Abstract: A spacecraft (8) includes a movable appendage such as solar panels (12) operated by a stepping motor (28) driven by pulses (311). In order to reduce vibration andor attitude error, the drive pulses are generated by a clock down-counter (312) with variable count ratio. Predetermined desired clock ratios are stored in selectable memories (314a-d), and the selected ratio (R) is coupled to a comparator (330) together with the current ratio (C). An up-down counter (340) establishes the current count-down ratio by counting toward the desired ratio under the control of the comparator; thus, a step change of solar panel speed never occurs. When a direction change is commanded, a flag signal generator (350) disables the selectable memories, and enables a further store (360), which generates a count ratio representing a very slow solar panel rotational rate, so that the rotational rate always slows to a low value before direction is changed. The principles of the invention are applicable to any movable appendage.

Abstract: A travelling-wave tube (TWT) for a satellite communication system includes a cathode current regulator which compares a sensed I.sub.K signal with a cathode current reference to form an cathode current error signal, and controls the cathode current in response to the cathode current error signal. Helix arc protection is provided by a helix current shut-down, which compares the helix current with a particular reference, and shuts the TWT supply down when helix current flows. Desirable high efficiency TWTs have a small input drive range between maximum TWT output and a defocussed condition, in which beam current flows in the helix. Sudden changes in path attenuation may result in an RF overdrive condition, and cause helix current flow, and undesirably shut down the TWT. A helix current limiter senses the helix current, and compares it with a reference which represents less helix current than that which causes shutdown, to produce a helix current error signal.

Abstract: A hinge includes a first clevis (30), defining a vee-block support (12) and a first limiter surface (44, 46, 48). A cylindrical hinge shaft is supported by the vee-block (12). A second clevis (140) is affixed to the shaft, whereby the second clevis may rotate in response to torques. The second clevis has a limiter surface (148) which coacts with the first limiter to limit rotation of the shaft and second clevis past the deployed state. The second clevis (140) has a cam follower (158) at a predetermined distance from the axis of rotation (2). A cam (100, 104) affixed to the first clevis (30) rotates about a cam axis (3'). The deployed-state engagement between the cam and the follower snugs one end of the shaft into the vee-block, while the rotation limiters snug the other end. A rotational driver (50, 54, 70, 75, 90, 110) urges the cam (100, 104) to rotate in a direction which tends to move the second clevis (140) from the stowed state toward the deployed state.

Abstract: A passive damping mechanism for a panel member in which a bonding layer is disposed on each side of the panel, a damping layer is disposed on each of the two bonding layers, and a constraining layer is disposed on each of the two damping layers. Each of the above-mentioned layers is coextensive with the respective panel surfaces to which it is attached. The bonding layers are preferably made of an aramid fiber material. The damping layers are preferably made of a viscoelastic material. The constraining layers are preferably made of an aluminum-graphite metal matrix composite material. In addition, the coefficient of thermal expansion of each bonding layer should match well with that of each constraining layer.

Abstract: A spacecraft (10) includes N paralleled power converters (24a, 24b . . . ) for producing current pulses in response to converter synchronizing pulses. The current pulses are integrated (28) to produce an operating direct voltage. The synchronizing pulses are produced in a recurrent cycle by a synchronizer (30). The synchronizer includes a voltage divider (320) with equal series-connected elements (324a, 324b . . . ), each of which is paralleled by a shorting switch (330a, 330b . . . ). A sawtooth signal generator (310) produces a sawtooth signal (312), which is applied in common to each of a plurality of comparators (332a, 332b . . . ). The voltage divider (320) is connected across a reference voltage (318), to produce a plurality of reference voltages at the taps (328a, 328b). Each comparator (332) compares the sawtooth ramp (314) with one of the reference voltages, so that the comparators trigger in sequence as the ramp rises. The comparators trigger at times corresponding to equal phase increments of 360.

Abstract: A Fabry-Perot etalon (310) has a cavity (12) occupied by an electric-field variable index of refraction material (22). The cavity is flanked by a pair of parallel mirrors (14), at least one of which is partially reflecting. At least one of the supporting substrates (18l) has deposited thereon a light polarizing field electrode (316l) which includes a plurality of mutually parallel, spaced apart electrical conductors (392) which define gaps (394). The electrical conductors are mutually interconnected by bus conductors (390), and may be connected by lead conductors (37) to an external voltage source, for generating the desired electric field across the cavity.

Abstract: An optical demultiplexer includes an electrooptic modulator (410) which modulates a beam of light (418) in response to frequency-multiplexed radio-frequency (RF) information signals, to produce diverging beamlets of light (420). The diverging beamlets are separated by a spatial separator arrangement (426), and each beamlet (420), including the information of its RF carrier, is coupled to a separate optical detector (428). The detector (428) can extract amplitude modulation from the signal. In order to reconstruct the RF signal as well as the amplitude modulator, an optical "local oscillator" signal (OLO) is coupled to each detector together with its information signal. Signal loss due to vibration or misalignment is avoided, and heterodyne mixing efficiency is maximized in an embodiment of the invention, by propagating the OLO and information signals through a single-mode optical fiber to the detector.

