541 results about "Space vehicle" patented technology

The invention discloses an attitude control system for space vehicle and a method thereof. The control system has only one biased momentum wheel, one set of tri-axial magnetic torquer and one attitude controller loaded with algorithm. The method comprises a step of rate damping controlling, a step of initially capturing controlling and a step of stationarity controlling. At the rate damping stage, geomagnetism change is used to control the magnetic control of three passages of a satellite by B-dot; at the initially capturing stage, the magnetic control is realized, PD control is performed by pitching and the passages are rolled and yawed to carry out nutation and precession composite control; at the stationarity controlling stage, the magnetic control is realized, PD control is performed by pitching and the passages are rolled and yawed to carry out nutation and precession composite control. The capturing stage and the stationarity controlling stage fully depend on magnetic torquer to perform positive magnetic control, thereby changing which a satellite only uses a magnetic torquer to carry out unload of the momentum wheel or auxiliary magnetic control, so as to refine system configuration to further improve reliability of the system. Momentum of a satellite is biased to rotate on the ground, so as to ensure stable separation of the satellite without performing air injection control. Therefore, the magnetic torquer can be used for realizing fast and stable initial rate damping.

The invention discloses a fault-tolerant sliding-mode control method for a near-space vehicle. According to the fault-tolerant sliding-mode control method for the near-space vehicle, for the situation that the order of the magnitude of external disturbance in a quick loop and slow loop system is greatly larger than the order of the magnitude of uncertain times of the system, the nonlinear disturbance observer technology is used for processing hybrid disturbance, and unknown hybrid disturbance is estimated by a disturbance observer through known system information; in order to solve the problem that saturation of the control surface of the near-space vehicle is limited, the upper bound of the deflection angle output of a steering engine is applied to design of a control law, it is guaranteed that the input is within a certain range, auxiliary variables are designed, the deflection angle output of the steering engine is automatically adjusted through the self-adaption law, and therefore the situation that when the upper bound of the deflection angle is too large, the output is too large is avoided; a compensator is established through a radial basis function neural network and is used for fault-tolerant compensation when the steering engine breaks down, and therefore the problem that the steering engine of the near-space vehicle breaks down is solved. By the adoption of the fault-tolerant sliding-mode control method for the near-space vehicle, under the conditions of system uncertainties, unknown external disturbance, limited input saturation and a fault of the steering engine, the near-space vehicle has good control performance.

A camera (1), in particular in a space vehicle, having a housing (2) which contains at least one optically sensitive exposure surface (7), and a base lens (9), having a first fixed focal length, connected thereto in each case, and which projects on the at least one exposure surface (7) and which is situated on a first optical axis (10) for the exposure surface (7). To provide the camera with various fields of view, in particular for the approach of two satellites toward one another over large distances, at least two afocal supplementary lenses (11, 12) which are each parallel with respect to their optical axes (15, 16) and spaced at a distance from the first optical axis (10) are situated in the housing (2), whose optical paths are alternately coupleable with the aid of a pivotable prism (17), to form further fixed focal lengths in an optical path of the base lens (9).

An automated docking system for space vehicles that includes a plurality of antennas on each of a target vehicle and a chase vehicle. A pseudo random code is transmitted via one of the antennas located on one of the vehicles and received by one of the antennas located on the other vehicle. The pseudo random code is then sent back to the original vehicle via transmissions from the plurality of antennas on the second vehicle and received by the plurality of antennas on the original vehicle. The distance from each of the antennas on the other vehicle to each of the antennas on the originating vehicle can be measured in this fashion. The antennas on each of the vehicles are located in a spaced-apart arrangement so that the angular orientation or attitude of each of the vehicles to each other can also be determined. A plurality of video cameras is provided on the exterior of one of the vehicles and video information from these cameras is transmitted to the other vehicle for display to operators in that vehicle. Commands between the vehicles can also be communicated.

A door frame component of cast titanium for a door opening in an outer skin of an aircraft or space vehicle is provided, the component including a contact surface for bearing against the outer skin, wherein the contact surface is formed to correspond to an inner contour of the outer skin in such a manner that the contact surface essentially fills the contour. From a further viewpoint the invention makes available a structural fuselage part for an aircraft or space vehicle which comprises an outer skin which has a carbon fiber reinforced plastic material together with a door frame component of the type described.

A new family of multifunctional smart coatings based on of stabilized diamond-like metal-carbon atomic scale composites and diamond-like atomic-scale composite (DL ASC) materials. Based on a unique combination of the coating fine structure, properties of the coating / substrate interface, and the mechanical and electrical properties of the coating, the disclosed smart coatings would integrate various high resolution sensors and interconnections, and the sensor would diagnose dangerous stress distribution in the coated subject with no distortion in real time, while these diamond-like coatings would simultaneously provide environmental protection of the coated surface and improve its aerodynamic quality.

