Abstract: A spacecraft pressure vessel has a tub member. A sidewall member is coupled to the tub member so that a bottom section of the sidewall member extends from an attachment intersection with the tub member and away from the tub member. The bottom section of the sidewall member receives and transfers a load through the sidewall member.

Abstract: In one embodiment, a wing-to-body fairing for reducing noise due to wing-to-body fairing vibrations in at least one of an aircraft and a spacecraft comprises a honeycomb center comprising a cavity which is substantially filled with foam material.

Abstract: In one aspect, a space vehicle includes a structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the structure. The structure includes at least one of an antenna mesh, a light-shielding mesh, an optical reflector mesh and a net. In another aspect, a space vehicle includes an antenna structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the antenna structure. At least one of the at least two equipment components include a solar panel, a propulsion system and an antenna feed. In a further aspect, a space vehicle includes a payload element and at least two spacecraft-support structures attached to the periphery of the payload element. At least one of the at least two spacecraft-support structures includes a propulsion system and a tracking, telemetry and control system.

Abstract: By sharing tasks between the CV and the KVs, the MKV interceptor provides a cost-effective missile defense system capable of intercepting and killing multiple targets. The placement of the acquisition and discrimination sensor and control sensor on the CV to provide target acquisition and discrimination and mid-course guidance for all the KVs avoids the weight and complexity issues associated with trying to “miniaturize” unitary interceptors. The placement of a short-band imaging sensor on each KV overcomes the latency, resolution and bandwidth problems associated with command guidance systems and allows each KV to precisely select a desirable aimpoint and maintain track on that aimpoint to impact.

Abstract: The present invention provides a MKV interceptor including multiple kill vehicles with autonomous management capability and kinematic reach to prosecute a large threat extent. Each KV can self-manage its own KV deployment and target engagement for a determined target volume assigned by a designated master KV. At least one KV is master capable of managing the post-separation of all of the KVs without requiring updates to the mission plan post-separation. The autonomous capability and increased kinematic reach provides for a more efficient use of boosters and more effective engagement of the threat.

Abstract: A system for single wire secondary distribution comprising a spacecraft platform; a central bus interface unit coupled to the spacecraft platform; a payload unit coupled to the central bus interface unit; and a centralized power supply for powering the central bus interface unit and the payload unit; wherein the spacecraft platform provides a command to the central bus interface unit; wherein the central bus interface unit interrupts the power to the payload unit in a manner corresponding to the commands received by the central bus interface unit; wherein the payload unit decodes the interruption to the power and executes the command from the spacecraft platform.

Abstract: An emulator, which is a computer program, for simulating 32 bit ARINC-429 data output by an Inertial Navigation System on board a P-3 aircraft. The value for each word output by the emulator is adjustable by user. The user adjusts each value by using the computer's mouse to engage slider controls located on a computer applications window. ARINC-429 data which the user adjusts includes aircraft latitude, longitude and altitude data as well as pitch and roll data for the aircraft, and time and distance to a destination by the aircraft.

Abstract: A system and method by which specialized components (tenants) are mounted to an assembly comprising a plurality of modular units. Each modular unit has one or more interface connecting locations for one or more of a power connection, a thermal connection and data base connection, along with one or more of power, thermal and data transmission paths. Modular units can be united in an assembly, and at least some of the modular units have a switching system to selectively transmit and/or receive one or more of power, thermal and data transmissions.

Abstract: Communication between a payload services unit (PSU) and a payload unit utilizes a single-wire interface used to power the PSU as well as to communicate telemetry and command signals. A telemetry and command (T&C) system includes a payload unit configured to respond to a plurality of commands and generate a plurality of telemetry data, an embedded service module (ESM) within the payload unit, and a payload service unit (PSU) configured to generate the plurality of commands and receive the plurality of telemetry data. A single-wire interface is coupled between the ESM and the PSU and configured to provide power to the ESM, wherein the payload unit and PSU are configured to communicate the plurality of commands and the plurality of the telemetry data over the single-wire interface.

Abstract: A method for assembling and landing a habitable module on an extraterrestrial mass is claimed. At least one inflatable module and a second module are placed into orbit about an extraterrestrial mass. Connecting nodes, propulsion busses, and landing pads are also placed into orbit. A habitable structure is constructed from the modules, busses, pads, and nodes. The structure can be robotically constructed. The habitable structure is then landed onto the surface of the extraterrestrial mass. The landing can be remotely controlled and the modules can be non-occupied.

