47 results about "Atmospheric drag" patented technology

An apparatus for lowering an orbit of a space object includes an envelope, an inflation system for inflating the envelope, an inflation control system for controlling the inflation system, and attachment hardware for connecting the apparatus and the space object. Inflating the envelope increases an effective drag area of the envelope for increasing atmospheric drag on the envelope.

The invention discloses a method and device for cleaning space debris. According to the method, electromagnetic force is applied to the space debris running on an original orbit; the speed and / or direction of the space debris are / is changed under the action of the electromagnetic force; the space debris of which the speed is changed is changed to run on a new orbit under the action of gravity; at least the height of the perigee of the space debris on the new orbit is smaller than the height of the perigee of the space debris on the original orbit; the space debris running on the new orbit finally falls into the atmosphere of the earth under the action of atmosphere resistance. The device comprises a satellite bearing platform which is provided with an electric field device cleaning the space debris or a magnetic field device cleaning the space debris or a combination of the electric field device and the magnetic field device. The method and device for cleaning the space debris do not have high requirements for the accuracy of position detection of the space debris, micro space debris can be effectively cleaned, solar energy can be fully utilized for power generation, and self-carried energy is saved. The device for cleaning the space debris has the advantages of being simple in structure, low in manufacturing cost, and better in pulse action effect.

The invention relates to robustness analysis method for a spacecraft orbit control strategy, and belongs to the technical field of the spacecraft orbit dynamics and control. The robustness analysis method comprises the following steps: modeling orbital motion of the spacecraft through a Gauss orbit element perturbation equation; analyzing the non-spherical perturbation of a low-orbit satellite and the atmospheric drag perturbation; designing an orbit maintaining strategy to execute the maintaining control to the spacecraft orbit elements; using a differential correcting algorithm to improve the precision of the orbit maintaining; building a position error, velocity error and engine thrust error model in the running process of the spacecraft; designing a calculation model of spacecraft control error mean, variance and error distribution proportion; and using a Monte Carlo simulation method to execute the simulation analysis to the spacecraft orbit control strategy with the error, and building a robustness evaluation system of the orbit control strategy. In the control strategy design, the actual situation of the satellite orbit control is considered adequately, and the feasibility of the method in the actual project is guaranteed under the precondition of simpliness and convenience and according with the actual situation.

An apparatus for reducing atmospheric drag on a vehicle. The apparatus includes an airfoil having a receiving feature, wherein the airfoil is attachable with a back portion of the vehicle. The apparatus includes at least one extender attachable with the back portion of the vehicle, and a slide arm attached to the at least one extender and mateable with the recess portion of the airfoil. The apparatus is configured such that when the slide arm is disposed in the receiving portion of the airfoil, the airfoil is in an extended position relative to the back of the vehicle. The apparatus is further configured such that applying a force on the airfoil in a direction substantially perpendicular to the back portion of the vehicle buckles the airfoil, thereby lifting the slide arm from the recess and moving the airfoil to a retracted position.

The invention provides an electro-dynamic tether based satellite deorbit device and a method thereof, relating to the satellite deorbit device and the method thereof. The device and the method solve the problems of high deorbit cost of a propellant deorbit method and low deorbit efficiency of an atmospheric drag deorbit method. In the system, a control signal output end of a control module is connected with a control signal input end of a tether releasing and reclaiming device; the tail end of an insulating tether section of the tether is fixed on the tether releasing and reclaiming device; and the tail end of a conductive tether section of the tether is connected with an electron collection transmitter. The method comprises the following steps: determining longitude, latitude and height of the position of a satellite at current time according to six factors of the initial orbit of the satellite to acquire the magnetic field strength of the current position; calculating the Lorentz force generated by the tether; resolving the Lorentz force along the radial direction, the tangential direction and the outerplanar direction to acquire the orbit six factors at next time, and determining a new height of the satellite orbit; and circulating the steps until the height of the satellite is less than or equal to 120km to complete the satellite deorbit process. The device and the method are applicable to the satellite deorbit process.

The invention discloses a mid-term orbit prediction method and device of a low-orbit spacecraft and a storage medium. The method includes the steps that target parameters are determined, wherein the target parameters at least comprise the atmospheric density parameter of the position where the target spacecraft is located, and the equivalent windward area and the atmospheric drag coefficient of the target spacecraft; and the mid-term orbit of the target spacecraft is predicted through the target parameters. The effect that the mid-term orbit prediction accuracy of the low-orbit spacecraft is improved is achieved through the method and device.

