, 16, 18) with no sharp transitions. To simultaneously both RHCP and LHCP sensitivity, an antenna feed is connected to both the center end and the outer end of each of the antenna arm elements (12, 14, 16, 18). A separate N-port transformer (74, 78) is connected to both the end feed and the center feed for all of the arms (12, 14, 16, 18) to provide impedance matching and cross-polarization compensation. An NXN port modeformer (76, 80) is connected to the N-port transformers (74, 78) to separate the various modes. The modeformers (76,80) separate the LHCP and RHCP modes to provide an accurate AoA estimation.

Abstract: A fault identification, isolation and fault recovery system autonomously controls spacecraft operational systems. A fault detection and isolation module monitors the operational systems, identifies faults via such monitoring and attempts to isolate component causing the fault. Isolation may constitute a plurality of hierarchically arranged techniques that enhance speed and maximize the likelihood of identifying a correct hypothesis regarding the actual failure. The fault isolation and fault recovery modules are bounded by a severe fault override module which places the spacecraft operational systems in a safe mode or state. Thus, the override module ensures rapid entry into the safe mode while preventing an erroneous hypothesis serviced by the fault recovery module from driving the system beyond acceptable limits. A high level command processor receives command sequences from remote ground support stations and from the on-board fault recovery module.

Abstract: A control scheme for operating an oscillator/counter A/D converter (10) so that it simultaneously provides frequency downconversion, band pass filtering and analog-to-digital conversion of an analog signal, where the analog signal includes a carrier wave modulated with information by any known modulation technique. The converter (10) uses a superconducting, Josephson single flux quantum circuit operating as a voltage controlled oscillator (12). The voltage controlled oscillator (12) receives the analog signal to be converted, and generates a series of sharp, high frequency pulses based on the characteristics of the carrier signal. The series of pulses are applied to a gate circuit (14) that either passes or blocks the pulses depending on a gate control signal. When the pulses are passed by the gate circuit (14), a counter circuit (16) accumulates the pulses during a sampling period.

Abstract: A method and system for deploying a multi-reflector antenna system (10). The antenna system (10) includes an antenna structure (12) mounted to a satellite (14), where an antenna feed array (16) is mounted to the antenna structure (12). A single articulated antenna arm assembly (26) is mounted to the antenna structure (12) by a first spring loaded hinge (28). The arm assembly (26) includes a first arm (30) on which is mounted a first reflector (38), and a second arm (32) on which is mounted a second reflector (40). The first and second arms (30, 32) are connected to each other by a second spring loaded hinge (34) such that the reflectors (38, 40) directly oppose each other and are substantially parallel when the arm assembly (26) is in the stowed position. A plurality of launch locks (44, 46, 50, 54) hold the arm assembly (26) in the stowed position against the bias of the hinges (28, 34) prior to deployment.

Abstract: A visible Stokes polarimetric imager (80) that separately and contemporaneously measures each of the four separate Stokes polarization parameters of a visible light beam reflected from a scene. The imager (80) includes a compound prism assembly (50) including five prism elements (52-58). The reflected beam from the scene is collected by an input lens (84) in the imager (80), and is directed to the prism assembly (50) at a certain location. The prism elements (52-58) split the beam (82) into four separate beams (62-68) of substantially equal intensity that are emitted from the prism assembly (50) in different directions. Four separate polarimetric filters (86) and imaging devices (88) are positioned relative to the prism assembly (50) to separately receive each of the four beams (62-68) emitted from the prism assembly (50).

Abstract: An optical analog-to-digital converter (10) which fully operates in the optical domain and utilizes an upward-folding successive approximation approach for conversion. The converter (10) includes a plurality of optical stages (14, 16, 18) where each stage (14, 16, 18) generates a digital bit. Each stage (14, 16, 18) includes an optical threshold switch (30, 56, 78) that sets the bit high when the switch (30, 56, 78) is closed. When a sample amplitude of the analog signal is compared to a threshold value and found to exceed the threshold value, the bit is set to "high" and the sample is passed directly onto the next stage (14, 16, 18). If the sample amplitude is found to be less than the threshold value, the bit is set to "low" and an intensity equal to the maximum signal intensity minus the threshold intensity is added to the sample amplitude. Each successive stage (14, 16, 18) compares the normalized signal sample to thresholds growing closer and closer to the maximum signal intensity.