Abstract: According to one embodiment of the invention, a stationkeeping method for a geostationary satellite includes determining a gravitational force of the sun on the satellite at a beginning of a stationkeeping operation and a gravitational force of the moon on the satellite at the beginning of the stationkeeping operation. An initial inclination vector of the satellite is determined at the beginning of the stationkeeping operation that accounts for a first set of one or more perturbations affecting the orbit of the satellite. A maneuver strategy is determined to correct for a second set of one or more perturbations affecting the orbit of the satellite without accounting for the first set of one or more perturbations affecting the orbit of the satellite. Finally, a maneuver is performed on the satellite according to the maneuver strategy.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: An annular fairing having aerodynamic, thermal, structural and acoustic attributes couples a launch abort motor to a space vehicle having a payload of concern mounted on top of a rocket propulsion system. A first end of the annular fairing is fixedly attached to the launch abort motor while a second end of the annular fairing is attached in a releasable fashion to an aft region of the payload. The annular fairing increases in diameter between its first and second ends.

Abstract: A spacecraft system that includes a primary space vehicle and a secondary space vehicle, both of which are designed to optimize payload capacity and launch weight of the primary space vehicle. The primary and secondary space vehicles combine to form an on-orbit space vehicle capable of performing functions and maneuvers that exceed the physical capabilities of the primary space vehicle at the time of its launch. The spacecraft system is designed to minimize propellant containment-related disturbances while maintaining a standard level of functionality. The primary space vehicle is designed to be incapable of independently performing a propellant-intensive orbit change maneuver. Instead the primary space vehicle is designed to couple to a secondary space vehicle having propellant and thrust capability sufficient to perform an orbit change maneuver when the primary and secondary space vehicles are coupled.

Abstract: A method and apparatus for controlling a gimbaled platform (108). The method comprises the steps of computing an acquisition phase gimbal angle rate command ?cmd—Acq (530) from a measured LOS angle error ??LOS (520) for an initial control period T while computing an estimated LOS angle rate {circumflex over (?)}LOS (524), computing a tracking phase gimbal angle rate command ?cmd—Trk (532) using a controller (512) having an output initialized with the estimated LOS angle rate {circumflex over (?)}LOS (524), and commanding the gimballed platform (108) according to an angle rate command ?cmd (522), wherein the angle rate command ?cmd (522) is the acquisition phase angle rate command ?cmd—Acq (530) during the initial control period T and the tracking phase gimbal angle rate command ?cmd—Trk (532) after the initial control period T.

Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: The invention relates to a thermal control device intended to dissipate the heat generated by a payload on a spacecraft, comprising a number of surfaces and including means for circulating a refrigerant. An evaporation zone (Z1) comprises means for circulating the refrigerant, a compression zone (Z2), a condensation zone (Z3) comprising at lEast one radiating panel, linked to a part of the means for circulating the refrigerant, including several branches and comprising means to allow or inhibit the circulation of the refrigerant within these branches so as to vary the area of the heat exchange surface in the condensation zone, a pressure reduction zone (Z4) comprising means for circulating the refrigerant. Such a device is particularly well adapted to thermal problems encountered in telecommunications satellites.

Abstract: In an exemplary embodiment, a pattern is recognized from digitized images. A first image metric is computed from a first digitized image and a second image metric is computed from a second digitized image. A composite image metric is computed as a function of the first image metric and the second image metric, and a pattern is identified by comparing the composite image metric against a reference image metric. The function may be a simple average or a weighted average. The image metric may include a separation distance between features, or a measured area of a feature, or a central angle between two arcs joining a feature to two other features, or an area of a polygon whose vertices are defined by features, or a second moment of a polygon whose vertices are defined by features. The images may include without limitation images of friction ridges, irises, or stars.

Abstract: The present invention provides an airframe structure of an aircraft or spacecraft, with: a hollow body section which can be subjected to an internal pressure; a membrane arrangement which has at least one single-part, multi-curved membrane component for sealing off the internal pressure from an external pressure which acts on the hollow body section and differs from the internal pressure; and a joining structure for a pressuretight joining of the membrane arrangement to a region of the hollow body section; wherein the membrane arrangement forms a receiving well which is accessible from outside the hollow body section, extends at least partially into the hollow body -section and is for receiving a component of the aircraft or spacecraft.