The invention discloses a deorbiting device for a micro / nano satellite. The deorbiting device for the micro / nano satellite comprises a sail storing device, a sail unfolding device and an unlocking device, wherein the sail unfolding device is of a centrosymmetric structure, and comprises an upper end cover, a lower end cover, a clack-shaped central shaft, two ring flanges, two strip elastic masts, four supporting prisms, four pairs of pilot sleeves, four hold-down mechanisms and four leaf springs. After a ground command is received, the unlocking device releases the clack-shaped central shaft of the sail unfolding device; the strip elastic masts wound on the clack-shaped central shaft drive a thin-film sail, which is stored in a sail storing chamber, to unfold through releasing elastic potential energy stored by the strip elastic masts. According to the deorbiting device for the micro / nano satellite, the normal sectional area, in which a satellite moves, is increased by utilizing the unfolding of the thin-film sail, the effect of atmospheric drag perturbation on the micro / nano satellite is enhanced, and thus, the passive deorbiting of the micro / nano satellite is prompted.

The present invention relates to an anti-interference filtering method of a drag-free satellite attitude channel. Aiming at the drag-free satellite attitude channel of multi-source interference consisting of an atmospheric drag torque, an solar radiation pressure torque, a gravity gradient torque, inertia uncertainty, actuating mechanism noise, measurement noise, model uncertainty and the like, the method comprises: firstly, building a mathematical model including the drag-free satellite attitude channel of multi-source interference; secondly, building a mathematical model of an external environment interference torque; thirdly, designing an anti-interference filter aiming at the drag-free satellite attitude channel; and finally, designing an optimal controller to perform feedback control. The anti-interference filtering method of a drag-free satellite attitude channel is high in anti-interference capability, is high in control precision and the like, and is able to be used in the drag-free satellite attitude channel.

The invention discloses a self-stabilizing torque-free CubeSat brake sail de-orbit device. The self-stabilizing torque-free CubeSat brake sail de-orbit device comprises a locking device, a storage mechanism, a mounting panel, a conical spring, a deployable mechanism and a polyimide film aluminized secondary surface mirror. The locking device is fixed to the top surface of the mounting panel, and the conical spring, the deployable mechanism and the polyimide film aluminized secondary surface mirror are all disposed in the storage mechanism. The self-stabilizing torque-free CubeSat brake sail de-orbit device is fixedly connected to the bottom of a satellite through the top mounting panel, and a whole brake sail de-orbit mechanism is located outside the satellite, so that no space inside thesatellite is occupied. After receiving a ground command, the locking device releases the deployable mechanism, the deployable mechanism pops out, and belt-shaped elastic masts drive the polyimide filmaluminized secondary surface mirror to deploy. The cross-sectional area of a CubeSat in the flight direction is increased by deploying the polyimide film aluminized secondary surface mirror, and theatmospheric resistance received by the CubeSat is improved, so that the rapid de-orbit of the CubeSat is accelerated. The self-stabilizing torque-free CubeSat brake sail de-orbit device is simple in structure and high in reliability, does not occupy the space inside the satellite, and has little dependence on a satellite platform.

A streamlined projectile is disclosed that includes a base or rear having a plurality of concentric sections located within each other at the aft of the projectile and in successively smaller sizes. Within each concentric section is a chamber having a wall perpendicular to expanding combustion gasses, with the total area of these walls being greater than an area of the rear of the projectile itself. This creates a larger projectile base for a propellant to push against and for that explosion to be held within the section chamber for a longer period than that of a standard boattail or flat-tail projectile. In addition, the concentric sections at least partially negate atmospheric drag by occupying the region behind the projectile where a partial vacuum forms, resulting in a projectile that will go faster and further for a given propellant charge.

The invention relates to an atmosphere collecting device for an inflatable unmanned helicopter. The device is connected to a frame of the unmanned helicopter. The device comprises a vertical inflatable tube opened at two ends and a gas collecting bag sleeving the lower end opening of the vertical inflatable tube; the vertical inflatable tube is internally provided with an inflatable fan which is controlled to start or stop through a ground remote controller; the vertical inflatable tube is fixedly connected with the frame through a horizontal connecting rod; the opening of the gas collecting bag is fixedly connected with the lower end opening of the vertical inflatable tube through an elastic sealing clip. In use, the unmanned helicopter carries the gas collecting device to collect atmosphere in air and remotely controls to open the inflatable fan, so that the gas collecting bag inflated is automatically separated from the unmanned helicopter under atmospheric resistance. The elastic sealing clip automatically closes the gas collecting bag and the gas collecting bag drops under the action of the gravity. The device provided by the invention is short in collecting time and high in efficiency, and the requirement on duration of flight of the unmanned helicopter is reduced. Moreover, the device is simple in structure, reasonable in design and low in cost and effectively solves the problems that in altitude gas collection, equipment is complex, inconvenient to carry, poor in automatic effect and the like.

