443 results about "Satellite bus" patented technology

The invention belongs to the satellite navigation and positioning technical field and discloses a low earth orbit satellite-based satellite-earth differential real-time precise positioning method. A low earth orbit satellite is utilized to broadcast the observation data and real-time orbit data of the satellite-borne GNSS (Global Navigation Satellite System) receiver of the low earth orbit satellite to the ground; and after receiving the differential information broadcasted by the low earth orbit satellite, a ground receiver generates a double-difference observation value consisting of the differential information and a local GNSS observation value and performs pseudorange-based moving base station DGNSS (Differential Navigation Satellite System) positioning and carrier phase-based moving base station RTK (Rea-time kinematic) positioning. According to the positioning method of the invention, the global mobile low earth orbit satellite platform is adopted as a reference station, so that real-time precision differential positioning service in the whole world can be realized, and the method does not depend on the distribution of ground reference stations; and a user can realize differential real-time precise positioning just through using a single receiver, and therefore, the method is free of operating range restrictions, and data communication links are not required to be considered.

The present invention provides an emulation system having a control system that allows the testing of a satellite control system with all of its hardware in place, i.e. fully integrated. The emulation methodology is applicable to the case of either a rigid spacecraft or a flexible spacecraft, provided that the spacecraft's sensors and actuators are stowed to the rigid part of spacecraft in the case of a flexible spacecraft. Practically, the latter condition is not restrictive, as the actuators and sensors are usually placed rigidly in the satellite bus, while the satellite solar panels constitute the flexible elements. The control system is used to tune the mass properties and dynamic behaviour of a rigid ground-spacecraft in a 1-G environment to those of a flight-spacecraft in 0-G. A six-axis force / moment sensor is placed at an interface of the ground-spacecraft and a manipulator. Signals received from the force / moment sensor, and in some cases signals relating to the position and velocity of manipulator joints, are received into the control system.

The invention belongs to the field of mechanical vibration isolation and particularly relates to a multiple-degree-of-freedom vibration isolator and a multiple-degree-of-freedom vibration isolating system for an effective load and a satellite. The multiple-degree-of-freedom vibration isolator provided by the invention consists of a casing, a motion guide rod, a metal wire network material member,flexible hinges, a fine tuning spring, an upper casing end cover and a lower casing end cover and has the advantages of compact structure, reliable performance, long service life and the like. The vibration isolating system consisting of the vibration isolator arranged according to a certain geometry configuration is arranged between the effective load and the satellite and can be used for isolating the vibration transmitted to the effective load from each direction of a satellite platform; and the vibration isolator is connected with the effective load and the satellite through the flexible hinges, and thus a gap, collision and the like are avoided.

A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and sub-orbital altitudes and velocities. The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance, navigation, and control unit (GNCU) that performs all guidance and control for the SLV from main stage ignition to orbital insertion. The GNCU can remain with the payload after orbital insertion to provide satellite station keeping and orbital maneuvering capability. The use of a single integrated avionics unit for all guidance, navigation, and control simplifies the SLV, reducing weight and significantly reducing cost. In addition, this architecture allows for a combined launch and satellite bus system as the GNCU can also be used as a satellite bus. This further reduces cost and increases the payload capacity to orbit by optimizing the use of launch vehicle and satellite bus subsystems and reducing non-instrument mass delivered to orbit. All support functions are provided by the IDMV. This approach represents a significant improvement over conventional systems, especially with respect to the orbital launch of payloads less than about 100 kg.

The invention relates to an integrated test simulation system for a spacecraft attitude control algorithm and belongs to the computer simulation technology and application technical field. The integrated test simulation system includes a controller, a measurement model, an actuator model, a spacecraft attitude dynamic model, an antenna model, a sailboard model and a load model, wherein the controller includes a mode management module, a path optimization module, a closed loop control module and an accessory monitoring module. According to the system of the invention, the components of the system are configured based on a universal satellite platform; modeling is performed on system equipment related to typical attitude control and dynamic characteristics; and a simulation on-orbit simulation environment supporting modularized applications can be provided for the control algorithm; near-earth orbit satellite mission modes are preset in the mode management module, and a user can select one kind of near-earth orbit satellite mission mode to fast establish a simulation scene, and test data is of practical significance; the path optimization module provides a function for planning paths of spacecraft attitude maneuver, and therefore, the performance of a closed-loop control mechanism can be improved.

