301 results about "Sun sensor" patented technology

The present invention includes a sun sensor assembly having an aperture that defines an area that is less than the area of the photo-detecting surface of a corresponding first photo-detector. According to another aspect, the present invention also includes a solar concentrator includes at least two sun sensor assemblies mounted on the concentrator in a manner to help the solar concentrator track the sun. According to another aspect, the present invention also includes a method of processing electrical signals from two or more photo-detectors. According to yet another aspect, the present invention includes a sun tracking system that includes a solar panel that includes a solar concentrator and control system.

A container for sunscreen having (a) an alphanumeric indication of a sun protection factor and (b) a solar detector for displaying whether accumulated ultraviolet radiation striking the skin of a user of sunscreen with the indicated sun protection factor has exceeded a threshold level. The solar detector has a sensor layer with a photochromic material that reacts to ultraviolet radiation by changing from a clear state to a colored one, and an ultraviolet absorbing or blocking material covering at least one portion of the sensor layer. The ultraviolet absorbing material covering the at least one portion being present in an amount or concentration such that, upon exposure of the solar detector means to the ultraviolet radiation striking the skin of the user of the sunscreen over the same period of time, blocks sufficient ultraviolet radiation from striking the photochromic material so as to prevent the photochromic material from changing color until the threshold level has been reached at which time the photochromic material changes from the clear state to the colored one.

The invention relates to a three-dimensional gesture determining and local positioning method of a lunar surface rover, which comprises the following steps: (1) ascertaining the rolling and pitching angles by use of a triaxial accelerometer with sensitivity while the rover is still; (2) determining the drift angle gesture by means of a sun sensor; (3) using the axial gesture and the gyro deviation as the state quantity, the rolling and pitching angles established by the triaxial accelerometer, the drift angle determined by the sun sensor as well as three gyro outputs as the measuring information, building a state equation and a measuring equation, and estimating the triaxial and gyro deviations by means of extended Kalman filter; (4) after compensating the gyro outputs by virtue of the estimated gyro deviations while the rover is in motion, calculating the gesture changes of the rover, finishing the preestimation of the gyro gesture, and fulfilling gesture update; (5) acquiring the information about the rotation speed of the driving wheels of the rover, the rotating angle of the steering wheel, the rotating angle of the left and right rocker arms, and getting the position increment of the rover in the body coordinate system by use of the forward kinematics relationship. The invention has the advantages of high precision of gesture determining and positioning, simple calculation and easy implementation of the engineering.

An integrated microelectromechanical system (MEMS) sun sensor includes a filter, microlens, aperture and a folded MEMS optical element combined with an active pixel sensor array to form an integrated spacecraft sun sensor in an integrated sealed package, offering lower power, smaller size and higher performance for use on spinning spacecraft useful in attitude determinations. Multiple like sun sensors can be disposed for increasing the reliability, spatial coverage or spatial resolution for a specific performance requirement.

A system for determining a new orientation and / or position of an object comprises a sky polarimeter configured to record image data of the sky, a signal processing unit, and logic configured to receive and store in memory the image data received from the sky polarimeter. The logic calculates the Stokes parameters (S0, S1, S2,), DoLP, and AoP from the image data, detects obscurants and filters the obscurants (such as clouds and trees) from the image data to produce a filtered image. The logic is further configured to find the Sun and zenith in the filtered image, and to determine the roll, pitch, yaw, latitude and longitude of the object using the filtered image. A method for determining a new position / orientation of an object comprises recording raw image data using a sky polarimeter, calculating S0, S1, S2, DoLP, and AoP from the image data, detecting obscurants and filtering the obscurants from the image data to produce a filtered image, obtaining last known position / orientation data of the object, finding the Sun and zenith in the filtered image, and determining the roll, pitch, yaw, latitude and longitude of the object using the filtered image.

The invention relates to a high-accuracy APS sun sensor, which comprises: a light lead unit, a picture sensor, a signal processing unit and an interface unit, wherein said light lead unit is connected to the picture sensor; the signal processing unit is connected to the picture sensor and the interface unit; said interface unit is connected to the main computer; the light lead unit is used to generate the speckle array picture; the picture sensor is used to output the pixel gray valve of speckle array picture; the signal processing unit, based on said gray valve, tests the state angle; the interface unit is used to transmit the state angle to the main computer. The invention has high speed, small volume, large view angle, high accuracy and low power consumption.