Abstract: A free space optical bus system and method are provided for connecting two units of circuitry across a free space gap. A first unit of circuitry includes a multiplexer that combines a plurality of bus signals to obtain an electronic combined signal and a laser that is driven by the electronic combined signal to transmit an optical signal. A second unit of circuitry is separated from the first unit of circuitry by a free space gap. The second unit of circuitry includes a photodetector that receives the optical signal and transforms it into an electronic composite signal and a demultiplexer that divides the composite signal into a plurality of constituent signals.

Abstract: Actuator for a device for the rotary positioning of a movable element for space equipment comprising a part attached to the fixed portion of said equipment, a bearing, a sleeve designed to actuate the movable portion of said equipment and mounted on an output shaft supported by a first race of said bearing, said actuator comprising a mechanically movable input means and being capable of converting a movement of the input means into a rotary movement of the output shaft about the axis of said bearing, characterized in that said second race of said bearing is connected in rotation to the part of said equipment, in that an adjustable stop so as to angularly limit the relative rotation of the two races of the bearing, and in that the rate of rotation of the output shaft in response to a movement of the input means is different depending on whether or not the sleeve is resting on the adjustable stop.

Abstract: The invention relates to a piggyback satellite, a device for and a method of attaching piggyback satellites with multiple payloads, including a primary satellite and at least one secondary payload, such as a deployable micro satellite, affixed instrument package or the like, wherein the secondary payload is mounted on the nadir end of the primary satellite, preferably employing a universal adapter between the primary satellite and the at least one secondary payload that releases at least one secondary payload or the like in a certain orbit in space for allowing the secondary payload to continue its journey independently in space for transmitting and receiving data.

Abstract: The invention relates to piggyback secondary payloads, a device for and a method of attaching piggyback secondary payloads, including a primary satellite and at least one secondary payload, such as a deployable microsatellite, affixed instrument package or the like, wherein the secondary payload is mounted on the nadir end of the primary satellite, preferably employing a universal adapter between the primary satellite and the at least one secondary payload.

Abstract: A truss structure for supporting a membrane that is to be deployed in space, where the truss structure maintains the membrane sufficiently flat with uniform tension. The truss structure includes a circular perimeter truss and a transitional truss. The transitional truss includes a series of ties that are coupled to the perimeter truss and the membrane. When the truss structure is deployed, telescopic parts of the perimeter truss that are stowed in expansion cause the transitional truss to expand and pull the membrane into the flat configuration.

Abstract: An experiment system with six different re-entry experiment locations for testing high temperature re-entry materials, creating new thermal protection systems, proving innovative new concepts for spacecraft exterior surfaces and the incremental development of next generation aerospace materials. A commercial transportation system to and from orbit provides a 24-hour return cycle for the experiments on a surface actually re-entering the earth's atmosphere. Now using existing doors, hatches and other points on the reusable launch vehicle's exterior, the actual re-entry environment is experienced by test specimens with quick turn around for a wide variety of different re-entry temperatures ranges for broad testing and development purposes. The reusable launch vehicle launches, remains in orbit for 24 hours and returns to provide an actual test environment for the exterior experiment system.

Abstract: The present invention provides a network data recording system particularly adapted for use in transportation systems. Cameras, microphones and a variety of sensors and existing vehicle systems are networked to a central controller, which receives and processes audiovisual and other data from the cameras and sensors. Raw and processed information is stored in a removable memory which may be mirrored to a fixed local memory. The fixed local memory can also be used to store programs and other system data. The controller may be ruggedized to meet disaster recovery requirements. Control panels can be placed throughout the vehicle for use by the crew to monitor conditions and respond to them. Information can be sent to a ground station, and the ground station may also exhibit control over the network data recording system. The system is suitable for use in all manner of transportation systems, such as aircraft, trains, and ships.

Abstract: A system and method to enhance attitude estimation for transfer orbit maneuvers of a spinning satellite includes a transfer orbit earth sensor that can receive and convert an infrared radiance to an analog signal. The analog signal is provided to a hardware chord processor and an analog-to-digital converter. The hardware chord processor determines an earth chord length based on the analog signal. The analog signal is sampled in the analog-to-digital converter to provide a digital signal to a digital earth sensor pre-processor, which reshapes the digital signal to provide a pre-processed signal having nearly distinct peaks. A software chord processor is provided with the pre-processed signal. The software chord processor locates the peaks in the pre-processed signal and determines an earth chord length.

