Abstract: A missile includes a separable shroud that covers a nose portion of the fuselage of the missile. The shroud covers and protects a seeker window and a seeker at the nose of the missile. The shroud is configured to remain coupled to the missile during and immediately after launch of the missile, and to separate during flight under the action of aerodynamic forces. Toward that end parts of the shroud are initially coupled together by a retainer, which allows the parts to separate during flight. The retainer may include one or more tension bands that break at a certain tension, and/or one or more weakened parts of the shroud, which break during flight. Parts of the shroud may include inward protrusions that make contact with an ogive portion of the nose of the fuselage.

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: A profiled part, for use, in particular, within an aircraft or spacecraft, the profiled part comprising at least one hollow profiled part portion and a T-shaped and/or L-shaped profiled part portion.

Abstract: A system and method for detection of anti-satellite vulnerability of an orbiting platform. An engagement volume processing unit receives a maximum impulse velocity of an SBI launched from a carrier platform of interest, a maximum time of flight (TOF) until intercept, and orbital data of the carrier platform of interest. A family of interceptor imparted velocities in a VNC frame is determined. The engagement volume processing unit applies the family of velocities over “N” release points using a carrier platform propagator to determine an engagement volume. A display and alert processing unit generates a visual representation of the engagement volume and sends the visual representation to the display device for display.

Abstract: The invention set forth herein describes propulsive guidance methods and apparatus for controlling and shaping an atmospheric skip reentry trajectory for a space vehicle. Embodiments of the invention may utilize a powered explicit guidance algorithm to provide a closed-loop control method for controlling a space vehicle during a skip reentry maneuver.

Abstract: A composite is produced by the steps of (a) blending a first mixture of metallic powders; (b) compacting the blended first mixture of metallic powders to a plurality of discretely shaped articles; (c) blending a second mixture of metallic powders; (d) mixing the plurality of discretely shaped articles with the blended second mixture of metallic powders to form a precursor blend; (e) compacting the precursor blend; and (f) sintering the precursor blend. The composite has a metallic matrix with embedded shapes dispersed throughout the matrix where the embedded shapes have an incipient liquid phase sintering temperature less than an incipient liquid phase sintering temperature of the matrix.

Abstract: A debris removal management system and a method of removing space debris. In one embodiment, the system includes: (1) a frame, (2) a plurality of net sections coupled to the frame and (3) a plurality of microvehicles coupled to the plurality of net sections and configured to be ejected relative to the frame to deploy the plurality of net sections, the plurality of net sections cooperating to form a net configured to capture space debris.

Abstract: This invention explains what anti-matter and black holes really are and how they can be used in a space ship to achieve space travel.

Abstract: System and method for inducing rapid reentry of orbital debris including determining a spatial extent of the orbital debris, and deploying dust to the orbital debris to enhance the drag on the orbital debris.

Abstract: A levitation system of a vehicle is provided. This system allows a vehicle to levitate and move in all directions. The vehicle can travel in Earth atmosphere. Also this system could be used in space for space travel allowing accelerations never reached before and cover large distances with great autonomy. If a miniaturization of the system is successful, then the system could be applied in a multitude of already existing transport systems, as a complement or as a principal source of movement. Although the purpose of this invention is to levitate and move a vehicle in a three dimensional space, the same system is perfectly applicable to move a vehicle in a two dimension space. The centrifugal force is applied in this system and an internal or external source of energy is necessary to allow the system to operate.

Abstract: A method is provided for controlling a set of at least two satellites, designed to provide a service on a geostationary orbit, rendering said service on a non-geostationary orbit, in which means involved in the performance of said service, installed on board a satellite, are deactivated when the sun can damage it, and means involved in the performance of said service installed on board another satellite of the set are activated, when this is necessary to the continuity of the service.

Abstract: A sky/space elevator vehicle for going from the surface of a planet or moon up into space and back again. The vehicle has a single fuselage comprising multiple floor sections including a top and bottom cone. The movement of the vehicle is controlled by helium filled envelops as well as a combined jet and rocket propulsion system.

Abstract: A high capacity satellite having multiple aspect ratios is configured to have at least two modules. Each module includes a first panel and a third panel facing, in an on-orbit configuration, respectively, toward north and south and having a width in an east-west direction (EW width); and a second and a fourth panel oriented, in the on-orbit configuration, respectively, east and west and having a width in the north-south direction (NS width). Each module has a respective aspect ratio of EW width to NS width, a first module is configured with a first aspect ratio, a second module is configured with a second aspect ratio; and the second aspect ratio is substantially larger than the first aspect ratio. At least one antenna reflector is disposed, during launch, proximate to at least one of the first and third panel of the second module.

