19 results about "Gravitation" patented technology

The invention provides a method for displaying a hotspot keyword. The method comprises the steps of obtaining the hotspot keyword, generating a display body used for displaying the hotspot keyword, implanting the hotspot keyword into the display body, and controlling the display body with the implanted keyword to move towards the bottom end of a rotated web page in a gravity falling mode after the web page where the display body with the implanted keyword is located is rotated. The invention further provides a device and a browser for displaying the hotspot keyword. The display body is generated by obtaining the hotspot keyword, the hotspot keyword is implanted into the display body, and after the web page where the display body with the implanted keyword is rotated, the display body with the implanted keyword moves towards the bottom end of the rotated web page in the gravity falling mode. Therefore, dynamic and stereoscopic display of the hotspot keyword is achieved, the display effect is improved, the visual sense is enriched, and the interaction performance is high.

The invention provides a multifunctional cycle power generation system and method by a gravity lever. The system comprises a supporting rack and a balance weight arranged on the supporting rack and slidably matched with the supporting rack, wherein engines are arranged on one side or two sides of the balance weight; the input shaft of each engine is connected to a transmission shaft; two or more single-directional clutches are fixedly connected to the middle of the transmission shaft; the locking directions of the single-directional clutches are same, gears are fixed to the peripheries of thesingle-directional clutches, and a rack matched with each gear to drive the transmission shaft to rotate is arranged next to each single-directional clutch; at least one pair of racks is arranged on the upper part and the lower part, separately, the racks and the balance weight which reciprocates are in driving movement and the balance weight is arranged on the supporting rack to swing through a lifting mechanism in a reciprocating manner. As the balance weight which swings in the reciprocating manner pulls the racks to drive the transmission shaft to rotate in the single direction, multiple engines can be driven synchronously to generate electricity synchronously, and the power needed by a supporting rack lifting device is alleviated greatly by means of a swing mode like to a lever, so that the energy sources are saved and the power generation efficiency is improved greatly.

The invention provides a satellite-borne high-precision orbit extrapolation method, a satellite-borne high-precision orbit extrapolation device and electronic equipment. The satellite-borne high-precision orbit extrapolation method comprises the following steps: providing an orbit extrapolation input quantity of a satellite at an initial extrapolation moment; calculating an extrapolation duration based on the initial extrapolation moment and a preset threshold value, and determining a termination extrapolation moment of the satellite based on the extrapolation duration; calculating the non-spherical gravitational perturbation acceleration of the earth according to the non-spherical gravitational perturbation model of the earth; calculating a day-moon three-body gravitational acceleration according to a day-moon three-body gravitational model; determining a satellite thrust acceleration according to the current working state of the satellite; adding the earth non-spherical gravitational perturbation acceleration, the sun-moon three-body gravitational acceleration and the satellite thrust acceleration to obtain an extrapolation resultant acceleration; performing orbit extrapolation calculation on the basis of the extrapolation resultant acceleration by adopting a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation speed at an extrapolation termination moment; and according to the target extrapolation position and the target extrapolation speed, updating the orbit extrapolation input quantity at the next extrapolation moment.

The invention discloses a moon rotation parameter resolving method based on earth-moon microwave distance measurement, which comprises the following steps of: S1, performing earth-moon distance measurement based on microwaves by utilizing a two-way distance measurement mode, establishing a distance measurement observation model, and resolving a distance measurement observation value; s2, constructing a microwave moon measurement model to obtain a distance measurement calculation value; the microwave moon measurement model takes a moon rotation parameter and a tide LOVE number as model parameters; s3, deriving a partial derivative of the microwave moon measurement model relative to a rotation parameter of a single moon and a tidal LOVE number; and S4, performing nonlinear least square solution by using the distance measurement observation value, the distance measurement calculation value and the partial derivative to obtain a moon rotation parameter. On the basis, microwave distance measurement and a microwave moon measurement model are fused, the advantages of all-weather and high-frequency sampling of microwave distance measurement are fully utilized, the limitation of a single technology is broken through, high-precision and real-time calculation of moon rotation parameters is achieved, and a new technical approach is provided for moon internal structure research and gravity theory verification.

