60 results about "Launch window" patented technology

A reusable space vehicle docked in an intermediate orbit for rescue missions allows a satellite to be serviced with less delay, energy expenditure, and cost than a space vehicle launched from Earth for each mission. The reusable repair vehicle can be moved from one orbit to another with minimal energy expense while not having to wait for a launch window. Once a servicing need is identified, a destination orbit is identified for the space vehicle and a minimum energy path is identified. If the time to the next launch window between the docking orbit and the destination orbit happens to be near enough to allow for a timely rendezvous, the space vehicle is moved directly to the destination orbit. The space vehicle can be a vehicle designed to be piloted by humans or telerobotically. In one implementation, the inactive space vehicle is docked in an Intermediate LEO orbit (altitudes of approximately 250 km to approximately 500 km) and is used to rendezvous with objects in High LEO orbits (altitudes of approximately 500 km to approximately 1500 km) or objects in Low LEO orbits (altitudes of approximately 250 km or less). The space vehicle can be a modified lunar lander.

The invention discloses a flying tongue capture mechanism and a space target acquisition method. The mechanism comprises a launch driving mechanism (1), a flying tongue capture mechanism (2), a flexible cable mechanism (3), an unlocking release mechanism (4), a wheel rotation reloading mechanism (5), an aiming and correcting system (6) and an aircraft platform mounting plate (7); the whole mechanism is in rigid connection with an aircraft body through the aircraft platform mounting plate (7); after the aiming and correcting system (6) determines a target to be captured, the launch driving mechanism (1) pushes the flying tongue capture mechanism (2) to drive the flexible cable mechanism (3) to eject from a flying tongue launch window and the flying tongue capture mechanism (2) captures the target; when the target is required to be abandoned, the flexible cable mechanism (3) is cut off by the unlocking release mechanism (4), the output shaft of a driving motor (53) is connected with a central shaft of a rotary inner wheel ring (52), and the driving motor (53) drives the rotary inner wheel ring (52) to rotate and further drives the next flying tongue capture mechanism (2) to be positioned at the flying tongue launch window and perform the next capture.

The invention relates to a method for improving the orbit-transferring safety of an inclined orbit satellite, wherein the problem of energy-source shortage of the inclined orbit satellite in the period of orbit-transferring motorization under the condition of high solar angles is solved through the reasonable adjustment of flight events by considering the condition of solar altitude angles in the designing process of a flight program, and meanwhile, flight-control operation can be simplified. The discharging time of a storage battery in the period of orbit-transferring ignition can be shortened by designing the flight program according to the method, so that the load power in the power-supplying period of the storage battery is reduced, therefore the discharging depth of the storage battery is reduced, the energy-source safety of the satellite is ensured, and a launch window is not limited by solar illumination angles.

The invention discloses an analysis method and analysis device for warning spacecraft launching collision risks. The analysis method includes the steps that S1, all intersection events of a spacecraft and a background object during launching are calculated; S2, for each intersection event, the possibility of collisions between the spacecraft and the background object is calculated; S3, the total possibility of collisions between the spacecraft and the background object is calculated. According to the analysis method and analysis device for warning the spacecraft launching collision risks, the collision risks which a certain launching window is going to encounter are evaluated through the total possibility of collisions, the multiple intersection events which the launching window is going to encounter at different intersection moments are fully taken into account, and the defect caused by one single intersection risk is overcome.

The invention relates to a method for returning payloads in a space station with assistance of an unpowered cable, in particular to the method for returning the payloads in the space station. The method provided by the invention is used for solving the problems that requirements for effectively returning the payloads in the space station in time with the low cost cannot be met because the returning cost of the payloads is high, and the high returning cost is caused by that the traditional technology for returning the payload in the space station is subjected to the limitation of a launch window. The method comprises the following main steps of: setting target parameters tn and L of a de-orbiting point; building an expected optimal expansion orbit of the cable; controlling current expansion status parameters of the cable to coincide with excepted values; controlling a tension; controlling a motion orbit of a returning cabin; judging whether the returning cabin arrives at a pre-set position; and expanding, releasing and finishing. The method can be widely applied to returning the payloads in the space station or a low-orbit space-based platform.

