212 results about "Solar sail" patented technology

A solar powered aerial vehicle includes an elongated airframe incorporating lifting and control surfaces, a mechanism for propelling the airframe through the air such that lift developed by the lifting surface is equal to or greater than the weight of the aerial vehicle, a planar solar sail coupled to the airframe and having at least one surface adapted to collect solar energy during the day and to power the propelling mechanism with a first portion of the energy collected, and an apparatus such as a fuel cell/electrolyzer for storing a second portion of the solar energy collected by the solar sail during the day and for powering the propelling mechanism with the second portion of energy during the night. The vehicle is capable of continuous operation at northern latitudes and during the winter months for extended periods without landing or refueling.

The invention discloses a method for designing a bias orbit of a spacecraft in a halo orbit in a sun-earth three-body gravitational field. The invention proposes a semi-analytic-solution-based orbit design method for a task demand of thrusting the spacecraft in the halo orbit to a heliocentric grave orbit by adopting a small thrust force in the sun-earth three-body gravitational field. According to the method, the calculation amount can be greatly reduced while the orbit precision is ensured. In addition, the invention proposes a method for realizing orbit maneuvering by adopting a minute continuous thrust system which mixes a variable reflectivity solar sail with electric propulsion. The method for realizing the orbit maneuvering by a mixed thrust force is capable of greatly reducing fuel consumption. The purpose of orbit maneuvering with minimum fuel consumption is achieved through optimization design of a posture angle of the solar sail. Compared with a conventional pulse bias orbit design method, the method proposed by the invention has the advantages that the calculation time can be shortened to a great extent on the premise of ensuring the solving precision.

The invention discloses a method for rapidly designing and optimizing a low-thrust transfer orbit. The method comprises the steps as follows: an inverse sextic polynomial method is introduced as an initial design, a high-performance optimal parameter initial value is provided for follow-up accurate optimization, and the calculating time of an entire optimizing process is reduced; a Radau legendre pseudospectral method is adopted to parameterize a low-thrust transfer orbit optimization problem into a nonlinear programming problem, not only is the number of the optimization parameters reduced, but also the accuracy of an optimization result is improved; and tedious first-order optimal necessary conditions are not needed to be deduced, even though the low-thrust transfer orbit containing various complicated constrains are processed, the operation is also convenient and simple. The method for rapidly designing and optimizing the low-thrust transfer orbit has an extremely important application value in the field of the design and optimization of transfer orbits of low-thrust propulsion spacecraft such as solar sail propulsion, solar energy electric propulsion, and the like.

The invention relates to the field of spaceflight and discloses a controllably and orderly inflated self-supporting type solar sail structure. A sail surface approximately triangle-shaped and formed by three curves is adopted, the sail surface can be unfolded in orbit stably and orderly, and distribution gradient of internal plane stress forced on the sail surface after unfolding is reduced by 20%; a support ring with the inner diameter equal to a winding shaft is arranged at the position of rectangular vertice of the sail surface, and concentrated stress of the sail surface can be lowered; a circular opening enclosed by 4n (n is larger than or equal to 3, and n is an integer) sawteeth, and no relative movement between the vertice of the sail surface and the winding shaft can be guaranteed; four sets of inflated extension arms used for supporting the sail surface are unfolded, each set is formed by two extension arms, straightness of the inflated extension arms and rigidity of the inflated extension arms after being unfolded can be improved, the rigidity of the structure is two times of that of a structure with a single extension arm, and unfolding stability of the extension arms is improved.

The invention discloses a deep space solar sail spacecraft driven by utilizing sunlight pressure. The deep space solar sail spacecraft driven by utilizing the sunlight pressure comprises a support frame, four support arms, four isosceles right triangle sail covers, a first solar cell, an etching antenna, two sliders and four roll axis stabilization mechanisms. The deep space solar sail spacecraft disclosed by the invention has the advantages that an open support packaging structure and a gird are adopted, and an outer packaging structure of a traditional satellite is removed, so that the mass of the solar sail spacecraft is reduced; inflation systems and air storage structures in traditional inflatable deployment support arms can be removed due to the fact that the self-deployment support arms are adopted, so that the mass of the solar sail spacecraft is reduced; by applying the wireless communication technology, the use of a communication line is reduced, so that communication and data transmission can be performed respectively by all parts of the large-area solar sail spacecraft, and the mass of the solar sail spacecraft is reduced; the sliders and the roll axis stabilization mechanisms are used for attitude control, and thus, the triaxial attitude control of the solar sail spacecraft can be realized; compared with spin stabilization attitude control, the triaxial attitude control can be used for executing broader tasks.

