43 results about "Orbital flight" patented technology

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 invention discloses a super-orbit speed flight technical scheme assisted by aerodynamic force, and belongs to the technical field of aircraft design. According to the basic principle of the technical scheme, aerodynamic lift force pointing to the earth center and generated by the aircraft with the high lift-drag ratio serves as supplement of centripetal force, super-orbital speed flight is achieved without centrifugal motion, meanwhile, thrust is used for offsetting aerodynamic resistance, and continuous flight at the super-orbital speed is maintained. The method specifically comprises thefollowing steps: firstly, establishing a flight dynamics model of a super-orbital-velocity aircraft based on a high-lift-drag-ratio aircraft; secondly, based on equal-height and equal-speed assumptions, calculating a super-orbit flight speed which can be reached by the aircraft; then establishing a burnup calculation method of the super-orbital-velocity aircraft; and finally, solving orbit parameters of the optimal fuel consumption flight scheme by adopting an optimization method. The ultra-high-speed flight technology has important application potential in the fields of global fast arrival,spaceflight emergency rescue and the like.

The invention belongs to a hypersonic aerodynamic thermal test technology, and particularly relates to an aerodynamic thermal combined test system and method. The system comprises an arc heater, a stretcher and a test chamber, the test chamber is arranged below the stretcher, the test chamber is fixedly connected with the stretcher, one side of the test chamber is connected with the arc heater, and the other side of the test chamber is connected with wind tunnel accessory equipment. According to the invention, the aerodynamic environment during orbit flight of an aerospace craft can be simulated as truly as possible, and the structural strength of a heat-proof material is examined in the specific environment, so that the problem of self thermal coupling interference test of test equipment is solved, and the system error of self thermal coupling interference of a test system in the test is reduced. According to the invention, aerodynamic thermal examination is completed by using the test piece and adopting a supersonic flat plate free jet test technology, so that the test result is more accurate.

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

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 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.

The embodiment of the invention discloses an aircraft with high orbit injection efficiency and strong maneuver capability and an orbit injection method thereof. The aircraft comprises a solid engine,an aircraft platform which is connected to the solid engine and a load which is installed on the aircraft platform, the load is intended to fly into a target orbit with the aircraft for performing tasks; wherein the solid engine is used for providing flight power required for the aircraft to fly from a ground launching site into a sub orbit, after the aircraft flies into the sub orbit, the solid engine disengages from the aircraft, the aircraft platform is provided with a liquid propulsion system, and the liquid propulsion system is used for providing the flight power required for the aircraftto fly from the sub orbit into the target orbit.

The invention belongs to the field of unmanned aerial vehicles, and discloses an unmanned aerial vehicle safety test system and method. The system comprises a central controller, a flight track, a safety connection assembly and an unmanned aerial vehicle. The safety connecting assembly is connected with the unmanned aerial vehicle and the flight track; the flight track is in an activity area limiting the unmanned aerial vehicle; the central controller generates aircraft test parameters; the unmanned aerial vehicle performs a flight test according to the aircraft test parameters to obtain flight test data; and the central controller generates a test result according to the flight test data. According to the scheme, an aircraft safety test track and a closed safety test site are constructed. And it can be ensured that the aircraft completes various key performance tests in an absolutely safe environment. The test safety is improved, the unmanned aerial vehicle cannot fall and be damaged when being out of control, the safety of the unmanned aerial vehicle for testing is guaranteed while the test effect is not hindered, and the potential safety hazard of the test is reduced.

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 provides a self-adaptive drag-free control method based on a set value identification algorithm, and belongs to the technical field of aerospace, control science and engineering. The method comprises the steps of processing measurement information inside and outside a measurement range of a gravity gradiometer to obtain a set value type measurement value of the gravity gradiometer for residual acceleration, wherein the measurement information is the residual acceleration of a gravity gradient satellite in the flight direction of an orbit; establishing an identification algorithmto estimate the resistance gain coefficient of the gravity gradient satellite along the orbital flight direction according to the obtained set value type measurement value of the gravity gradiometer to the residual acceleration; and according to the estimated resistance gain coefficient, constructing a self-adaptive controller of the gravity gradient satellite along the flight direction of the orbit. According to the invention, drag-free control under saturation constraint measurement can be realized.

