31 results about "Orbital flight" patented technology

A ground-based system for simulating dynamic suborbital-flight environments occurring during substantially complete suborbital-flights by a flight vehicle.

The invention relates to a satellite group orbit design method for geostationary orbit satellite distributed co-orbital flight. The method includes the following steps that: the position relationship model of a reference satellite and orbiting satellites is built; an orbiting satellite orbit model is built through using the position relationship model; the number of orbiting satellite orbits is calculated according to orbiting satellite orbit requirements; and the number of orbiting satellite group orbits is generated according to the number of the orbiting satellite orbits. Relative distance fluctuations of the orbiting satellite group orbits designed by using the method of the present invention and the reference satellite is small; the azimuth angle, elevation angle and distance relationships of the orbiting satellites and the reference satellite meet multi-satellite co-orbital spacing design requirements; the azimuth angles, elevation angles and distances of adjacent orbiting satellites satisfy inter-satellite link acquisition and tracking design requirements; defects such as weak function, poor anti-interference ability, unfavorability to miniaturization of a single-satellite node condition can be eliminated; and the orbiting satellites and the reference satellite collaborate with each other, and therefore, the anti-destruction ability and self-organizing ability of a system are improved.

A turbojet is combined with a co-axially integrated rotary rocket to form a propulsion system called a Rotary Turbo Rocket that can function as a turbojet, as an afterburning turbojet, as an Air Turbo Rocket, or as a rotary rocket. The Rotary Turbo Rocket can operate in any of these propulsion modes singularly, or in any combination of these propulsion modes, and can transition continuously or abruptly between operating modes. The Rotary Turbo Rocket can span the zero to orbital flight velocity speed range and / or operate continuously as it transitions from atmospheric to space flight by transitioning between operating modes.

The present invention provides a boundary control method for bounded accompanying flight of a satellite in a circular reference orbit. The boundary quantitative analytical relationship between the boundary of the satellite's bounded accompanying flight under orbit and the satellite's orbit element difference is solved to obtain the expected orbit element difference corresponding to the given boundary accompanying flight task; using a predetermined quantitative analytical control model Calculate and obtain the pulse speed increment that needs to be applied to the tracking satellite at the initial moment and the terminal moment. Using the boundary analysis model to solve the difference in expected orbital elements expands the space for the design of accompanying flight configurations and brings certain flexibility to the selection of formation or cluster flight mission configurations. The pulse velocity increment requirement with a given boundary is accurately calculated by the quantitative analytical control model, which has the advantages of high control precision and simple control process.

The invention relates to a glass panel for an aircraft suitable for a suborbital flight and an aeronautical flight. The glass panel comprises an outer panel (1) made of polycarbonate or aluminosilicate for temperature resistance, a main panel (2) for pressure resistance, sized according to standard aeroplane safety factors, and an inner panel (3), providing redundancy for the main panel, sized with a minimum pressure margin, the outer, main and internal redundancy panels being separated from one another by spaces (5, 6). The invention also relates to an aircraft comprising windscreen elements and portholes made using the glass panel.

The invention discloses a recyclable and reusable combination structure for the end of a carrier rocket, which includes a barrel-shaped rocket end and a recovery runway. Splicing, and an electromagnetic circuit is arranged on the plane end of each half-tube rocket end, a controller and a power supply connected to the controller are arranged inside each half-tube rocket end, and the power supply is connected to the electromagnetic circuit. In the present invention, the terminal rocket is divided into two designs, and the small satellites are respectively installed in the end of the half-tube rocket. The end of each half-tube rocket can independently fly according to the orbit and release the internal satellite, which solves the flight burden and control difficulty of the traditional terminal rocket. Big question. The magnetic fields generated by the respective electromagnetic circuits between the ends of the two half-tube rockets change the mutual force, repeatable aggregation and separation, simple structure and high control stability, which is conducive to promoting the development of the aerospace industry.