Abstract: A device for constructing a synthetic image of an aircraft environment may include information sources having information relating to the aircraft and to its environment. An information processor constructs a synthetic image of the environment of the aircraft on the basis of information emanating from the information sources. A display presents, on at least one part of a display screen, a synthetic image, which is constructed with the aid of a conical projection onto a plane of projection that is orthogonal to a line of aim that forms angles of lateral separation and angles of vertical separation with the course of the aircraft.

Abstract: A satellite equipment panel on which may be mounted an installation equipment of a predetermined function, connectable by an interface, constitutes an artificial satellite. The satellite equipment panel includes a panel body, at least two interfaces for an interconnection to another satellite equipment panel, a line transfer switch disposed within the panel body for line transferring of the interconnection between the interfaces, a wiring disposed within the panel body and connected at one end to one of the interfaces and at the other end to the line transfer switch for interconnecting the interfaces on that panel body. The interfaces and the wiring, in cooperation with a support structure, constitute a wiring module disposed within the panel body. The interface may comprise at least one of interfaces of electric communication, optical communication, and optical sensor.

Abstract: Optical surface reflector, for a spacecraft such as a geostationary satellite. The exterior surface (10) of the reflector comprises a plurality of facets (14) inclined to each other and to the interior surface (12) of the reflector. The facets (14) advantageously form a pyramid with a square base in which the angle at the apex is equal to 90°. However, other arrangements are possible without departing from the scope of the invention. The invention increases radiative capacity without significantly increasing mass and overall size.

Abstract: An asymmetrical bulkhead for a pressure vessel, tank or other storage facility is disclosed. The bulkhead is configured so that there is a high region extending past a cylindrical portion of the vessel on one side of the bulkhead, with a contour of the bulkhead transitioning from the high region to a low region within the last cylindrical region. Such a bulkhead configuration allows two cylindrical pressure vessels to be nested together in end-to-end relation so as to occupy a total smaller volume than similarly sized vessels having convex ends, or to construct pressure vessels that contain a larger volume for storage than similarly sized vessels having convex ends.

Abstract: The invention generally relates to a flexible heal exchanger comprising a pair of flexible multilayer thermoplastic polyimide films each of which comprises an aromatic polyimide substrate film showing no glass transition temperature or a glass transition temperature of 340° C. or higher and a thermoplastic aromatic polyimide surface film showing a glass transition temperature in the range of 190 to 300° C. in which the surface films of the flexible multilayer thermoplastic polyimide films face each other and are in part fused together, whereby producing between the flexible multilayer thermoplastic polyimide films a conduit pattern through which a fluid passes. The flexible multilayer thermoplastic polyimide films preferably have a thickness in the range of 10 to 125 ?m and comprise a linear expansion coefficient of MD, a linear expansion coefficient of TD and an average of linear expansion coefficients of MD and TD, in the range of 10×10?6 to 35×10?6 cm/cm/° C. at 50-200° C.

Abstract: A device for indicating at least a first vibration level VIB1 derived from a first acceleration signal ?1 includes: first analysis element FT1 for resolving the first acceleration signal ?1 into a plurality of axial analysis signals ?1i each occupying an identified frequency band Bi; a first memory RM1 having recorded therein a plurality of axial weighting coefficients KAi, each corresponding to one of the identified frequency bands Bi; and a first control circuit CC for summing the pth powers of the axial analysis signals ?1i weighted by the axial weighting coefficients KAi, thereby producing the first vibration level VIB1.

Abstract: A small, substantially spherical, free-flying utility platform for use in space can be docked, inspected and serviced in an apparatus including a container with a chamber therein mounted on the outside of a space vehicle or space station. A motorized closing mechanism drives a door to selectively open or close an opening of the container into the chamber, and another mechanism automatically latches and locks the door. The door operations can be automatically triggered as the platform breaks the beams of light beam devices arranged in the container. The proper positioning of the platform in the container can be detected by a photoelectric switch. Capture and docking of the platform is facilitated by sloping funnel-shaped interior walls of the container, and properly shaped cushions and/or a docking cradle. When docked, the platform has its batteries recharged, its fuel tank refueled, and its electronics evaluated and diagnosed.