In the current space navigation, communication, and defense technologies, it is difficult to integrate local control of time, gravity operation, use of dark matters, and communication with the outer space (the Buddhist monk Tenkai's gyrus), and intergalaxy communication, ultrahigh-speed movement, and creation of a time acceleration / deceleration space cannot be realized by modern technologies.SOLUTION: The present invention is a future note related to an ultra-future space technology capable of integrally operating time, gravity, and material behavior by performing local and large-scale dark matter control by a BHc / WHc gate based on a dark matter fluid space model and a space gate theory.SELECTED DRAWING: None

The application discloses a lunar orbit and libration parameter solving and forecasting method and system, and the method comprises the following steps: considering the interaction among the mass point gravity model based on the post-Newtonian framework, the moon figure, other planet mass points in the solar system and the influence of the earth figure, a high-precision lunar orbit dynamics model is established under the general relativity framework; considering the lunar moment of inertia and the lunar torque, a double-layer lunar libration attitude dynamics model is established; when the initial conditions of integration meet the preset accuracy, an adaptive step integrator and an adaptive step integrator with time delay effect are used to respectively forecast the orbit parameters and the libration parameters, and initial products of the orbit and the libration parameters with time sequence labels are obtained; when the initial conditions of integration do not meet the preset accuracy, an automatic differentiation-based nonlinear solver is used to process historical observation data, and the optimal solution of the to-be-estimated parameters in the initial conditions is obtained, and re-forecasting is performed.

A kind of relative navigation reference deviation calibration method of star and pulsar information fusion is disclosed, which comprises the following steps: configuring a star angular distance measuring sensor and an X-ray detector on a spacecraft; observing stars by the star angular distance measuring sensor to extract navigation information contained in two kinds of relativistic effects, i.e., star aberration caused by the movement of the spacecraft and light deflection caused by celestial gravitational field; observing pulsars by the X-ray detector to obtain pulse arrival time difference observations; combining spacecraft and earth orbit dynamics model to predict the state; processing star angular distance and pulse arrival time difference observations by an extended Kalman filter to correct the state prediction value, and obtaining the estimation of the spacecraft position and velocity, the geocenter position and velocity, and the sensor system error parameters, so as to realize the calibration of two kinds of reference deviations, i.e., earth ephemeris error and sensor system error. The present application provides a new way to improve the performance of relativistic navigation system.

The geosynchronous orbit debris inclination vector evolution calculation method disclosed by the application comprises the following steps: firstly, calculating the influence of the J2 term of the earth's non-spherical perturbation on the change rate of the inclination debris, calculating the angular frequency of the geosynchronous orbit debris inclination vector, the period of the precession and the angular rate of the precession under the influence of the J2 term; then, calculating the frequency of the geosynchronous orbit debris inclination vector precession, the period of the precession and the angular rate of the precession caused by the sun's gravity and the moon's gravity respectively; finally, synthesizing and calculating the different angular frequencies, the periods of the inverse precession and the angular rates of the precession. The geosynchronous orbit debris inclination vector evolution calculation method disclosed by the application separately calculates the effects of the J2 term of the earth and the sun and the moon, and compared with the traditional geosynchronous orbit debris inclination vector evolution calculation method which adopts complete numerical calculation, the method has the advantages of fast calculation speed and can be used for the analysis and calculation of the long-term evolution of the geosynchronous orbit debris orbit.

The invention discloses a high-precision space target orbit forecasting method based on numerical fitting, and relates to the technical field of space target orbit forecasting, frequency domain decomposition is combined with a harmonic compensation model, and seasonal system drift caused by sun-moon gravity is thoroughly eliminated; high-frequency components peel off attitude noise through light pressure parameterization and smooth difference, and perturbation coupling error transmission is avoided; a harmonic order adaptive mechanism captures a main disturbance term with the minimum calculation amount; carrying out least square batch estimation to realize coefficient primary solution; the template library pre-loading technology effectively compresses the time consumed for new target initialization; in addition, a geomagnetic modulation factor is introduced to dynamically suppress compensation distortion during a magnetic storm period, a sliding window mechanism guarantees real-time updating of a coefficient, and stable forecasting is still maintained under an extreme event; attitude measurement is replaced by luminosity curve inversion, so that non-cooperative target forecasting becomes possible; tangential component directional compensation ensures the stability of an orbital plane, and reduces the fuel consumption of satellite position protection.

