11 results about "Orbital speed" patented technology

The invention belongs to the technical field of ocean surface flow velocity remote sensing measurement, and particularly relates to a Luoga No.2 SAR (Synthetic Aperture Radar) flow velocity correction method based on Bragg orbit velocity, which comprises the following steps of: verifying the physical correctness of an orbit velocity mechanism relative to a phase velocity mechanism; under a DopRIM framework, a physically consistent sea wave error correction model is established; calibrating parameters of the sea wave error correction model; applying the calibrated sea wave error correction model to along-orbit interference SAR observation data of the Luoma No. 2 satellite, and calculating the radial surface flow velocity and sea wave error of each resolution unit; comparing an output result with a matched reference truth value to evaluate the effectiveness of the sea wave error correction model; comparing the sea wave error correction model with a KaDOP model, and respectively counting and analyzing simulated sea wave errors under different methods; the sea wave error simulation result is visually displayed, and the improvement effect is quantitatively evaluated through in-situ observation data and a KaDOP simulation result.

Disclosed is a method for ejecting at least two satellites (2; 2a, 2b) from a launcher (3), said at least two satellites (2; 2a, 2b) being oriented along a central axis coinciding with a longitudinal axis (A) of the launcher and secured together side by side, the method comprising: a. an initial orientation of the direction of the longitudinal axis (A) of the launcher, so as to form a predetermined angle (θ) with an orbital speed axis (X), the predetermined angle (θ) being between 20° and 60°, b. a first separation comprising the simultaneous separation from the launcher (3) of said satellites (2; 2a, 2b) along the longitudinal axis (A) with a first predetermined differential speed (dV1) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining secured together, then c. a second separation comprising the separation of said satellites (2; 2a, 2b) from each other along an axis perpendicular to the central axis and included in the orbital plane, with a second predetermined differential speed (dV2).

In one embodiment, an air-scoop includes an air inlet that air molecules enter the air-scoop through at an orbital speed when the air-scoop is moving through an atmosphere at the orbital speed. The air-scoop also includes a rotor that is rotated by a motor at a rotational speed, and the rotor includes multiple rotatable blade stages. A first one of the rotatable blade stages has a blade configuration that maximizes transparency of the first one of the rotatable blade stages to air molecules entering the air-scoop through the air inlet at the orbital speed when the rotor is rotating at the rotational speed. A last one of the rotatable blade stages has a blade configuration that maximizes opacity of the last one of the rotatable blade stages to air molecules in the air-scoop flowing directionally toward the air inlet when the rotor is rotating at the rotational speed.

The application discloses a high-precision time offset autonomous calculation method and system based on satellite speed v and gravity acceleration g. By establishing a closed correlation between the gravity field strength and the orbital speed, the traditional multi-parameter relativistic time correction model is simplified to a closed-form calculation depending on only two in-orbit measurable physical quantities, without the need for input of orbital radius, universal gravitational constant, earth mass, external ephemeris, etc. The application has outstanding originality, simple calculation, strong autonomy and reliable precision, can realize a daily cumulative time error of less than 1 ns, is suitable for low-orbit satellites, navigation satellites, constellation systems and various high-reliability time service scenarios, has a wide application range and high industrial value, and can significantly improve the autonomous time service capability of a spacecraft and the stability of a time system.

The present invention relates to the field of astronomical calculation technology, specifically a method and system for calculating the relationship between the sun and planets, which includes the following steps: collecting solar activity observation data, analyzing the time series, intensity, and frequency information of the data, extracting the characteristics of solar activity in different cycles, summarizing the change trends, and calculating the correlation between the cycle and the activity intensity to obtain solar activity correlation data. In the present invention, by extracting the characteristics of solar activity in different cycles and combining with the magnetohydrodynamics method, the velocity and density parameters of electromagnetic radiation and particle flow are deeply analyzed, the perturbation characteristics of the planetary magnetic field are scientifically evaluated, the interference of solar activity on the planetary orbit deviation is clarified by analyzing the gravitational perturbation, the dynamic changes of the orbital velocity and orbital intersection are traced, the correlation between magnetic storms and orbital deviation is incorporated into the prediction analysis, and the analysis of the orbital change trend is further improved by combining the rotation period and gravitational perturbation data, providing a more refined reference basis for predicting planetary motion and astronomical phenomena.

