243 results about "Space vehicle" patented technology

Various products, systems, and parts or components of such products and systems include a binder matrix (such as a polymer matrix) reinforced with carbon fibers and graphene particles. Inclusion of the carbon fibers and graphene particles substantially improves mechanical strength and utility of the resulting parts, components, and corresponding products and systems. The improvements facilitate substantial advances in various practices of fabrication using additive manufacturing techniques, which lend to applications in biomedical, industrial, agricultural, mechanical, energy generation, transmission, and storage, military, transportation, recreational, and other fields. The inventive materials are particularly well suited for manufacturing of parts for terrestrial vehicles, manned or unmanned aerial vehicles, and space-borne vehicles. Substantial improvements to mechanical strength are achieved even at very low levels of graphene loading (e.g., at about 0.1 wt %), such that the resulting products are substantially stronger than corresponding products lacking said graphene loading, while yet having substantially the same mass.

The invention provides a space carrier intelligent bus GLINK interface bridging design method based on a domestic FPGA, belongs to the field of FPGA application, can provide two paths of GLINK data transmission channels and two paths of GLINK data relay / forwarding channel high-speed interfaces at the same time by utilizing FPGA chip resources and combining with a GLINK chip, and can meet the requirements of switching networking and annular networking at the same time. An external memory access interface (EMIF) of a GLINK chip is configured for an international standardized PCIe protocol, and meanwhile, an FIFO synchronous read-write interface in a DMA mode is adopted, so that the communication safety is guaranteed, malicious attacks and accidental interruption are prevented, the communication reliability is improved, and the communication transmission rate is increased.

The invention provides a Hall and ion hybrid electric propulsion system. The Hall and ion hybrid electric propulsion system comprises a composite material gas cylinder, an inflation valve, a pressure adjusting module, a Hall flow adjusting module, an ion flow adjusting module, a Hall thruster, an ion thruster, an electric propulsion control unit, a power processing unit and a power supply processing and control unit, the pressure adjusting module supplies gas to the Hall flow adjusting module and the ion flow adjusting module in a time-sharing manner; the Hall flow regulation module and the ion flow regulation module are used for carrying out flow control on the Hall thruster and the ion thruster respectively; the electric push control unit performs time-sharing power supply control on the power processing unit and the power supply processing and control unit; the power processing unit and the power supply processing and control unit respectively provide electric energy for the Hall thruster and the ion thruster; the Hall thruster and the ion thruster provide thrust or impulse for the spacecraft; according to the Hall and ion hybrid electric propulsion system, the integrated design of air supply, power supply and autonomous process control of the Hall and ion thrusters is achieved, system air path and circuit configuration is simplified, and the requirement for a spacecraft with high specific impulse and high thrust electric propulsion systems at the same time is met.

The invention discloses a thermomechanical treatment method for manufacturing a high-performance ribbed shell section, and belongs to the technical field of high-performance precision plastic forming manufacturing. According to the method, firstly, through high-temperature multidirectional forging cogging, medium-temperature rectangular ring rolling and high-temperature in-mold forging and rolling composite forming, near-net forming of the ribbed shell section, coarse crystal phase crushing and refining and grain refining are achieved; and finally, through the design of solid solution, in-mold cold rolling and aging, the problem that cold deformation is discontinuous and uneven due to a split mold structure in an existing ring piece cold bulging process is solved, accurate shape correction, residual stress reduction and performance improvement are achieved, and the product quality is improved. And the high-performance near-net plastic forming and high-performance manufacturing requirements of a key structural part, namely a ribbed shell section, of an aerospace carrier are met.

A control surface system for a body located in an air and / or space vehicle. The control surface system having at least one control surface which is located on the body and which, by moving relatively to the body, enables air flow to be controlled and the air vehicle to maneuver accordingly. The control surface system having at least one actuator which is made of an electroactive polymer material which is located on the body and which changes its form depending on electrical energy and based thereon, triggers the control surface.

A radar absorbing structure has a body that forms an air and / or space vehicle. At least one aerodynamic surface is located on the body. At least one resistive layer is located on the at least one aerodynamic surface and allows substantial damping of electromagnetic waves on the at least one aerodynamic surface by means of the destructive interference. A plurality of fibers are located in the resistive layer. At least one binding agent holds the fibers together.