The invention relates to the technical field of satellite orbits, in particular to a sun-synchronous orbit dip angle offset calculation method and device, equipment and a medium. According to the method, an orbit perturbation equation containing earth non-spherical gravity and atmospheric damping perturbation is constructed, and the influence of semi-major axis attenuation caused by atmospheric resistance on the inclination angle offset is considered. According to the method, the influence of atmospheric damping on the orbit height and different orbit height maintaining strategies are comprehensively considered, more accurate orbit injection parameters are calculated in combination with the prior art, and a more reliable guarantee is provided for subsequent on-orbit normal operation of a spacecraft. In the initial stage of orbit design, more accurate spacecraft orbit injection parameters can be calculated by introducing more perturbation models. The orbit height maintaining strategy is also considered, which means that the later-stage spacecraft in-orbit operation and maintenance cost is also considered, the engineering implementation is more reliable, the result is more accurate, and meanwhile, flexible adjustment can be realized by fully combining actual conditions.

A planetary learning device is provided. The device is comprised of a body that may take various geometric shapes resembling celestial objects, such as planets or stars, and may incorporate phosphorescent materials and LEDs for simulating natural celestial phenomena. The device also includes at least one speaker for providing audio content related to the celestial object, with facts or scientific data being played via a removable memory card or streamed wirelessly. The body can be charged through a docking station, which may include a rotating platform to simulate planetary motion. Additionally, the device features sensors for interactive functionality and may include an augmented reality application for enhanced learning. A vibration motor can further simulate seismic activity or gravitational forces, offering an immersive educational experience.

The invention relates to a method for calculating the east gradient of a sea area disturbing gravity prime component, and provides a method for calculating the east gradient of the sea area disturbing gravity prime component aiming at the disturbing gravity prime component auxiliary navigation requirement of an inertial navigation system under the constraint of a sea area motion carrier position error. And constructing a sea area perturbation gravitational prime component east gradient calculation model based on a completely normalized global gravitational potential coefficient by utilizing a physical geodetic boundary value theory to realize sea area perturbation gravitational prime component east gradient calculation. The invention aims to improve the disturbance gravity prime unitary component aided navigation capability of the inertial navigation system under the constraint of the position error of the motion carrier and improve the navigation precision of the inertial navigation system.

The application discloses a high-precision space target orbit prediction method based on numerical fitting, and relates to the technical field of space target orbit prediction.The frequency domain decomposition of the application is combined with a harmonic compensation model to completely eliminate seasonal system drift caused by the gravitation of the sun and the moon; high-frequency components are stripped of attitude noise through light pressure parameterization and smooth difference to avoid the transmission of perturbation coupling error; a harmonic order adaptive mechanism captures the main disturbance term with the least amount of calculation; least square batch estimation realizes one-time solution of coefficients; template library preloading technology effectively compresses the time consumption of new target initialization; in addition, the application introduces a geomagnetic modulation factor to dynamically suppress compensation distortion during magnetic storms, and a sliding window mechanism guarantees real-time updating of coefficients, and stable prediction is still maintained under extreme events; photometric curve inversion is adopted instead of attitude measurement to make non-cooperative target prediction possible; tangential component directional compensation ensures the stability of the orbit plane and reduces satellite position fuel consumption.

The present application discloses a small celestial body offset efficiency evaluation method based on state transition tensor, and belongs to the field of spacecraft orbit mechanics. The method of the present application is as follows: a multi-body system dynamics model is established, the impact geometry of the small celestial body relative to the earth is calculated, and the state transition tensor of the small celestial body orbit is derived, so as to quickly calculate the influence of the initial state change of the small celestial body on the impact geometry, and obtain the orbit offset caused by the impact of the impactor when the earth's gravity is dominant after the small celestial body approaches the earth; the second-order state transition tensor is used to obtain the orbit offset of the small celestial body after a period of time, and the high-efficiency quantitative evaluation of the offset efficiency of the small celestial body after being impacted by the impactor is realized. By fixing the velocity deviation and changing the position and direction of the applied deviation, the optimal change direction corresponding to the optimal offset efficiency of the small celestial body orbit can be determined. The present application has the advantages of high evaluation accuracy and high efficiency, and is suitable for offset strategy selection and efficiency evaluation of small celestial body defense tasks.

