154 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 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 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 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 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 invention provides an anti-radiation high-entropy alloy nitride hard film and a preparation method thereof, the molecular formula of high-entropy alloy nitride in the hard film is TiaCrbSicWdTaeGdfNx, a = 8-15, b = 8-15, c = 6-12, d = 6-12, e = 6-8, f = 6-8, x = 30-50, and a + b + c + d + e + f + x = 100. The preparation method comprises the following steps: preparing a pure Ti target material and a TiCrSiWTaGd high-entropy alloy target material; preparing a substrate and pretreating to a to-be-coated state; performing sputter coating on the substrate by using a Ti target material to obtain a TiN transition layer on the surface of the substrate; and a TiCrSiWTaGd high-entropy alloy target material is used for conducting sputtering coating on the TiN transition layer, and the high-entropy alloy nitride hard thin film is obtained. The problem that radiation resistance, high hardness, high wear resistance and high stability of the spacecraft material surface cannot be considered at the same time is solved, and the requirement for radiation protection of the spacecraft surface is met.

The invention discloses a non-singular tandem pyramid configuration control moment gyroscope and a design method thereof, and belongs to the technical field of control moment gyroscope design, the design method comprises the following steps: step 1, establishing a tandem control moment gyroscope system of a pyramid configuration; step 2, establishing attitude dynamics models of the tandem control moment gyroscope and the spacecraft in the tandem control moment gyroscope system with pyramid configuration; 3, performing singular analysis on the torque output model of the tandem control torque gyro; 4, carrying out anti-saturation characteristic analysis on the torque output model of the tandem control torque gyro; 5, designing the control rate of the tandem control moment gyroscope with the pyramid configuration; according to the method, by increasing the degree of freedom, the designed non-singular tandem pyramid configuration control moment gyroscope can effectively avoid singularity under the condition of minimum energy consumption, and has the characteristics of simple system implementation, low power consumption, high efficiency and saturation resistance.

The invention provides a double-rib cylindrical metal diaphragm storage tank which comprises a storage tank shell and a metal diaphragm, an air cavity is formed between the storage tank shell and the periphery of the metal diaphragm, and a liquid cavity is formed between the storage tank shell and the inner periphery of the metal diaphragm; the metal diaphragm comprises a connecting section, a bent edge section, a conical section, a transition section, a column section and a ball section which are sequentially connected, one end of the connecting section is fixedly connected with the inner wall of the storage tank shell, the other end of the connecting section is fixedly connected with the bent edge section, and a double-rib structure is arranged on the column section; the metal diaphragm is sequentially turned over from the bent edge section, the conical section, the transition section, the column section and the ball section under the action of the pressure of the inflated air cavity to continuously discharge liquid in the liquid cavity, the minimum bent edge radius R2 ranges from 0.7 R1 to 1R1 in the turning process of the metal diaphragm, and R1 is the initial bent edge radius. The problem that the cylindrical metal diaphragm cannot be effectively overturned is solved, the application scene of the metal diaphragm storage box is expanded, and layout optimization of a spacecraft is facilitated.

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

A system and method for the liquefaction and densification of oxygen for use in space vehicle applications is provided that uses high pressure air or synthetic air as the refrigerant source. The disclosed system and method employs a heat exchanger arrangement comprising a first heat exchange device configured to liquefy the high pressure gaseous oxygen stream and at least a portion of the high pressure gaseous air stream via indirect heat exchange with a refrigerant stream to yield a liquid oxygen stream and a liquid air stream. The heat exchanger arrangement also includes a second heat exchange device configured to densify the liquified oxygen stream via indirect heat exchange with the liquid air stream which yields the densified liquid oxygen and a cold vaporized air stream. The refrigerant stream comprises a mixture of the exhaust streams from one or more turbines with the cold vaporized air stream.

The utility model discloses a narrow space vehicle auxiliary reversing device which comprises a rotary table (1), a chassis (2) arranged below the rotary table (1), a thrust bearing (3) arranged between the rotary table (1) and the chassis (2), a gear motor (4) arranged at the bottom of the chassis (2), a driving shaft (5) upwards penetrating through the chassis (2) arranged at the output end of the gear motor (4), and an upper friction ring (6) arranged at the bottom of the rotary table (1). A lower friction ring (7) in sliding connection with the driving shaft (5) is arranged at the bottom of the upper friction ring (6); a pressure spring (8) is arranged between the lower friction ring (7) and the chassis (2); the driving shaft (5) is a square shaft, and a square hole (16) matched with the driving shaft (5) is formed in the lower friction ring (7). The utility model has the advantage that faults are not easy to occur.

The invention relates to an in-situ test device and method for simulating space multi-temperature-range environment irradiation, and belongs to the technical field of material tests.The device comprises an annular pipe, a sample table, a refrigerator, a heater and an irradiation source, the sample table is arranged in the annular pipe, the sample table can periodically move in the annular pipe, and the sample table can also rotate; according to the invention, a multi-temperature-zone design is adopted, and conversion of different temperature zones is realized by using a periodic rotation orbit, so that the temperature field effect of the spacecraft is simulated; and the irradiation uniformity of the ground point radiation source is realized by utilizing the autorotation of the irradiated sample, so that the authenticity and effectiveness of the ground test are further improved. Ground equipment is utilized, the method is simple, the cost is low, in-situ test simulation of various space environments such as vacuum, high and low temperature, irradiation and the like can be realized, and the method has important application value for improving selection and reliability tests of space materials.