The invention relates to a satellite multiple attitude control mode test system based on a double gimbal control moment gyroscope (DGCMG) structure. Performance of three satellite attitude control systems which are based on a DGCMG, a single gimbal control moment gyroscope (CMG) and a flywheel is tested and verified through a DGCMG actuating mechanism. The satellite multiple attitude control mode test system based on the DGCMG structure comprises a platform system, the satellite attitude control system, a space environment simulation system and a ground station system. The platform system is composed of a tri-axial air bearing table, a satellite service comprehensive management system, a power source and a wireless bridge and used for simulating satellite dynamic characteristics and information management. The satellite attitude control system is composed of a jet propulsion system, the DGCMG, a fiber-optic gyroscope, a star sensor, a sun sensor and a global position system (GPS) receiver and used for determination of an attitude and an orbit and control of a satellite platform. The space environment simulation system is composed of a GPS simulator, a sliding block which is of a pyramidal structure, a sun simulator and a star simulator and used for simulating space interference torque, and part performance of a GPS satellite and a celestial body. The satellite multiple attitude control mode test system based on the DGCMG structure can provide ground testing and verification for multiple attitudes of a satellite.

A method is presented for utilizing excess satellite bus power comprising (1) sending a broadband signal in at least one feeder beam in a forward direction from a gateway terminal to a bent pipe repeater satellite for relay to at least one subscriber terminal, wherein the satellite is operable to provide a total amount of bus power, wherein an existing payload consumes an occupied portion of the total bus power, wherein an additional payload consumes a remaining portion of the total bus power and comprises a plurality of satellite-based transmission amplifiers, (2) receiving the broadband signal and amplifying the broadband signal using one of the plurality of satellite-based transmission amplifiers, (3) sending the amplified broadband signal, as one of a plurality of service spot beams, to the at least one subscriber terminal, and (4) receiving and retrieving data from the amplified broadband signal at the at least one subscriber terminal.

The invention provides an eight-rod-connecting type non-contact satellite platform configuration and an assembling method. The eight-rod-connecting type non-contact satellite platform configuration comprises an upper connecting plate, a lower connecting plate, four same vertical mounting force actuators and four same horizontal mounting force actuators; both the vertical mounting force actuators and the horizontal mounting force actuators are connected between the upper connecting plate and the lower connecting plate and are arranged in a circumference manner; the vertical mounting force actuators and the horizontal mounting force actuators are arranged along the same circumference in a staggered and orthorhombic manner; the vertical mounting force actuators are vertical to the horizontal mounting force actuators. The eight-rod-connecting type non-contact satellite platform configuration implements complete isolation of an effective load on platform vibration and interference response, can be applied to spacecrafts with high-resolution sensitive effective load, such as a very-high accuracy remote sensing satellite, a deep space exploration astronomical telescope and the like, and can greatly improve exploration performance of spacecrafts.

A spacecraft for removing space debris is disclosed. The spacecraft includes a satellite bus, a shield member foldable on an outer side face of the satellite bus and disposed facing towards space debris to reduce a movement speed of the space debris, and a support member configured to support the shield member with respect to the satellite bus, in which the shield member includes a central panel configured to overlap one face of the satellite bus, a plurality of first panels connected to peripheral sides of the central panel and radially extended, and a plurality of second panels located between the first panels.

A satellite solar battery array in-orbit test circuit is used for a satellite solar battery array in-orbit IV curve test, and comprises an on-satellite electronic device, an on-satellite information processing unit and a satellite-ground communication unit. After an IV test circuit is switched on, in-orbit real-time data including voltage, current, temperature and the like of a solar battery array can be acquired through the on-satellite electronic device in real time, then is processed by the on-satellite processing unit and is downloaded by the satellite-ground communication unit. A complete solar battery array IV characteristic curve is drawn. The satellite solar battery array in-orbit test circuit overcomes the defects of an existing method that only the voltage and current of the working point of a satellite bus can be counted and the IV curve cannot be visually and comprehensively reflected, improves the accuracy and comprehensiveness of performance monitoring of the solar battery array, provides complete IV curve information for the performance evaluation of the solar battery array, improves the evaluation efficiency of the satellite life, and provides the in-orbit management of the satellite solar battery array with simpler and more visual data support.

