64 results about "Orbit correction" patented technology

The invention provides a Halo orbit on-orbit keeping method considering amplitude constraint, and belongs to the aerospace technology field. The method comprises the steps that by establishing a kinetic equation under a limited three-body model composed of two main celestial bodies and a detector, the position of a balance point of a three-body system composed of the two celestial bodies and the detector is determined; a Halo orbit, near the balance point, in the three-body system composed of the big and small celestial bodies and the detector is determined; a differential correction algorithm for Halo orbit keeping is designed according to the disturbance variable; a Halo orbit on-orbit keeping strategy is designed according to the differential correction algorithm; in a real ephemeris environment, measurement errors and execution errors are considered, the speed increment of each time of orbit correction is obtained according to the orbit keeping strategy through the differential correction algorithm, orbit correction control is conducted according to the speed increments. Accordingly, Halo orbit on-orbit keeping considering amplitude constraint can be achieved, and consumption of fuel needed in orbit keeping can be reduced as much as possible.

The invention relates to a method for selecting time of deep space probe approach orbit correction maneuver, which belongs to the field of spaceflight and aviation and solves the problems of deep space probe approach orbit correction maneuvering time including low control precision, great calculation amount and poor independency in the prior method for selecting time of deep space probe approach section orbit correction maneuver. The method comprises: 1, determining miss distance rhof passing through the B plane of an object celestial body; 2, determining the acceptable maximum miss distance rhomax of a deep space probe; 3, judging whether the miss distance rhof is more than the acceptable maximum miss distance rhomax, if yes, executing step 4; otherwise, returning to step 1; and 4, executing deep space probe approach section orbit correction maneuver. The method not only can save fuel at a rate of 6m / s, but also can ensure that final passing-through precision is increased by more than 100m as compared with ground station designated maneuver time precision under the same condition and reaches within 250m.

The present invention relates to a solar blind ultraviolet camera absolute radiometric calibration equation coefficient in-orbit correction method, wherein a camera is subjected to laboratory absolute radiometric calibration to obtain the absolute radiometric calibration coefficient and the non-uniformity correction coefficient of the camera, the atmosphere background radiation value calculated based on the atmosphere radiation transmission model can be considered as the true background radiance when the camera in-orbit operating meets a certain constraint of the imaging mode, the location and the illumination condition, the background radiance is inverted according to the laboratory absolute radiometric calibration coefficient and the image local gray mean, and a linear regression method is used to analyze the inverted radiance value and the true radiance value so as to achieve the in-orbit correction of the calibration equation coefficient. According to the present invention, according to the solar blind ultraviolet spectrum atmosphere background radiation characteristic, the high-frequency and operational in-orbit absolute radiometric calibration of the solar blind ultraviolet spectrum remote sensor can be supported, and the blank of the in-orbit absolute radiometric calibration method of the domestic camera at the spectrum is filled.

The invention discloses a delay chain temperature drift on-orbit correction device and a delay chain temperature drift on-orbit correction method based on FPGA. The device comprises a temperature sensor, a delay chain, a D trigger array, a decoding unit, a lookup table storage unit and a temperature drift correction control unit, wherein the temperature sensor is used for collecting real-time temperature; the delay chain is used for reading out time information of an input signal relating to a clock edge; the D trigger array is connected to the delay chain, for locking an output state of each delay unit in the delay chain when arriving the clock edge and outputting temperature coding data; the decoding unit is connected to the D trigger array, for converting the temperature coding data into binary code data and outputting; the lookup table storage unit is used for storing integral nonlinear lookup table data of the delay chain marked at a preset temperature point; the temperature drift correction control unit is connected to the temperature sensor, the decoding unit and the lookup table storage unit respectively, for achieving on-orbit correction of delay chain temperature drift. According to the delay chain temperature drift on-orbit correction device and method based on FPGA, correction of the delay chain temperature drift is achieved by fewer sources, and high time resolution capacity is ensured.

Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

The invention belongs to the satellite navigation field and provides a DGNSS satellite orbit deviation correction method. The method includes the following steps that: broadcast ephemeris and precise prediction ephemeris are utilized to calculate double-difference station-satellite geometric distances; a double-difference orbit correction number of a mobile station and a reference station is calculated according to the double-difference station-satellite geometric distances; an orbit error is corrected according to the double-difference orbit correction number. With the DGNSS satellite orbit deviation correction method adopted, the separation of the orbit error from other distance related errors such as tropospheric and ionospheric delay errors can be realized, and the quality of the orbit deviation correction number can be monitored and controlled more effectively; and correction accuracy and stability are significantly improved. The method has a wider application range. The DGNSS satellite orbit deviation correction method can satisfy the requirements of the correction of orbits of baselines of different lengths such as medium-short baselines, medium-long baselines and ultra-long baselines.

The present invention relates to an orbit correction method for a charged particle beam, and aims to solve problems inherent in conventional aberration correction systems and to provide a low-cost, high-precision, high-resolution optical converging system for a charged particle beam. To this end, employed is a configuration in which a beam orbit is limited in ring zone form to form a distribution of electromagnetic field converging toward the center of a beam orbit axis. Consequently, a nonlinear action outwardly augmented, typified by spherical aberration of an electron lens, can be cancelled out. Specifically, this effect can be achieved by an electron disposed on the axis and subjected to a voltage to facilitate the occurrence of electrostatic focusing. For a magnetic field, this effect can be achieved by forming a coil radially distributed-wound on a surface equiangularly divided in the direction of rotation to control convergence of a magnetic flux density.