The invention provides an estimation method of an output current of a solar cell array at a satellite orbit-injection stage. The estimation method is used for realizing prediction and estimation of the output current of the solar cell array during the sun positioning process from satellite and rocket separation to satellite attitude after the launching of a satellite. In the process, the estimation method comprises a satellite and rocket separation to solar cell array unfolding mode, a solar cell array unfolding hold mode, a solar searching mode and a solar cell array return-to-zero mode. In each mode, an incident angle and illuminated area of the solar cell array change continuously along with the change of whole-satellite attitude, and the output current of the solar cell array in each mode is estimated. The estimation method introduces a panel coordinate system to solve the difficult that the output current of the + / -Y solar cell array cannot be calculated independently in the solar cell array return-to-zero mode, improves the calculation precision of the output current of the solar cell array, and provides an important reference for selecting a reasonable launching window of the satellite and improving an in-orbit flight control plan.

The invention relates to a storing-taking express cabinet for unmanned aerial vehicle delivery and a working method thereof. The storing-taking express cabinet comprises a cabinet body and a solar cell panel. Vertical electric pushing rods are fixedly connected to the four corners of the top face of the cabinet body correspondingly, the upper ends of the electric pushing rods are hinged to the solar cell panel, and a launching window facilitated passing of an unmanned aerial vehicle is formed in the front end of the solar cell panel. A sorting car for receiving express is arranged at the frontend of the top face of the cabinet body, and an ejector plate for ejecting falling of the express is arranged on the top face of the sorting car. A plurality of vertical sliding rails for receiving the express are formed in the middle of the top face of the cabinet body in the horizontal direction, a plurality of cabinet grids distributed in a matrix shape are arranged on the front and rear sidefaces of the cabinet body correspondingly, the number of columns of the cabinet grids are the same as the number of the vertical sliding rails, and the positions of the cabinet grids correspond to thepositions of the sliding rails, the cabinet grids located on the corresponding same columns communicate with the vertical sliding rails corresponding to the cabinet grids correspondingly, and receiving plates which can be shifted in the longitudinal direction and are used for receiving the express are arranged on the inner bottoms of the cabinet grids correspondingly. According to the storing-taking express cabinet for unmanned aerial vehicle delivery and the working method thereof, rapid sorting of unmanned aerial vehicle delivery can be realized, the structure is reasonable, the inventory amount is large, and the working efficiency is high.

The invention relates to the technical field of nuclear radiation measurement calibration, and provides a gamma ray collimator for a reference gamma radiation field. The gamma ray collimator comprises a collimating diaphragm, a light path of the collimating diaphragm is in the shape of a truncated cone, an incident window of the collimating diaphragm is flared, a launch window of the collimating diaphragm is extended based on the edge of the light path, a plurality of annular grooves are formed in the inner surface of the collimating diaphragm, the sections of the annular grooves are triangular, vertexes of all triangles in the same section are positioned on the same horizontal plane, bottom edges of all the triangles are positioned on the same edge of the light path, the triangular sections of the annular grooves are varied with energy of radioactive source gamma ray energy, and different energy dispersion gamma photons are scattered backwards. Scattered gamma photon reduction capacity of the reference gamma radiation field collimator can be improved, and uniformity and accuracy of the radiation field are improved.

The invention discloses a target asteroid selection method for manned asteroid exploration. The method comprises four steps as follows: Step 1: determining constraint conditions for target asteroid preliminary selection; Step 2: screening an asteroid data base according to the constraint conditions which are determined in Step 1 and used for target asteroid preliminary selection to obtain primarily-selected targets; Step 3: performing primarily-selected asteroid orbit design and related calculation, including design of task orbits, calculation of velocity increments, search of launch windows and calculation of task periods; Step 4: performing target asteroid optimization, including a target asteroid optimization model and solutions. The method involves a strict target asteroid preliminary selection process and a target asteroid optimization mathematical model, so that a target asteroid which has the project realizability and the exploration value can be provided.