Fuel less ACS for solar sails, using furl- and unfurl able ballast-sail-foil-segments for simultaneously displacement of center of mass and center of solar radiation pressure into opposite directions to each other.

Solar-Sail-Launch-System for direct launch of the System-Sail including already docked in daughter units and payload in the sailcraft's central docking station.

Unlike todays launch able solar sail designs the System-Sail features ample solar cell arrays and additional SEP-thruster-units for steering and propulsion, while the solar sail rather serves for longtime fuel less attitude controls and station keeping.

The SEP-Sailcraft may also serve as a carrier-ship for daughter-units in asteroid exploration missions and is able to deliver prospector landers back to LEO with furled in foils.

Disclosed is a solar sail structure finite element modeling and post-processing method. The solar sail structure finite element modeling and post-processing method includes: 1), setting a solar sail configuration mode and a connecting format of a support arm and a sail surface and reserving an area for the support arm; 2), establishing a solar-sail geometric model with the preserved area; 3), setting a support arm model and generating a support-arm geometric model in the reserved area; 4), establishing finite element models of the solar sail surface and the support arm; 5), subjecting the solar sail surface and the support arm to finite element mode assembly according to the generated infinite element models of the solar sail surface and the support arm; 6), setting statics boundary conditions and loading conditions on the finite element models; 7), calculating buckling load of the support arm, simulating reverse acceleration speed released by the inertia, and exerting the reverse acceleration speed on the infinite element models; 8), analyzing statics of a solar sail to acquire maximum deformation and axial pressure of the support arm; 9), calculating buckling safety factors and structuring feasibility of a solar sail structural scheme according to the buckling safety factors.

The invention discloses a coiling type telescopic arm capable of achieving segmental, sequential and progressive expanding. The coiling type telescopic arm comprises a locking and releasing mechanism, a telescopic arm truss, a collection barrel and a mechanical and electrical transmission mechanism, and the telescopic arm truss, the collection barrel and the mechanical and electrical transmission mechanism are connected in sequence. The telescopic arm truss is fixedly mounted in the collection barrel and arranged to be capable of being coiled and drawn back. The collection barrel is arranged to be capable of storing the telescopic arm truss which is coiled and drawn back. The locking and releasing mechanism is arranged around the barrel wall of the collection barrel and used for pressing and locking the expanded telescopic arm truss which is drawn back and expanding and releasing the telescopic arm truss in the rebounding and expanding process. The mechanical and electrical transmission mechanism is used for driving the telescopic arm truss and the locking and releasing mechanism. The telescopic arm is simple in structure, small in expanding impact and reliable, the top end of the arm is kept directional, the overall mass is low, rigidity is large, and the telescopic arm can be used for space-expandable antennas, solar cell arrays, solar sails, space platforms, detecting arm mechanisms and the like.

The invention discloses a supporting packaging structure oriented to a solar sail adopting an inflatable auxiliary self-rebounding supporting arm. The structure comprises a center body (1), the inflatable supporting arm (2), a supporting arm flange (3), a guide rod (5), a supporting arm guide structure (6), a sail surface (7), a sail surface rolling shaft (8), a supporting structure (9), a sail surface rolling shaft rotating bearing (10), a connecting mechanism (11), an external packaging box body side plate (12), an external packaging box body upper plate (13), an external packaging box body bottom plate (14), an external packaging upper plate unlocking mechanism (15) and an external packaging side plate unfolding mechanism (16). A solar sail spacecraft adopting the inflatable auxiliary self-rebounding supporting arm is specifically designed, and the inflatable auxiliary self-rebounding supporting arm and the flexible sail surface are folded and supported simultaneously. Due to the adoption of the supporting packaging structure, the solar sail spacecraft can be stably and sequentially unfolded by being stressed in only one direction.