Abstract: A method of establishing a formation of a plurality of directionally fixed satellites, such as primary formation of the satellites with in-track velocity disparity. The plurality of satellites includes an anchor satellite and at least one non-anchor satellite, each satellite of the plurality of satellites including at least one drag flap, each satellite of the plurality of satellites being free of one of a propulsion system and an expendable mass propellant. Control is applied to deploy the at least one drag flap of the at least one non-anchor satellite based on the current spacing and the current separation rate of the each non-anchor satellite, and the predicted final spacing of the plurality of satellites. The final spacing predicting and the drag flap control applying is repeated until the current separation rate has been nulled to within a threshold value.

Abstract: The invention relates to a satellite having a structure that includes a circular launching interface ring and at least two main planar apparatus-holding walls that are substantially parallel to the axis of the ring and rigidly connected to one another and to the ring. Each wall directly bears on the ring by a base of the wall, and each wall is rigidly and directly attached to the ring by at least one somewhat point-by-point connection on at least one point or area of tangency or intersection with the base of the wall and the ring.

Abstract: A pivotal connection and a substantially rigid yet disruptable mechanical connection mount a respective foot pad on a distal end of each one of plural landing legs of a lander spacecraft. The mechanical connection includes at least one predetermined rated breaking point designed to break when the foot pad makes contact with a landing ground surface. At least two sensor elements, preferably electrical conductors or switches, are incorporated in the mechanical connection(s) of one or more foot pads. When the foot pad makes ground contact, the mechanical connection breaks at the rated breaking point(s), which also disrupts or actuates the sensor element(s), which thus provides a signal to an evaluating unit, which triggers the shut-down of at least one braking retrorocket or retrothruster.

Abstract: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: A system for deploying a plurality of independent payloads with a launch vehicle, the launch vehicle being longitudinally oriented about an axis, includes a plurality of launch modules. Each respective launch module of the plurality of launch modules includes a plurality of package elements. A respective predetermined set of package elements of the plurality of package elements comprises a respective payload of the plurality of payloads. Each respective launch module of the plurality of launch modules is situated in an abuttingly supporting relation with at least one adjacent launch module in a facing relation at an interface in a launch orientation. The interface is generally parallel with the axis. At least one restraining member cooperates with the plurality of launch modules to effect compressingly maintaining the plurality of launch modules in the launch orientation.

Abstract: Provided are an attitude control system and method of a spacecraft of an artificial satellite that may enhance a maneuverability and a controllability by simultaneously applying a reaction wheel and a thruster among drive units used to maneuver an attitude of the spacecraft of the artificial satellite. The attitude control system may include: a thruster-based attitude controller which control firing time of thrusters mounted on the spacecraft; and a reaction wheel-based attitude controller controlling driving of a reaction wheel mounted on the spacecraft. The spacecraft may include a plurality of reaction wheels. When a defect occurs in the spacecraft due to a partial malfunction of the reaction wheels, an attitude maneuverability of the spacecraft may be corrected by simultaneously applying the thruster-based attitude controller and the reaction wheel-based attitude controller.

Abstract: A single center elevator, circle inner-ring, circle outer-ring elevator, vertical fuselage flight, vertical plan multi-floor section, bottom escape cockpit, top escape cockpit, mass-levitate, ascend, descend flight, first, second, third, fourth section, control tower, bed, food service, lavatory, landing pad, winch machine, jet/rocket fuel tank, circle passage hole, rope anchor hook, earth terminal station, center-hanger, up nozzle, down nozzle, Jet/rocket power plants, inner-docking bays, outer docking bays, multi robotic arms, cone shaped circular, glue, cylindrical multi-soft envelopes, cylindrical thick outer wall envelope, ceramic fiber fabric, geometric form, web-mesh carbon wire, upper chute-flap, down chute flap, helium gas, float, circle center post pole, truss bridge, top circle truss, hover, anchors, parking zone, V-shaped circle, laser, floor gate, test-gauge, repair/inspection, guy-wires, heavy grade mesh-ceramic fiber, launch/docking platforms, light grade web-mesh ceramic fiber, cargo-transport

Abstract: In a method for forming a structural component of an aircraft or spacecraft a preform comprising at least one band is formed by means of braiding around a core. A severing cut is introduced at least in portions into the at least one band of the preform to form two band portions of the at least one band. The two formed band portions of the at least one band are then deformed into a predetermined shape by means of a deformation tool which can be guided along the severing cut. A deformation device is used to carry out the method. A structural component is produced by means of the method and/or the deformation device.