A system and method provide for the collection of interferometric synthetic aperture radar (IFSAR) data. In the system, a first space vehicle configured for emitting electro-magnetic energy and collecting the reflection from a region of interest (ROI), may be directed along a first orbital track. The collected image data may be stored and later provided to a ground station for image and interferometric processing. A second space vehicle may also be configured for emission and collection of electro-magnetic energy reflected from the plurality of ROI's. The second space vehicle is positioned in an aligned orbit with respect to the first space vehicle where the separation between the vehicles is known. In order to minimize decorrelation of the ROI during image processing, the lead and trail satellite are configured to substantially simultaneously emit electromagnetic pulses image data collection. In order to avoid interference in the collection of this image data, each system is configured to control the emission of pulses such that the receipt of direct and bistatic pulses by the other vehicles does not interfere with data collection.

An adaptor system includes a space vehicle separation plate operably coupled to a space vehicle, a launch vehicle adaptor plate operably coupled to a launch vehicle capable of carrying the space vehicle into space for release of the space vehicle from the launch vehicle, an actuator release mechanism assembly, a bend / shear restrain assembly that is non-coaxial with the actuator-release connector, and a biasing element. The actuator-release mechanism assembly may be configured to separably couple the space vehicle separation plate to the launch vehicle adaptor plate. The actuator-release actuator-release mechanism assembly may pass through the launch vehicle adaptor plate to engage the space vehicle separation plate to hold the space vehicle separation plate substantially parallel to the launch vehicle adaptor plate prior to release of the space vehicle separation plate.

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.

A method and associated system for multi-disciplinary optimization of various parameters associated with a space vehicle that experiences aerocapture and atmospheric entry in a specified atmosphere. In one embodiment, simultaneous maximization of a ratio of landed payload to vehicle atmospheric entry mass, maximization of fluid flow distance before flow separation from vehicle, and minimization of heat transfer to the vehicle are performed with respect to vehicle surface geometric parameters, and aerostructure and aerothermal vehicle response for the vehicle moving along a specified trajectory. A Pareto Optimal set of superior performance parameters is identified.

A joint in a space vehicle to hold together a first part of the vehicle and a second part of the vehicle that can be separated from the first part, in which the joint includes a first flange member (1) attached to the first part of the vehicle and a second flange member (2) attached to the second part of the vehicle. A number of clamps (4) equipped with two clamping lips (5, 6) are distributed around the perimeter of the flange members (1, 2) that transfer radial force from a tensioning means (7), arranged around the clamps (4), to both of the flange members (1, 2) in order to press these together, whereby rolling members (8) are arranged between at least one of the clamping lips (5, 6) of the respective clamps (4) and the adjacent first (1) or second (2) flange member.

The invention belongs to the technical field of space propulsion. The technical problem to be solved is to provide a realization method of the space propulsion system, which can effectively overcome the defects of the existing technology, and can adapt to the development requirements of long service life and high safety of space vehicles in orbit. The present invention has designed a "well" structure gas path supply system and a cross-supply propellant supply system, and can realize repeated filling / leaking of the propellant on the ground metal membrane box type storage tank, and the present invention brings the following beneficial effects: 1) The air cavities of the two propellant tanks are physically isolated and not connected to each other; 2) Solve the fault isolation of gas leakage, and the system will not fail to work due to the leakage of a gas cylinder; 3) Meet the long-term on-orbit aircraft to adapt to multiple times Supplemented tank design requirements.

Simulation and analysis of conceptual flight vehicles in an early design phase is conducted in a framework for evaluating performance and control effectiveness that is captured by multiple design parameters modeling the effects of control and disturbance moments and forces acting upon a conceptual flight vehicle along a specified trajectory. Characteristics of multiple types of flight vehicle effectors are modeled as a system and adjusted in such a framework to converge to a controllable air frame configuration for conceptual flight vehicle under static design considerations prior to performing dynamic analyses and control system simulations.

Each of a first and a second navigation satellite system (NSS) are adapted to operate according to a first and a second specification, respectively, and each includes a first and a second plurality of space vehicles (SV), respectively. Each of the first and the second plurality of SVs are adapted to be identified by a first and a second plurality of unique corresponding identifications (IDs), respectively. A processor is adapted to receive and identify a first plurality of corresponding signals transmitted from the first plurality of SVs in response to the first plurality of unique corresponding IDs. The processor is adapted to receive and identify a second plurality of corresponding signals transmitted from the second plurality of SVs in response to the second plurality of unique corresponding IDs. The processor is adapted to determine position location information in response to receiving and identifying the first plurality of corresponding signals and the second plurality of corresponding signals.