The present invention belongs to the field of complex equipment system design and verification. It discloses a model-based integrated design and verification method for complex equipment systems. With the four levels of "demand-behavior-structure-parameter" as the framework, combined with mission scenario requirements, a full-process design and analysis method for reusable space vehicle systems is constructed; starting from the demand analysis, the complex system is decomposed, and the model is calibrated and performance verified by combining theoretical and experimental data to construct a reusable space vehicle system model, and carry out system comprehensive analysis and verification; combined with the model's three-dimensional virtual environment and configuration optimization algorithm, various configurations of the space vehicle are quickly optimized to evaluate the feasibility and performance of the design scheme; through the two-way interaction of design and verification models, and the integration of system-level design and professional-level design, the problems of design and verification separation and insufficient model collaboration in the existing technology are solved, and the efficiency and accuracy of complex equipment system design and verification are improved.

Systems, devices, methods, and computer-readable media for space vehicle sensor management. A method includes receiving, at a mission operations center (MOC), respective regional requests from respective regional schedulers, the respective regional requests indicating mission windows (MWs) and corresponding sensors to be operated in associated MWs, receiving, at the MOC, respective sensor plans from corresponding space vehicle operation centers (SVOCs), each sensor plan indicating MWs for which a given sensor is unavailable, generating, based on the regional requests and the sensor plans, a MW apportionment for each regional scheduler, the MW apportionment indicating MWs and corresponding sensors that a user associated with the regional scheduler has authorization to command the sensor, and providing the MW apportionment for the regional scheduler to the regional scheduler.

A method includes providing a quantity of sheet metal, press-forming a first panel from the sheet metal to form a first stamped panel, press-forming a second panel from the sheet metal to form a second stamped panel, fixing at least one stiffener to the first stamped panel such that the stiffener extends across a width of the first panel, where a first end of the stiffener is attached to the first stamped panel at a first fixation location, and where a second end of the stiffener is attached to the first stamped panel at a second fixation location, fixing at least one onboard equipment component to the first stamped panel, where at least one of the at least one onboard equipment component is fixed to the stiffener, sealing the first panel to the second panel to form a space vehicle.

The invention relates to the technical field of satellite orbit prediction, and provides a space craft behavior prediction method and system, and the method comprises the steps: S1, obtaining the continuous space states of a space craft at a plurality of moments; s2, selecting a space state as a current space state; s3, calculating a reward value for executing a flight action in the current space state; s4, acquiring a next space state after the flight action is executed in the current space state; s5, calculating income values of executing all the next flight actions in the next space state; s6, updating the income value; s7, taking the next space state as the current space state; s8, repeating the step S4 to the step S7 until the income value is greater than a specified income value; and S9, selecting the flight action with the maximum profit value as the predicted action of the space craft behavior according to the current space state and the profit value of the space craft. According to the scheme, real-time monitoring and high-precision prediction of the space aircraft can be realized.

The invention provides a multifunctional modular combination valve for a spacecraft cold air pressurization system, which comprises a valve body and a part assembly mounted on the valve body, the part assembly comprises an electric explosion valve, a cushion valve, a pressure reducing valve, a safety valve and a pressure sensor, the valve body is a structural body with a plurality of cavities inside, an air inlet connector and an air outlet connector are formed in one side face of the valve body side by side, installation parts corresponding to the components are arranged outside the valve body, and the components are connected to the valve body in a sealed mode through threads. The structure is compact, the size is small, the structural layout of a pressurization system is simplified, the space envelope size is reduced, the assembly manufacturability is improved, the maintainability is improved, and the requirements for miniaturization, light weight and integration of a new-generation aircraft are met; the number of high-pressure pipelines and weld joints in a pressurization system is reduced, the weight of the system is reduced, and the sealing reliability and the mechanical environment resistance are improved.