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 provides a detection sensor for spatial atmospheric surface density and a manufacturing method thereof. The detection sensor can realize direct detection of the atmospheric surface density of a spatial fixed point; the resonant sensor is employed for detection of spatial atmospheric surface density, realizes direct measurement of atmospheric surface density via atmospheric substancesadhering onto the surface of a probe of the sensor and better serves research on the resistance of atmospheric surface density to spacecrafts; and the resonant sensor is low in power consumption, canmeet the networking requirements of micro-nano satellite constellations, and realizes spatial large-scale distributed measurement of atmospheric surface density and real-time measurement of orbit atmosphere simultaneously in virtue of the networking micro-nano satellite constellations. According to the invention, the resonant sensor is small in size, light in weight and capable of monitoring atmospheric surface density at different positions and different angles in virtue of a sensor array mode. A data processing unit of the invention can be separated from a gas acquisition module, so distributed measurement can be realized, and the atmospheric surface density of the spacecraft during orbital flight can be accurately determined.

The application discloses a kind of geological surveying and mapping coordination operation methods for Beidou patrol unmanned plane, it is related to image optimization technology, for improving the problem of low image acquisition accuracy of single orbit flight unmanned plane, including setting multiple unmanned plane flight height, and collecting the flight environment data of each group unmanned plane at different height, according to the attribute data of each group unmanned plane, the precision of the region to be measured detected by each group unmanned plane is analyzed by selecting different flight environment data combination, and the flight height of unmanned plane is screened and marked, set retrieval scoring mechanism is used to retrieve and score unmanned plane, select geological surveying and mapping unmanned plane in each marked group to carry out geological surveying and mapping, real-time detect the real-time power of unmanned plane and image transmission time stamp interval, analyze the image effect of the geological surveying and mapping unmanned plane of each marked group, record the surveying and mapping time of the geological surveying and mapping unmanned plane and divide into multiple surveying and mapping time periods, select the surveying and mapping image of each marked group in each surveying and mapping time period according to the image effect analysis result.

The invention provides a commercial spaceflight and civil aviation operation cooperation system based on digital twinning. The system comprises a physical commercial spaceflight and civil aviation scene module, a virtual commercial spaceflight and civil aviation scene module, a commercial spaceflight and civil aviation twinning big data module and a commercial sub-orbit flight and civil aviation operation cooperation module. The system can be applied to a civil aviation air traffic control department, and aims at verifying the feasibility of a decision-making scheme for a flight to avoid a sub-orbital debris danger area and a cooperation mode of commercial space flight and civil aviation operation by using a digital twin-based commercial space flight and civil aviation operation cooperation system. And under the condition of dynamic change of the sub-orbital debris danger area, designing a civil aircraft avoidance decision scheme in real time, and re-planning an air route which conforms to an air traffic control rule, so that the safety of the civil aviation airspace is improved.

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 provides a battle system double-encircling detection method and system, and the method comprises the steps: carrying out the emission in a form of a combination of a polar orbit encircling device and a medium-low inclination angle encircling device, jointly completing the flight from the earth to a battle transfer orbit, carrying out the selective separation of the two encircling devices before approaching a battle, carrying out the capturing, braking and ignition, and enabling the two encircling devices to enter different orbits, the two surrounding devices are each provided with a magnetic field measurement device, an inter-device communication device and a device-ground communication device, the two surrounding devices carry out star magnetic field and gravitational field measurement respectively, a machine is selected to carry out star atmosphere occultation and plasma ring measurement, and when resources are sufficient, the middle-low dip angle surrounding device can adjust a track to carry out ring wood and satellite detection. The star system double-surrounding detection system comprises the links of transfer orbit design, assembly resource analysis, star system inner orbit design, resource analysis, load addition and distribution demonstration, environment-wood-satellite extended task analysis and the like. According to the invention, two surrounding devices can be launched at a time to obtain the star data, the resource utilization rate and the result output are improved, and the engineering value is high.

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 a skip reentry aerial vehicle online trajectory planning guidance method based on a linear pseudo spectrum. The method comprises the following steps of initializing, performingstandard control on a flight segment, performing skip segment tilting counter-rotating determining, performing skip segment predictive trajectory integrating, performing precision determining, correcting skip segment rolling counter-rotating time correction, re-establishing parameter control, performing skip segment predictive trajectory integrating, performing precision determining, performing skip segment control parameter correction, and performing Kepler trajectory flight. According to the skip reentry aerial vehicle online trajectory planning guidance method, a roll angle modulus is updated through iteration and furthermore a terminal range deviation is corrected. When a segmentation condition is not decided by time, control compensation is introduced, thereby supplying a good initialcondition for terminal guidance. Furthermore a roll angle modulus value control rule and roll angle counter-rotating time are synchronously updated in one guidance period. Furthermore terminal transverse range and longitudinal range precision are simultaneously ensured. Through an analysis relation between a control amount correction value and a terminal offset, no iterative integration is required in updating the control amount at each time. Only 2ms are required in calculation at each time. The skip reentry aerial vehicle online trajectory planning guidance method is particularly suitable for online application.