The invention relates to a laser projectile simulation launching device and a laser projectile simulation launching control method. The launching device comprises a laser projectile launcher and a laser projectile projecting screen; a laser launching window is formed in a front panel of the laser projectile launcher; the laser projectile projecting screen is arranged just oppositely to the laser launching window; an insulation board is arranged on the bottom surface inside the laser projectile launcher; an optical element backing plate is arranged on the insulation board; an X-Y laser high-speed scanning galvanometer, a diffraction grating and a laser launcher are sequentially arranged on the optical element backing plate; and the X-Y laser high-speed scanning galvanometer is opposite to the laser launching window. The laser projectile simulation launching device and the laser projectile simulation launching control method disclosed by the invention have the characteristics of economy, lightweight, stable structure, flexible operation and safety in use; and in addition, a simulated projectile has the advantages of wide ranges of firing rate and flying speed and continuous adjustability.

The invention discloses an internet of things terminal capable of automatic adjusting launch power by using an ultrasonic measured distance. The internet of things terminal capable of automatic adjusting the launch power by using the ultrasonic measured distance comprises a main control unit, a radio frequency unit and a memory unit, wherein the internet of things terminal capable of automatic adjusting the launch power by using the ultrasonic measured distance has an ultrasonic ranging program, a ultrasonic measured distance probe is installed in a housing on the back of the internet of things terminal, a launch window and a receiving window of the ultrasonic measured distance probe are exposed on the outer face of the housing on the back of the internet of things terminal, a radio frequency antenna is installed in the housing on the back of the internet of things terminal, and located below the ultrasonic measured distance probe, a main radiation direction of the radio frequency antenna is perpendicular to the surface of the back of the housing, and a launch direction of the ultrasonic measured distance probe is parallel to the main radiation direction of the radio frequency antenna. According to the internet of things terminal capable of automatic adjusting the launch power by using the ultrasonic measured distance, the ultrasonic measured distance probe is added on a prior hand type internet of things terminal, a whole machine can adjust the launch power automatically according to a target distance after matched with software, and labor intensity of operation personnel is reduced. And simultaneously, misreading to a distant target is avoided, and therefore usability of the whole machine and reliability of a system are increased.

The invention relates to an orbit accelerating airplane capable of completing lift-off of a two-stage rocket spacecraft with two-stage moving platforms, which comprises an orbit, an orbit moving platform, an airplane platform and the two-stage rocket spacecraft, and is characterized in that the orbit moving platform and the airplane platform complete start, acceleration and lift-off by means of different power sources and pass through a troposphere to enter a stratosphere, and a relaying launch means of sending the spacecraft by a rocket in the stratosphere is realized. Compared with the prior art, the orbit accelerating airplane has the notable advantages that fuels of a jet engine are saved, the rocket is small in size and low in cost, launch is less limited by a launch site, the weather and the climate, a launch window can be selected, the orbit moving platform and the airplane platform can bear various spacecrafts and is reusable and the like, thereby having quite good use value.

The invention provides a multi-orbit design method for carrier rocket fire-striving launching and a fire-striving launching control method. The fire-strike launching multi-orbit design method comprises the steps that task data of P Mars orbit entrances specified by a fire-strike launching task are obtained, the task data of each Mars orbit entrance comprises a launching window where a detector is located at the Mars orbit entrance, and P is a natural number larger than 2; an iterative optimization method is adopted, P fire-strike launching orbits where the carrier rocket carries the detector to fly are designed, and target points of the P fire-strike launching orbits are in one-to-one correspondence with the P Mars orbit entrances; and when the carrier rocket is controlled to fly according to any one of the P fire-striving launching orbits, the carrier rocket carries the detector to the Mars orbit entrance corresponding to the fire-striving launching orbit. According to the design method, the design efficiency is improved, the complexity of flight data is reduced, and reliable launching and accurate orbit injection of the large low-temperature carrier rocket can be realized.

