66 results about "Sun-synchronous orbit" patented technology

The invention discloses an acquisition method of inclination biased quantity for sun synchronous orbit. The method comprises the following steps: 1, correcting an initial value i0 of an orbit inclination by virtue of a correction value delta I of the initial value of the orbit inclination; 2, establishing an ordinary differential equation to acquire an orbit inclination at moment t and orbit right ascension of ascending node, and variation delta LTDN of local time of descending node at moment t; 3, judging the orbit drifting direction, wherein the biased quantity of the orbit inclination is negative if the direction is toward east, while the biased quantity of the orbit inclination is positive if the direction is toward west; and 4, traversing inclination biased quantities with step length of 0.0001 degree in a range from 0.0 to 0.1 degree, and respectively adding the corrected initial value of the sun synchronous orbit inclination in the step 1 to serve as an orbit initial value with inclination biased quantity, and carrying out the step 2 to obtain the variation of the local time of descending node at the moment t, arranging the maximum values of absolute values of variations into vectors according to a sequence, and the minimum value in the vectors is the inclination biased quantity. According to the method, the defect of low calculation accuracy of inclination biased quantity of the conventional sun synchronous orbit inclination biased quantity is overcome, and the orbit control capability is effectively improved.

A method for obtaining satellite solar angle and time on the basis of satellite orbit characteristics comprises the following steps: (1) according to the satellite sun-synchronous orbit characteristics, calculating the orbit inclination angle from the height of the satellite orbit; (2) according to the number of satellite orbits, calculating the longitude and Beijing time of sub-satellite point of the position where a satellite descending orbit intersects the equator, and converting the Beijing time into the local time so as to obtain the time of the position; (3) calculating to obtain the local times of any latitude where the satellite sub-satellite point pass; (4) according to the orbit injection error of satellite orbit inclination angle and the daily change rate of satellite orbit inclination angle during the operation process, calculating to obtain the change of satellite local time and actual local times of any latitude; (5) calculating to obtain the actual local time after satellite side-sway; (6) calculating to obtain the solar altitude, solar azimuth, and the Beijing time of an observation site. The provided method can precisely obtain the solar altitude, solar azimuth, and imaging time of a satellite according to a remote sensing satellite nominal orbit and changes thereof without knowing the satellite ephemeris data and specific imaging time.

The invention provides an optimization design method of a common track emergency reconnaissance constellation based on a sun synchronous return orbit. Firstly, the longitude and latitude coordinates of a target center position, the satellite overhead time and the reconnaissance frequency requirements are inputted. Secondly, the orbital altitude, eccentricity and orbital inclination of the satellite are calculated. Thirdly, the Greenwich star time angle is calculated according to the satellite overhead time. Then, the true nearest point angle, the right ascension of the reference satellite andthe number of satellites in the constellation are calculated. Then, the right ascension of each satellite in the constellation is optimized according to the right ascension of the ascending intersection point of the reference satellite, the true nearest point angle and the reconnaissance frequency. Finally, the orbital altitude, eccentricity, orbital inclination, perigee amplitude, right ascensionand true proximity of each satellite in the constellation are output. By using the method, the emergency reconnaissance constellation which meets the requirement of target area coverage can be designed quickly.

The invention provides a satellite temperature control method allowing rapid responding and multi-orbit adaption, wherein by adopting a heat pipe network, reinforcing contact thermal conduction, and spraying a high-emissivity thermal control coating inside a satellite to reinforce in-cabin thermal conduction and radiation heat exchange, a temperature field is distributed uniformly in a cabin; a whole satellite temperature level is adjusted by a compensative electric heater; a nightside of a side plate of the satellite is taken as a major heat dissipation face, the outer surface of the nightside is sprayed with a white paint, and all the other outer surfaces of the satellite body, except for the major heat dissipation face, are covered by multiple layers of thermal insulation assemblies; and components outside the satellite are sprayed with a white paint thermal control coating. A propelling cabin has independent design, wherein the entire outside of the cabin is covered by the multiple layers of the thermal insulation assemblies, a baseplate inside the cabin is covered by the multiple layers of the thermal insulation assemblies, and the propelling cabin is finally integrated with a platform cabin. By integral design of machinery, heat and electricity, the satellite temperature control method allowing the rapid responding and the multi-orbit adaption provided by the invention realizes thermal control requirements of the whole satellite under conditions of a sun-synchronous orbit and an inclined track by the temperature control method.