The invention discloses an atmospheric drag torque compensation method and system for controlling the attitude of an ultra-low orbit satellite. The attitude is guided to adapt to the atmospheric incoming flow of the satellite at different orbit positions under the condition that the hardware configuration of a satellite system is not additionally increased, thereby reducing the atmospheric drag and drag torque to the maximum degree, and improving the reliability; extra hardware cost of the satellite is not increased; atmospheric resistance is resisted without increasing propellant consumption,a large angular momentum flywheel system does not need to be adopted to absorb resisting moment, the method and system are completely achieved based on existing hardware configuration of a satellite,and the method and system are good in economical efficiency and particularly suitable for small satellites with low cost. Secondly, the control effect is good, test simulation shows that the aerodynamic resistance and the resistance torque can be reduced to one third of those before compensation, and good realizability is achieved.

A projectile launcher comprising an elongate barrel provided with a bore, a magnetic detent, an associated discarding sabot, and an associated subcaliber projectile. Pending launch, the magnetic detent exerts a magnetic restraining influence to hold the subcaliber projectile inside the bore within proximity of the sabot. Launch thrust overcomes the magnetic restraining influence of the detent upon the projectile, and the sabot separates from the projectile after launch under the influence of atmospheric drag.

The invention discloses an atmospheric resistance perturbation modeling and calculating method for a low earth orbit satellite. The method is specifically implemented according to the following steps:1, establishing a Box-Wing model of a satellite; 2, establishing a piecewise linear atmospheric resistance perturbation model; 3, according to the Box-Wing model established in the step 1 and the piecewise linear atmospheric resistance perturbation model established in the step 2, calculating the folded atmospheric resistance perturbation acceleration at the moment t. By utilizing the method to calculate the atmospheric resistance perturbation of the satellite, atmospheric resistance perturbation calculation errors caused by the fact that a traditional single atmospheric resistance factor isdifficult to accurately describe atmospheric changes can be overcome, the orbit determination precision is effectively improved, and technical support is provided for precise orbit determination of alow-orbit satellite, especially precise orbit determination under special conditions such as magnetic storm.

The invention provides a high-precision Mars high-rise atmospheric wind field and density inversion method. The specific implementation method comprises the following steps: constructing a relationship between atmospheric resistance acceleration entering a ball and parameters such as atmospheric wind speed and density, determining trajectory attitude parameters entering the ball by using inertialcomponents such as an accelerometer and a gyroscope, realizing atmospheric wind field parameter inversion based on an atmospheric resistance direction measured by the accelerometer, and performing modeling based on the atmospheric resistance modulus value and the atmospheric resistance coefficient to realize atmospheric density inversion. The method has the advantages of high precision and low cost; the effective load only comprises one three-axis accelerometer and one gyroscope, a high-precision result can be obtained based on high-precision data given by a precision instrument; the method can be realized based on a micro-nano satellite platform, a micro-nano satellite detection satellite group is established on the basis, and the full-coverage detection of the Mars high-rise atmosphere can be realized.

The invention discloses a low-orbit satellite constellation configuration keeping method. The method comprises the following steps: acquiring satellite orbit parameters of an initial low-orbit satellite constellation; determining a first relative drift amount of the inter-satellite configuration caused by the initial satellite orbit parameters in the constellation life period; determining a second relative drift amount of the inter-satellite configuration caused by atmospheric drag perturbation under the initial satellite orbit parameters during the constellation lifetime; according to the first relative drift distance and the second relative drift distance, determining a third relative drift distance of the inter-satellite configuration caused by other perturbation under the initial satellite orbit parameters in the constellation life period; determining an offset amount of a satellite orbit parameter required for offsetting the third relative drift amount; and adjusting satellite orbit parameters of the initial low-orbit satellite constellation according to the offset to obtain offset satellite orbit parameters so as to maintain the configuration of the low-orbit satellite constellation. By adopting the satellite constellation configuration keeping method, the problem of long-term keeping of the low-orbit constellation configuration can be solved.