The invention relates to an integrated satellite system for measuring a global gravity field. The system comprises two satellites which fly on the same nearly polar and nearly circular orbit in a following mode, wherein each of the two satellites is loaded with a double-frequency global navigation satellite system (GNSS) receiver besides a satellite platform guarantee subsystem; a high-low satellite tracking satellite link consists of the two satellites and a high orbit navigation satellite system and is used for determining a satellite orbit at high accuracy and acquiring long wave band information of an earth gravity field, which is contained in satellite orbit information; meanwhile, the two satellites are loaded with high-accuracy laser interference distance measurement devices, and a low-low satellite tracking satellite link is formed through laser interference and used for acquiring medium-long wave band information of the earth gravity field, which is contained in a laser interference signal; and in addition, one satellite is loaded with a high-accuracy gravity gradiometer for satellite gravity gradient measurement, and a measurement output signal contains the medium-long wave band information of the earth gravity field, so that the global time-variant and static gravity field measurement of long waves, medium-long waves and medium-short waves is finished.

The invention discloses a system of an attitude and orbit control actuator for an all-electric propulsion satellite platform. The system of the attitude and orbit control actuator for the all-electric propulsion satellite platform comprises a satellite orbit control system, a satellite attitude control system, an electric propulsion system, a control torque gyroscope system, a reaction flywheel system and a satellite measurement and feedback system. In view of the working characteristics of a GEO satellite, the system of the attitude and orbit control actuator for the all-electric propulsion satellite platform provides an actuator system which adopts the control torque gyroscope, the reaction flywheel and the electric propulsion system, and aims at providing technical assurance for orbit transfer, orbit keeping and agile maneuvering of the GEO satellite; and through cooperation of the system, the following objectives on GEO satellite that the launch and orbit quality are reduced, the launch cost is reduced, the satellite position maintains accurate, the maneuverability and agility of the satellite attitude are improved, and the service life of the satellite is prolonged are achieved.

The invention discloses a system for cross backup of telemetering and digital transmission data channels for aerospace applications. The system comprises an integrated electronic subsystem, a spaceborne measurement control transponder, a digital transmission subsystem and spaceborne data equipment which are locally arranged on a satellite platform, wherein data transmission channels exist between the spaceborne data equipment and the integrated electronic sub-system and the digital transmission subsystem, and a data transmission channel is established between the integrated electronic sub-system and the digital transmission subsystem through the spaceborne data equipment to achieve the interconnection and multiplexing between a telemetering data download channel and a digital transmission data download channel; the spaceborne data equipment achieves the exchange of the telemetering / remote control data through a measurement control channel and the integrated electronic subsystem, unidirectionally transmits the large-capacity test data to the digital transmission subsystem at high speed through an LVDS data interface, and also achieves a two-way data connection through an RS-422 data interface and the digital transmission subsystem. According to the system disclosed by the invention, the advantages of the telemetering channel and the digital transmission channel are fully combined, the advantage complementation and cross backup of the telemetering channel and the digital transmission channel can be achieved, and the real-time performance and security of the data can be increased.

The invention discloses a CubeSat satellite platform based on a mobile phone. The CubeSat satellite platform comprises the smartphone, an interface switching module and an external module, wherein the external module comprises one or more of an attitude control module, a communication module, a maneuvering module and an extendable loading module, the smartphone is used for performing interaction with the attitude control module, the communication module, the maneuvering module and / or the extendable loading module respectively through the interface switching module to complete satellite platform tasks, the attitude control module is used for achieving attitude control of a CubeSat, the a CubeSat is used for achieving information interaction between the CubeSat and a ground station or other spacecrafts, the maneuvering module is used for controlling the CubeSat to operate in a set orbit, and the extendable loading module is used for extending load and / or adding corresponding functional units so as to achieve extendibility of the satellite platform under the control of the smartphone. The coupling problem of satellite design, tasks and progress is effectively solved, the period of study and development is short, and expansibility and reconfigurability are high.

