98 results about "Elliptic orbit" patented technology

A first satellite is placed practically directly by a launcher on a first final orbit. A second satellite is placed on the same launcher is initially transferred to a highly elliptical waiting orbit whose semi-major axis points to intercept the torus formed by the sphere of influence of the moon on its orbit, and then during a maneuver at the perigee of the highly elliptical orbit, the second satellite is transferred to lunar transfer orbit. Changes in the perigee altitude and the inclination of an intermediate orbit on which the second satellite is to be found are obtained mainly by gravitational reaction in the sphere of influence of the moon, and during a last maneuver, the second satellite is placed on a final orbit having orbital parameter values that are quite different from those of the orbital parameters of the first final orbit.

Seamless handover of a communications signal from a first satellite to a second satellite is provided when the satellites are at orbital positions which coincide. Timing marks are inserted simultaneously in signals transmitted through the satellites, and signals received from the satellites compared to determine the difference in path length. Handover occurs when the path length difference is zero and the two signals are perfectly synchronized. Interference between the signals transmitted through the two satellites is avoided by using different transmission modes, such as different carrier frequencies, orthogonal senses of polarization, or digital signals with uncorrelated spreading codes. Using these different transmission modes in the right- and left-leaning orbits of a Cobra Teardrop system also permits overlaying multiple teardrop patterns, at longitudinal spacings comparable to the Basic Cobra system, as well as closer in-track spacing of satellites. The result is over an order of magnitude increase in global system capacity.

The invention discloses a pulse and pneumatic assistance combination-based low-orbit orbit plane transfer method, relates to a large-range orbit plane transfer method for an earth low-orbit spacecraft, and belongs to the field of aerospace. The method comprises the steps of firstly establishing the number of orbits and a dynamic equation of a flight process in atmosphere; changing an orbit of the spacecraft to a highly elliptic orbit in a maneuvering manner by applying a pulse, and enabling the spacecraft to enter the atmosphere by applying an de-orbit pulse at an apogee; selecting an optimization target as a maximum change amount of an orbit plane, giving constraints and obtaining an optimal control rate and a terminal state variable meeting aerodynamic requirements, thereby finishing pneumatic assisted orbit plane transfer; and enabling the spacecraft to fly out of the atmosphere and run to a target orbit height along a transfer orbit, and enabling the spacecraft to enter a target orbit by applying an orbit determination pulse. According to the method, the orbit plane transfer of the low-orbit spacecraft can be finished with relatively low fuel consumption; and the method is good in robustness, high in repeatability, small in influence of spacecraft orbit orientation, high in flexibility of a pneumatic assistance process, and wide in application range for the target orbit.

The invention relates to a gravity satellite formation orbital stability optimization design and an earth gravity field precision inversion method on the basis of disturbing inter-satellite distance principle, in particular to an orbital stability optimization design method for a four-satellite system (FSS). In order to guarantee the stability of the four-satellite system, the quantity of satellite orbits is optimally designed, orbital semi-major axes, orbital eccentricities, orbital inclinations and right ascensions of ascending nodes keep unchanged, the difference of arguments of perigees each pair of satellites and the difference of mean anomalies of the pair of satellites are 180 degrees respectively, an initial argument of perigee of each satellite is arranged at the equator, an initial mean anomaly of each satellite is arranged at a pole, and the ratio of the semi-major axis of each elliptical orbit of the four-satellite system to a semi-minor axis of the elliptical orbit of the four-satellite system is 2:1. The gravity satellite formation orbital stability optimization design and the earth gravity field precision inversion method have the advantages that an earth gravity field is precisely and quickly inverted on the basis of a disturbing inter-satellite distance process; and the orbits are high in stability owing to the method, the earth gravity field computation precision is effectively improved, the gravity field inversion speed is increased to a great extent, and requirements on the performance of a computer are low.