The invention discloses an ultra-low orbit resistance coefficient determination method, device and equipment and a storage medium, and the method comprises the steps: firstly calculating the resistance coefficient of incoming flow atmosphere at a preset speed based on the obtained aerodynamic force of the incoming flow atmosphere forming a preset incident angle with a measurement plane; and calculating the resistance coefficients at different incoming flow speeds based on the expressions of the resistance coefficients at different incoming flow speeds and the resistance coefficients of the flowing atmosphere to the measurement plane at the preset speed. The resistance coefficient of the flowing atmosphere of the measuring plane at the preset speed is convenient and simple to measure, the resistance coefficient under the incoming flow speed condition which cannot be achieved by the test condition can be conveniently obtained by popularizing a formula, and the ultralow-orbit atmosphericresistance actually borne by the aircraft can be easily obtained.

An ultra-low orbit satellite orbit autonomous maintenance method is provided. The method comprises: 1, setting the working orbit range of a satellite, and estimating the magnitude of atmospheric resistance; 2, analyzing the magnitude of the noise of the inertial acceleration measurement system according to the magnitude of the atmospheric resistance to obtain a noise analysis result of the inertial acceleration measurement system; 3, setting parameters of a small-thrust execution system according to the noise analysis result of the inertial acceleration measurement system, and performing on-orbit calibration on the inertial acceleration measurement system and the small-thrust execution system to obtain a calibrated inertial acceleration output result; and 4, setting an orbit control smallthrust output algorithm of the small thrust execution system according to the calibrated inertial acceleration output result.

The invention discloses a water drop diameter measuring method. The water drop diameter measuring method is as follows: obtaining the maximum length L and the maximum width W of the projection of a water drop to be measured on a light intensity sensing array according to the refraction and attenuation of light caused by water drops; obtaining an array element with the maximum light intensity attenuation in the projection of the water drop to be measured on the light intensity sensing array; and obtaining the thickness H of the water drop to be measured according to the thickness calculating formula of the water drop. As the water drop suffers the atmospheric drag in air, the water drop becomes ellipsoidal, and then the volume of the ellipsoidal water drop to be measured can be obtained according to the L, the W and the H, the volume of the ellipsoidal water drop to be measured can be converted into the volume of a sphere, and then the diameter of the water drop to be measured can be obtained. The water drop diameter measuring method has the advantages that the simplicity is achieved, the model of the water drop is more similar to various water drops in real life, and the measuringis more accurate. The method can be applied to the fields of observing the influence on freezing caused by the water drop, observing the influence on growth of crops caused by the size water rainwater, and the like, so that more accurate data can be provided for the research of the fields.

The invention relates to a method for quickly estimating thrust of an aircraft on a real-time basis. Without relying on the information of aircraft-loaded sensor equipment and only relying on an external measurement data processing system, the parts that can be accurately modeled such as the gravity and the atmospheric drag of the aircraft are described with the physical model by constructing a hybrid dynamic model of the aircraft, the thrust part that cannot be accurately modeled is described with a mathematical model, and the online real-time estimation of the thrust of the aircraft is realized by the unscented filtering (UKF) estimation algorithm, which provides the test identification and accuracy evaluation of the aircraft with very important support.

A method for re-entry prediction of an uncontrolled artificial space object, the method including: calculating an average semi-major axis and an argument of latitude by inputting two-line elements or osculating elements of an artificial space object at two different time points; calculating an average semi-major axis, argument of latitude, and atmospheric drag at a second time point; estimating an optimum drag scale factor while changing the drag scale factor; predicting the time and place of re-entry of an artificial space object into the atmosphere by applying the estimated drag scale factor. Here, orbit prediction is performed by using a Cowell's high-precision orbital propagator using numerical integration from the second time point to a re-entry time point.

A method for obtaining the residual orbital lifetime of a space object in a low earth orbit include obtaining a semi-major axis change rate, an eccentricity change rate, an orbital element number at acurrent time and an atmospheric area mass ratio of the space object according to an atmospheric drag action; using a numerical integration algorithm to numerically integrate the rate of change of thesemi-major axis and the rate of change of the eccentricity with a predetermined integration step size, and obtaining the atmospheric density, and further integrating to the nth predetermined integration step from the current time to obtain the semi-major axis and the eccentricity of the orbit of the space object when the nth predetermined integration step is taken, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate; calculating the product of the semi-major axis and the eccentricity and determining the difference between thesemi-major axis and the product to obtain the difference between the perigee altitude of the space object and the radius of the earth; when the perigee height is less than the preset perigee height, obtaining the difference between the time corresponding to the nth preset integration step and the initial time for the remaining orbital lifetime of the space object.