Abstract: A method for making a pressure tank includes procuring a metallic feed pipe (12) with a substantially cylindrical outer surface (312), a bore (21), and, at an attachment end thereof, an annular ring (350) protruding from the cylindrical outer surface. A dissolvable mandrel (410) in the shape of the desired tank or pressure vessel is made. The mandrel includes a neck portion (412, 414) dimensioned to fit within the bore of the feed pipe. If necessary, the outer surface of the mandrel is rendered electrically conductive, as for example by metal sputtering or evaporation, or by electroless deposition. A thin layer (328) of ductile metal is then electrolytically deposited over the outer surface of the mandrel, extending onto the outer surface of the feed pipe in the pipe-to-tank coupling region, closely fitting over the annular ring, to thereby form a tank liner. A composite reinforcing shell (322) is formed around the liner, and extending over the liner in the coupling region.

Abstract: A method for making a storage tank with an integral vane starts with a feed pipe. A thin, flat vane has a shape which defines an inner and outer curve extending between attachment and second ends. The outer curve of the vane is dimensioned to lie just within the inner periphery of the tank, spaced therefrom by a particular gap. The attachment end of the vane is affixed to a first part of a coupling region of the feed pipe, with the pipe axis lying in the plane of the vane. The vane-pipe combination is placed in the cavity of a mold having the desired interior shape of the tank, with the edge of the outer curve of the vane spaces away from the wall of the mold by the desired gap. A dissolvable, hardenable liquid is injected into the mold and hardened. The mold is removed, leaving a soluble mandrel in the shape of the cavity of the tank. A copper coating is applied to the mandrel and the pipe, to form a metal liner including a neck extending over the feed pipe.

Abstract: A Fabry-Perot etalon 10 is located in the light path 30 between a light sourcesinks 40 and 44. The etalon is polarization insensitive, and its optical path length and transmission wavelength(s) is controlled by a bias source 60. When light at the desired wavelength is propagating from source 40, the bias is set to establish transmission through the etalon at that wavelength, whereupon the light propagates to the sink 44, and the switch is transmissive. By contrast, if the switch is to be nontransmissive, the bias sets the transmission wavelengths of the etalon on either side of the wavelength produced by source 40, whereupon the light from source 40 is reflected by the etalon, and does not reach sink 44. In a second embodiment, a third sourcesink 48 is coupled in a second light path 330a, for being coupled to the first sourcesink 40 when the etalon is set for reflection, to thereby form a single-pole, double-throw switch.

Abstract: A modulator or demodulator has a transparent nonlinear dielectric material (crystal), and an optical fiber for coupling light thereinto. RF is coupled into the crystal by a dielectric waveguide which is matched to the crystal. The nonlinear crystal causes the light and RF to interact, to cause products which can be detected at the dielectric waveguide. When used as a modulator or demodulator, the modulated light may be coupled out by a second optical fiber, and or the modulated RF may be coupled out by a second dielectric waveguide. The dielectric waveguide may be transitioned to conventional hollow conductive waveguide.

Abstract: A power supply for a thruster for space applications (e.g., an arcjet thruster) comprises a series/parallel resonant converter wherein the load quality (Q) factor of the resonant tank circuit inherently varies with arc voltage so as to accommodate changes in arc voltage from initiation thereof through steady-state operation. Effectively, the resonant tank circuit acts as a ballast by matching the arc during initiation thereof through steady-state operation. Furthermore, the resonant converter operates in a soft-switching, low-voltage switched mode so as to maximize efficiency.

Abstract: A multi-pole magnetic component, such as an inductor, includes z-folded conductive film windings and a core having a plurality of pairs of spaced apart core posts extending between a base plate and a top plate of the core. The core includes an air gap that is distributed substantially evenly along the flux path. Magnetic flux flows through the core posts in a series manner so as to have an opposite flux direction in adjacent poles. Preferably, the air gap is determined such that the ratio of the distance between each adjacent core post and the distance between the base and top plates results in magnetic fields that are substantially tangential to the surface of the conductive film winding. Furthermore, the core posts are preferably shaped to have a larger cross sectional area at the base portion of the posts than at the top portion thereof. Alternatively, the core posts are attached to the bottom plate and are inserted into suitably shaped cut-out portions of the top plate.

