79 results about "Space craft" patented technology

This invention provides a coupling method for coupling a first and a second stiffening profile element for an outer skin of an aircraft or space craft. In a first step the stiffening profile elements are arranged on the outer skin in such a manner that the stiffening profile elements oppose each other with their respective front sides and enclose within themselves a cavity. Furthermore, a fixing hole is formed through a wall of at least one of the stiffening profile elements in the cavity inside a coupling zone of the stiffening profile element. An access opening is formed through the wall in the cavity along the coupling zone. A coupling strap, which couples the stiffening profile elements together, is fastened to the coupling zone by means of a fastening element guided through the fixing hole.

The invention discloses a non-cooperative target satellite acquisition device and an acquisition method, and belongs to the technical field of aerospace. The non-cooperative target satellite acquisition device includes an outer shell, two four-link capturing pawl mechanisms, four horizontal locking mechanisms, a screw rod and a controller; the two four-link capturing pawl mechanisms are arranged in the outer shell face to face, and the lower parts thereof are fixedly connected with an installing bottom plate; the lower end of the screw rod crosses through the installing bottom plate, and the lower end thereof is connected with a driving mechanism; four horizontal locking mechanisms are respectively located at outside of the two four-link capturing pawl mechanisms, and every two horizontal locking mechanisms are installed on the upper part of the outer shell; both sides of the upper part of the four-link capturing pawl mechanisms are respectively equipped with pressure lever mechanisms; when the four-link capturing pawl mechanisms are descended, the pressure lever mechanisms compress the horizontal locking mechanism to horizontally glide towards inside; the top part of the outer shell is equipped with a photoelectric sensor, and the photoelectric sensor and the driving mechanism are electrically connected with the controller. The non-cooperative target satellite acquisition device is widely applied to the acquisition of various space crafts, and the non-cooperative target satellite acquisition device is wide in application range, big in tolerance, quick in response speed, high in reliability, and others.

The present airship-shaped space craft has a middle fuselage extending in a fore-and-aft direction, and a pair of two outer fuselages extending in the fore-and-aft direction located symmetrically on both sides of the middle fuselage. In the above fuselages, gas of a specific gravity of which is smaller than that of air is filled, and the middle fuselage is connected with the outer fuselages by a horizontal wing. The horizontal wing is provided with propelling devices supported in a gimbal fashion for generating thrust in any optional direction, jet engines with backwards directed nozzles to be controlled within a range from a slantwise upward direction to a slantwise downward direction, and rocket engines with ejection nozzles to be controlled in right-and-left and up-and-down directions. During ascent of the space craft, at first the propelling devices, then the jet engines and at last the rocket engines are actuated so as to make the space craft reach and fly along a satellite orbit. Upon return of the space craft, the rocket engines, the jet engines and the propelling devices are actuated in reverse order, and aerodynamic heating at reentry into the atmospheric space can be reduced by making the space craft descend slowly. Also, control of the flight in the atmospheric space becomes easy, so that the space craft can safely and easily land on a predetermined narrow area of the ground.

A system (30) for adjusting the orientation of a spacecraft adapted for use with a satellite (10). The system (30) includes a first control circuit (32, 38, 40) for canceling any momentum of the spacecraft via a counter-rotating spacecraft bus (16, 18). A second controller (32, 42, 44, 46, 48) orients the spacecraft via the application of internal spacecraft forces. In a specific embodiment, the spacecraft bus (16, 18) serves a dual use as storage section and includes a mass (16) having a moment of inertia on the same order as the moment of inertia of the satellite (10). The satellite (10) includes a bus section (16) and a payload section (14). The mass (16) includes the bus section (16). The first control circuit (32, 38, 40) runs software to selectively spin the mass (16) to cancel the momentum of the satellite (10). The software computes an actuator control signal, via a computer (32), that drives a first actuator (38) that spins the mass (16). The first control circuit (32, 38, 40) further includes a circuit for determining the inertial angular rate of the satellite (10) that includes a gyroscope sensor package (34) in communication with the computer (32). The gyroscope sensor package (34) provides a rate signal to the computer (32) that is representative of the momentum of the satellite (10). The computer (32) runs software for generating the actuator control signal in response to the receipt of the rate signal from the gyroscope sensor package (34). The second controller (32, 42, 44, 46, 48) includes a first reaction wheel (20) having an axis of rotation (26) approximately perpendicular to an axis of rotation (28) of a second reaction wheel (22). The first and second reaction wheels (20, 22) are rigidly mounted to the spacecraft bus (18, 16) and are free to spin about their respective axis. The first and second reaction wheels (20, 22) are selectively spun via first and second actuators (44, 48), respectively, in response to the receipt of first and second steering control signals, respectively.