Abstract: Disclosed is a method for clearing space debris in geocentric orbit, wherein, in geocentric orbit, a spatial region having a greater resistive force than atmosphere in the geocentric orbit is provided, space debris passing through the spatial region is decelerated by the resistive force, and the decelerated space debris is dropped from the geocentric orbit.

Abstract: Disclosed herein, in certain embodiments, is a method of altering the stability of unstable space debris. In some embodiments, the method further comprises changing the orbit of the unstable space debris.

Abstract: A method comprises removing space debris having a relatively low ballistic coefficient by hastening orbital decay of the debris. A transient gaseous cloud is created at an altitude of at least 100 km above Earth. The cloud has a density sufficient to slow the debris so the debris falls into Earth's atmosphere.

Abstract: A system and method for assessing the risk of conjunction of a rocket body with orbiting and non-orbiting platforms. Two-body orbital dynamics are used to initially determine the kinematic access for a ballistic vehicle. The access may be represented in two ways: as a volume relative to its launcher and also as a geographical footprint relative to a target position that encompasses all possible launcher locations.

Abstract: A polygonal support structure is disclosed herein. The polygonal support structure has a plurality of rigid segments configured to be joined at longitudinal ends by a joint affixed to at least one end of each rigid segment. Each joint is configured to rigidly, rotatingly or hingedly join each rigid segment to an adjacent rigid segment to form the polygonal support structure. A plurality of flexible members are configured and disposed to provide tension between each joint and joints spaced two segments therefrom, in each direction, about the polygonal support structure. One or more joints may be configured to join and un-join two adjacent rigid segments and each rigid segment and each joint may be configured to enable stacking of rigid segment into a compact configuration for transport. In at least one aspect, a polygonal support structure may be configured to support assemblies in a substantially weightless environment.

Abstract: The disclosure provides in one embodiment a method of insulating a structural frame member of a transport vehicle. The method comprises providing at least one flexible insulation member. The method further comprises providing at least two self-retaining foam insulation members. The method further comprises wrapping the flexible insulation member over a first end of a structural frame member of a transport vehicle and over at least portions of two sides of the structural frame member. The method further comprises positioning one self-retaining foam insulation member on each side of the wrapped structural frame member and in compression fit between each side of the wrapped structural frame member and adjacent structural frame members, such that the self-retaining foam insulation members impart one or more compression forces on the flexible insulation member to secure the flexible insulation member in place without use of any fastener device.

Abstract: [OBJECT] It is an object to solve a problem that a light-emitting element and a large-capacity battery required for allowing light generated from a satellite to be visually observed from the earth lead to enormously high costs. [SOLUTION] The present invention provides a satellite which comprises a reflecting mirror for reflecting sunlight toward the earth, and a transceiver, wherein the transceiver is operable to receive information including at least a position of the sun, a position of the satellite, and an irradiation point on the earth surface to be irradiated with the reflected light, so as to set a direction of a reflecting surface of the reflecting mirror based on the received information, and wherein the satellite is adapted to orient the reflecting surface of the reflecting mirror in the direction set based on the received information.

Abstract: An air-based vertical launch system is described by means of which ballistic missile defense can be achieved effectively from a large aircraft. A method for ensuring safe missile egress is proposed. A method for ensuring that the missile strikes the ballistic missile payload section is also proposed. Together, the air basing method employing vertical (or near-vertical) launch and semi-active laser guidance yield an affordable and operationally effective missile defense against both tactical and long-range ballistic missiles. The affordability of missile defense is enhanced by the ability of an aircraft equipped with a vertical launcher to simultaneously carry out several defensive and offensive missions and to provide other capabilities such as satellite launch at other times.

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: A high capacity satellite having multiple aspect ratios is configured to have at least two modules. Each module includes a first panel and a third panel facing, in an on-orbit configuration, respectively, toward north and south and having a width in an east-west direction (EW width); and a second and a fourth panel oriented, in the on-orbit configuration, respectively, east and west and having a width in the north-south direction (NS width). Each module has a respective aspect ratio of EW width to NS width, a first module is configured with a first aspect ratio, a second module is configured with a second aspect ratio; and the second aspect ratio is substantially larger than the first aspect ratio. At least one antenna reflector is disposed, during launch, proximate to at least one of the first and third panel of the second module.

Abstract: The invention relates to a database and a method of generating a database containing topographical information on obstacles distributed over a geographic area subdivided into cells. The method includes reading topographical data obtained from heterogeneous sources. Topographical data are generated which relate to pointlike obstacles from the information collected from the topographical data. Topographical data are generated which relate to linear obstacles from the information collected from the topographical data. The obstacle database is created. Each cell contains a link to the obstacles whose coordinates in the geographic area covered by the cell are recorded with an uncertainty value. In particular, the invention applies to the generation of a database on board an aircraft.