The invention relates to a method for processing the interference in an infrared earth sensor from the moon through a medium and high orbiting satellite. The method comprises the following steps: calculating the orbit of the moon appearing in an earth sensor field when the earth is captured by the satellite according to the loci of the moon before or after the earth is captured, predicting the condition of the interference in all probes of the earth sensor from the moon and the attitude angle outputted by the earth sensor, remotely measuring the data in real time according to the altitude angle outputted by the earth sensor, judging whether the cause of an EP signal is the earth or the moon, choosing proper time to safely transfer the satellite to an earth pointing mode to avoid the interference from the moon in the earth sensing period. The method has the advantages that the choosing range of the satellite launching window is increased, and the operation safety of transferring the orbit of the satellite is improved.

The invention discloses an efficient data transmission method under a narrowband communication condition, and belongs to the field of wireless communication data transmission. The method comprises the following steps: issuing wireless channel resources and user information needed to participate in data transmission services by a base station; after receiving and parsing the information by a terminal, automatically calculating an allocated service channel and a launch window, automatically completing the migration to the service channel and performing window matching on time synchronization information issued by the base station broadcast in real time; and finishing data transmission on the corresponding launch window, and after the data transmission is finished, returning to a control channel by the terminal itself. The whole process of business execution does not need base station intervention, so that the frequent signaling interaction between the base station and the terminal is avoided, the consumption of air interface resources is reduced, and the air interface resources are sufficiently used for the transmission of valid data. Especially in the case of multi-terminal concurrent data transmission application scene, the speed and efficiency of data transmission are improved remarkably.

The invention relates to a launch window planning method for near-earth target interception, and belongs to the technical field of aerospace. The method is particularly suitable for calculation of a launch window for intercepting a low-mid-orbit near-earth target by an aircraft taking off on the ground. The method comprises the steps of firstly determining a total flight duration range of an interceptor from taking off on the ground to interception according to an orbit height of an intercepted target; secondly determining an outer boundary of a range which the interceptor can reach accordingto a maximum flight duration of the interceptor; thirdly determining a quasi launch window based on the outer boundary of the range which the interceptor can reach and the flight duration range; fourthly dividing the flight duration range, and solving an annular range which the interceptor can reach and corresponds to each flight duration point; and finally obtaining an accurate launch window according to a crossing relationship between a target sub-satellite point and each annular range.

The invention provides a launch task allocation method under a resource scheduling constraint, which can realize rapid allocation of launch tasks under the resource scheduling constraint. According to the method, task feasibility analysis, multi-point concurrent emission time window recheck analysis and the like are carried out on the basis of deduction models such as emission styles, emission platforms and emission rules, conflicts and potential defects in the aspects of time, space, capacity, interfaces, resources and the like possibly occurring in emission plans are avoided, and support is provided for perfecting the emission plans. Conflicts of emission tasks are eliminated in dimensions such as callable emission point positions and emission window sequences, consumption is reduced, and overall optimization of scheduling of various resources and coordination of emission time sequences is achieved.

A deferred placement prompt is provided in a user interface when an application window is launched. Responsive to the presentation of the deferred placement prompt, the user can provide a directional placement instruction (e.g., a “drag” of the prompt, a selection of a screen region, a keyboard input) to indicate the user's preference for placement of the launched window. The user interface can give cooperative feedback to the user during the deferred placement prompt operation, such as highlighting an application window to be replaced, responsive to a directional placement instruction, animating the deferred placement prompt, etc.