The application discloses a kind of lunar space optical observation orbit determination precision estimation method, solve the problem that orbit determination precision evaluation needs to be calculated by a large number of Monte Carlo simulation, causes the problem of calculation resource and time consumption, belong to orbit determination precision estimation field;Including: using circular restricted three-body problem to describe the gravitational model of lunar space, construct the dynamics equation of target in lunar space under conjunction coordinate system;Linearized dynamics equation and the differential equation that state transition matrix satisfies are obtained;Construct target imaging signal-to-noise ratio model and apparent magnitude model to obtain target observation standard, construct the Jacobian matrix of observation model according to the relative position relationship of target and sensor;Based on posterior Cramer-Lo bound theory, combined with the differential equation that state transition matrix satisfies and the Jacobian matrix of observation model, construct orbit determination precision estimation index model, the lower bound of position error and the lower bound of velocity error are used as precision estimation index in the application, reduce the amount of calculation, effectively reflect the observation orbit determination effect.

The application discloses a kind of small celestial body overall stress analysis method, comprising the following steps: (1) for small celestial body, its boundary is, a spherical surface is set to its complete encirclement outside small celestial body, the three-dimensional space surrounded by spherical surface is recorded as;(2) the spherical surface is discretized, generates triangular grid, the area and unit normal vector of each triangle element are calculated, and are recorded as and;(3) the gravity vector of triangle element center is obtained, and is recorded as;(4) is substituted into formula (), and the resultant force that small celestial body is received is calculated.

This invention discloses a precise orbit determination and astronomical ranging method based on equivalent optical path bending correction using celestial body motion position offset, relating to the fields of aerospace telemetry and control and astronomical observation technology. This invention is based on the premise that light always travels in a straight line, abandoning the traditional approach of modeling light beams by gravitational self-deflection and spacetime curvature. It utilizes the fixed duration of light propagation, during which high-speed motion of celestial bodies generates misalignment of transmission and reception points, forming a real physical mechanism of equivalent optical path bending at the measurement level. The method models and calculates the offset and time delay corresponding to the equivalent bending, corrects errors in the original measured data, calculates the true absolute position of the target, deduces the orbit, and implements pre-orbit compensation for the satellite. This invention conforms to objective physical laws, has high correction accuracy, significantly reduces the frequency of satellite orbit fine-tuning, saves fuel and manual maintenance costs, requires no hardware modifications, and can be widely applied in aerospace telemetry and control, deep space exploration, and astronomical observation.

The embodiment of the invention provides a device for judging the motion balance state of an object based on attractive force and magnetic force, which analyzes the interaction among a plurality of objects in periodic motion, particularly uniform circular motion, based on the attractive force and magnetic force, so as to predict the information of the objects, such as the motion direction, the motion plane and the motion period when the objects reach the balance state. In a specific embodiment, the formation reason of a galaxy disc shape is analyzed, the inclination angle of earth rotation is calculated, and the possibility of developing a corresponding gravitational magnetic device by referring to an existing electromagnetic device is analyzed.

The invention discloses a potential threat asteroid-based virtual impact earth asteroid data set generation method, system and product. The method comprises the following steps: inputting orbit data of a real potential threat asteroid group, and setting impact time; solving a corresponding earth position, and calculating and generating initial orbit elements of a virtual impact body including ascending node longitude, near point argument and true anomaly under a heliocentric two-body dynamic model on the basis that the impact occurs at the ascending node or the descending node; with the initial orbital element as an initial value, correcting the orbital element of the virtual impactor by adopting a differential correction algorithm in a high-precision dynamic model of the solar system, so as to accurately impact the earth at the impact time; a planetary gravity capture model is adopted to optimize an impact orbit, and the perigee height of a virtual impact body is corrected to a proper range; carrying out reverse recursion on the optimized perigee orbit parameters to the outside of the earth gravity influence ball, converting the position speed into orbit elements, and obtaining the sun center orbit elements before the virtual impactor impacts the earth; and obtaining a data set.