Novel transparent composites composed of single wall carbon nanotubes incorporated into the matrix of a polymer are utilized in services wherein the composites are exposed to ionizing radiation, including galactic cosmic radiation. Accordingly, the composites are useful in deep space applications like space vehicles, space stations, personal equipment as well as applications in the biomedical arts and atom splitting research. The composites can be modified with organic dyes containing at least one phenyl ring and the resulting doped composite is useful as a radiation detector. The preferred polymer is poly(4-methyl-1-pentene).

A vehicle detector system having a number of individual vehicle detectors each capable of sampling a plurality of vehicle loops in mutual synchronization. One detector operates as a master detector for synchronization purposes; the other detectors are operated as slave detectors. The system can be configured for series or parallel synchronous operation. The system is particularly advantageous in installations requiring a large number of closely spaced vehicle loops each operated by a detector set to high sensitivity.

A thermal protection system suitable as a heat shield for a space vehicle includes a flexible insulation mat that is adhesively bonded onto the outer skin of the space vehicle, a cover layer made of at least one ply of a ceramic fiber composite material secured onto the insulation mat, and a coating of an inorganic material applied onto an outer surface of the ceramic cover layer. The cover layer is secured to the insulation mat by adhesive bonding, stitching with ceramic threads or the like, or by a snap fastener button connection. The ceramic fiber composite materials of the cover layer include inorganic fibers which are preferably oxides, carbides or mixtures thereof embedded in a matrix of the same compositional components, whereby the components can be selected appropriately depending on the requirements of the particular application. The overall thermal protection system remains flexible, lightweight, and highly effective for thermal shielding, so that it may be used not only on the relatively cooler portions of the space vehicle, but also on the most highly thermally loaded areas such as the nose and the leading edges of wings and the like.

The invention relates to a separation mechanism of on-board equipment, comprising a mounting base body, a plurality of compressing clamping blocks, a connecting flange, a compressing and releasing mechanism and an unlocking device; the compressing and releasing mechanism comprises a connecting shaft and a connecting spring, the unlocking device comprises a locking dog, a locking dog shaft, a tension spring, a pre-tightening rope and a flame cutter; the locking dog is adopted to lead the direction of the pre-tightening force of the mechanism, directly acted on the compressing clamping block, to be changed, so as to lead the direction of the pre-tightening force to be downwards; the force is transmitted to the compressing clamping block by the connecting shaft, and the structure is simpler and is easier to realize compared with the prior art; in the invention, a compression matching way of the compressing clamping blocks is adopted without high precision of a space vehicle or target load at the connecting position for separation, thereby effectively reducing the machining difficulty of the space vehicle or the target load at the connecting position for separation; a way that multiple groups are matched is adopted, so as to achieve the using requirements of universalization, modularization and serialization.

The invention provides a deployable radiator for a space vehicle such as a satellite. The radiator of the invention has two radiating faces, one of which is insulating over a portion of its area so that during launch and while the satellite is being moved onto station, with the radiator in the stored position, it naturally serves to keep the highest possible temperature on board the vehicle.

The method serves to place a space vehicle, such as a satellite, on a target orbit such as the orbit adapted to normal operation of the space vehicle and starting from an elliptical initial orbit that is significantly different from, and in particular more eccentric than the target orbit. The space vehicle is caused to describe a spiral trajectory made up of a plurality of intermediate orbits while a set of high specific impulse thrusters mounted on the space vehicle are fired continuously and without interruption, thereby causing the spiral trajectory to vary so that on each successive revolution, at least during a first stage of the maneuver, perigee altitude increases, apogee altitude varies in a desired direction, and any difference in inclination between the intermediate orbit and the target orbit is decreased, after which, at least during a second stage of the maneuver, changes in perigee altitude and in apogee altitude are controlled individually in predetermined constant directions, while any difference in inclination between the intermediate orbit and the target orbit continues to be reduced until the apogee altitude, the perigee altitude, and the orbital inclination of an intermediate orbit of the space vehicle have substantially the values of the target orbit.

An automatic identification system (AIS) communications payload for a space vehicle traveling in earth orbit includes an antenna; a receiver connected to the antenna to receive an AIS signal including a reported vessel position transmitted from a vessel; a processor to process the received AIS messages and determine the accuracy of the reported vessel position; and a transmitter to transmit the processed AIS messages and the determined accuracy to a ground earth element. The processor may be configured to determine a position of the space vehicle; compare the reported vessel position to the space vehicle position; and determine the accuracy of the reported vessel position based on the comparison. The processor also may be configured to flag the AIS message as suspect when the reported vessel position is determined to be inaccurate.

