Abstract: In a multi-cell battery, an open-circuit failure of any one cell renders the entire battery unusable. An automatic electronic bypass switch is connected across each battery cell to provide a low-resistance conductive path around such battery cell should it fail. Contrary to conventional usage in which a fuse is employed to "clear" or "open" a short-circuit condition, the electronic switch of the present invention employs a fuse to activate a highly conductive path, which ideally would approach a short circuit, around an open-circuit condition. The electronic switch includes a field-effect transistor connected across the battery cell and a fuse to hold the field-effect transistor in the "off" condition.

Abstract: A spacecraft of a set of spacecraft has an antenna support frame defining square apertures, which are the same size in all spacecraft of the set of spacecraft. The set also includes a plurality of radiating tiles for various operating frequencies. The tiles all have the same dimensions, so that any combination of tiles may be mounted in the apertures of the frames of any spacecraft to form antenna arrays of various sizes, depending upon the customer's requirements. Each tile includes plural beamformers and power amplifiers, and remotely controllable phase shifters and level controls, for ready reconfiguration of the antennas while the spacecraft is in operation. The preferred dimension of each radiating tile is a one-foot square.

Abstract: A spacecraft includes a body defining a center of mass, and an attitude control system, for generating attitude error signals representing deviation of the body from a desired attitude. A plurality of thrusters mounted on the body have nominally parallel thrust components, in a particular direction. The thrusters are arranged in pairs, which produce opposed torques. A velocity change controller is coupled to the thrusters, for generating pulsatory firing command signals for the thrusters for a particular total time duration, for achieving a desired velocity change in the desired direction.

Abstract: In the described embodiments, a spacecraft is powered by paralleled solar panels, which supply power to a load. The panels are ORed together, and each is controlled by a switched shunt, which provides coarse control of the load voltage. A control loop senses the load voltage, and logic selects those shunts which are conductive at any particular time in a manner which tends to maintain the load voltage constant. The control loop may undesirably become unstable because of gain changes caused by changes in solar panel insolation. A further controller adjusts the clock frequency at which the shunts are switched in a manner which tends to stabilize the loop. The further controller responds (a) to the current of one or more of the panels, (b) to the frequency of oscillations across the load, or (c) to both.

Abstract: A spacecraft (10) has an attitude sensor (20) mounted on its body (12). The sensor must be maintained near a temperature setpoint. Each sensor (20) produces its own temperature-indicative signal. Each attitude sensor is coupled to a thermally conductive instrument platform (18). A standoff (21) supports the platform (18) away from a baseplate (16) and the spacecraft body (12). The standoff (21) includes a thermally conductive portion (22) adjacent the platform (18), and a nonconductive portion (24) remote from the platform. An electric heater (26) is connected to the thermally conductive portion (22) of the standoff (21). A temperature sensor (28) thermally coupled to the platform (18) generates platform temperature signals. A filter (212) high-pass filters either the platform temperature signals or the attitude sensor temperature signals, to form filtered signals. A combining circuit (218) combines the filtered signals with the other temperature signals, to make composite temperature signals.

Abstract: A spacecraft includes a plurality of solar panels interconnected with a power coupler and an electrically operated device to provide power to the device when the solar cells are insolated. The solar panels are subject to bending distortion when entering or leaving eclipse. Spacecraft attitude disturbances are reduced by mounting each of the solar panels to an elongated boom made from a material with a low coefficient of thermal expansion, so that the bending of one panel is not communicated to the next. The boom may be insulated to reduce its bending during changes in insolation. A particularly advantageous embodiment mounts each panel to the boom with a single mounting, which may be a hinge. The single mounting prevents transfer of bending moments from the panel to the boom.

Abstract: A spacecraft includes heat-radiating panels. Heat-generating boxes such as TWTs are mounted on shelves which are inside the spacecraft, where they cannot radiate to space. The shelves are coupled by bent heat pipes to the exterior radiating panels.

Abstract: A spacecraft includes a three-axis attitude control system. When velocity change thrusters are fired, their plumes impinge on a solar array, at angles which vary with the solar array position. This causes disturbance torques which vary with the solar array position. Disturbance torque information signals or torque bias signals which depend upon the solar array angle are summed with the torque demand signals which control the attitude control system during firing of the velocity change thrusters, to modify the attitude correction torques. The bias torque signals are generated by a Fourier processor based upon stored Fourier coefficients together with signals from a solar array angular position sensor.