The lighter-than-air vehicle for shading is an airborne, movable system that provides shade from the sun for an open area. The vehicle is a substantially flat flying device having upper and lower surfaces, and which is filled with a lighter-than-air gas or gases, such as helium or hydrogen. The vehicle may be tethered to the ground through control wires, or may be fully autonomous and controlled by propeller fans. The required power may be provided by wire from ground, or through a portable power source, such as solar cells, mounted atop of vehicle. The vehicle has sun sensors that allow it to track the sun, and provides around the clock shaded area underneath its surface. The vehicle can also have pressure sensors and altitude sensors that may trigger an alarm in case of an emergency descent to the ground.

The invention discloses a zero momentum magnetic control sun capture device and method of a satellite. The device comprises a 0-1 type sun sensor, a magnetometer and a magnetic torquer, wherein the 0-1 type sun sensor and the magnetometer are matched, and are used for performing satellite attitude measurement; the output of the magnetic torquer acts on the satellite body after delay, and is used for providing control moment for the satellite in the satellite rate damping process and controlling the attitude angle deviation to be less than a preset threshold value, so as to realize zero momentum magnetic control sun capture; when the 0-1 type sun sensor judges that the sun appears in the field of view, the double-vector attitude determination FORM algorithm is adopted, and a three-axis attitude matrix of the satellite is determined through the 0-1 type sun sensor and the magnetometer; when the 0-1 type sun sensor judges that the satellite enters the shadow zone, the satellite elevation angle is determined through the magnetometer. According to the zero momentum magnetic control sun capture device and method, the single machine configuration of the control system is simplified, the reliability of an attitude control subsystem is improved, and the resources on the satellite are saved.

The invention discloses a sun sensor and a measuring method thereof, relating to a device of an attitude sensor for measuring the relative positions of an aerocraft such as a satellite and the like and the sun and a measuring method thereof. The device comprises a light path introducer, a photosensor and a signal processor. The light path introducer consists of a reference optical fiber and introducing optical fibers, wherein the reference optical fiber is arranged at the center of the light path introducer, 2-2n pairs of introducing optical fibers are uniformly distributed on a circular track taking the reference optical fiber as the center of the a circle, wherein n is 2, 4 or 8, each pair of the introducing optical fibers comprise two optical fibers which are symmetrically distributed based on the center of the circle, and the end face normal of each introducing optical fiber and the end face normal of the reference optical fiber form an included angle of 15-60 degrees. The photosensor consists of photoelectric sensing elements with the same number as that of optical fiber, and the output end of each optical fiber is in butt joint with a photoelectric sensing element. The invention has the characteristics of high-precision resolution ratio, small volume, light weight, simple structure, low price and suitability for aerospace environment, and has wide generalization and application prospects.

The invention discloses a calibration method for digital sun sensor, comprising following procedures. According to outer and inner parameters of sun sensor calibration system, unified mathematical model of sun sensor imaging is built. Two rotary shafts of rotary table are rotated to different angle so as to obtain data of calibration point and they are transmitted to processing computer by interface circuit. Processing computer sends said data of calibration point to the unified mathematical model and calibration parameters are calculated by two-step method. The invention also relates to the device applying the method. The device supplies illumination by sun simulator and data of different points can be obtained by rotating two-shaft rotary table. Processing computer is used to record data of calibration point and calculate corresponding calibration parameter. The method and the device are suitable for multiple digital sun sensors. The invention adopts unified building method of inner and outer parameters and improves calibration accuracy. At the same time because installing aligning accuracy is not required strictly the whole calibration process is simple.

The invention discloses a two-step on-orbit iterative calibration method based on multiple parameters for a multi-error model of a microsatellite magnetic sensor. The method comprises the following steps: at first, calculating by utilizing a sun sensor and a top in an illumination area to obtain a satellite pitch angle; then, only calibrating zero-offset errors by using ground calibration or last-measurement scale factors as standard; and finally, re-building a multi-error calibration model, and re-calibrating the zero-offset and scale factor errors by using a best-fit method. According to the method, the zero-offset and scale factor errors of a magnetometer are calibrated and compensation in an attitude algorithm is carried out, so that the navigation accuracy of the satellite is effectively improved; a new filter does not need to be designed for a calibration algorithm, real-time calculation is also not needed, so that the operation cost and the load of the satellite are effectively reduced.