An apparatus includes at least one boom including a tip and a foot and being designed and arranged to be rolled up to form a roll. The tip is located in a radial inward region of the roll and the foot being located in a radial outward region of the roll. The apparatus further includes a unit being designed and arranged to unfold the at least one boom and including a bearing being designed and arranged to cooperate with the foot in a way that the boom in the region of the bearing at the beginning of the unfolding process has an enlarged cross section. Especially, the apparatus may be used in space technology for unfolding a solar sail.

The invention provides a sliding block executing mechanism used for attitude control over a solar sail spacecraft. The sliding block executing mechanism comprises a central box, a pre-tightening regulating unit, a control unit, a driving wheel unit, a driven wheel unit and an adjustable auxiliary driven wheel unit; a telescopic arm penetrates through a hole in the center of a front panel and a hole in the center of a rear panel of the central box; the pre-tightening regulating unit comprises a pre-tightening screw and a pre-tightening nut, and the pre-tightening screw penetrates through an upper panel of the central box; the control motor is driven by obtaining electric energy from a solar battery; the driving wheel unit comprises two protruding idler wheels which are connected with a motor shaft of the control motor through connection shafts, and the outer walls of the two protruding idler wheels make tight contact with the periphery of a stretching arm; two protruding idler wheels of the driven wheel unit penetrates through a rotary shaft and are arranged between two polytetrafluoroethylene sliding rings, the outer walls of the two protruding idler wheels make tight contact with the periphery of the stretching arm, and the rotary shaft is fixed to a support of a lower panel of the central box; the adjustable auxiliary driven wheel unit comprises four protruding idler wheels and four triangular supports, and the triangular supports are regulated so that the outer walls of the protruding idler wheels can make tight contact with the periphery of the stretching arm.

The invention relates to an asteroid orbital transfer method. According to the asteroid orbital transfer method, an asteroid orbit is changed through a tether formation flight system composed of a solar sail and an asteroid. The asteroid orbital transfer method comprises the steps that the solar sail backwards and horizontally throws out a tether lock connected with a tether, the solar sail captures the tether lock after circling the asteroid for a circle, and is in butt joint with the tether lock, and then the tether is retracted so that the tether can be bound to the surface of the asteroid; the solar sail is thrown off through autorotation of the asteroid; the light pressure borne by the solar sail is adjusted by adjusting the posture of the solar sail, acting force is transmitted to the asteroid through the tether, and the autorotation speed of the asteroid is gradually reduced through the acting force; after the autorotation speed of the asteroid is reduced, the asteroid is transferred to a low-energy transfer orbit through adjustment of the posture of the solar sail. The asteroid orbital transfer method has the advantages that the threat caused by the asteroid to the earth is removed, the structure and the ingredients of the asteroid cannot be damaged, long-term orbital transfer can be carried out, and therefore the asteroid can be controlled to serve as a transfer station for interplanetary exploration of the human.

This invention relates generally to method and apparatus for solar-boost assist and solar-sail assist by a mini-optics light concentrator system utilizing a dynamic ensemble of mini-mirrors. In the boost phase of a rocket ship launch, the system focusses solar energy into the rocket chamber to raise the temperature of the propellant and increase the impulse. In the mid-course phase, the system focusses solar energy onto the rocket's solar sail to increase the thrust. In both cases, the system reduces the weight of the rocket by providing a weightless source of thrust.

The invention provides a spacecraft heliocentric ellipse suspension track design method based on the solar sail propulsion technology. The method comprises the steps that the characteristics of track suspension height are analyzed, and the suspension height is decomposed into the suspension height with a constant value, and the suspension height changing in direct proportion to the sun-planet distance; on the basis, a track dynamic model of a solar sail spacecraft under a rotation inconsistent fit coordinate system is established, conditions for forming a heliocentric ellipse suspension track are given according to the balance of a two-level dynamic model, the balance conditions are met by controlling posture changes of a solar sail and changes of reflectivity, and then the heliocentric ellipse suspension track of the solar sail is obtained. The spacecraft heliocentric ellipse suspension track design method based on the solar sail propulsion technology provides a worthy reference general method for heliocentric ellipse suspension track design based on the solar sail propulsion technology.