The invention discloses a two dimentional attitude control system of a synthetic aperture radar satellite, which includes a satellite position analysis module, a satellite position analysis module and an expected attitude determination module. The two dimentional attitude control system is applied in the situation that the synthetic aperture radar satellite runs on an elliptic orbit, and echoed signal doppler center frequency can be controlled to be zero by applying yaw-pitch-two-dimensional attitude control technology, thus effectively weakening the degree of coupling between distance vector and direction vector and lowering the difficulty and operand of imaging process.

The invention provides a low-orbit microsatellite formation system suitable for medium / high-latitude region coverage. According to the low-orbit microsatellite formation system, a surface projection principle is designed based on Flower constellation, all satellites constitute a compact geometric space configuration at the initial moment of a regression period, and the topological structure is ensured to have a periodical repeat property relative to all points on the ground; highly elliptical orbits which are completely the same in shape are adopted for all the satellites in the formation; and a plurality of permanent inter-satellite links are established under the condition of guaranteeing no link overlap inside the formation. According to the low-orbit microsatellite formation system provided by the invention, the complexity of a ground tracking control system can be reduced greatly, the optimization of inter-satellite links and the increase of communication reliability are facilitated, a favorable ground multi-coverage characteristic of medium / high-latitude regions is guaranteed, and the low-orbit microsatellite formation system has wide application values and development prospects in the field of space communication.

The invention discloses an initialization method of an elliptical orbit satellite formation configuration under multiple constraints, relates to satellite formation configuration initialization methods, and belongs to the fields of spacecraft guidance and control. The method comprises that an analytical solution is derived from a T-H equation of a real anomaly domain via typical track dynamics andcoordinate transformation, dynamic constraint and periodic fly-around constraint are analyzed, monitoring camera visual field constraint and thrust amplitude constraint are established, a multi-constraint optimal control model of formation is established and solved in a Gaussian pseudo-spectral method, a co-state value in random time is obtained, the completeness of the spacecraft structure is monitored in real time, and the fuel consumption and highest amplitude of a propelling system are reduced by increasing the eccentricity and initialization time. Thus, the elliptical orbit satellite formation configuration can be initialized under multiple constraints, and effective optimal control is realized under the multiple constrains by initialization.

The present invention provides a method for the attitude control of satellites in elliptic orbits or satellites initially placed in circular orbits perturbed to elliptic orbits due to environmental disturbances. The method relies on the application of solar radiation pressure to provide the desired torque for the satellite attitude control. The satellite is equipped with two-oppositely placed light-weight solar panels extending away from the satellite along a predetermined direction (satellite body fixed Y-axis). By rotating one of these solar panels or both of them through desired angles about their axis using the respective driver motors as per the simple open-loop control law, the torque about the satellite axis is developed to achieve the desired attitude performance. The open-loop control law is derived using an analytical approach to neutralize the excitation caused by eccentricity and it is implemented via analog logic based on the information of sun angle and satellite position provided by the sensors. The present invention significantly improves the performance of the satellite by a factor of more than 20 times approximately in general and it only requires the rotation of the solar panels by fraction of a degree for particular system parameters. Thus, the semi-passive nature of the present invention makes it attractive for future space applications.

The invention discloses a small elliptic target reentry forecasting method under a sparse data condition. N circles of data collected in the past 5 days are processed, a numerical method is adopted tocarry out track determination on the single circle of data, the Kepler root number is adopted as a root number system, a semi-numerical method is adopted to carry out track integration, and a least square method is adopted to fit a ballistic coefficient of a reentry target. Compared with the prior art, the invention provides a reentry forecasting method combining the numerical method and the semi-numerical method for the characteristics of the sparse data and the small elliptic orbit. The problems that under the condition of sparse data, the multi-circle data joint orbit determination residual error is too large or not convergent, and the ballistic coefficient is difficult to determine through single-circle data orbit determination are effectively solved.