The invention relates to a satellite multiple attitude control mode test system based on a double gimbal control moment gyroscope (DGCMG) structure. Performance of three satellite attitude control systems which are based on a DGCMG, a single gimbal control moment gyroscope (CMG) and a flywheel is tested and verified through a DGCMG actuating mechanism. The satellite multiple attitude control mode test system based on the DGCMG structure comprises a platform system, the satellite attitude control system, a space environment simulation system and a ground station system. The platform system is composed of a tri-axial air bearing table, a satellite service comprehensive management system, a power source and a wireless bridge and used for simulating satellite dynamic characteristics and information management. The satellite attitude control system is composed of a jet propulsion system, the DGCMG, a fiber-optic gyroscope, a star sensor, a sun sensor and a global position system (GPS) receiver and used for determination of an attitude and an orbit and control of a satellite platform. The space environment simulation system is composed of a GPS simulator, a sliding block which is of a pyramidal structure, a sun simulator and a star simulator and used for simulating space interference torque, and part performance of a GPS satellite and a celestial body. The satellite multiple attitude control mode test system based on the DGCMG structure can provide ground testing and verification for multiple attitudes of a satellite.

The invention discloses a three-axis sun-oriented control method of a satellite for a guaranteeing satellite-earth link. The three-axis sun-oriented control method comprises the following steps: solving a projection of a sun vector under a satellite orbit coordinate system according to the sun vector calculated according to a satellite orbit; establishing a sun-oriented reference coordinate system according to the satellite orbit and the definition of the satellite polarity, and calculating an attitude quaternion of the sun-oriented reference coordinate system relative to the satellite orbit coordinate system; calculating an error quaternion between an attitude quaternion of a star relative to the satellite orbit coordinate system and the attitude quaternion of the sun-oriented reference coordinate system relative to the satellite orbit coordinate system; carrying out attitude reference tracking according to a symbol of the error quaternion; and when continuously meeting that the error quaternion is smaller than a preset threshold value, establishing a star sensor sun-oriented sign, carrying out three-axis stabilization sun-oriented control by virtue of a star sensor, and otherwise, continuing to carry out two-axis sun-oriented control based on sun sensor. By virtue of the three-axis sun-oriented control method, the smoothness of the satellite-earth link after the sun orientation of a sailboard can be guaranteed, and a near-optimal maneuvering path is guaranteed when a sun-oriented maneuvering manner is changed into a ground-oriented maneuvering manner.

The invention which discloses a precision compensation method for an area APS digital sun sensor belongs to the spacecraft attitude control measure system field. The precision compensation method comprises the small-angle incidence angle precision compensation and the wide-angle incidence angle precision compensation. The precision compensation method comprises the following steps: 1, carrying out primary rotation correction on the solar faculae center coordinate value which is obtained through the calculation of a sun sensor image processing chip and is input by the sun sensor image processing chip; 2, carrying out secondary correction on the coordinate value of the wide-angle incidence angle, and utilizing the rotarily-corrected coordinate value to do the coordinate mapping of a refraction model; 3, calculating through an uniaxial higher-order polynomial fitting process to obtain the solar angle tangent value, utilizing the rotarily-corrected coordinate value to calculate if the incidence angle is a small angle, and utilizing the secondarily-corrected coordinate value to calculate if the incidence angle is a large angle; and 4, finding the arc tangent to obtain the sun angle value. The precision compensation method which is suitable for the precision compensation of the area APS array digital large-field sun sensor has the advantages of major error factor compensation and high precision.

A lighter-than-air aircraft includes a gas envelope for containing a buoyant gas, and a solar panel is carried by a predetermined portion of the gas envelope. A solar sensor is used for determining a direction of the sun. A propulsion system carried by the gas envelope orients the gas envelope so that the solar panel is oriented in the direction of the sun based upon the solar sensor. A navigation controller cooperates with the propulsion system to move the lighter-than-air aircraft along a desired flight path while the solar panel is oriented in the direction of the sun.

Under one aspect of the present invention, a monolithic sun sensor includes a photosensor; a spacer material disposed over the photosensor; and a patterned mask disposed over the spacer material and defining an aperture over the photosensor. The spacer material has a thickness selected such that the patterned mask casts a shadow onto the photosensor that varies as a function of the monolithic sun sensor's angle relative to the sun. The sun sensor may further include a substrate in which the photosensor is embedded or on which the photosensor is disposed. The spacer material may be transparent, and may include a layer of inorganic oxide, or a plurality of layers of inorganic oxide. The patterned mask may include a conductive material, such as a metal. The aperture may be lithographically defined, and may be square. The sun sensor may further include a transparent overlayer disposed over the patterned mask.