The invention relates to a test system for impact analysis of imaging quality of TDICCD (Time Delayed Integration Charge Coupled Device) by multi-source interference, which can be divided into an error source interference and an equivalent sub system thereof, a TDICCD imaging sub system, an optical target, gesture error and an interference measuring sub system thereof, a signal collecting and processing sub system, and a wireless communication and ground comprehensive processing sub system. The error source interference and the equivalent sub system thereof is used to simulate interference generated by a solar sailboard/a control torque gyro system (CMGs) and a satellite platform in a spatial environment for the satellite. The measuring sub system is used to measure the gesture angle of the satellite platform and the vibrating linear velocity. The signal collecting and processing sub system collects sensor information and processes and analyzes and calculates the information correspondingly. According to the invention, main interference by motion errors of the satellite to the imaging quality of the TDICCD is simulated in the spatial environment on the ground, and the impact analysis of the imaging quality of the TDICCD by the multi-source interference is realized by combining a floating turntable for dynamical test and analysis of satellite gesture.

The invention provides a manufacturing method of a solar sail capable of being unfolded controllably and orderly based on shape memory polymers, and relates to the manufacturing method of the solar sail, in particular to the manufacturing method of the solar sail for a spacecraft. The invention aims to solve the problems that a driving mechanism is complex, an unfolding process is instable, a structure is easy to damage and air leakage occurs existing in the conventional solar sail with an inflatable type structure, and solve the problems that the structure is complex, the unfolding reliability is poor, the total weight is large, the folding efficiency is low and the emitting volume is large existing in the solar sail with a mechanical unfolded and supported structure. The manufacturing method disclosed by the invention comprises the following steps of firstly synthesizing shape memory polymeric materials into shape memory thin polymer films, and covering electrothermal films on the shape memory thin polymer films; and fixing flexible material light reflecting films on the shape memory thin polymer films covered with the electrothermal films, then energizing for heating the electrothermal films until the shape memory thin polymer films are softened, orderly folding the sail surface of the solar sail, and after power cut, cooling the electrothermal films so as to obtain the solar sail which is folded. The manufacturing method disclosed by the invention is suitable for manufacturing the solar sail.

The invention discloses a method for transferring between planetary circular displaced orbits of a solar sail spacecraft. According to the method, firstly dynamic models of a planetary circular displaced orbit and an elliptical displaced orbit are analyzed, and a solution method of a solar sail attitude parameter and a reflectivity parameter to achieve the two orbits is given; meanwhile, a boundary condition to conduct splicing and transferring between original and targeted planetary circular displaced orbits and the planetary elliptical displaced orbits is analyzed; and then, according to mission requirements, the orbit parameters of the original circular displaced orbit and a final circular displaced orbit to conduct transferring on the spacecraft are determined, the planetary elliptical displaced orbit used for orbit transferring is solved according to the boundary condition, and further the solar sail attitude parameter and the reflectivity parameter are worked out by combining the dynamic balance point of the planetary elliptical displaced orbit with a sunlight pressure model. The method can give an appropriate planetary elliptical displaced orbit aiming at the problem of transferring between the planetary circular displaced orbits.

The invention provides a plane storage ratio folding and inflation unfolding structure for a spacecraft. The plane storage ratio folding and inflation unfolding structure comprises a plurality of folding units; each folding unit comprises a main folding line (2), an auxiliary folding line (3) and a plurality of triangular folding lines (4); the multiple triangular folding lines (4) are located onthe side parts of the main folding lines (2) correspondingly; and the auxiliary folding lines (3) are located on one sides of the triangular folding lines (4). Based on the simplicity and practicability of a product, the high-rigidity large plane high-storage-ratio folding and inflation unfolding structure for the spacecraft is provided. According to the plane storage ratio folding and inflation unfolding structure for the spacecraft, through application of a simple folding line folding method or structure and thermocuring inflation supporting pipes, high-storage-ratio folding close-up of large planes, such as a solar sail, an antenna and a reflector, in the spacecraft can be realized, the unfolded planes have high rigidity and the high shape-preserving ability, and the high rigidity of the thermocuring inflation supporting pipes do not need to be maintained through the pressure difference in the pipe (that is, after the planes are unfolded in place, gas can be exhausted at any time, and the rigidity is not affected by gas leakage).