The invention discloses a satellite-ground integrated high-precision satellite multi-beam calibration method. On-satellite calibration beam forming equipment is used for achieving multi-beam calibration signal generation and beam forming. A ground calibration station is located at the theoretical pointing center of a multi-beam calibration beam, and real-time beam pointing error information is obtained by capturing and receiving calibration signals and measuring power. Measured data are transmitted to the control center through a ground link or a satellite link; wave beam pointing deviation compensation parameters are generated in the control center and sent to the satellite measurement and control station; and the satellite measurement and control station converts the beam pointing deviation compensation parameter into an upload instruction and injects the upload instruction into the satellite platform. According to the method, a ground calibration mode is adopted, the requirement foron-satellite operational capability can be reduced, an on-orbit operation processing flow is simplified, measurement errors caused by satellite power fluctuation are reduced, the precision of a calculation result is improved, and the method has the advantages that the system structure is clear, the calibration process is clear in division of work and positioning, and the operation complexity of adeviation calculation method is low.

The present invention provides a measurement device used for space ions and neutral atoms. The measurement system comprises an ion deflection system, a neutral atom ionization deflection system, an electrostatic analyzer, a flight time system, an electronics processing unit and a device shell. The back-end electrostatic analyzer, the flight time system and the electronics processing unit are shared to measure the ions generated by ionizing the neutral atoms and the ions radiated from the space directly, the measurement of the space ions and the neutral atoms is integrated in a probe, and the as much as possible measurement parts are shared, so that the weight and the power consumption of the device are reduced furthest, the requirements for the satellite platform resources are also reduced, and the fields to which the device is applied are expanded. The measurement device of the present invention has the wide application demands both in the terrestrial space field and the deep space exploration field.

The invention relates to a reconstruction method of a satellite-borne control computer. The reconstruction method includes the steps of 1, subjecting a FPGA (field programmable gate array) of each computer unit to system partitioning; 2, allowing a surface control center to generate FPGA configuration files of the computer units; 3, allowing the surface control center to transmit the FPGA configuration files to a data management subsystem; 4, allowing the data management subsystem to transmit the FPGA configuration files to a satellite bus; 5, allowing the computer units to perform reconstruction in the on-duty state. The reconstruction method has the advantages that the original central processing unit of the satellite-borne computer is replaced by the SRAM (static random access memory) type FPGAs, a reconstruction control circuit and a reconstruction configuration data storage designed on the original satellite-borne computer are connected to a configuration interface of the SRAM type FPGA, each SRAM type FPGA includes a constructible area and a fixed area, reconstruction of the satellite-borne control computer can be achieved on orbit, the satellite-borne control computer is under control of the surface during an on-orbit phase, and functions are changed in real time through the configuration files injected from uplink.

The invention discloses a computer of integrated satellite-borne SAR loads. The computer comprises a main control module, a telemetry and telecommand module, an interface expanding module, a power module and a backboard. The main control module is used for data processing, timed controlling and wave beam controlling; the telemetry and telecommand module is used for receiving a satellite platform instruction, analyzing the instruction to all single units of the load system to instruct production and OC gate control, and collecting the temperature, voltage analog quantity and the digital quantity telecommand signal of the load system; the interface expanding module has the functions of communication management and interface expanding; the backboard is used for power supply and signal connection of the main control module, the interface expanding module, the power module and the telemetry and telecommand module. Through integrated design of functions of time sequence control, wave beam control, and interface control when the SAR loads work and the telemetry and telecommand of all single units of the load system, the integration level and the reliability of the SAR system are improved,the weight, size and the power dissipation are lowered, data transmission among all machines of the SAR load system is reduced, and the reliability of the SAR load system is improved.

The invention provides an integrated electric propulsion satellite platform position holding electric thruster redundant configuration method. Firstly, a satellite ground board rectangular coordinate system is built; secondly, the side, close to a south board, of a satellite ground board is provided with three electric thrusters, and the side, close to a north board, of the satellite ground board is provided with three electric thrusters; thirdly, work areas of the electric thrusters are generated respectively under the condition that all the electric thrusters work or some electric thrusters break down, ignition is conducted according to the work areas of the electric thrusters, and the electric thruster redundant configuration method is implemented. By arranging the six electric thrusters to keep the south, north, east and west positions of the satellite on-orbit period and complete momentum unloading together, requirements of an integrated electric propulsion satellite for thrust strength and the service life of the thrusters can be effectively reduced, adaptability of the integrated electric propulsion satellite to the electric thrusters is improved, the problems that a single principal part thruster makes an application strategy complex if losing efficacy and propellant consumption demand is increased are solved, and the method has the advantage of being good in adaptability to failure of the single thruster.