The present application belongs to the technical field of remote sensing measurement of ocean surface current velocity, and in particular to a Lujia-2 SAR current velocity correction method based on Bragg orbital velocity, comprising the following steps: verifying the physical correctness of the orbital velocity mechanism relative to the phase velocity mechanism; and under the DopRIM framework, establishing a physically consistent sea wave error correction model; calibrating the parameters of the sea wave error correction model; applying the calibrated sea wave error correction model to Lujia-2 satellite along-track interferometric SAR observation data to calculate the radial surface current velocity and sea wave error of each resolution unit; comparing the output results with the matched reference true value to evaluate the effectiveness of the sea wave error correction model; comparing the sea wave error correction model with the KaDOP model, respectively, to statistically analyze the simulated sea wave error under different methods; visualizing the sea wave error simulation results, and quantitatively evaluating the improvement effect through in-situ observation data and KaDOP simulation results.

A method is provided for efficient orbital launch trajectories. A payload (e.g., satellite) is launched as high as a first radius with respect to the center of the Earth. The method decreases the payload altitude in response to a gravitational pull of the Earth, and ultimately the payload attains a stable orbit around the Earth at a second radius with respect to the center of the Earth. If the first radius is twice the second radius, the payload acquires a gravitational first potential energy at the first radius, and in the stable orbit the payload has a second (potential and kinetic) energy equal to the gravitational first potential energy, with the second kinetic energy being equal to the energy required to maintain a stable orbital velocity. Advantageously, the stable orbit can potentially be at any orbital inclination angle in the range between 0 and 360 degrees.

In one embodiment, a method includes collecting and compressing air with an air-scoop that includes an air inlet that air molecules enter the air-scoop through at approximately an orbital speed when the air-scoop is moving through an atmosphere at approximately the orbital speed and a rotor configured to be rotated by a motor at approximately a rotational speed. The method also includes providing the collected and compressed air for supply to a thruster as propellant.

A passive landing system for decelerating a payload delivered from orbit or at orbital velocity onto a non-atmospheric celestial body, including a catcher assembly positioned to intercept incoming payloads at orbital speeds, and a deceleration system mechanically coupled to the catcher assembly and extending along a sloped terrain, wherein the system decelerates the payload by exchanging momentum between the descending payload and the extended deceleration system, without the use of propulsion. A landing system for oversized payloads on a non-atmospheric celestial body, which can be used in combination with the passive landing system, including a structure configured to house and secure large-volume cargo during descent and deceleration, a surface interaction mechanism configured to establish frictional contact with regolith of the celestial body, and a dynamic deployment system that controls the extent and duration of surface contact during descent, wherein the landing system decelerates primarily by converting kinetic energy into heat and mechanical resistance through friction with the regolith.

The present invention discloses a residual phase error correction coefficient estimation method based on a sub-aperture cancellation combined quadratic fitting method, comprising the following steps: (1) using the satellite orbit velocity and the beam movement velocity to preliminarily estimate the approximate value of the radar velocity, using the wKA algorithm to roughly image the SAR echo data, and generating a complex image; (2) windowing and intercepting the complex image, setting the initial correction parameters, dividing the sub-aperture and pre-processing, constructing the objective function E1(X), and calculating the image difference after sub-aperture imaging; (3) gradually approaching the optimal solution of the phase error through reflection, expansion, and contraction optimization strategies; (4) after reaching the termination condition, constructing the objective function #imgabs0#, and performing fine optimization based on quadratic fitting to obtain the optimal phase correction value x; (5) using the wKA algorithm to azimuthally compress the obtained optimal phase correction value x to achieve azimuth imaging; the present invention can work stably under different scenarios and radar configurations.

The invention discloses an atomic oxygen generation device for ground simulation of an ultra-low orbit atmospheric environment. The atomic oxygen generation device comprises an atomic oxygen generation module, a speed screening module, a neutralization module, a shell and an electronic gun, the atomic oxygen generation module comprises an ionization source device, one side of the shell is connected with the ionization source device through a flange and a matched nut, and pure oxygen forms high-speed oxygen ions in the ionization source device and then is guided into a cavity of the shell; the speed screening module comprises parallel electrode plates arranged in the cavity and an external magnetic field arranged on the outer side of the shell, and the module realizes high-speed oxygen ion screening and directional transmission; the screened high-speed oxygen ions enter a neutralization module on the right side of the cavity, an electron gun is located at an outlet of the cavity, electrons are injected into the neutralization module, and atomic oxygen with the real orbital velocity and neutral state characteristics is formed. The high-speed atomic oxygen generator can generate high-speed atomic oxygen with the orbit equivalent speed of 7800m / s, improves the atomic state retention capability and the flow field stability, and has the characteristics of modularization and low power consumption.