The invention discloses a solar panel unfolding system based on a hot knife, and belongs to the technical field of spacecraft solar panel unfolding and locking. According to the invention, the problems of large separation impact and large distribution forbidding area required by the solar wing in the existing compression release and expansion locking mode of the micro-nano satellite solar panel are solved. Each sailboard is connected with the satellite body through a micro hinge, a pressing embedded part is embedded in each sailboard, the hot knife structure is installed on the satellite body,the two pressing embedded parts are connected through a strong horse wire, and the unfolding of the sailboards is realized by cutting the strong horse wire through the hot knife structure. At least one supporting base is arranged between each sailboard and the satellite body, and each supporting base is fixedly connected with the satellite body. The system is used for compressing, releasing, unfolding and locking the micro-nano satellite solar panel. The system has the characteristics of low cost, simple structure, easiness in assembly, light weight, small impact, high piece distribution rate, easiness in standardization and modularization and the like.

The invention provides a method for driving a solar wing to face the sun through a low-inclination-angle orbit uniaxial SADA. The method comprises the following steps: S1, determining coordinates of asolar vector in an orbit system by an on-board sensor; S2, calculating a satellite yaw angle according to the projection of the sun vector on the XoOYo plane of the orbit; S3, calculating expected yaw angle and angular velocity information in satellite attitude motion; S4, calculating the rotation angle of a solar wing driving device according to an included angle between the normal coordinate ofthe solar panel and the solar vector when the solar wing driving device is in a zero position; and S5, controlling the satellite to maneuver by the on-board actuating mechanism, and rotating the solar wing driving device in a matching mode to ensure that the solar wing points to the sun. The invention further provides a system for driving the solar wing to face the sun through the low-inclination-angle orbit single-shaft SADA. The method has the advantages that the single-axis SADA is used for replacing the double-axis SADA in the low-inclination-angle orbit to achieve sun orientation of thesolar wing, cost is reduced, reliability is improved, and control is easy.

The invention relates to a solar sailing vehicle, which comprises a vehicle body, a sail, a steering system, a sail steering system, a solar panel, a sensor and a control system, wherein the sail, the steering system, the sail steering system, the solar panel, the sensor and the control system are arranged on the vehicle body. The solar sail vehicle is characterized in that the sail is arranged on the vehicle body, the steering system comprises a steering engine, a parallelogram steering frame and four wheels and is arranged at the lower part of the vehicle body, the sail transmission system comprises an inclined pulling rope, a motor and a shaft coupler, and is connected with the sail, the solar panel covers the surface of the sail and supplies electricity to the motor, the steering engine and the control system, the sensor is arranged on the top end of the sail and is used for detecting the wind direction and the wind speed, the control system is connected with a steering engine controller, a motor driver and the sensor, the solar sail vehicle is driven to advance by wind energy, the motor controls the rotation of the sail to obtain the optimum power, and the steering engine controls the advancing direction and the angle of the vehicle body, so that the sail vehicle can move according to a preset track.

The invention discloses a modular solar sail assembled and reconfigurable space-based debris removal platform. The modular solar sail assembled and reconfigurable space-based debris removal platform can change the status of a single-scale debris removal mission, and cover a plurality of removing missions such as large-scale space debris and abandoned satellites over ten centimeters in a high orbit, small satellites such as low-orbit failure cubic satellites and a large number of low-orbit millimeter-class micro-space debris through flexible on-orbit reconfiguration. According to specific debris removal tasks, the required sail area is determined, the number of cell modules is determined by the area, a certain number of the cell modules are released from the platform for on-orbit connectingreconfiguration and perform a corresponding task after on-orbit unfolding.