The present invention relates to an anti-interference filtering method of a drag-free satellite attitude channel. Aiming at the drag-free satellite attitude channel of multi-source interference consisting of an atmospheric drag torque, an solar radiation pressure torque, a gravity gradient torque, inertia uncertainty, actuating mechanism noise, measurement noise, model uncertainty and the like, the method comprises: firstly, building a mathematical model including the drag-free satellite attitude channel of multi-source interference; secondly, building a mathematical model of an external environment interference torque; thirdly, designing an anti-interference filter aiming at the drag-free satellite attitude channel; and finally, designing an optimal controller to perform feedback control. The anti-interference filtering method of a drag-free satellite attitude channel is high in anti-interference capability, is high in control precision and the like, and is able to be used in the drag-free satellite attitude channel.

The invention discloses a method for extracting atmospheric density based on a SWARM-C satellite. Accelerometer data and GPS orbit data of the SAWRM-C satellite are obtained and then subjected to resampling and coordinate conversion; then a peak signal and a step signal in the accelerometer data are then removed; the accelerometer data are calibrated after data check is finished, deviation coefficient, proportional coefficient, and temperature coefficient in a calibration formula are estimated by a generalized least square method to obtain calibrated accelerometer data; then illuminating radiation pressure of a satellite received on an operating orbit of the satellite is subjected to modeling, and the acceleration caused by the illuminating radiation pressure is calculated, and the calibrated accelerometer data subtracts acceleration caused by the illuminating radiation pressure to obtain atmospheric drag acceleration; and in the end a product of damping coefficient of satellite operating in the orbit of the satellite and the effective area is calculated, and atmospheric density near the satellite orbit is calculated based on the atmospheric drag acceleration and the product of thedamping coefficient and the effective area.

The invention discloses an on-orbit calibration method for micro electric propulsion thrust. The on-orbit calibration method comprises the steps: calculating change of a semi-major axis in the rail ascending and descending processes; theoretically analyzing the influence of earth shape perturbation and atmospheric resistance perturbation in the orbit ascending and descending processes; obtaining the thrust acceleration by combining a semi-major axis variation formula obtained through calculation in the rail lifting process and the rail descending process; and obtaining electric thrust according to the thrust acceleration and the satellite mass. According to the scheme, the thrust does not need to be obtained through posture or angular velocity change conversion, that is, a thruster does not need to be eccentrically installed to generate eccentric torque to change the posture or the angular velocity, and the influence of earth shape perturbation and atmospheric resistance perturbation which are close to the magnitude of the thrust of a micro thruster on thrust calibration is eliminated, and on-orbit thrust calibration of the single thruster mounted through the mass center can be realized.

The invention provides an ultra-low orbit satellite orbit autonomous accurate maintenance method, which adopts a Kalman filtering algorithm to realize accurate acquisition of an orbit flat semi-major axis, and adopts an average method to eliminate a flat semi-major axis fluctuation item caused by earth high-order perturbation to acquire a meter-scale precision flat semi-major axis; under the condition that atmospheric resistance acceleration measurement is carried out without an accelerometer, the atmospheric resistance is determined through the change of the flat and semi-major axis, and meanwhile, the orbit control thrust is corrected and compensated in real time according to the determined atmospheric resistance, so that the problems of accurate acquisition of the meter-scale precision flat and semi-major axis of the ultra-low orbit satellite and autonomous and accurate orbit maintenance control of real-time correction of the orbit control thrust are solved.

The invention relates to the protection of spacecraft from debris and to de-orbiting devices of the atmospheric drag type, and to debris sweeping apparatus for the removal of debris from the space environment. The debris shielding apparatus for a spacecraft has a shield unit including a shielding surface for impeding incident debris. The shield unit is attached to the spacecraft body and has a drive mechanism for positioning the shield unit in relation to the spacecraft body. The drive mechanism is capable of moving the shield unit between a stowed first position and a deployed second position. In the deployed second position the plane of the shielding surface is at an angle to the spacecraft body.