A conical region in space with a base at a geosynchronous distance from the earth and the apex of the cone at a point on the ground may be termed a space needle. A multiplicity of small satellites in elliptical orbits located within such a space needle may establish timing of their radio frequency (RF) transmissions forming what may be termed a needle beam downlink having an apparent origin that may be thousands of kilometers to the North or South of a Kepler geosynchronous satellite parking orbit. A noise-like RF signal may be transmitted synchronously from multiple transmitters in space forming a spatially distributed spread spectrum RF needle beam. Applying the method of space-based needle beams to a multiplicity of transmitters on the ground, a network of ground stations may form a ground-based needle beam uplink that may be pointed a given satellite at a given time.

A satellite system provides geosynchronous satellites in elliptical orbits in respective elliptical orbital planes separated by 120 degrees. The satellites traverse a common figure-eight ground track comprising northern and southern loops. The satellites are controllably switched to operate the satellite currently traversing the northern loop to deliver a selected signal (e.g., a selected frequency signal) to satellite receivers.

When a satellite is orbiting the earth in an elliptic orbit, it has a certain inclination with respect to the earth's equator. The usual way to change the inclination is perform a maneuver by firing the rocket engines at the periapsis of the ellipse. This then forces the satellite into the desired inclination. There is a substantially more fuel efficient way to change the inclination. This is done by an indirect route by first doing a maneuver to bring the satellite to the moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so called fuzzy boundary or weak stability boundary. A negligibly small maneuver can then bring it back to the earth on a reverse BCT to the desired earth inclination. Another maneuver puts it into the new ellipse at the earth. In the case of satellites launched from Vandenberg AFB into LEO in a circular orbit of an altitude of 700 km with an inclination of 34°, approximately 6 km / s is required to change the inclination to 90°. This yields a savings of approximately 13% in Delta-V as compared to the standard approach which could translate into a significant increase of payload or perhaps a smaller launch vehicle. This may have applications to commercial satellite launches for the Iridium or Teledesic networks and others.

A communication system having a plurality of artificial satellites each having a communication unit traveling on elliptical orbits with an orbital period of 24 hours. Arguments of perigee are set to values with a predetermined allowance in a setting process of the plurality of artificial satellites.

A system and method for producing electrical energy utilizing horizontal axis water turbine assemblies, completely submerged and secured to a foundation on the seafloor beneath ocean surface waves in water depths corresponding to a nominal range of 1 / 30th to 1 / 10th of the wave lengths of prevailing waves. The water turbine assembly captures kinetic energy of vertically compressed elliptical orbital water particle motion having accentuated horizontal velocity for production of electrical power delivered to an onshore terminal. A group of multiple water turbine assemblies may be disposed at selected distances apart relative to one another and at selected distances from the coastline to form an array spanning a range of surface wave lengths in the shallow water depth to reduce temporal variability and collectively amplify and optimize the electrical power produced by the facility ashore.

The invention discloses an elliptical orbit based multi-beam low orbit satellite channel simulation method. The method includes steps of configuring satellite communication working frequency points, satellite orbit parameters and user terminal parameters and modifying universal parameters according to precision requirements; generating three-dimensional coordinates of different users randomly in asingle process or configuring three-dimensional coordinates of different users manually by utilizing a terminal coordinate system; selecting a multi-beam satellite antenna module and configuring parameters; calculating beam gain of the satellite antenna; calculating free space loss, Doppler frequency shift and time delay; calculating channel coefficients; and calculating valid communication timeof the satellite to users. According to the invention, related problems occurring when the low orbit satellite is in an elliptical orbit are taken into account; after simulation, multiple parameters from different beam antennas of the satellite to the ground multiple terminals can be output.