The invention provides a method for attitude control of an orbital return and reentry air vehicle over satellite communication, and relates to the field of measurement and control of in-orbit segmentsand relay satellites. The method comprises the following steps that step 1, an earth-fixed coordinate system o1x1y1z1 is established; according to the longitude L, latitude B and height H of the aircraft; the geocentric radius vector corresponding to the air vehicle is represented as red in the earth-fixed coordinate system o1x1y1z1; step 2, the semi-geocentric-angle delta corresponding to the earth area covered by a single relay satellite is calculated; step 3, the geocentric radius vectors corresponding to n relay satellites in the earth-fixed coordinate system o1x1y1z1 are represented as resi; step 4, the relay satellite corresponding to the air vehicle is selected from the n relay satellites; step 5, after the relay satellite is selected, the attitude of the air vehicle is adjusted toachieve that a measurement and control antenna on the air vehicle is aligned with the selected relay satellite and achieve bidirectional communication between the air vehicle and the relay satellite.The method for attitude control of the orbital return and reentry air vehicle over the satellite communication avoids the problems of repeated binding caused by the change of a launch window and thestar loss caused by the deviation of an actual flight trajectory, and ensures the continuous space-based measurement and control capability of the in-orbit segments.

The invention discloses a method for designing an earth-moon L2 point task launching window. The method comprises the steps of obtaining window design constraint conditions, tracks, task important parameters and the like according to engineering constraints, task requirements and track design conditions; simultaneously carrying out constraint condition-related characteristic analysis on the transfer section and the mission section orbit, and respectively screening out launching window date sets meeting task requirements; and finally, solving a union set of the two to generate a complete nominal flight path under a corresponding date for confirmation, thereby obtaining a final launching window design result. The method can meet the multi-constraint requirements of engineering tasks on the launching window.

An energy distribution method of a GEO satellite under an evening launch window during a transfer orbit relates to the field of geosynchronous orbit satellite energy distribution. The method comprises: step 1, switching on a communication cabin in the satellite to replace a heater; step 2, at reaching the moment t1 before the first orbit change of the satellite, raising the temperature value of the communication cabin replacing the heater to A2; step 3, at reaching the moment t2 before the first orbit change, reducing the temperature value to A3; step 4, raising the temperature value to A4 until the first orbit change of the satellite is completed; step 5, at reaching the moment t3 before the second orbit change of the satellite, raising the temperature value to A5; step 6, reducing the temperature value to A6 at reaching the moment t4 before the first orbit change; and step 7, raising the temperature value to A7 until the second orbit change of the satellite is completed. The method is applicable to the geosynchronous orbit satellite under the evening launch window, the satellite reasonably distributes the energy of the whole satellite during the transfer orbit, and the energy security for various missions of the satellite during the transfer orbit is ensured.

The invention discloses a method for generating a safe launching window of a weak-gravitation small celestial body surface lander, and belongs to the technical field of deep space exploration. The implementation method comprises the steps of: firstly, defining a landing point tangent normal coordinate system of a small celestial body, and deriving the analytical relation among the launching state,the landing state and the landing time of a lander near the small celestial body; secondly, establishing a launching constraint meeting the safety rising requirement of a main detector and a landingconstraint meeting the safety landing requirement of the lander on a tangent system, and defining a launching window; then, deducing a two-dimensional projection boundary analytic expression of a window of the landing speed constraint on a two-dimensional plane according to the landing dynamics analytic relation and the launch landing constraint; and finally, designing evaluation indexes representing landing safety to obtain a safe launching window with index distribution. The generated safe launching window can provide a basis for rapid selection of the launching state of the lander, and thelanding safety and the landing success rate are improved.

The invention relates to a method for selecting a net capture load launching window based on waste satellite attitude nutation characteristics, and belongs to the technical field of satellite net capture of waste satellites. The method comprises the steps of: 1, extracting a maximum nutation angle theta max, a minimum nutation angle theta min and a nutation period T; 2, acquiring a net capture load launching moment tint and a capture target moment tinf; calculating the time width Tnet according to the formula that Tnet=tinf-tint; 3, setting an adaptive nutation angle threshold theta net directly launched by a net capture load non-selection window; 4, calculating the time width T theta=t theta s-t theta j when the nutation angle theta is less than or equal to theta net in a single period; and 5, determining the satisfactory conditions of time width delta T of a net capturing load launching window, the launching moment tint and the nutation angle theta during launching; according to thewaste satellite attitude nutation rule, the net capture load nutation angle adaptive capacity and the time width of the net capturing load from the launching moment to capturing, the invention provides a launching window selection method, and reliable capturing of the large-attitude nutation waste satellite is achieved.