The invention provides a gesture maneuvering trajectory calculation method for a synchronous orbit SAR (synthetic aperture radar) satellite and relates to the field of gesture transition between a yaw guide gesture and a yaw bias gesture of a satellite. Through planning a posture motion track in a yaw angle phase plane, the problem of the gesture transition between the yaw guide gesture and the yaw bias gesture is solved. Only two points among all yaw guide gestures, with the gesture angular velocity thereof to be zero, can be adopted as the cut-in or cut-out point of a state for the gesture transition between the yaw guide gesture and the yaw bias gesture. In this way, the transition between gestures can be conducted at any moment.

The invention relates to a method for acquiring the ground surface temperature through NPPVIIRS data. By means of the method, appropriate ground surface temperature data can be provided for weather, agriculture drought and environment monitoring. The method overcomes the defect that due to the fact that a VIIRS sensor is not provided with a moisture channel, the inversion precision is limited. According to the method, MODIS sensor moisture data carried by an Aqua satellite operating at a sun-synchronous orbit with the descending node being 13: 30 are used for acquiring the moisture data of VIIRS data observation swath. The ground surface temperature inversion algorithm suitable for 11 microns and 12 microns of the VIIRS sensor is derived, an atmospheric radiation transfer model MODTRAN4 is used for simulating and calculating the transmittance of an M15 channel and the transmittance of an M16 channel of the VIIRS sensor under different atmospheric moisture contents (0.4-6.0 g/cm<2>, and increasing is carried out with 0.1 g/cm<2> as the step size), the emissivity of ground features corresponding to the M15 channel and the emissivity of ground features corresponding to the M16 channel of the VIIRS sensor are estimated according to International Geosphere Biosphere Program (IGBP) data, and finally the inverting of the ground surface temperature is obtained. According to the method, study is carried out on the inversion of the ground surface temperature based on the VIIRS sensor, and stable transmission from the MODIS data to the VIIRS data in the agriculture environment data monitoring is facilitated.

The invention discloses a designing method for installation orientation of a three-field star sensor on a sun-synchronous orbital satellite. According to the method, vector relation between stray light and a coordinate system of a satellite body is analyzed on the basis of a vector model for stray light and the coordinate system of a satellite body during in-orbit operation of the satellite, and areas free of influence by stray light are resolved by establishing a stray light boundary curve mathematical equation; on the basis of included angle relation of boundary curves, the method determines workable conditions of three star sensors and the scope and equation of installation orientation and calculates relation between a maximum sunlight shielding angle and a maximum earth-atmosphere radiation shielding angle of the light shields of the star sensor for suppressing stray light; and the method brings forward optimal schemes for installation orientation of star sensors to allow at least two star sensors to be in operation under a single condition when installation conditions for double star sensors are not met in areas free of influence by stray light under a variety of working conditions. The method has the advantage that a design scheme of practical application value is provided for installation orientation and designing of star sensors to prevent influence of stray light.

The invention discloses a space debris orbit determination traversal observation method and system. A camera installed in a +Y-axis direction of a satellite arranged on a terminator sun synchronous orbit is controlled to rotate around a +Z axis; large-range scanning from a starting deflection angle to an ending deflection angle is achieved, then space debris orbit determination traversal observation of the synchronous belt is achieved; and in the observation process, through real-time calculation, the rotating angle of a satellite around the +X axis of a satellite body is controlled to ensurethat the visual axis of a camera is always perpendicular to the geosynchronous belt at the specified latitude.

The invention relates to an apogee ignition high-precision analytical orbit autonomous prediction method. Orbit information of each phase of a transfer orbit including apogee ignition can be accurately calculated. The second type of nonsingular orbital elements are selected for recurrence so as to be suitable for the situations of small orbit inclination angles and low eccentricity. Therefore, the apogee ignition high-precision analytical orbit autonomous prediction method is also suitable for orbit calculation of each phase of a synchronous orbit and has great significance.