Abstract: A lift vehicle has an elongated support structure affixed to its upper end, with the axis of elongation of the support structure parallel to or coincident with the longitudinal axis of the lift vehicle. The support structure defines a plurality of attachment planes, each of which includes the axis of elongation, which together divide the region about the axis of elongation into N angular sections or regions, each with an included angle of 360.degree./N. A plurality MN of spacecraft are organized into a plurality of M sets or layers, each containing N spacecraft. Each spacecraft defines first and second lateral support planes, skewed relative to each other, which intersect at a line of intersection, is parallel to the axis of elongation when the spacecraft are mounted to the support structure. Two or more spacecraft are mounted in end-to-end position in each angular portion. Each spacecraft stiffens the support structure, so that the stiffened support structure, can sustain the launch forces.

Abstract: A thermal control surface for a spacecraft includes flexible OSR tiles (310) 0.0015 inches thick. Each tile includes a transparent substrate (312) with a reflective second surface (314), such as a coating of silver. The front or space-facing side is coated with a transparent, electrically conductive layer (316) of indium oxide (IO) or indium-tin oxide (ITO), which laps over the edges of the OSR tile, and is in electrical contact with the reflective coating.The IO or ITO coating, or a separate electrically conductive coating (610) such as nichrome, forms a protective coating for the reflective surface, and maintains contact with the front surface of the tile. The IO or ITO layer (316) in conjunction with the back side coating provides front-to-back electrical conduction. The OSR is mounted to an electrically conductive surface (110) of the spacecraft, such as an aluminum or graphite-fiber-reinforced epoxy resin, by a sheet of transfer adhesive (410).

Abstract: A spacecraft uses monopropellant arcjets for velocity change such as for north-south stationkeeping. It has been discovered that, while an arcjet cannot be modulated by pulsing the fuel supply, the amount of thrust can be varied by modulating the power applied to the arc, without extinguishing the arc. While the specific impulse (I.sub.SP) of the arcjet is thereby reduced from the maximum I.sub.SP of which the arcjet is capable, the resulting I.sub.SP may still be larger than the combined I.sub.SP of an unmodulated arcjet in conjunction with a modulated chemical thruster in a typical scenario. According to the invention, attitude control is provided in conjunction with north-south stationkeeping or other velocity change by, in response to an error signal generated by an attitude control system, modulating the arc power(s) of an arcjet thruster(s), which provides the velocity change. The arc is not extinguished during the stationkeeping maneuver, but is varied in magnitude.

Abstract: A power supply system for thrusters (e.g., arcjet thrusters) in a spacecraft system comprises lightweight, redundant power supplies that share the power distribution function to the thrusters. Each power supply comprises a plurality of parallel-connected power supply subunits of which the combined power output capability exceeds the required maximum power demand by at least one subunit capacity for each arcjet thruster in the system. Each power supply subunit comprises a lightweight, high-frequency, soft-switching power supply. For arcjet systems comprising arcjet thrusters that do not operate simultaneously, relays are employed to switch between thrusters.

Abstract: A satellite communications system includes a feed array and reflector receiving antenna coupled to a dual-mode network, for generating two channels, with independent signals representing two beams, one for each channel. Each output channel of the dual-mode network contains redundant information relating to a plurality of information channels. In the absence of an interfering signal or intrusion, the combined signals are amplified and block frequency-converted, then de-multiplexed into separate channels, which are each applied through selection switches to a transmitter multiplexer for retransmission. An alternate signal is tapped from each of the two received signal channels, phase controlled, and combined, to produce a combined signal representing a receive antenna beam with a null controllable in position in response to phase. The combined signal is demultiplexed into separate information channels.

Abstract: A 3-axis stabilized spacecraft includes roll and yaw magnetic torquers, and a momentum wheel oriented with its spin axis orthogonal to, and pivotable about, the roll axis. Roll control may be applied by pivoting the wheel. Secular increases in pivot angle may result in loss of control authority when the mechanical limits of the pivot are reached. The pivot angle is sensed, and an unloading control loop is closed, by which magnetic torquers are energized to torque the spacecraft, to return the pivot angle toward zero. The unloading control loop includes a bandpass filter, which eliminates constant components of pivot angle offset. This prevents the unload control loop from attempting to maintain the pivot at a position in which the wheel axis is offset from the desired normal to the orbit plane. Consequently, the magnetic torquers do not expend system energy attempting to maintain an undesirable attitude.

Abstract: Equipment support and thermal radiation panels for a spacecraft include a transverse header heat pipe and a set of a plurality of "spreader" heat pipes each including an elongated longitudinal portion and an "L" portion lying in thermal contact with the header heat pipe. The elongated longitudinal portions of the spreader heat pipes are spaced apart and mutually parallel. A further set of spreader heat pipes, similar to the first set, is provided. At least part of the elongated longitudinal portion of each further heat pipe is physically parallel, and in thermal contact, with a heat pipe of the first set, and the "L" portions of the further heat pipes are in thermal contact with a second header heat pipe. A make-break thermal joint juxtaposes a header heat pipe of each of two panels lying in different planes.