The invention relates to the technical field of spacecrafts, in particular to an attitude and orbit control system and an aircraft. The system is applied to the aircraft, the wall of an aircraft body of the aircraft is formed by connecting a main body on the inner side and a skin on the outer side, a flow channel extending from the head of the aircraft body to the tail of the aircraft body is arranged between the main body and the skin, and the system comprises a gas storage bottle, an attitude controller and a track booster, the outlet end of the gas storage bottle is communicated with the flow channel at the head position of the machine body, and the gas storage bottle is used for providing heat-conducting gas into the flow channel; the posture controller is arranged at the position between the head and the tail of the machine body, the posture controller communicates with the flow channel, and the posture controller is used for discharging gas in the flow channel out of the machine body; the rail booster is arranged at the tail of the machine body, communicates with the flow channel and is used for discharging gas in the flow channel out of the machine body. According to the technical scheme, the maneuverability of the aircraft can be improved.

The application relates to an on-orbit maintenance method for a self-repairable satellite-borne DSP, which is realized by an on-orbit maintenance system and is used for maintaining and managing a DSP chip; the on-orbit maintenance method comprises the following steps: S1, powering on the on-orbit maintenance system; S2, performing three-mode redundancy comparison error correction and CRC check word checking to judge whether the checking is passed; S3, mapping the DSP chip to a target chip area address and performing DSP program data loading; S4, judging whether the monitoring function of the DSP chip is normal; and S5, downloading the chip area address of the DSP program data which is not passed in the telemetry and executing satellite platform instructions to repair the DSP program data. The application has simple architecture, low implementation difficulty, can effectively prolong the on-orbit life of a satellite and other space vehicles while keeping low cost.

The satellite supporting cabin comprises an upper end frame, a truncated cone shell and a lower end annular flange which are integrally formed by carbon fiber composite materials, the upper end frame is formed by wrapping a first foam sandwich with the carbon fiber composite materials and solidifying the first foam sandwich, and the upper end frame is used for being connected with a satellite support. The lower end annular flange is used for being connected with a substage of an aerospace vehicle. The satellite supporting cabin can achieve the purpose of reducing the weight and has high rigidity.

A space vehicle includes a space vehicle body and a radiator, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein the space vehicle body includes a panel including a sheet metal; wherein the radiator is positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, wherein the corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of: sunlight illuminating Earth in a manner of an equinox date, the space vehicle positioned at a low Earth orbit altitude in an equatorial plane, and an observer positioned on the equator, the plurality of corrugations reflect sufficient sunlight away from the Earth such that a minimum magnitude of the space vehicle is in a range of from 6.5 to 9.5.

A system, including: a satellite; a first solar array including a first solar panel; a second solar array including a second solar panel; a boom having a first end and a second end opposite the first end, where the first end connects to the satellite; and a bracket assembly, where the bracket assembly has a first, second, and third brackets, where the first bracket connects the first solar array to the third bracket, where the second bracket connects the second solar array to the third bracket, and where the third bracket connects the first bracket and the second bracket to the second end of the boom; where the bracket assembly is configured to reorient the first solar array and the second solar array between a stowed position to a deployed position, and where in the deployed position, the first and second solar arrays are oriented at a predetermined angle.

Aspects presented herein may enable a network entity to determine whether there is any error in broadcast data from one or more spacecrafts (SVs) based on crowdsourcing information from a set of roadside units (RSUs). In one aspect, a network entity obtains at least one measurement of a satellite associated with a set of RSUs. The network entity estimates a location of the satellite based on the at least one measurement and a location of the set of RSUs. The network entity identifies whether the estimated location of the satellite is associated with an error based on a comparison of the estimated location of the satellite to ephemeris or almanac data of the satellite.

A space vehicle includes a space vehicle body and a radiator, the space vehicle body including an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, wherein at least a portion of the space vehicle body includes a panel including a sheet metal; the radiator being positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations configured such that, when the space vehicle is evaluated under certain conditions, the corrugations reflect sufficient sunlight away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle lacking the plurality of corrugations and otherwise equivalent to the space vehicle, when evaluated under the same conditions

Systems, devices, methods, and computer-readable media for space vehicle security are provided. A space vehicle can include a navigational controller configured to determine a mode of the space vehicle, a power supply configured to provide power to components of the space vehicle, a transceiver configured to determine whether a communications channel to a ground station is open, and an intrusion detection monitor configured to monitor for non-nominal outputs from components of the space vehicle, the components that are monitored are variable based on a power level of the power supply, a connectivity state of the transceiver, and the mode.