A reusable space vehicle docked in an intermediate orbit for rescue missions allows a satellite to be serviced with less delay, energy expenditure, and cost than a space vehicle launched from Earth for each mission. The reusable repair vehicle can be moved from one orbit to another with minimal energy expense while not having to wait for a launch window. Once a servicing need is identified, a destination orbit is identified for the space vehicle and a minimum energy path is identified. If the time to the next launch window between the docking orbit and the destination orbit happens to be near enough to allow for a timely rendezvous, the space vehicle is moved directly to the destination orbit. The space vehicle can be a vehicle designed to be piloted by humans or telerobotically. In one implementation, the inactive space vehicle is docked in an Intermediate LEO orbit (altitudes of approximately 250 km to approximately 500 km) and is used to rendezvous with objects in High LEO orbits (altitudes of approximately 500 km to approximately 1500 km) or objects in Low LEO orbits (altitudes of approximately 250 km or less). The space vehicle can be a modified lunar lander.

A null slit antenna pattern receives GPS signals from at least three GPS satellites functioning as pseudo stars for spin rate and spin axis attitude determination of a spinning space vehicle in an earth orbit below the GPS satellites.

A method of launching space vehicles by towing them aloft, then twirling them around a large transport aircraft (40) at the center of a formation (AA) of other tow aircraft (28, 34) and other devices of the invention. A lengthy, semi-rigid tow pipeline (14) serves as a conduit for the transfer of fuels and oxidizers, as the tow cable, and as an energy storage device that reflexes efficiently when it is flexed. The flexing of tow pipeline (14) is caused by a parachute (22) acting in conjunction with all the aircraft making the tighest turn they are capable of doing. Tow aircraft in certain arrays (28) are joined to tow pipeline (14) by sliding trollies (26) that also host canard rotor wings for the aerodynamic support of the main tube (12). The tow trollies (26) aid the sliding tow aircraft arrays (28) in gaining mechanical advantage to accelerate the space vehicle. The space vehicle may also burn its own motors for a longer-than-usual time as it gains in angular velocity because its fuels are replenished by a pipeline that automatically increases the pressure and volume of fuel flow as the angular acceleration increases. The tow pipeline (14) features a micro-hole laminar lift foil (50) on the top surface of the wing and on the bottom. By balancing the vacuum level in the plurality of chamber underneath the plurality of lift foil 50, the pilot or autopilot has an effective way to control the flight characteristics of the tow pipeline (14). In some embodiments the micro-hole laminar lift foil (50) on the top and bottom will made of transparent hollow wire segments or other material so arranged as to pass ambient light through to photo-voltaic cells below that convert sunlight to electrical energy.

The invention discloses a rolling compounding method for an aluminum / magnesium / titanium three-layer composite plate. The characteristics of high strength of titanium alloy, low melting point and good vibration reducing performance of magnesium alloy, good plasticity and space adaptation of the aluminum alloy, and the like, are utilized and the excellent characteristics of respective component metals are combined by the aluminum / magnesium / titanium three-layer composite plate provided by the invention, so that the high-speed impact damage resisting property which the component metals do not own is acquired. Through the control on the ratio and interface of the aluminum alloy, the magnesium alloy and titanium alloy, the degree of fragmentation, melting and atomizing of shred cloud after the buffer composite plate is subjected to high-speed impact is changed and adjusted, so that the form of the shred cloud is changed, the degree of impact damage to a backplate is reduced, and the high-speed impact damage resisting property is increased. The aluminum / magnesium / titanium three-layer composite plate is applied to an outer layer buffer plate in an outer layer protecting structure of an aerospace vehicle and can be used for replacing a traditional single aluminum alloy material. The space environmental effect and protecting level of the space vehicle are increased; the design level of the space vehicle are greatly promoted, and the service life of the space vehicle are greatly prolonged.

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.

A precursor of a ceramic adhesive suitable for use in a vacuum, thermal, and microgravity environment. The precursor of the ceramic adhesive includes a silicon-based, preceramic polymer and at least one ceramic powder selected from the group consisting of aluminum oxide, aluminum nitride, boron carbide, boron oxide, boron nitride, hafnium boride, hafnium carbide, hafnium oxide, lithium aluminate, molybdenum silicide, niobium carbide, niobium nitride, silicon boride, silicon carbide, silicon oxide, silicon nitride, tin oxide, tantalum boride, tantalum carbide, tantalum oxide, tantalum nitride, titanium boride, titanium carbide, titanium oxide, titanium nitride, yttrium oxide, zirconium, diboride, zirconium carbide, zirconium oxide, and zirconium silicate. Methods of forming the ceramic adhesive and of repairing a substrate in a vacuum and microgravity environment are also disclosed, as is a substrate repaired with the ceramic adhesive.