Abstract: Capture devices are mounted on a space platform such as a spacecraft or utility satellite. Each capture device includes a folded capture net stored in a housing, guide weights connected to the net and received in spring-loaded ejection tubes oriented at an acute angle relative to the axis of the housing around the perimeter thereof, and a spring-loaded releasable cover that is connected to the net, covers the net in the housing, and holds the weights in the ejection tubes. When the cover is released, it is spring-ejected from the housing while pulling out and unfolding the net, and the weights are spring-ejected out of the ejection tubes while spreading out the net. After the net wraps around a free-flying target object in space, net closure mechanisms reel-in a net pursing line to close the net, which is connected to the space platform by a tether line on a winch.

Abstract: A device for verifying the calibration of an infrared sensor includes a body configured to be placed in orbit. The body defines a spin axis and an outer surface. The outer surface has an infrared radiant intensity that is substantially independent of the viewing angle between the spin axis of the body and the infrared sensor. The device may be used for on-orbit verification of the calibration coefficients associated with each pixel of an infrared sensor, including linearity calibration and verification of the resolution capability of the sensor. The calibration coefficients may be verified over the entire sensing capability range of the infrared sensor. After a thermal model of the device is developed, the device operates passively and provides a known source of infrared radiation observable during the device's entire orbital lifetime. The device may be designed to automatically de-orbit from space after a predictable time in orbit.

Abstract: A closed-loop LRF pointing technology to measure the range of a target satellite from a chaser satellite for rendezvous is provided that includes: LOS angle measurements of the target, a relative navigation Kalman filter, attitude determination of the visible sensor with gyros, star trackers and a Kalman filter, pointing and rate commands for tracking the target, and an attitude controller. An analytical, steady-state, three-axis, six-state Kalman filter is provided for attitude determination. The system provides improved functionality and precision for relative navigation, attitude determination, pointing, and tracking for rendezvous. Kalman filters are designed for the closed-loop system to allow for pointing the laser rangefinder to a target even if a visible sensor, a laser rangefinder, gyros and a star tracker are misaligned and the LOS angle measurements from the visible sensor are interrupted.

Abstract: This disclosure presents a communication receiver system for spacecraft that includes an open loop receiver adapted to receive a communication signal. An ultrastable oscillator (USO) and a tone detector are connected to the open loop receiver. The open loop receiver translates the communication signal to an intermediate frequency signal using a highly stable reference frequency from the USO. The tone detector extracts commands from the communication signal by evaluating the difference between tones of the communication signal.

Abstract: A method and apparatus for estimating a slave payload attitude is disclosed. The method includes accepting a plurality of slave payload attitude measurements, deriving a model of the relative attitude of the slave payload and a master payload attitude at least in part from the plurality of slave attitude measurements, predicting the relative attitude between the slave payload attitude and the master payload attitude using the derived model, and estimating the relative attitude between the slave payload attitude and the master payload attitude at least in part from the predicted relative attitude between the slave payload attitude and the master payload attitude. Furthermore, the absolute attitude of the slave payload is computed using the relative attitude and the master payload attitude.

Abstract: A spacecraft having facilities for carrying astronauts and payloads into the earth satellite orbit and space having a composite fiber body with a Kevlar exterior skin. Said craft having a substantially round shape with a portion covered with solar cells. Said solar cells powering electrical motors and a plurality of rockets using a quantity of hydrogen fuel.

Abstract: The invention provides methods, systems, and products for communicating between engine computers without the participation of a flight deck computer. The invention allows the engine computers to coordinate activities for real-time control of the engines for improved efficiency using an existing ARINC 429 communications bus. The invention facilitates cross-engine communications over the same ARINC 429 communications bus used for communicating with the Flight Deck Computer, without installation of a second, dedicated communications bus for cross-engine communications. The invention thus provides a cost effective way to allow the engines to coordinate activities without expensive installation and recertification of airframe modifications.

Abstract: An apparatus for entering a flight plan into an aircraft navigation system, the apparatus comprising: an acoustic sampler adapted for sampling a microphone signal and generating an acoustic signal; a waypoint identifier adapted for generating an identified waypoint from the acoustic signal and the flight plan; and a navigation interface adapted for incorporating the identified waypoint into the flight plan and for transmitting and receiving navigation data to and from the aircraft navigation system.

Abstract: A Hall effect thruster for propelling spacecraft and satellites includes at least two acceleration channels, each of the channels has a closed end and an open end, and a plurality of flux guides adjacent each of the channels. The plurality of flux guides includes an innermost flux guide, an outermost flux guide, and at least one intermediate flux guide. Each intermediate flux guide helps provide a magnetic field to each of two adjacent acceleration channels.