The invention provides a sun positioning satellite star sensor mounting orientation acquiring method and system. The method comprises the following steps: acquiring a geogas light boundary curve according to a boundary curve and a geogas light half corn angle of an earth core-satellite vector of an orbit satellite to be mounted in a satellite body coordinate system, and taking a stray light interference free area surrounded by the geogas light boundary curve as a mounting orientation area of an orbit satellite star sensor to be mounted. By analyzing distribution of a 'satellite-earth core/sun'vector in the satellite body coordinate system in an in-orbit operation process of a satellite, an analysis model of a spatial stray light boundary curve in a satellite sun positioning mode is established, then star sensor mounting orientation and stray light inhibition angles can be confirmed, and mounting orientation of a sun-synchronous orbit satellite star sensor with the sun positioning modeis instructed.

A system which comprises a sun-synchronous orbit satellite (3) revolving around the sun (5) and apparently revolving around the earth at a period of approximately one year while keeping the direction and the geometry to the sun and the earth (4) constant at an ultrahigh altitude of several millions km keeping clear of the influence zone of the earth, and which provides any of astronomical or global observation service and communication service. An observation satellite, which meets such conditions that a vector normal to a virtual orbital plane centering around the earth is substantially fixed regardless of revolution, can sweep the substantially entire celestial sphere during a single revolution while avoiding the effect of light and heat from the earth, and can keep a local solar time at a sub-satellite point on the surface of the earth constant, can be attained. A communication satellite, where the substantially hemisphere of the earth can be covered at a time simultaneously with one satellite and the substantially entire region of the earth can be covered with a small number or three satellites, is attained. Geometrical relations with the sun and the earth can be frozen approximately by the adjustment of the i/e ratio or by the use of perturbation by the attraction of the earth gravity. Operation from launch to insertion into an intended orbit can be achieved with significantly low fuel consumption by utilizing the perturbation owing to the attraction of the earth gravity.

The invention relates to a star sensor installation and layout method based on a remote sensing satellite gaze gesture, and relates to the field of application to a remote sensing satellite on a sun-synchronous orbit. The problems that the existing star sensor layout method is complicated, the repeated iteration is used; in addition, the effectiveness at the limited gesture of the satellite can only be met are solved. A space spherical surface is built; a region for effectively avoiding stray light on the built space spherical surface is determined; a region for effectively avoiding sun lighton the space spherical surface is determined; the determined region for effectively avoiding the stray light and the region for effectively avoiding the sun light are subjected to intersection taking;the obtained intersection region is an effective region for star sensor effectiveness in the satellite gaze gesture; the sphere center of the space spherical surface is used as a starting point of the star sensor optical axis vector; a random spare point in the effective region is selected to be used as a terminal point of the star sensor optical axis vector; the installation layout position of the star sensor is determined. The star sensor layout method is simple, convenient and fast. The requirement for meeting the star sensor effectiveness under all gesture conditions can be directly solved in one step.

The invention discloses an orbit control method based on a wheel-controlled mode. According to high-precision orbit freezing characteristics of a satellite in a sun-synchronous orbit, the control method combining a double-frequency global position system (GPS) and DORIS orbit determination, high-precision star sensitivity posture determination and high-precision wheel controlling is adopted in orbit controlling, satellite orbit control precision is improved, satellite high-precision orbit freezing tasks are achieved, and the orbit control method is a satellite orbit control method which is never adopted in the past. From 2011, the orbit control method is successfully applied to orbit controlling of a plurality of satellites in the sun-synchronous orbit, the orbit control precision reaches advanced levels in the world, and good effects are obtained.

The application relates to an earth gas light suppression method and system for a synchronous band target optical observation satellite. The system includes a satellite, an optical observation moduleand a baffle, wherein the orbit of the satellite selects a dawn-dusk sun-synchronous orbit, and maintains the attitude to be earth-oriented; the optical observation module is installed on the satellite and is perpendicular to the main axis of the satellite; and the baffle is installed on the optical observation module, and during the operation of the satellite, the baffle is arranged between the optical observation module and a ground plane, wherein the satellite is operated according to an attitude maneuver rule model to achieve the traversal observation of latitudes that need to be stared. By adopting the technical scheme of the application, the suppression of the earth gas light can be realized only through a simple structural design, and the operation is simple and the cost is saved.