The invention relates to a carrier rocket orbit parameter reconstruction method based on iterative guidance, and belongs to the field of carrier rocket guidance control. The method comprises the following steps of 1, collecting carrier rocket state information; 2, calculating the consumption of a propellant entering a target orbit according to the state information, if the consumption of the propellant is less than the available amount, normally flying to an original target orbit, and if the consumption of the propellant is greater than the available amount, performing orbit parameter reconstruction; 3, searching parameters of a reconstructed orbit on line, and taking the reconstructed orbit obtained by searching as a new target orbit; and 4, controlling a carrier rocket to fly towards the new target orbit. The method is based on iterative guidance, the principle is simple, engineering implementation is convenient, whether the carrier rocket can enter the target orbit or not can be judged according to the state information, orbit parameters are searched for and reconstructed on line when necessary, the carrier rocket has the autonomous orbit reconstruction capacity, economic losses are avoided or reduced under the fault condition, and the launch task performance capacity is improved.

An ultra-low orbit satellite all-electric propulsion orbit maintenance system comprises a GNSS receiver, an atmospheric density measurement unit, a satellite service management unit and an electric propulsion system. The GNSS receiver provides a state vector of the ultra-low orbit satellite in an earth-fixed coordinate system. And the atmospheric density detection unit obtains atmospheric density data and inverts the orbit height of the ultra-low orbit satellite. And the electric propulsion system completes distribution of thrust parameters and generation of a thrust according to an orbit maintenance control law transmitted by the satellite service management unit. The satellite service management unit comprehensively determines orbit information of an ultra-low orbit satellite, generates a satellite space trajectory error after processing, further determines an orbit maintenance control law through an orbit maintenance algorithm according to the satellite space trajectory error, and sends the orbit maintenance control law to the electric propulsion system to complete orbit maintenance control. According to the invention, the problem of rapid attenuation of the satellite in the ultra-low orbit flight is solved, stable flight of the satellite in the ultra-low orbit is realized, and the operation duration of the ultra-low orbit satellite is greatly prolonged.

The present invention discloses a recoverable aircraft and control method carried by an ultra-low-orbit satellite. The recoverable aircraft is carried on an ultra-low-orbit satellite platform and connected through an unlocking device. The ultra-low-orbit satellite platform supplies power to the recoverable aircraft through a power supply path. The recoverable aircraft relies on the electric thrust of the ultra-low-orbit satellite platform to fly in orbit and carry out on-orbit experimental tasks. At the separation point, in response to ground commands, the ultra-low-orbit satellite platform and the recoverable aircraft are separated, and the power supply path between the two is synchronously disconnected; after the combination is separated, the ultra-low-orbit satellite relies on the propulsion subsystem to return to the ultra-low-orbit orbit and continue to perform long-term on-orbit missions; the recoverable aircraft relies on the propulsion subsystem to aim at the 100km altitude orbit re-entry point for adaptive guidance, while carrying out short-term experiments, and returning to the ground after completing the experiments. Using the ultra-low-orbit satellite platform, the recoverable aircraft can achieve low-cost, high-performance mission flights.

The application provides a control system of a space vehicle's thrust chamber, comprising: an actuator, which provides power to control the thrust chamber to swing according to a predetermined tilt angle; an angle monitor, which monitors the actual tilt angle of the thrust chamber; a reset processing module, which controls the thrust chamber to reset through the actuator; a correction module, which corrects the reset processing module; wherein the control of the thrust chamber to reset through the actuator comprises: if the actual tilt angle is equal to the predetermined tilt angle, the reset processing module makes the thrust chamber reset according to the predetermined tilt angle; if the actual tilt angle is not equal to the predetermined tilt angle, the correction module calculates a correction offset through the relationship between the actual tilt angle and the predetermined tilt angle, and the reset processing module makes the thrust chamber reset according to the predetermined tilt angle plus or minus the offset. The control system can correct the flight orbit and maneuver in flight in real time, and is beneficial to improve the accuracy of the space vehicle flying according to the predetermined orbit.