Abstract: An array antenna especially adapted for spacecraft use includes a support frame made up of intersecting beams which form an "eggcrate" of square openings. A plurality of subarrays or radiating tiles are dimensioned to fit within the openings. Distortion of the support frame due to external forces, or due to expansion and contraction of the radiating tiles, is accommodated andor resisted by a four-point mounting arrangement for each tile. Each mounting is located on the center of the sides of the radiating tile, and the corresponding center of the side of the opening in which it fits. Each mounting consists of a thin, flexible mounting beam, extending parallel to the associated support beam, and fastened to the support beam at the ends of the mounting beam. The center of the side of the tile is fastened to the center of the flexible support beam, so expansion and contraction of the tile are compensated by bending of the flexible beam.

Abstract: A three-axis stabilized spacecraft includes a plurality of primary attitude control thrusters, the torque vectors of which lie in, or parallel to a primary plane. It also includes at least two more secondary attitude control thrusters, the torque vectors of which lie in a secondary plane which is not parallel to the primary plane. The control system produces attitude error signals, which are processed with a PID characteristic to produce impulse demand signals, all in known fashion. The impulse demand signals are transformed into an auxiliary coordinate system, in which two of the three auxiliary axes lie in the primary plane, and the third is orthogonal thereto. One of the secondary thrusters is selected, which has, along the third auxiliary axis, the largest torque magnitude and the same sign as the transformed impulse demand.

Abstract: A zero-momentum spacecraft's attitude is controlled by determining the torque required about a control axis to maintain the desired attitude, and, during each of recurrent control cycles, enabling a magnetic torquer if the torque demand exceeds a threshold. During each of the control cycles, thruster(s) are enabled to make up the difference between the torque demand and the estimated torque produced by the magnetic torquer. In determining the torque demand, the attitude rate signal is low-pass filtered to reduce noise, and the control loop bandwidth is maintained by totalling the estimated torque applied by the magnetic torquer and thrusters, integrating and high-pass filtering the estimated torque signals, and adding the filtered estimated torque with the filtered attitude rate signals to generate low-noise attitude rate signals. A three-axis system is described.

Abstract: Two opaque sheets made from the same electrically conductive polymer material have adjoining edges. Each sheet is carbon loaded, electrically conductive, opaque black polymer on a first broad side thereof. The sheets are joined along a seam region by means of an opaque, electrically conductive adhesive tape. The electrically conductive adhesive tape includes an elongated opaque sheet of electrically conductive polymer material of the same kind as that of the two opaque sheets which the tape joins. The opaque, electrically conductive adhesive tape also includes a layer of electrically conductive adhesive, which is connected to the electrically conductive polymer material. In a particular embodiment of the invention, the polymer material is polyimide. In a preferred embodiment, the two opaque sheets are thermal coverings for a spacecraft, and each includes a sheet of opaque polyimide, carbon loaded which is electrically conductive, on the exterior of the thermal coverings.

Abstract: A spacecraft is controlled by a composite attitude signal including two components with different passbands. In one embodiment, a spacecraft attitude control system includes an attitude sensor and a controller which provides a time derivative function. High frequency noise components of the sensed attitude signal are enhanced by the derivative, and tend to cause attitude jitter or excess power consumption. The jitter is reduced by low pass filtration of the sensor signal, but this undesirably reduces the high frequency response of the attitude sensor. The high frequency response is restored by high pass filter coupled to a reaction or momentum wheel tachometer, which produces a signal representative of the high frequency components of the body rate. A summing circuit couples together the filtered attitude sensor signals with the high frequency components of the wheel speed signal to produce a relatively noise free broadband body rate signal.

Abstract: An antenna array for use in space includes an optical solar reflector (OSR)-like structure (310, 510), with sheet of transparent dielectric material (320, 520) and a metallic mirror (322, 522) formed on the sheet. In order to allow the individual antenna elements to radiate electromagnetic energy, the metallic mirror contains apertures (330, 530) which, in one embodiment, define the antenna elements, and which, in another embodiment, are registered with the antenna elements of the array. Where the mirror exists, insolation is reflected, and the heat gain from insolation is minimized. At the locations of the apertures in the mirror, insolation passes through the OSR-like structure, which may tend to increase the temperature of the structure. According to the invention, the apertures are covered with a thin layer (370, 570) of semiconductor material such as germanium or silicon, which does not affect the electromagnetic energy flow.