The invention provides a real-time on-orbit correction method of slow varying errors among a plurality of star sensors. The slow varying errors between every two head parts of the star sensors are corrected. The method comprises the following steps of S1, calculating the coordinate conversion quaternion filtering practical measuring value Q[RotOHiOHjCAL] between every two head parts of the star sensors; S2, specifying a certain head part on the star sensor as a reference head part OHref by a ground control center; S3, calculating the coordinate conversion quaternion ground precise measuring value Q[OHiOHref] from the reference head part OHref to each head part OHi of the star sensors; S4, calculating the coordinate conversion quaternion practical measuring value Q[RotOHiOHref] from each head part OHi of the star sensors to the reference head part OHref; S5, calculating each head part posture quaternion q[OHiRc] after the correction of the slow varying errors between the head parts according to the Q[OHiOHref] and the Q[RotOHiOHref] obtained in the S3 and S4. The method can realize the on-orbit real-time estimation and correction; by using a simple calculation mode, the posture determination precision of the combined fixed posture of the plurality of star sensors is improved; the slow varying errors between the non-reference head parts and the reference head parts is simply and effectively corrected; the head part posture quaternion after the correction of the slow varying errors between the head parts is obtained.

The invention relates to a method for selecting time of deep space probe approach orbit correction maneuver, which belongs to the field of spaceflight and aviation and solves the problems of deep space probe approach orbit correction maneuvering time including low control precision, great calculation amount and poor independency in the prior method for selecting time of deep space probe approach section orbit correction maneuver. The method comprises: 1, determining miss distance rhof passing through the B plane of an object celestial body; 2, determining the acceptable maximum miss distance rhomax of a deep space probe; 3, judging whether the miss distance rhof is more than the acceptable maximum miss distance rhomax, if yes, executing step 4; otherwise, returning to step 1; and 4, executing deep space probe approach section orbit correction maneuver. The method not only can save fuel at a rate of 6m / s, but also can ensure that final passing-through precision is increased by more than 100m as compared with ground station designated maneuver time precision under the same condition and reaches within 250m.

To provide a scroll compressor capable of reducing a noise occurring in a pin-and-ring type self rotation preventing mechanism and improving compression performance. In a scroll compressor comprising: a fixed scroll member and a revolving scroll member; a driven crank mechanism for driving the revolving scroll member to revolutionary turn; and a pin-and-ring type self rotation preventing mechanism provided in plural places for preventing self rotation of a revolving scroll member, at least one of the self rotation preventing pin, the self rotation preventing ring and the self rotation preventing ring hole, which form the pin-and-ring type self rotation preventing mechanism, is provided with an orbit correction part for reducing a maximum displacement R in a direction of self rotation of the revolving scroll member to smooth a change of an orbit of the revolving scroll member in changing a pin and a ring in a section of prevention of self rotation by means of a corresponding pin and ring part.

In a method for injecting a plurality of spacecraft into different circum-earth or interplanetary orbits individually in a single launch, a plurality of spacecraft coupled to an assist module are injected into an interplanetary orbit having a periodicity synchronous with the earth's revolution period. Then, in a maneuver allowing the assist module to re-counter with and pass near to the earth (Earth swing-by), the assist module appropriately performs an orbital change maneuver and a separation maneuver for each of the spacecraft in a sequential order. Through these maneuvers, a closest-approach point to the earth is changed for each of the spacecraft so as to drastically change a subsequent orbital element for each of the spacecraft. The assist module takes a sufficient time to determine a target orbit for each of the spacecraft with a high degree of accuracy until a half month to several days before a closest-approach time in the Earth swing-by. Based on the determined orbit, the assist module makes an orbit correction of about several m / sec in a sequential order, and then separates the spacecraft therefrom in a sequential order. In this process, an inertia navigation is performed based on an accelerometer mounted in the assist module and information about attitude.

The invention discloses an orbit control strategy on-orbit correction method and system based on thruster deviation estimation. The method includes: separately obtaining the positions of the center of mass of the entire satellite, the first orbital thruster and the second orbital thruster relative to the layout system information; and, obtain the thrust vector directions of the first orbital control thruster and the second orbital control thruster; perform on-orbit identification, and determine the generation of the first orbital control thruster and the second orbital control thruster according to the on-orbit identification results The torque; construct the disturbance torque evaluation function; solve the disturbance torque evaluation function, determine the minimum value of the disturbance torque evaluation function, and calculate the jet pulse width; perform orbit control according to the calculated jet pulse width. The invention adopts the energy optimal principle to update the orbit control thrust distribution strategy on the orbit, and realizes the minimization of the orbit control interference.

The invention discloses a low-orbit space target orbit correction method, a low-orbit space target orbit correction device and electronic equipment. The method comprises the following steps: selecting a plurality of reference targets which have similar orbits with a main target in a low-orbit space; wherein the similar orbit is determined according to the orbit characteristics of the main target in the low-orbit space; calculating a first track along-track direction error of each reference target; according to the first track along-track direction error of each reference target, calculating a second track along-track direction error of the main target corresponding to the reference target; calculating a final track direction error of the main target according to the second track direction error of the main target corresponding to each reference target; and track correction is carried out on the main target according to the final track along-track direction error of the main target. According to the method, the orbit correction of the main target is realized by utilizing the plurality of reference targets with the similar orbits to the main target, the orbit prediction precision is improved, and the method can be applied to the aspects of spacecraft high-confidence collision early warning, reentry target guide correction and the like.