Recent schemes have been accepted by NASA in its TPF program for simple sparse arrays. But these methods require a great deal of interferometric control of tip tilt and other motion. But with this invention we just need (actuated) flats on a parabola with a central focal point module. The mirrors will be on the order of a kilometer distant from that module so that tip tilt can be sensed (and then controled) by sensors on the module that also detect beam pointing. The actual (high SNR) star the planet is orbiting can be used for rapid Strehle ratio maximization so that NO laser interferometry is needed at all. Fine mirror control can be done by variable reflectivity of the backside of the mirrors. The mirrors can be placed at L2 so that their orbits are all the same. A lot of this multi space craft control has already been accomplished with the ESA "Cluster".

The invention provides a rod pin type blind alignment space butt joint mechanism. An outer truncated cone and an inner truncated cone are installed on two space crafts respectively, one end of a lead screw pushing rod is an outer thread and penetrates through a central hole of a flange to be in threaded connection with a straight thread sleeve, and a boss of the side wall of the lead screw pushing rod is matched with an axial groove formed in the inner wall of the central hole of the flange. The other end of the lead screw pushing rod is a conical face. A radial through hole is formed in the conical face of the outer truncated cone, and a pin is installed in a through hole through a compression spring. The inner wall of the inner truncated cone is provided with an annular groove in the axial direction, the flange is fixed to the bottom face of the outer truncated cone, and the straight thread sleeve pushes the lead screw pushing rod to perform axial motion on the outer truncated cone and pushes the pin to get over the spring and extend out of the outer conical face of the outer truncated cone to be in clearance fit with the annular groove during rotation. The rod pin type blind alignment space butt joint mechanism can perform bidirectional initiative butt joint, is high in butt joint reliability and butt joint speed and can work repeatedly.

The invention provides a rail-changing method which reduces the gravity loss of an aircraft, relating to aircraft rail-changing technical field. The invention comprises the steps of (1) before the rail of the aircraft is changed, rail power starting time, initial gesture and gesture angular speed are calculated on the ground; (2) the three parameters of the rail power starting time, the initial gesture and the gesture angular speed are optimized, thus leading the direction of the thrust to be always close to the speed direction of the aircraft in the rail-changing process of the aircraft; (3) the calculated parameters are input to the star; (4) the aircraft uniformly rotates according to the angular speed determined by the ground instruction parameters. The rail-changing method of the invention leads the rail-changing fuel consumption to be extremely close to the fuel consumption of the rail-changing in the speed direction and can effectively reduce the fuel consumption compared with the fixed thrust direction rail-changing method.

The invention discloses a low-damping suspension hanging type deployment test device which is suitable for horizontal two-dimensional deployment tests of solar wings and antennas of space crafts. The low-damping suspension hanging type deployment test device comprises two air floating long guide rail assemblies (1) fixed to a truss (2), two longitudinal air floating assemblies (3) make contact with the upper surfaces of air floating long guide rails, air floating short guide rail assemblies (4) are connected to the lower ends of the longitudinal air floating assemblies, transverse air floating assemblies (5) make contact with the upper surfaces of air floating short guide rails, and the lower ends of the transverse air floating assemblies are connected with a solar wing (7) through hanging connecting assemblies (6). Clean compressed air with the pressure intensity ranging from 0.5 MPa to 1.0 MPa is input into the transverse air floating assemblies and the longitudinal air floating assemblies through an air distribution system (8) and hoses (9), and tested equipment is supported to be suspended on the air floating long guide rail assemblies and the air floating short guide rail assemblies. By the adoption of the low-damping suspension hanging type deployment test device, the beneficial effects that the space weightlessness condition is simulated and friction resistance in the direction of the air floating long guide rails and friction resistance in the direction of the air floating short guide rails are both smaller than 0.2 N under a 10 kg load condition are obtained.

Disclosed is an attitude change control system that is designed to efficiently output a torque for attitude change of a space craft using CMGs and realizes a real time CMG driving rule. A CMG gimbal steering law 15 generates target profiles for setting angle and angular velocity for each gimbal by applying an anisotropic weighted gradient method based upon the necessary torque calculated by a feed back controller 13 and a feed forward controller 14 from the angle and angular velocity in the target direction from the attitude navigator 12 and the current angle and angular velocity of the space craft estimated by the attitude estimator 11 as well as the current condition of each gimbal from the CMG 40, thereby controlling the CMG 40 for changing the attitude of the space craft dynamics 50 to the target direction.

The present invention provides one micro space generation module with photoelectric converter, thermoelectric converter and plasma-to-electricity converter integrated together, and the module has deposited photoelectronic film and hot ion emitting film, semiconductor photoelectronic element and other micro cell module units. The micro space generation module can realize excellent stepped energy utilization and may be used in various fields, especially as the high efficiency and long life cell for micro space craft and other effective load after being provided with isotope fuel. It may be used in desert and other places with rich solar energy, and may be miniaturized for use in micro electromechanical system.

A model space vehicle having a fixed circular wing, streamlined fuselage, fins with booms attached thereto all efficiently configured to result in an article possessing suitable aerodynamic properties to enable it to glide smoothly through the air for a substantial period of time and substantial distance relative to its size and weight.