The invention relates to an autonomous orbit reprogramming method of a spaceflight aircraft. The method comprises the following steps that spacecraft state information is acquired; out-of-tolerance of an orbit is calculated according to the state information; whether reprogramming is performed is judged according to the out-of-tolerance situation, flying according to the normal default orbit is performed if the out-of-tolerance situation does not occur, and ballistic reprogramming is performed if the out-of-tolerance situation occurs; the minimum i value corresponding to an accessible target orbit parameter in an orbit parameter library is judged according to the order of priority, and the preselected orbit corresponding to the minimum i value acts as a new target orbit; and the aircraft is controlled to fly according to the new target orbit. The spaceflight aircraft is creatively enabled to have the autonomous orbit reprogramming capacity so that self-rescuing under the fault state can be realized, the expected target is completed under the fault situation and the economic loss can be reduced and the safety risk can be reduced; and the manpower and material resource cost can be reduced through the autonomous mode without depending on ground equipment and personnel.

A satellite data processing method, apparatus, and a satellite backup subsystem belongs to the technical field of satellites. This method is applied to the satellite backup subsystem. The method comprises: receiving a data task, wherein the data task comprises data backup or data restoration; splitting the data task into a plurality of single-orbit tasks; and executing respective single-orbit task in each orbital flight.

The invention provides an adaptive drag-free control method based on a set-value identification algorithm, which belongs to the technical fields of aerospace, control science and engineering. The method includes: processing the measurement information inside and outside the measurement range of the gravity gradiometer to obtain a set-valued measurement value of the residual acceleration by the gravity gradiometer, wherein the measurement information is the residual acceleration of the gravity gradient satellite along the orbital flight direction; According to the obtained set-valued measurement value of the residual acceleration by the gravity gradiometer, an identification algorithm is established to estimate the drag gain coefficient of the gravity gradient satellite along the orbital flight direction; adaptive controller. By adopting the present invention, drag-free control under saturation constraint measurement can be realized.

The present invention relates to a method for analyzing the projected area of ​​a cooperating target of a rendezvous and docking laser radar. Firstly, according to the installation relationship of the cooperating target in the space station, the pose of the cooperating target in each docking plane coordinate system is obtained. The relationship of the system is obtained, and the pose of the cooperation target under the system of the space station is obtained. Then, according to the orbit data of the space station, the relationship between the space station's own system and the space station orbit system can be obtained, and the pose of the cooperative target can be expressed under the space station orbit system. In the same way, according to the installation relationship of the lidar on the spacecraft, the relationship between the lidar coordinate system and the spacecraft itself can be obtained. Finally, by linking the orbit system of the space station and the spacecraft itself through the J2000 inertial system, the relative position and attitude relationship between the cooperative target and the lidar can be obtained according to the orbital flight trajectory, and finally the projection of the cooperative target of the rendezvous and docking lidar can be analyzed during the orbiting flight. area. The present invention can ensure smooth docking in the fly-around process.

The invention relates to a rail type flight simulator having continuous overload simulation capability. The rail type flight simulator having continuous overload simulation capability is characterized by comprising two ring-shaped bottom rails connected to the ground, a vertical ring-shaped wall is arranged on the periphery of the bottom rails, the inner wall of the ring-shaped wall is connected to two horizontal ring-shaped side rails, the bottom rails and the side rails are slidingly connected to a vehicle, the vehicle comprises a vehicle body with opened front end and closed rear end, the bottom surface of the vehicle body is provided with a rolling wheel set in sliding connection with the bottom rails and a driving unit, the rear end surface of the vehicle body is provided with a rolling wheel set in sliding connection with the side rails and a driving unit, the outside of both left and right side walls of the vehicle body are connected to two motors symmetrically, output shafts of the two motors penetrate through the side wall of the vehicle body to support a rectangular middle frame together, two motors are symmetrically arranged on both front and rear sides of the middle frame, and the output ends of the two motors support a closed hanging basket together. According to the rail type flight simulator having continuous overload simulation capability, the high-speed centrifugal motion is performed in a closed track structure, and completely-enclosed rectification treatment can be performed on the vehicle, so that the resistance generated during operation is reduced.