Abstract: Disclosed herein, in certain embodiments, is a method of altering the stability of unstable space debris. In some embodiments, the method further comprises changing the orbit of the unstable space debris.

Abstract: An apparatus and method for estimating future pitch and roll angles of an aircraft to compensate for a time delay caused by a flight instrumentation system calculating the aircraft's pitch and roll. The apparatus may output signals to an aircraft autopilot system. The apparatus may comprise various inputs and outputs, at least one compensation module, and at least one addition module. The compensation module may multiply a gain value by the pitch rate of the aircraft and by the roll rate of the aircraft to determine a pitch compensation amount and a roll compensation amount. Then the pitch compensation amount may be added to the pitch of the aircraft, and the roll compensation amount may be added to the roll of the aircraft to solve for a compensated pitch and a compensated roll to send to the autopilot system.

Abstract: A cryogenic propellant depot and sunshield are provided for operation in earth orbit to fuel or refuel other space vehicles. The sunshield is deployed to effectively mitigate solar radiation emanating from the earth and the sun thereby providing a long term storage solution for cryogenic fluids prone to boil-off. The depot has supporting subsystems to include station keeping equipment and communication equipment so that the depot can be independently controlled. Inflatable booms are used to deploy the sunshield in a desired pattern around the depot.

Abstract: A cryogenic propellant depot and sunshield are provided for operation in earth orbit to fuel or refuel other space vehicles. The sunshield is deployed to effectively mitigate solar radiation emanating from the earth and the sun thereby providing a long term storage solution for cryogenic fluids prone to boil-off. The depot has supporting subsystems to include station keeping equipment and communication equipment so that the depot can be independently controlled. Inflatable booms are used to deploy the sunshield in a desired pattern around the depot.

Abstract: This invention presents a missile defensive system based on existing technology in which numerous non-explosive, meteor-like projectiles are launched in a shotgun manner from an orbiting space station at an enemy's missiles while on earth or in flight. Collateral damage is minimized with this system. This Spaced-based Anti-Missile System (SAMS) requires international cooperation. An International Orbital Treaty Organization (IOTO) is proposed.

Abstract: A space-based occulter is selectively movable between stowed and deployed configurations. A base hub has a longitudinally oriented center axis. A plurality of petals are provided, each petal extending radially outward from, and spaced radially apart from, the base hub when in the deployed configuration. Each petal has a hypergaussian edge made up of a plurality of hypergaussian edge panels. Each hypergaussian edge panel defines a portion of the hypergaussian edge of the petal. Each hypergaussian edge panel is made substantially of a material having a near zero coefficient of thermal expansion. A circular central portion is located radially between the base hub and the plurality of petals. A covering skin extends laterally between the hypergaussian edges of each petal and over the circular central portion. The covering skin defines a substantially planar surface oriented substantially perpendicular to the center axis when the occulter is in the deployed configuration.

Abstract: A fairing includes three sections: an aft section, a middle section, and a nose section. The aft section includes a first surface having a generally constant conic angle (i.e., a “boat-tail angle”) with respect to the longitudinal axis such that the aft section tapers to a first end configured to attach to a generally cylindrical body. The middle section, which intersects and is axially aligned with the aft section, has a second surface characterized by a constant elliptical cross-section along a plane orthogonal to the longitudinal axis. The nose section intersects and is axially aligned with the middle section. The nose section is further defined by four generally concave trianguloid surfaces, each extending from the middle section to a common apex intersecting the longitudinal axis, wherein each adjacent pair of trianguloid surfaces intersects at an edge (e.g., an incurvate edge).

Abstract: A space probe including a descent module with a mobile exploration vehicle, wherein the descent module is a landing module inside which the mobile exploration vehicle is fastened, the landing module being provided with landing legs which can be deployed under the lower level of the mobile exploration vehicle.

Abstract: Fiber-based debris interceptors are used to intercept and/or contain space debris. The debris interceptors may be made up of fibers that are formed in space from a material supply on a space vehicle. The fibers for the debris interceptors may be formed by extrusion, with a heat source, such as a mirror to focus sunlight, used to heat the material of the material supply that is carried on the space vehicle. The debris interceptor may be separated from the space vehicle, and used to remove debris from an orbit, or otherwise prevent debris entering an orbit from damaging a satellite or other spacecraft that travels in that orbit. The debris interceptor may be deployed prior to later launch of a valuable spacecraft, in order to “cleanse” the intended orbit of debris. Debris objects may pass through the debris interceptor, but in so doing may lose energy so as to de-orbit.