The invention relates to the technical field of spacecrafts, in particular to a cooling system and an aircraft. The system is applied to an aircraft, an aircraft body of the aircraft is formed by connecting a main body on the inner side and a skin on the outer side, a flow channel extending from the head of the aircraft body to the tail of the aircraft body is arranged between the main body and the skin, and the aircraft body is provided with an air film hole penetrating through the main body and the skin; the cooling system is arranged in an inner cavity of the machine body and comprises a storage assembly, a cooling assembly and a circulating assembly, the storage assembly is used for storing a cooling medium, and the cooling assembly is used for cooling the cooling medium; the circulating assembly comprises a power pump and a turbine, and the power pump is used for pressurizing a cooling medium and pumping the pressurized cooling medium into the flow channel from the head of the machine body; an air inlet of the turbine is communicated with the runner between the tail part and the neck part of the machine body; and an air outlet of the turbine is communicated with one end of the air film hole in the inner cavity of the machine body. The system can quickly reduce the temperature of the head of the aircraft.

The invention provides an attitude control energy optimization method for a spacecraft. The attitude control energy optimization method comprises the following steps: firstly, obtaining a smoothed attitude angle instruction and an attitude angle rate instruction by using a tracking differentiator; estimating impact disturbance through a nonlinear extended state observer; forming an impact disturbance sign signal according to the estimated value; obtaining a state error, establishing an attitude control law under a non-interference condition and an attitude control law under impact disturbance, and obtaining a corresponding control quantity and a sliding mode surface; according to the state error and the two attitude control laws, the switching compensation control quantity is calculated, the attitude control law under the impact disturbance containing integral switching compensation control is designed, and the smoothed control quantity under the impact disturbance is obtained; according to the impact disturbance mark signal, the control quantity under the interference-free condition or the control quantity under the smooth impact disturbance is selected and sent to the attitude control engine. The method effectively solves the energy optimization problem of the direct force control system when the parameters of the control system change drastically under large impact disturbance.

The utility model relates to novel straddle type monorail three-section track changing equipment which can meet the requirement for efficient track changing of multi-turnout vehicles in a station yard and save space and comprises a bottom plate, a track and a locking device, the track is located on the bottom plate, a walking trolley is located on the track, and a plurality of turnout beams are located on the walking trolley. One end of the turnout beam is provided with a piled beam and a slewing device, the other end of the turnout beam is provided with a plurality of piled beams, a slewing crank arm mechanism is arranged at the bottom of the turnout beam, rollers on the slewing crank arm mechanism do linear motion in a fixed bottom plate and a guide groove in a track, and the slewing crank arm mechanism is driven by a driving device. The method has the advantages that the problem that the vehicle base space is limited is solved, and the requirement for small-space vehicle base route planning is met; signal receiving and safety feedback requirements are met, and the characteristics of linkage interlocking, flexible action, safety, reliability and the like are achieved; narrow space of a parking lot is fully utilized, the structure is compact, and functions are complete; and fourthly, the part design has good manufacturing manufacturability, and the manufacturing cost is favorably controlled.

A system and method for preparing and supplying a densified liquid oxidant to a spacecraft launch facility is provided. Liquid oxygen streams obtained from co-located air separation devices for liquid preparation are densified in an integrated densification system of two refrigeration stages. The first refrigeration stage is a nitrogen-based inverse Brayton cycle refrigeration cycle that provides refrigeration to the second refrigeration stage. The second refrigeration stage is a refrigerant circuit comprising helium and / or neon that densifies the liquid oxygen to a temperature between about 70 Kelvin and 57 Kelvin. The integrated densification system may also be configured to densify liquid methane or other propellants used in a spacecraft.