A method for transporting people and cargo from the surface of the Earth to orbit around the Earth is disclosed. The present invention uses a series of lighter-than-air vehicles to allow for a far safer and less strenuous trip to orbit than using current rocket-based technology. The high altitude atmospheric airship is flown from the ground to the upper atmosphere, where it docks with the buoyant transfer station, and from there, the people and cargo transfer to an orbital airship for the remainder of the trip to orbit. The orbital airship returns to the station for another transfer back to the atmospheric airship for the return back to the surface of Earth.

The invention discloses a low-orbit spacecraft quasi-real-time orbit maneuver detection method based on actual measurement data. The invention proposes a low-orbit space vehicle detection method based on actual measurement data, aiming at the characteristics of low-orbit spacecraft in-orbit operation rules and actual measurement data. The quasi-real-time orbit maneuver detection method, according to the pre-set detection threshold, can detect in real-time whether there is orbit maneuver during the orbital flight of the low-orbit spacecraft after the tracking of each arc segment of the low-orbit spacecraft. It effectively solves the problem that iterative calculation is difficult to converge when there is orbit maneuver in the orbit determination process of low-orbit spacecraft and the spacecraft tracking is difficult after orbit change.

The invention discloses a fixed-wing unmanned aerial vehicle, which belongs to the field of unmanned aerial vehicle surveying and mapping. The fixed-wing unmanned aerial vehicle includes a single lens, a flying platform, and a lens stabilization module for adjusting the attitude of the single lens. The flying platform There is a stabilizer on the top, and the single lens is fixed on the action end of the stabilizer. When the fixed-wing unmanned aerial vehicle is flying around according to the flying circle route, the attitude of the single lens is aligned with the flying circle. Hover the center and keep the tilt angle of the single lens constant. The present invention also discloses a single-lens oblique photography system and a corresponding single-lens oblique photography method for the fixed-wing unmanned aerial vehicle. The fixed-wing unmanned aerial vehicle disclosed in the present invention does not need to expand the route to adjust the attitude or obtain the image of the edge of the survey area through the expansion, and has the ability to surround the shooting. Its ability to cover the effective shooting area with a single lens is stronger than that of the prior art. The fixed-wing UAV equipped with multiple lenses has a shorter total route and fewer photos.

The invention discloses a distributed aerodynamic fusion orbit coupling attitude perturbation analysis method, which relates to the field of spacecraft flight research. pose motion, and establish the perturbation analysis equation of the coupled attitude of the aerodynamic fusion orbit based on the attitude dynamics equation and the perturbation analysis model of the orbital number of the orbital flight; obtain the multi-source data of the spacecraft in-orbit flight process, the multi-source data includes: The position information of the spacecraft, the speed information of the spacecraft, the attitude information of the spacecraft and the sailboard information of the spacecraft are processed by using the perturbation analysis equation of the aerodynamic fusion orbit coupling attitude to process the multi-source data to obtain the tracking results of the spacecraft state. The parameter identification of the solar panel angle of the spacecraft is carried out to obtain the identification results; the error factors of the identification results are quantitatively analyzed, which improves the calculation accuracy of the spacecraft on-orbit flight and re-entry prediction.

The invention discloses a limited-time formation method for unmanned aerial vehicles using a directed communication topology, and relates to the technical field of unmanned aerial vehicles. The obtained state information is sent to the distributed finite-time formation controller; the distributed finite-time formation controller forms maneuvering commands through transformation and calculation according to the formation configuration information, the acquired state information of adjacent UAVs and its own state information. Under the request of the maneuvering command, the UAV changes its course or speed and flies along the predetermined track. Compared with the traditional method, the position error e in the method of the present invention p，i Without using formation trajectory information, the controlled system can be stabilized in a limited time, and has the advantages of continuous controller signal, small calculation amount, no need for the controlled system to satisfy the homogeneity assumption, and the use of directed communication topology.