A semi-physical simulation test system for sun-oriented control comprises a satellite-borne computer system, a gyro, an infrared horizon instrument, a star sensor, a three-axis magnetometer, an actuator and a target simulator. The target simulator simulates the space environment. A motion simulator simulates the attitude motion of an aircraft. The satellite-borne computer system performs attitude and control quantity calculation according to measurement data acquired by single measurement machines, and outputs the attitude and control quantity to the actuator. The semi-physical simulation test system further comprises a sun sensor equivalent device, a dynamic geomagnetic field simulator, a target simulator, a motion simulator, a PXI acquisition control computer, a dynamics simulator, a data distribution unit, a telemetry, remote control and remote injection machine, a database, a display terminal, a CAN bus network and an optical fiber reflective memory network. The problem that ground simulation verification of sun-oriented control cannot be carried out under abnormal attitude rolling of an aircraft in the prior art is solved. The semi-physical simulation test system has the beneficial effect that the test authenticity and effectiveness are improved.

The invention provides a stable control system and method for a satellite spinning around the sun. A simulation sun sensor is used for measuring a sun vector under a satellite body system; a triaxial magnetometer is used for measuring a terrestrial magnetism vector under a satellite body; a controller is electrically connected with the simulation sun sensor and the triaxial magnetometer and used for obtaining the sun vector and the terrestrial magnetism vector, judging the current mode of the satellite and generating and sending a control command according to the current mode; and the magnetic torquer is used for outputting the corresponding magnetic torque according to the control command sent by the controller. By the adoption of the system and method, the satellite can stably spin around the sun through the high-reliability and low-power-consumption execution component magnetic torquer, and the satellite is passively stabilized and obtains energy. A guarantee is provided for the reliability and the safety of the satellite, and the defect that large-power-consumption and consumptive components such as a propeller and a flywheel are adopted in a satellite in the prior art is overcome.

The invention relates to a multi-mode attitude determination method for a remote sensing micro nano satellite, and belongs to the technical field of spacecraft attitude control. The method comprises the steps that a magnetometer, a sun sensor and three star sensors serve as original attitude measurement data sources, according to a telemetry command received by the satellite and by combining withthe working state of the sensor, two attitude determination modes are designed, and the two modes are the sun pointing attitude determination mode and the remote sensing imaging posture determinationmode. For the sun pointing attitude determination mode, the magnetometer and the sun sensor serve as measurement input, in combination with the Calman filtering method, the sun pointing attitude determination mode has the advantages of being low in power consumption and high in reliability, and certain precision attitude determination can be provided to meet the control requirements of the micro nano satellite sailboard sun pointing. The remote sensing imaging attitude determination mode uses the three star sensors as the measurement input, in combination with the three-star sensitive attitudefusion method, high-precision micro nano satellite attitude and angular velocity determination can be achieved, and the high-precision attitude control requirements of micro nano satellite remote sensing imaging are met.

The invention relates to a sun track system based on GPS and GIS and a relative method, which uses a lighting control terminal and a detection control center to control the sun track system and automatically track sun. The lighting control terminal controls variable control modules, a GPS terminal, a sun sensor module, a terminal communication module, a memory module and a detection module which are all connected with a terminal control module and mounted on the lighting control terminal. The detection control center comprise a plurality of control terminals, a main communication module, a layer memory module and a layer management module which are all connected with the main control module and mounted on the detection control center. The invention can enable the sun track device to accurately track sun, effectively improve solar energy application efficiency, reduce traditional energy consumption and reduce environment pollution, with automatic positioning, correction, detection and error report.

The invention discloses an optical imaging minisatellite attitude control system and an in-orbit working mode switching method, belongs to the field of satellite attitude control and aims at solving the problem that existing optical imaging satellites cannot be simply and effectively switched in the orbit. The optical imaging minisatellite attitude control system comprises an attitude measurement sensor, an actuating mechanism and an attitude controller, wherein the attitude measurement sensor comprises a sun sensor, a star sensor and a gyroscope; the actuating mechanism comprises a counteraction flywheel and a magnetorquer. According to an in-orbit working mode switching method of the optical imaging minisatellite attitude control system, the optical imaging minisatellite attitude control system has six working modes, namely a rate damping mode, a sun capture mode, a sun-orientated three-axis stabilization mode, an earth-orientated three-axis stabilization mode, a data transmission mode and a safety mode. The optical imaging minisatellite attitude control system and the in-orbit working mode switching method are suitable for all the optical imaging minisatellites.