A propellant budge-based low orbit elliptic track satellite successful injection determining method is provided. Through acquisition of relationships between propellant waste and apogee, perigee, eccentricity ratio and inclination angle adjustment amount, an injection success determination formula containing the apogee, perigee, eccentricity ratio and inclination angle adjustment amount can be achieved via total propellant waste provided by satellite orbital transferring; primary orbital transferring propellant amount allowed by a satellite can be determined according to limitations of each party; the apogee, perigee, eccentricity ratio and inclination angle deviation value of the satellite can be determined after separation of the satellite and a rocket; and whether the satellite successfully enters an orbit is determined via the determination formula. An orbital maneuver theory and formula are employed to achieve an analytic calculation formula, so accuracy, rationality, simplicity and high efficiency can be achieved; easy operation is provided and the method is in particular suitable for quick determination of success carry launch; and basis and instruction are provided for implements of emergency measures of the satellite upon problems during the carrying launch.

The invention relates to a reconfiguration method for optimal formation under J2 perturbation and adopting relative navigation information, and belongs to the field of space-flight trajectory control.Aiming at the formation reconfiguration problem of formation flying spacecraft considering the J2 perturbation, under the condition that information of a primary spacecraft is unknown and a followingsecondary spacecraft only adopts the relative navigation information, a continuous thrust optimal formation reconfiguration control strategy which does not need an initial predictive value, can perform rapid calculation and is suitable for being used on satellites is provided. The formation flying spacecraft using the reconfiguration method can automatically complete formation reconfiguration without ground guidance information. According to the reconfiguration method for the optimal formation under the J2 perturbation and adopting the relative navigation information, a relative motion equation is not subjected to linear treatment, meanwhile, the influence of earth J2 perturbation on formation reconfiguration is considered, and the reconfiguration method is suitable for continuous thrustformation reconfiguration under an elliptical orbit; and only relative motion information of the primary spacecraft and the secondary spacecraft is used, the formation reconfiguration of formation flying is realized, and ground support is not needed in the reconfiguration process, so that high in-orbit application value is achieved.

The embodiment of the invention provides a Mars exploration method and a Mars probe. The Mars exploration method comprises the steps that the Mars probe is launched from the earth, lifts off and enters an earth Mars transfer orbit, orbit trimming is conducted, and the Mars probe enters a Mars atmospheric layer; pneumatic braking is conducted through the Mars atmospheric layer, and the Mars probe enters and parks on a large elliptical orbit around the Mars; when the Mars probe parks on the large elliptical orbit around the Mars, dust storm on the Mars is monitored; and when the situation that the dust storm on the Mars is formed is determined, the Mars probe enters the Mars atmospheric layer to sample the dust storm on the Mars. According to the scheme, the cost can be reduced during Mars exploration.

The invention discloses a miniature piezoelectric monocrystal linear motor. The miniature motor comprises a piezoelectric monocrystal driver, a slide block and a drive power supply, wherein the piezoelectric monocrystal driver is a square-columnar piezoelectric monocrystal; a friction block is arranged in the centre of the lower surface of the piezoelectric monocrystal; the friction block is in frictional contact with the slide block; four electrodes which are respectively distributed on four edges symmetrically are arranged on the piezoelectric monocrystal along four axial directions; and a voltage generated by the drive power supply excites the piezoelectric monocrystal to perform high-frequency small-amplitude (nano-grade to micron-grade) directional vibration or synthetic elliptic orbit motion so as to drive the slide block to perform linear motion. The invention provides a compact linear drive structure, which has the characteristics of capability of greatly reducing a working voltage, simple structure and small size, and is particularly suitable to be used for driving ultra-miniature optical lens to zoom and the like. A piezoelectric monocrystal material has extremely high low-temperature piezoelectric property, so the miniature piezoelectric monocrystal linear motor also can be used for low-temperature or spatial precise driving.

Deorbiting of an earth-orbiting satellite is accomplished by executing an orbit transfer maneuver, the orbit transfer maneuver resulting in transference of the satellite from an operational orbit to a disposal orbit, where the disposal orbit is substantially circular and has a nominal radius of approximately, 31,000 kilometers. The operational orbit may be substantially geosynchronous and have at least one of (i) an inclination of greater than 10 degrees and (ii) a nominal eccentricity greater than 0.1. Alternatively, the operational orbit may be a medium earth orbit.

