207 results about "Space vehicle" patented technology

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.

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 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.

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 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 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.

The invention provides a controllable distance approaching control method and system for the tail end of a spacecraft. The controllable distance approaching control method comprises the following steps: step 1, determining a sight line conversion rate corresponding to a controllable approaching distance; 2, correcting the sight line conversion rate according to the controllable approaching distance sight line conversion rate; and 3, according to the proportional guidance law, obtaining the required longitudinal channel and lateral overload instruction requirements from the corrected sight line rotation rate. According to the invention, the purpose of controllable distance approaching control of the tail end of the aircraft to the target aircraft is realized.

Space vehicles with paraglider re-entry, and associated systems and methods are disclosed. A representative system includes a re-useable space vehicle, a collapsible, deployable and re-stowable re-entry heat shield carried by the space vehicle, and a collapsible, deployable and re-stowable flexible paraglider wing also carried by the space vehicle. The space vehicle can accordingly carry out repeated space-based missions, and can be refurbished and replenished on Earth and / or at an orbiting dock between missions.

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.

The regional composite bearing retainer comprises a metal force bearing part and a porous oil storage part, the porous oil storage part is matched with a mechanical embedding groove structure through a connecting structure on the porous oil storage part to realize physical interlocking of the metal force bearing part and the porous oil storage part, and the bonding surface of the mechanical embedding groove structure of the metal force bearing part is subjected to surface roughening treatment and mechanical anchoring effect. According to the scheme, through function partition design and structural innovation, the technical problems that the oil storage capacity of a metal retainer is poor, the strength of a polymer retainer is insufficient, and a composite interface is prone to failure are solved. The retainer is formed by physically interlocking a high-strength alloy steel metal force bearing part and a gradient pore polyimide-based porous oil storage part through a mechanical embedding structure, and the oil storage part adopts porosity gradient distribution and a surface densification layer to synergistically optimize the oil storage capacity and directionally release a lubricant, and is matched with a vacuum impregnation and oleophobic modification process; emergency lubrication can be provided when lubricating oil is interrupted, the emergency lubrication operation time of the bearing is remarkably prolonged, and the bearing is suitable for extreme environments such as high-speed aero-engines and spacecrafts.

A first space vehicle (fig.1,101) has two or more latch arms 151 and a moving probe (fig.10,163) carried by a first valve (fig.10,160). The arms are rotated and translated to couple the spacecraft and

The application discloses a space vehicle heater control and state detection system, which comprises a safety switch, n overcurrent protection circuits, n heaters, n heater safety switches, n voltage division circuits, a multipath buffer / bus driver, an internal bus and a CPU module; wherein the safety switch is connected with the power supply positive terminals of the n heaters; the power supply positive terminal of each heater is connected with one overcurrent protection circuit; the power supply negative terminal of each heater is connected with the corresponding heater safety switch of each heater; a voltage division circuit is connected between the power supply negative terminal of each heater and the corresponding heater safety switch of each heater; the n voltage division circuits are connected with the multipath buffer / bus driver; and the multipath buffer / bus driver is connected with the CPU module through the internal bus. The application has the advantages of small volume, light weight, low cost and high reliability, and can work in a complex and severe space environment.

Independently moving deployment and positioning space vehicles that are configured to deploy and / or position structures are disclosed. The deployment and positioning vehicles work together to deploy and / or maintain the position / shape of a space structure. The deployment and positioning vehicles may include one or more thrusters, an attitude determination and control system (ADCS), a precision vehicle-to-vehicle location determination system, processing circuitry, etc. Two or more deployment and positioning vehicles are configured to coordinate the deployment and / or positioning of the space structure themselves or in concert with the other deployment and positioning vehicles in the deployment and positioning vehicle network, as well as achieve precision positioning.

The present invention belongs to the technical field of pulsar autonomous navigation for space vehicles, and provides a pulsar autonomous navigation pulse phase estimation system and method. Taking a satellite system with a distributed layout as a carrier, it includes a high-precision time synchronization timing system, a pulsar detection payload unit, a starlight vector sensitive payload unit, a Earth sensitive payload unit, and an inter-satellite link payload unit. The present invention also proposes a new fusion method for pulse observation profiles and starlight angular distances. This method is a deep fusion method for pulsars and starlight angular distances, which can effectively utilize various types of satellites with inter-satellite links, carry payloads with the functions required by the present invention, and achieve pulse phase estimation. By introducing multiple starlight angular distance measurement values at the same moment, the present invention can achieve bin filtering of profile amplitudes and realize high-precision pulse phase fusion estimation.