The invention discloses a calculation method for evaluating rendezvous between a carrier rocket and a space target before launching. The method comprises the following steps: in a given launch window time period, screening a calculation method of a safe launch window time meeting a condition, wherein the screening condition is that an intersection distance between a carrier rocket and all space targets is smaller than a given safety threshold. The invention provides a calculation method for carrying out rendezvous evaluation with a space target by utilizing a theoretical flight path of a carrier rocket before launching according to the theoretical flight path characteristics of the carrier rocket and the on-orbit operation rule of the space target. According to the method, the requirements of high timeliness and high precision of rendezvous evaluation of the carrier rocket and the space target before launching are effectively met.

The invention relates to the technical field of nuclear radiation measurement calibration, and provides a gamma ray collimator for a reference gamma radiation field. The gamma ray collimator comprises a collimating diaphragm, a light path of the collimating diaphragm is in the shape of a truncated cone, an incident window of the collimating diaphragm is flared, a launch window of the collimating diaphragm is extended based on the edge of the light path, a plurality of annular grooves are formed in the inner surface of the collimating diaphragm, the sections of the annular grooves are triangular, vertexes of all triangles in the same section are positioned on the same horizontal plane, bottom edges of all the triangles are positioned on the same edge of the light path, the triangular sections of the annular grooves are varied with energy of radioactive source gamma ray energy, and different energy dispersion gamma photons are scattered backwards. Scattered gamma photon reduction capacity of the reference gamma radiation field collimator can be improved, and uniformity and accuracy of the radiation field are improved.

The invention relates to a space station logistics supply planning method for on-orbit demand traction. The method comprises the steps of obtaining an on-orbit event execution plan, and obtaining a daily material consumption list according to the on-orbit event execution plan; according to the space station orbit dynamics model, calculating the occurrence time of orbit maintenance maneuver in the planning period and the consumption amount of a propellant for orbit maintenance; according to the on-orbit material demand, the preset space station on-orbit initial inventory, the preset space station on-orbit inventory capacity upper limit and the preset space station on-orbit inventory minimum warning value, determining the earliest supply time, the latest supply time and the cargo list of each cargo ship; and according to the earliest supply time and the latest supply time of each cargo ship, determining a cargo ship launching scheme by combining actual launching window constraints. The space station logistics supply planning method takes the on-orbit demand as the basis of material supply planning, quickly obtains the space station logistics supply planning scheme at a small calculation cost in a very short time, and has the advantages of high calculation efficiency, strong scheme pertinence, capability of simultaneously meeting multiple constraints, good practical engineering applicability and the like.

The invention discloses a laser target spot calibration device of a laser obstacle clearing system. The laser target spot calibration device comprises a distance measuring machine and a calibration support, wherein the distance measuring machine is mounted on the calibration support; the calibration support is mounted on the laser obstacle clearing system; the distance measuring machine comprisesa shell as well as a distance measuring laser transmitting module, a distance measuring laser receiving module, a computing display module and a power module arranged in the shell; the distance measuring laser transmitting module is provided with a distance measuring light launching window; the distance measuring laser receiving module is provided with a distance measuring light receiving window;the computing display module is provided with a display window; and a bull's-eye indicating area is arranged in the center of the display window. With adoption of the laser target spot calibration device, all that is needed is to adjust the obstacle position in the bull's-eye indicating area, that is the actual position of the obstacle coincides with the laser target spot of a laser, at the moment, the focus point of the laser is precisely adjusted to the actual position of the obstacle, so that the calibration precision of the laser target spot is improved.