The invention provides a space-based network intelligent satellite development system based on task requirements. The space-based network intelligent satellite development system comprises a constellation configuration module, a satellite assembly module, a component design module, a system mapping module, a simulation evaluation module and a compiling and downloading module. The task demand is analyzed through constellation. Hardware resources and software resources required for solving task requirements are extracted. Load modularization is realized from the aspects of hardware and software.Separation of hardware resources and software definition is realized; a required hardware environment and a required software environment are built according to task requirements; storing the data asexecutable files and downloading the executable files to all intelligent satellites; the intelligent satellite realizes function transformation according to the change of constellation task requirements; according to the space-based network node, intelligent satellite function solidification of special research and development is overcome, the problem that the space-based network node depends onan immobilized satellite platform is solved, and the space-based network node has multifunctional bearing, network collaboration, resource optimization and intelligent service capabilities to provideefficient, active and reliable space-based information precision service for users.

A middle and high layer atmosphere density separation type detector comprises a universal module with a microsatellite platform, and is characterized by further comprising a spherical structural body shell used for simplifying an atmosphere resistance analysis model, an accelerometer, a GPS receiver and an attitude control system, the surface of the special structural body shell is covered with a solar cell, the accelerometer is arranged in the spherical structural body shell, the accelerometer is formed by distributing 4 to 6 static suspension type accelerometers around the axis, the accelerometer is used for measuring the atmosphere resistance, the GPS receiver is arranged in the spherical structural body shell and used for tracking and monitoring the track speed and the position of a microsatellite, the attribute control system is arranged in the spherical structural body shell and used for controlling the attribute of the microsatellite so as to keep the spinning of the satellite stable and ensure the correct directing of a GPS receiver antenna and a data transmission antenna.

The invention discloses a satellite platform flutter detection method based on a TDICCD splicing zone image. The method comprises the following steps: 1) obtaining overlapped images, and using a timedelay and integration-charge coupled device splicing zone to obtain the overlapped images; 2) calculating a difference of relative imaging positions, performing accurate and dense coupling treatment on two overlapped images to obtain a homonymous matching point, and calculating the difference of the relative imaging positions of a same target in two overlapped images; and 3) estimating the satellite platform flutter according to the difference of the relative imaging positions. The method can effectively solve the problems of low flutter detection accuracy and only detection on several isolated frequency points, provides the flutter detection method with high precision and full coverage in a wide frequency range for a satellite platform, and increases the satellite attitude detection capability and extraction accuracy for the remote sensing image information.

In order to provide a satellite information autonomous distribution satellite capable of collecting attribute information for satellites after failure or damage occurs at or to the satellite, an information collection device is operated by a second power supply provided independently from the first power supply operating the satellite bus system, is provided separately from the satellite bus system, and collects and stores attribute information. The light receiver is provided separately from the satellite bus system and forms a prescribed control signal when light sent from the information collection station is received. The reflecting type light modulating device is operated by the second power supply, is provided separately from the satellite bus system, and transmits reflected light superimposed with a modulation signal containing attribute information in the reflected waves using this modulation signal when a modulation signal is formed based on attribute information stored in the information collection device according to the shape of the control signal.

The invention discloses a device for unfolding a surface of a space rope system, and is mainly used for solving the problems that the prior art is difficult to mold and has bad stability. The device comprises a plurality of reflectors or transmission surfaces (1), a plurality of flexible pulling cables (2), a satellite platform (4) and a plurality of control mechanisms (3) of the pulling cables, wherein each reflector or transmission surface (1) is composed of a folding ring (4) and a ring connecting joint (5), and is stacked according to the caliber size; one end of each flexible pulling cable (2) penetrates through a corresponding hole of the folding ring (4) of the reflector or transmission surface with the biggest caliber to be tied fixedly, and then penetrates through the folding ring (4) of the other reflector or transmission surface to be tied fixedly; the other end of each flexible pulling cable (2) is connected with the control mechanisms (3) on the satellite platform; and the reflectors or transmission surfaces are connected hierarchically to form a stacked structure composed of the satellite platform and all reflectors or transmission surfaces so as to form a space rope system. The device can achieve effective molding structure, autonomous orbit and pose stability of the whole system, so that the device can be applied in a large antenna or telescope in space.