Abstract: A bidirectional power converter array has overall voltage feedback, an error signal generator, and plural bidirectional power converter modules. Each module of the array includes a series inductor and local feedback based upon the inductor current. Each module also has a resistive current sensor, providing wideband local feedback. A low-pass filter is coupled to the current sensing resistor to reduce switching noise. The low-pass filter also reduces the bandwidth of the local feedback signal, so interactions with the overall feedback occur, which may result in instability. A winding transformer-coupled to the series inductor generates AC signal representing high-frequency components of the inductor current, relatively free of switching noise. The AC current sample is added to the low-frequency components derived from the resistor and low-pass filter, to form a low-noise local feedback signal, with sufficient bandwidth to reduce unwanted interactions with the overall system feedback.

Abstract: A spacecraft (10) includes and attitude sensor (16) for generating sensed attitude signals. A controllable drive arrangement (238) is coupled to a reaction wheel (46), for driving it in response to speed error signals. The drive results in changes in reaction wheel speed. The spacecraft has a digital speed sensor (248) coupled to the reaction wheel, for generating quantized speed signals. A reaction wheel speed error signal generator (234) produces speed error signals in response to the difference between the attitude command and the sensed attitude signals, corrected by signals responsive to the speed of the reaction wheel, whereby at low reaction wheel speeds, the quantization causes attitude errors.

Abstract: A process and apparatus for making high purity aluminum nitride from organometallic aluminum, such as an alkylaluminum compound. A gaseous alkylaluminum compound and gaseous ammonia are introduced into a heated reaction zone where the gases are mixed and high purity aluminum nitride is produced. The high purity aluminum nitride is collected in the form of a powder or deposited as a thick, dense layer on an appropriate substrate mounted in the reaction chamber. A carrier gas such as hydrogen gas, may be used to conduct the alkylaluminum compound from a suitable reservoir containing liquid alkylaluminum compound to the reaction chamber. A preferred alkylaluminum compound is triethylaluminum.

Abstract: A dual-polarized antenna includes a planar line or surface array of reradiating elements of two different polarizations. Each reradiating element in one embodiment is a short-circuited vertical or horizontal dipole. The vertical and horizontal dipoles may be collocated on the array to form crossed short-circuited dipoles. The elements of each polarization form an array separate from the elements of the other polarization. Within each of the two separate arrays, the 1/.lambda. of each reradiator is adjusted to provide a phase shift which causes a collimated beam incident on the array to result in reradiation or "reflection" of energy in the form of a beam converging at a focal point. Therefore, each planar reradiator array acts as a parabolic reflector with a particular focal point. The focal points for the vertical and horizontal arrays are different, and a feed of the appropriate polarization is located at each of the two focal points.

Abstract: A color television camera includes a panchromatic light sensor such as a CCD imager. Light from a scene to be imaged is filtered by an electronically controlled light filter such as an etalon, which is scanned from color to color at the field rate, so that the image for each field is generated by a different color. The periodic spectral line or spectral band response of a single etalon is improved by a cascade of two etalons with different periodicity, which pass only a single spectral line or band. When an RGB color sequence is used, a complete color signal frame is generated in three fields. The imager may be operated at a field rate higher than the standard field rate, such as 90 Hz. for use with conventional 30 Hz. color frame rate, and a memory may be coupled to the output of the imager to store the three-field sequence of signals which occurs during each standard color frame, so that a complete color signal is available during each standard color frame.

Abstract: A notch filter includes a directional coupler (16) with an input port (17a), coupled output port (17b), coupled 0.degree. (17d) and direct 90.degree. (17c) ports. A cylindrical dielectric resonator (40) is supported by a spacer (58) above a ground plane (14) and dielectric substrate (12). A first microstrip transmission line includes a strip conductor (22) coupled at one end to the coupled 0.degree. port, and extending parallel to a tangent to the edge of the resonator at a central plane (44), terminating in an open-circuit (226). A second transmission line includes a strip conductor (24) coupled to the direct 90.degree. port (17c) and extending parallel to the first transmission line, on the other side of the resonator. The first and second transmission lines each have an electrical length .lambda./4 between the central plane (44) and their open-circuit terminations, to reflect a high current to the plane.