The invention provides an attitude and orbit control method based on fore and after arrangement of an engine. In the method, the front end and the rear end of a spacecraft are respectively provided with a translation engine; starting-up of the engines is combined to control the position and attitude of the spacecraft at the transverse moving direction, pitching direction and yawing direction of the spacecraft respectively. According to the barycentre position of the spacecraft, startup factor ki1 of positions of the front and rear engines is calculated by the translating control force according to a principle that magnitudes and directions of thrust generated by the front and rear engines at the same direction of the spacecraft are the same as the requirement of control force, and the magnitudes of torques generated by the front engine and the rear engine are the same; startup factor ki2 of attitudes of the front engine and the rear engine is calculated by the attitude control torque requirement according to the principle that the magnitudes and directions of torques generated by the front and rear engines at opposite directions of the space craft are the same as the control torque requirement, and the magnitudes of the thrust generated by the front and rear engines at the opposite directions are the same; and finally, the total startup factor ki equaling to ki1 plus ki2 of the translation engines are calculated, and the engine is started to complete attitude and orbit control after normalization processing.

The invention concerns a space craft, such as a telecommunication geostationary satellite, comprising a body, at least one main telecommunication antenna having a specific orientation relative to the body, at least a omnidirectional antenna having a field of view opposite to that of the main antennae and at least a radiator extensible by tilting about an axis linked to the craft body between a stowing position wherein it is pressed against the body and an extended position. The axis is positioned substantially in the plane of one surface of the body parallel to the orientation of the omnidirectional antenna and to the surfaces or the surface bearing the main antenna. The radiator tilts at an angle close to 180° from a position wherein it is pressed against one surface bearing a solar panel to a position wherein it extends substantially in the plane of the surface which bears it when it is stowed.

An actuatable emergency exit door with a door actuating device that comprises at least two connection rods that are pivotally mountable to an associated structural frame that is adapted for accommodating the actuatable emergency exit door in closed state, the at least two connection rods being provided for enabling an opening movement of the actuatable emergency exit door with respect to the associated structural frame during opening, wherein at least one door-mounted goose neck-shaped structure is mountable to the associated structural frame for enabling an initial translational opening movement and subsequently a swiveling opening movement of the actuatable emergency exit door with respect to the associated structural frame during opening. The invention is further related to an aircraft or space craft with a pressurized cabin having such an actuatable emergency exit door.

An automated communication system for communicatively connecting a caller with a communications device to an space craft. The system includes a port for accepting a call and a call management program that uses a spacecraft identifier to automatically effect the communication to the space craft via a ground earth station and a satellite. The satellite retransmits the communication to a dedicated spacecraft receiver in the space craft. The system provides Internet-based call monitoring and billing capabilities. The system reduces cost and overall connection time by using the Internet to verify the spacecraft's flying status through the ground station databases before initiating a call.

Methods and an apparatus for a lifeboat of a space station in orbit to be docked with and used in combination with a separately launched logistics module having an upper stage propulsion capability as a space craft for human pilots and passengers to be flown to orbits and trajectories in deep space beyond low Earth orbit including the Moon, at the lowest practical cost.

The invention discloses an autonomous navigation system and an autonomous navigation method thereof based on a laser space communication transceiver. Based on the traditional laser space communication transceiver, a star capturing system is added, so that fixed stars and planets or satellites thereof can be recognized and the azimuth information can be obtained; by calculating the included angle among the fixed stars-space crafts-planets (or planets thereof), the position information of the space crafts can be obtained finally so as to achieve the navigation process. The laser space communication transceiver can obtain the navigation information without a set of additional autonomous navigation device, so that the relevant research, development, launch and running expense is saved, and the volume, weight and power consumption of the effective load of the aircrafts are reduced; the integral performance of the autonomous navigation system is improved; approximate real-time flight track correction is provided for deep space exploration according to the navigation result, so that not only can the signal round-trip delay between satellites and the ground be eliminated, but also ground data processing can be omitted.

The invention discloses a rapidly-inflatable-deployment thin-film support tube for a space craft, and relates to an inflatable-deployment thin-film support tube for a space craft, for solving the problems of high self-weight, limit on the length after deployment, and long deployment time of the present inflatable-deployment thin-film support tube for a space craft due to a mechanical deployment structure. According to the invention, an upper end cover is arranged at the upper end of a barrel body and in threaded connection with the barrel body; an airflow guide tube is vertically arranged at the centre of the upper end surface of the upper end cover; the inner cavity of the airflow guide tube is connected with an air vent; one end of a folding thin-film tube is sleeved on the upper end cover; one end of the folding thin-film tube is fastened with the upper end cover via a hoop, and the other end of the folding thin-film tube is a closed-type top; a direction indicator is vertically arranged on the inner wall of the closed-type top; the section shape of the direction indicator is fitted with the section shape of the inner cavity of a direction indicator guide tube; at least three connection tubes are arranged on the outer end surface of a flange; and each lower thread hole is in threaded connection with one connection tube. The rapidly-inflatable-deployment thin-film support tube disclosed by the invention is used for supporting various micro cameras and various magnetometers on satellite detectors and lunar detectors.