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.

A spacecraft power system relates to the field of aerospace engine technology. The YRCW aerospace engine comprises an air intake system, a central RCW engine, and two side YRC engines. The YRCW aerospace engine can operate in four modes: ramjet turbine, ramjet, superramjet, and rocket. The ramjet combustor can accommodate four combustion modes. An inlet airflow separator randomly distributes the RCW and YRC airflows. The combustion gas impacts the turbine, driving the axial-flow intake compressor and centrifugal compressor to inhale and compress air. The compressed air is distributed through four bypass ducts, entering the cold air, hot exhaust gas, and vectored airflow hoods, respectively. This provides compressed cold air and cooling gas, as well as vectored injection air, for closed-loop combustion of the RCW and YRC engines. Adjusting the geometric positions of the conical ramjet intake valve, the compressed cold air, hot exhaust gas, the fuel injection plate, and the oil pressure and oil volume support channel valves ensures smooth transitions between the engine's multi-mode operating modes, stable flames, and sustained combustion, enabling the aerospace vehicle to achieve single-stage orbital flight within a full envelope.

The invention relates to a ground test method for simultaneously examining inertial navigation and simulated flight. The ground test method comprises the following steps: determining simulated flight initial parameters; acquiring initial navigation parameters of a test site at an initial moment; simulating orbit flight through a thrust curve; according to actual measurement of data output by a master-slave inertial measurement combination in real time, initial speed, position and attitude information of a test point as initial values, and gyro data and table data measured by master-slave inertial navigation equipment in real time, enabling a master inertial measurement unit and a slave inertial measurement unit to respectively use a set of navigation to carry out inertial navigation calculation; and enabling the first set of navigation to obtain a master inertial measurement integrated navigation result, and the second set of navigation to obtain a slave inertial measurement integrated navigation result for checking the navigation precision of the master-slave inertial measurement unit; therefore, three sets of navigation results are obtained, and aircraft thrust curve flight simulation completed by the third set of navigation is used for flight control. Thrust curve simulation is adopted, a real track can be simulated, real inertial measurement unit characteristics are reflected, and a ground test is more real and reliable.

The invention discloses a three-pulse near-moon braking fixed-time and fixed-point landing aiming method. The method comprises the steps that input parameters are obtained; determining a power descending point moment initial value, a second-time near-moon braking moment, a four-device separation point moment initial value, a third-time near-moon braking moment and a lunar surface take-off moment; a finite thrust model and an accurate forecasting model are adopted for orbit control strategy solving, and orbit parameters at the four-device separation point moment are obtained through calculation; inner and outer ring strategy solving is carried out according to four-device separation point moment orbit parameters; and outputting an orbit control strategy calculation result. According to the method, the characteristic of perturbation rate difference of lunar elliptical orbit planes of different periods is utilized, a 4h lunar parking elliptical orbit is designed, the parking orbit flight time corresponding to the orbit plane adjustment amount is determined by establishing the drift rate difference between the 4h lunar parking elliptical orbit and a 2h target lunar orbit plane, the flight time of the nearly one month from capturing to descending is fully utilized, and the flight speed of the target lunar orbit plane is determined. On the premise of not additionally consuming a propellant, high-precision aiming of a landing point is 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 discloses an aircraft restraining system restrained by an I-shaped track and a flight restraining method used for limiting the track. The I-shaped track comprises a track top plate, a track bottom plate and a middle vertical plate. Each set of control surface adjusting device comprises a height constraint adjusting mechanism and a direction constraint adjusting mechanism; horizontal control surfaces on two sides of the aircraft at the same spanwise position are correspondingly connected with a set of height constraint adjusting mechanism, and a vertical control surface is connected with one or two sets of direction constraint adjusting mechanisms; the height and direction restraining and adjusting mechanism and a track structure on one side of the I-shaped track form a floating connecting structure; in the process that the aircraft flies along the I-shaped track, the horizontal control surface is driven by the height constraint adjusting mechanism to deflect, and the vertical control surface is driven by the direction constraint adjusting mechanism to deflect, so that reverse acting force is generated, and the relative distance between the aircraft and the I-shaped track is kept within a preset range. The problems existing in a traditional ground traffic transportation form and an eVTOL transportation form are solved.