The converting method for measuring datum of sun sensor includes calibrating the inside and outer parameters of sun sensor with a sun simulator and a high precision two-axis turntable. A diaphragm receiving screen is set between the sun simulator and the sun sensor, and a CCD image locating and aligning system is used to judge the verticality of the simulating sunlight to the reflecting surface of the cubic mirror, so as to complete the measurement on the vector relation of two reflecting surface normal directions of the cubic mirror and simulating sunlight. By means of the vector relation and the outer parameters of the sun sensor, the converting matrix for conversion from the measuring coordination system of the sun sensor to the coordination system of the cubic mirror is determined. Corresponding conversion is completed. The device for the converting method is also disclosed. The present invention has simple method and high precision.

The invention discloses a novel space station extravehicular inspection maintenance unit (EIMU), and belongs to the design field of aerospace crafts. The novel space station EIMU comprises intravehicular monitoring equipment, an antenna and an EIMU device, wherein the intravehicular monitoring equipment and the EIMU device are in communication through the antenna. The intravehicular monitoring equipment comprises a video display terminal, an inspection device control platform and a manual remote control device. The EIMU device comprises solar cell panels, a micro-mechanical arm, sun sensors, nitrogen propellers, a binocular stereo vision camera, a three-axis magnetorquer, a momentum wheel, a satellite-borne computer board and a communication board. The novel space station EIMU can provide extravehicular inspection and maintenance service for large spacecrafts such as space stations and ensure that the spacecrafts can achieve in-orbit operation safely and fulfill tasks smoothly.

A method for determining the attitude of a satellite based on a visible light earth sensor and a sun sensor uses combination of the visible light earth sensor and the sun sensor to determine the earth triaxial attitude of the satellite. The method comprises the following steps: processing an earth image formed by a CMOS imaging device, extracting the effective edge of the earth image, and fitting the center of the extracted effective edge by using a least square technology; approximately calculating the earth roll angle and the pitch angle under the assumption that the earth attitude angle of the satellite is a small angle; and calculating by using the sun sensor according to the earth roll angle and pitch angle attitude information obtained after approximate calculation to obtain the earth yaw angle of the satellite. Compared with traditional methods for determining the earth triaxial attitude angle of the satellite through using earth infrared sensor and sun sensor combination, the method provided by the invention has the advantages of improvement of the measurement precision of the attitude angle of the satellite, suitableness for attitude measurement of small satellites, and conformation to the current development trend of the small satellites.

The invention relates to a thick sun sensor for sun azimuth angle measurement, which comprises a main structure, a base, a battery pack, a standard mirror, an electric connector and an earthing assembly, wherein the main structure is mounted on the base; the main structure is rectangular pyramid shaped; mounting grooves are formed in both the four side surfaces and the top surface of the rectangular pyramid; the mounting grooves are used for mounting the battery pack; a ring band surface is arranged on the periphery of the four side surfaces of the rectangular pyramid; a light blocking ring is arranged on the outer side of the ring band surface; the light blocking ring adopts a thin-walled ring band structure; the height of the light blocking ring is not larger than the height the undermost position of a sensitive surface of the battery pack; the standard mirror, the electric connector and the earthing assembly are mounted on the base. The thick sun sensor has the advantages of simple structure, small size, light weight, large measurement viewing field (2 pi solid angle), stray light interference resistance and space irradiation resistance, and has a main backup measurement mode.

The invention discloses a micro-satellite platform multi-sensor data dynamic fusing system which comprises a spinning top, a star sensor, a sun sensor, a magnetometer and a posture modifying module. The spinning top is connected to a posture predicting module. The posture angle which is outputted by the star sensor, the posture parameter error value which is outputted by the spinning top and a first error state variable estimated value which is outputted by a first information distribution system through a first sub-filter are transmitted to a data fusing module. The sun vector which is outputted by the sun sensor, the posture parameter error value which is outputted by the spinning top and a second error state variable estimated value which is outputted by a second information distribution system through a second sub-filter are transmitted to the data fusing module. The geomagnetism vector which is outputted by the magnetometer, the posture parameter error value which is outputted by the spinning top and a third error state variable estimated value which is outputted by a third information distribution system through a third sub-filter are transmitted to the data fusing module. The data fusing module is connected with the first sub-filter, the second sub-filter and third sub-filter. The posture modifying module is connected with the posture predicting module and the data fusing module.