The invention discloses a satellite based on a compound-eye camera. The satellite comprises a compound-eye camera system and a satellite platform system for carrying the compound-eye camera. Simultaneously, the invention also provides a space-based observation system thereof. The space-based observation system comprises a plurality of satellites based on the compound-eye camera, which are networked. The satellite platform system adopts a 1000kg platform, the viewing field of the compound-eye camera system is 18*18 degrees, the viewing field of each satellite is 180*360 degrees, and the satellite flies on a sun-synchronous twilight orbit to form the space target monitoring capability; the satellite platform system adopts a 3000kg platform, the viewing field of the compound-eye camera system is 1*1 degree, the viewing field of the satellite is 60*60 degrees, and the satellite flies on a sun-synchronous orbit to form the earth reconnaissance capability; the satellite platform system adopts a 9000kg platform, the viewing field of the compound-eye camera system is 0.1*0.1 degree, the viewing field of the satellite is 12*12 degrees, and the satellite flies on a stationary orbit to form the capability of monitoring a target moving back and forth between the space and the earth.

The invention relates to a solar tracking system posture adjustment method being able to prolong the solar measuring time of a space instrument, and belongs to the technical field of solar space measurement. The method comprises for making a space solar measuring instrument working on a sun-synchronous orbit to acquire more solar measuring time, acquiring the orbital elements of a spacecraft and the current system time and taking the orbital elements of a spacecraft and the current system time as the input parameters to calculate the position of the sun within the certain time frame before the solar measuring instrument enters the sun illumination zone; according to the input parameters of the first step, estimating the right ascension and the declination of the sun under an earth's core inertial coordinate system; calculating a coordinate transformation matrix from the earth's core inertial coordinate system to an orbital coordinate system, after coordinate transformation, acquiring the estimated coordinate position of the sun in the spacecraft orbital coordinate system; acquiring the estimated coordinate position of the sun in the spacecraft body coordinate system; acquiring the estimated coordinate position of the sun in the solar tracking system coordinate system; and taking the estimated coordinate position as the reference input for the control system, and driving a yaw oscillation sub system and a luffing sub system to move to a target position.

The invention relates to a radiator suitable for a space optical remote sensor in a sun-synchronous orbit and aims to solve the problem that the radiator cannot be arranged on a non-irradiation surface of the remote sensor during actual work. The working efficiency of the radiator can be improved through effectively shielding external heat flow of the sun-synchronous orbit. The radiator is composed of a radiant panel, an earth screen, a solar screen and a side screen, wherein the solar screen shields the solar radiation heat flow and shields the earth infrared heat flow on the side concurrently; the earth screen shields the earth infrared and solar reflection heat flow; and the triangular side screen is used for shielding the earth heat flow from the direction. During actual application, the shielding screens are arranged according to specific gestures and illuminated angle of the remote sensor, the whole or most part of the external heat flow from the sun and the earth directions can be shielded, the structure comprises no driving parts, and the radiator is simple and reliable.

The invention belongs to the field of optical remote sensing of satellites, and discloses a global multi-target all-day real-time monitoring system and method. The system comprises a motorized carriermodule, an image acquisition module, and a ground center module, wherein the motorized carrier module consists of 24 satellites uniformly distributed on 4 orbital planes; and the local time of a satellite ascending node is 10:30, and the height of an orbit is a sun-synchronous orbit. The system can simultaneously monitor multiple targets in real time and obviously improve the time resolution of photographic observation, so that a real-time monitoring effect similar to a video is achieved; and furthermore, by optimal arrangement of remote sensing constellations, customization of multiple observation targets by a user, and motorized control of an attitude of a satellite covering all the targets, all-day video-similar real-time monitoring can be simultaneously performed on global multiple targets.

The invention discloses a power supply capability prediction method for a sun-synchronous orbit satellite power supply system. The method comprises the following steps that 1, n historical telemeasuring parameter data related to the output power of a power supply array and a charging array on the satellite are acquired; 2, the telemeasuring parameter data are ordered according to the time identifier i; 3, the historical output power E of n effective data of the satellite power supply system is calculated; 4, a satellite motion period, including the power supply phase in which the satelliteis in a light area and the no-power-supply phase in which the satellite is in a shadow area, is defined, and the length T of each satellite motion period and the average period T<-> are calculated; and 5, the average period T<-> is taken as the next prediction period and output power prediction is performed on the time points within the next time period of which the prediction length is T<->. Thesatellite motion period is extracted through the historical telemeasuring data, and the output power of the satellite power supply system in the following time periods is predicted by using the powersupply periodic characteristics of the satellite power supply system.