A space vehicle includes a first aluminum panel, a second aluminum panel fixed to the first aluminum panel around at least a portion of their respective perimeters to form a closed structure with an interior volume, a plurality of slots formed in at least one of the aluminum panels, either the first or the second, at least one heat transfer element located in at least one of the slots, and a plurality of onboard equipment elements, wherein at least one of these elements is located next to the heat transfer element, wherein at least some of the elements lack a casing, and wherein the heat transfer element is configured to transport heat away from the other elements.

The invention provides a method for correcting wind before launching of an aerospace vehicle. The method comprises the following steps: acquiring high-altitude wind data actually measured in minus 1 hour; taking minus 1 hour high-altitude wind data as input of trajectory correction, and outputting guidance parameters; carrying out six-degree-of-freedom Monte Carlo shooting simulation, and calculating a flight load; and determining whether to launch the space launch vehicle at least according to the flight load envelope. The method can improve the launching release probability and reduce the pneumatic attack angle and the flight load.

The utility model relates to the technical field of spacecraft thermal control multi-layer manufacturing, in particular to a distributed thermal control multi-layer material cutting device which comprises a supporting shell, a film supporting set, an X-axis track, a Y-axis track, a cutting machine and a vacuum unit, the film supporting set is arranged at the upper end of the supporting shell, the X-axis track is arranged on one side of the upper end of the film supporting set, and the Y-axis track is arranged on the other side of the upper end of the film supporting set. An X-axis track is arranged on the supporting shell in a sliding mode, a Y-axis track is arranged on the X-axis track in a sliding mode, a cutting machine is arranged at the upper end of the Y-axis track, the vacuum unit is located on a bottom plate of the supporting shell, the film supporting set comprises a fixed surrounding plate, a telescopic rod, a first film roll, a second film roll, a telescopic connecting rod and a film, and the telescopic rod is located above the supporting shell. The material consumption is obviously reduced, and the utilization rate of the material is improved.

The application provides a space refrigeration equipment vibration composite control system and method, comprising: a passive vibration isolation module and an actuation module; the passive vibration isolation module performs preliminary attenuation on the vibration of a compressor in the space refrigeration equipment in a passive vibration isolation mode, reduces the transmission of the vibration of the compressor from a base plate to a cold finger end detector; the actuation module performs attenuation on residual vibration transmitted from the compressor to the cold finger end through a connecting pipe in an active vibration suppression mode, so as to suppress the micro-vibration of the gas in the connecting pipe from the compressor to the cold finger end caused by uneven flow; and performs attenuation on residual vibration transmitted from the compressor to the cold finger end through the base plate in an active vibration suppression mode. The application is suitable for guaranteeing that the space vehicle load refrigeration equipment outputs minimal disturbance, and avoiding affecting the resolution and stability of the load.

Aspects presented herein may enable a UE to eliminate low accuracy “subset positions” and / or “all-in-view position” from a solution separation (SS) receiver autonomous integrity monitoring (RAIM) (SS RAIM) evaluation, thereby improving the computation efficiency for the SS RAIM. In one aspect, a UE estimates a plurality of positions of the UE using all space vehicles (SVs) in a list of SVs and a plurality of subsets of SVs in the list of SVs. The UE excludes a set of positions in the plurality of positions from an SS evaluation if at least one of: (1) a weighted sum of squared residual (WSS) of the set of positions is greater than a chi-square threshold, or (2) a computed residual of the set of positions indicates a negative value. The UE performs the SS evaluation based on rest of the plurality of positions.

The invention is suitable for the technical field of aerospace craft curved surface body machining, and provides a preparation method of an aerospace craft complex curved surface body thin-wall shell, and the preparation method comprises the steps of model modeling, digital-analog decomposition, mold manufacturing, plate selection, plate welding, appearance treatment, heat treatment and compression molding treatment, so that the hollow thin-wall shell is obtained. On one hand, the fitting degree of the two plates is better improved, so that the precision between the welded plates is guaranteed, and the pressure bearing performance of the curved surface body can be guaranteed while the machining difficulty is reduced; and on the other hand, curved surface welding can be changed into plane welding during production of the complex curved surface body, the problem that multiple pieces cannot be welded at a time during production and welding of existing complex curved surface bodies is effectively solved, a large number of materials are saved while the machining period is greatly shortened, meanwhile, batch production can be achieved, and the yield can be increased.