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 discloses a control method of a geosynchronous earth orbit (GEO) synthetic aperture radar (SAR) for covering all over China, which belongs to the technical field of synthetic aperture radars and overcomes the defect that the GEO SAR is capable of realizing one-dimension phase scanning in direction only rather than covering all over China under limitation of engineering implementation capacity of a phased array antenna at the present stage. In the control method, a satellite platform with the SAR has at least 2-degree pitching maneuver capacity and the SAR has the one-dimension phase scanning capacity in distance. By means of maneuver of pitching angle of the satellite platform and phase scanning in direction of the SAR, any locations within China can be observed as needed, the defect of one-dimension phase scanning of GEO SAR for covering all over China is overcome, and engineering implementing is facilitated.

The invention discloses an aerosol optical thickness inversion method of a domestic multi-angle polarization satellite sensor. The method comprises the following steps: carrying out satellite multi-angle and multi-band data matching preprocessing; identifying cloud pixels of the multi-angle polarization satellite image data; based on surface reflection contribution estimation of an empirical orthogonal function, constructing a scattering matrix under multi-angle observation to perform surface radiation item estimation; constructing an aerosol model and establishing an inversion lookup table; performing aerosol optical thickness inversion based on optimal model identification. The invention provides a global land overhead aerosol optical thickness inversion method suitable for a high-resolution No.5 satellite multi-angle polarization imager (DPC) for the first time, and particularly relates to a global land overhead aerosol optical thickness inversion method suitable for a high-resolution No.5 satellite multi-angle polarization imager. The algorithm is stable, efficient and verifiable, and the method can be applied to other satellite platforms with multi-angle detection capability in a business manner.

The invention provides a stationary orbit ultra-large type assemblable satellite platform configuration, which belongs to the technical field of spacecrafts, and comprises a mother platform, a transition sub-platform I and a transition sub-platform II, a functional sub-platform I and a functional sub-platform II, a functional sub-platform III and a functional sub-platform IV, an energy module I and an energy module II, and a mechanical arm I and a mechanical arm II, wherein the mother platform is a control center of a whole satellite; the transition sub-platform I and the transition sub-platform II are respectively connected with the mother platform through connecting devices; the functional sub-platform I and the functional sub-platform II are respectively connected with the transition sub-platform I through connecting devices; the functional sub-platform III and the functional sub-platform IV are respectively connected with the transition sub-platform II through connecting devices; the energy module I and the energy module II are respectively connected with the transition sub-platform I and the transition sub-platform II through connecting devices to supply power; the mechanicalarm I and the mechanical arm II are installed on the transition sub-platform I and the transition sub-platform II correspondingly to execute an on-orbit assembling task. The platform configuration issuitable for expansion, can adapt to on-orbit construction of large satellites with the on-orbit flight weight of 10-20 t, is applied to various types of satellites and service stations, and is wide in application range.

A satellite communication system, infrastructure, and method of providing same, involves inventorying certain types of pre-selected communication satellite payload modules and communication satellite bus modules based on analysis of anticipated communication satellite needs and providing each payload and bus module with a pre-configured standard interface coupling. When communication satellite parameters for a customer are received, a selection is made of one or more of the inventoried payload modules and also of the inventoried communication satellite bus module that is sized and configured to accommodate the selected payload module or modules. The selected payload module or modules are coupled to the selected bus module via the standard interface coupling, and the selected bus module is, in turn, coupled to a launch vehicle.

The invention discloses a full-link photon counting laser altimeter point cloud generation method. A full-link simulation model including ground light spot coordinate calculation, atmospheric refraction offset correction, target geometry and radiation characteristics, detector response and the like in the height measurement process of a satellite-borne laser altimeter is established, and high-precision single-photon point cloud is generated through simulation in combination with satellite platform parameters, laser altimeter system parameters, environmental parameters, target characteristic parameters and the like. The laser altimeter point cloud generation method can accurately adjust various system parameters and environmental parameters, and has the advantages of being rapid, easy to simulate various conditions and good in expansibility. ICESat-2 system parameters can be used as input to simulate and detect randomly generated terrain surfaces, and high-quality photon point clouds are generated.