The invention provides a stability control method for unfolding of a three-dimensional space tether system operating in a low earth elliptical orbit in an initial stage. A tether is released to a desired length and the swing of the tether is restrained only by adjusting the limited tether tension without relying on other thrusters. The method comprises the following steps that S1, a two-body spacetether system dynamic model considering tether mass is established; S2, under the conditions such as disturbance, uncertainty, underactuation, restricted tether swing angle and tether tension, basedon a nonlinear model predictive control (NMPC) algorithm, a pseudo-spectral algorithm is used for discretizing an original nonlinear model, and an open loop optimal problem in predictive control is transformed into a nonlinear programming problem for solving; and S3, according to the solution of the nonlinear programming problem, the control quantity of the next control period is determined to form the closed-loop control. Finally, the three-dimensional space tether system operating in the low earth elliptical orbit is stably released and unfolded in the initial stage.

The invention provides an elliptical orbit spacecraft relative position regressing control method based on a cascade connection equation. Firstly, an elliptical orbit relative dynamical model of the cascade connection equation is set up. Secondly, a relative position regressing controller is designed based on a cascade connection equation manner. The elliptical orbit spacecraft relative position regressing control method based on the cascade connection equation has the advantages of structuralization and systematization. The design process is made to be more flexible. Subsystems of all ranks of a system can be designed respectively. In addition, the special problems existing in the subsystems can be considered independently. The system can obtain the good overall or local stability, the good following characteristic and the good parameter robustness.

The invention discloses an orbit planning method for a large elliptical orbit and a small-inclination-angle circular orbit, belongs to the technical field of orbit planning, and aims at solving the problem of orbit planning according to selected target area information, constraint conditions of orbit regression characteristics, constraint conditions of reconnaissance loads, launching deployment parameters, an accurate earth motion model and other data. And planning and designing of a large elliptical orbit and a small-inclination-angle circular orbit for detection of different types of target areas are completed. According to the design of the large elliptical orbit and the small-inclination-angle circular orbit, traversal is carried out at specific intervals for angle values within an effective inclination angle range, and longitude values of sub-satellite point trajectories of a satellite orbit corresponding to each inclination angle value i on latitude circles of multiple transit target points are calculated; and the difference between the geocentric longitude of the target point and the geocentric longitude value is calculated when the sub-satellite point trajectory passes through the latitude circle of the punctuation. According to the invention, orbit scheme design can be carried out on a large ellipse (small-dip-angle circle) orbit, and rapid and effective detection of a target area is realized.

Methods of launching a vehicle using impulsive force are disclosed. In one instance, a vehicle is launched easterly with impulsive force in a plane corresponding to the vehicle's elliptical orbital path. In another instance, a method of closing a timing difference is disclosed. The vehicle undergoes a series of forces after impulsive launch. The first force establishes an orbit having a period significantly different from the orbital period of a satellite or desired vehicle location, closing the difference in an integer number of orbits. The second force establishes the vehicle in circular orbit with the satellite or desired vehicle location. In another instance, the vehicle launched impulsively from a first celestial body travels a first path, and the vehicle experiences a second force along a hyperbolic path about the second celestial body and enters circular orbit about the second celestial body.

The invention provides a hovering cooperative control method and system for an electromagnetic spacecraft fleet. The method includes calculating the electromagnetic force of each electromagnetic spacecraft and obtaining an electromagnetic force calculation equation; taking a round track in the same cycle of an oval track of a reference spacecraft as a nominal track, establishing a target dynamicsequation of relative motion of the electromagnetic spacecraft and the reference spacecraft, and designing a robustness control rule by adopting a slide mode control method based on the target dynamicsequation; distributing magnetic torques of all the electromagnetic spacecrafts according to the electromagnetic force calculation equation and the robustness control rule. According to the invention,in a condition that the electromagnetic spacecraft cannot acquire reference track real time information and is subjected to different interference, hovering of the spacecraft fleet in any position inthe space can be realized; good robustness is achieved. By adopting the optimization method for realizing optimization configuration of magnetic torques of the electromagnetic spacecrafts, decouplingof electromagnetic spacecraft tracks and attitudes can be facilitated.