Abstract: A method is adapted for slewing a body from an initial state to merge with a desired state profile. In an embodiment, the method includes the steps of (a) obtaining data indicative of a time profile of desired future states for the body; (b) searching sequentially through the time profile of desired future states; (c) calculating several slews for the body from the initial state to each of the desired future states; (d) determining the earliest future state to which the body can actually slew; (e) choosing the minimum-energy slew to that state; and (f) causing the body to perform that chosen slew.

Abstract: The invention is an aircraft having an inflatable wing connected to a base unit, with the inflatable wing inflated with a lifting gas such as helium. The inflatable wing has a series of cell structures, and may be configured with ballonets to selectively introduce and expel outside air within the inflatable wing to vary the buoyancy and/or airfoil properties of the inflatable wing. The aircraft is particularly useful at low speeds and in thin atmospheres (such as at high Earth altitudes and on Mars), and can be used for interplanetary missions to explore planetary bodies, such as moons and planets, having atmospheres.

Abstract: An improved approach to satellite-based navigation (e.g., GPS) is provided. In one embodiment, a method includes determining a nominal orbital path of a navigation satellite. The method also includes transmitting ephemeris data corresponding to the nominal orbital path from the navigation satellite to a plurality of navigation devices. The method further includes determining an actual orbital path of the navigation satellite locally at the navigation satellite. In addition, the method includes determining a deviation between the actual orbital path and the nominal orbital path locally at the navigation satellite. The method also includes autonomously adjusting the actual orbital path locally at the navigation satellite to reduce the deviation between the actual orbital path and the nominal orbital path.

Abstract: A system for displaying throttle settings to an aircraft operator includes, but is not limited to, a display unit, a data storage unit configured to store aircraft and airport data, an operator input unit configured to receive thrust related information from the aircraft operator, an aircraft sensor configured to detect an aircraft condition, and a processor. The processor is communicatively connected to each of the above components and is configured to obtain therefrom the aircraft data, the airport data, the thrust related information, and the aircraft condition. The processor is also configured to calculate a first thrust requirement based on the aircraft data, the airport data, the thrust related information, and the aircraft condition, and to provide commands to the display unit to display either the first thrust requirement.

Abstract: A rotary casing of a satellite antenna having an angle adjustable display screen includes a supporting arm, a hollow casing, a rotating means and an adjusting and controlling assembly. The casing has a panel provided with two keys and a top surface provided with a window. The rotating means is provided in the casing and controlled by the keys to drive the rotation of the supporting arm. The adjusting and controlling assembly is provided in the casing and includes a control unit and a display unit. The two keys are electrically connected to the control unit to drive the supporting arm. By this arrangement, the control unit receives an adjustment instruction inputted by the keys and transmits the instruction to the rotating means to generate a corresponding amount of rotation, while the control unit converts the instruction into a numerical value to be displayed on the display unit.

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 either a short-band imaging sensor and headlamp or a MWIR 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. An implicit divert and attitude control system (DACS) using tow or more divert thrusters performs KV divert and attitude maneuvers to respond to the command guidance pre-handover and to maintain track on the aimpoint to terminal intercept post-handover.

Abstract: An apparatus for trapping high-speed particles, such as space debris, is provided. The apparatus includes a first front tapping surface and at least one rear trapping surface. The front and rear trapping surfaces are arranged substantially parallel to each other and are spaced apart from each other in the normal direction.

Abstract: An artificial satellite system 200 comprises at least two panel type artificial satellites 100 closely connected with a plurality kind of multi networks and the minimum necessary essential tools for one artificial satellite installed on a substrate of each panel type artificial satellites 100, wherein the multi networks are formed by a communicating network for communicating data among data processing tools installed on each the panel type artificial satellites 100, a heat pipe for providing thermal energy among tools installed on each the panel type artificial satellites 100 and a battery supplying line for connecting battery controlling devices installed in the each panel type artificial satellite so that the artificial satellite system 200 can maintain higher and various controlling performance with a back-up effect of the multi network in the case that any function tool is damaged in one of the panel type artificial satellite 100.

Abstract: The present invention is a space launch system and method to propel a payload bearing craft into earth orbit. The invention has two, or preferably, three stages. The upper stage has rocket engines capable of carrying a payload to orbit and provides the capability of releasably attaching to the lower, or preferably, middle stage. Similar to the lower stage, the middle stage is a reusable booster stage that employs all air breathing engines, is recoverable, and can be turned-around in a short time between missions.