The invention relates to a carrier rocket elliptical orbit online planning method, in particular to a rocket online trajectory planning method, and belongs to the field of aerospace guidance control.According to the method, the nonlinear constraint, namely the elliptical orbit injection constraint, can be converted into the convex constraint so that the elliptical orbit planning problem can be constructed into the convex planning problem, and online solving and implementation are facilitated. By updating the orbit injection point conjecture and establishing a new elliptical orbit convex programming model through the sequence, the orbit injection precision can be ensured, and the terminal constraint of elliptical orbit planning is accurately met.

The invention discloses a sun earth moon motion instrument, which comprises a central fixing shaft. The top end of the central fixing shaft is fixed with a central main gear; a sun model is arranged on the central main gear; a rotation arm is arranged, in a rotating mode, on the central fixing shaft and located below the central main gear; the center of the rotation arm is overlapped with the center of the central fixing shaft; the rotation arm is provided with a main motor; the output shaft of the main motor is provided with a driving motor gear; the driving motor gear and the central main gear are engaged; one end of the rotation arm is provided with an earth-moon system sliding assembly, and the other end of the rotation arm is provided with a balance weight; the earth-moon system sliding assembly comprises a sliding fixing plate and an earth autorotation mechanism and a moon revolution mechanisms arranged on the sliding fixing plate; and an elliptical orbit control mechanism and a belt wheel tensioning mechanism are arranged between the earth-moon system sliding assembly and the central main gear. The sun earth moon motion instrument provided by the invention can accurately demonstrate sun earth moon motion rules.

The present invention discloses a swing inhibition method of a local space elevator system in an elliptical orbit. A working body is a local space elevator system formed by three spacecrafts connected through two tethers. The lengths of the tethers are decided through a concrete condition, and are about 10 kilometers in a general way. When the connection of each portion of the system is completed, the whole system can be taken as a three-dimensional space tether system. A local space elevator system model is established, and a control strategy is provided aiming at the features of the system; an attitude control rule aiming at a target system is provided; and the experiment shows that: even if the system is operated on an orbit having a large eccentricity ratio, the swing inhibition method of local space elevator system in the elliptical orbit can avoid the occurrence of chaotic motion to cause out of control. The swing of the system can be effectively inhibited, and the configuration of the system can be maintained on the state required by the operation of the system. The size of the thrust required by control accords with the actual situation of the engineering, and the control effect is good.

The invention relates to a single-row reciprocating piston type rotor engine, belonging to the field of power generating machines. The engine comprises a main shaft (1), an elliptic orbit groove disc body (2), a transmission connecting rod (3), a positioning U-shaped fork (4), a cylindrical rotary cylinder body (5), a rotary cylinder body end cover (6), a turning lever (7), a turning lever shaft (14), a traditional single cylinder engine nose (8), a piston connecting rod (9), a piston (10), a cylindrical enclosure (11) and the like, wherein the lower end of the transmission connecting rod (3) is provided with a bearing (21) inserted into an elliptic orbit groove (12); a horizontal elbow of the transmission connecting rod (3) passes through the tube wall of the cylindrical rotary cylinder body (5); a perforation is sealed by using a shaft seal (13); the horizontal elbow is connected with one end of the turning lever (7) by using the bearing; the piston connecting rod (9) is connected with the other end of the turning lever (7) also by using the bearing; the turning lever shaft (14) is fixed on the bottom of the rotary cylinder body (5) and is in positioning connection with the turning lever (7) by using the bearing. The invention has the advantages of simple and compact structure, few components, light weight, small volume, high efficiency, energy saving, convenience of maintenance and long service life and not only can be made into a gasoline engine but also can be made into a diesel engine.