75 results about "Deployable structure" patented technology

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

An apparatus includes a structure configured to receive thermal energy and to reject the thermal energy into an external environment. The structure includes a lid and a body. The structure also includes (i) multiple inline and interconnected thermomechanical regions and (ii) one or more oscillating heat pipes embedded in at least some of the thermomechanical regions. Different portions of at least one of the lid and the body form the thermomechanical regions. The one or more oscillating heat pipes are configured to transfer the thermal energy between different ones of the thermomechanical regions. At least one of the thermomechanical regions includes one or more shape-memory materials configured to cause a shape of the structure to change. Each of the one or more oscillating heat pipes includes at least one channel in the structure.

The invention relates to the technical field of aerospace, in particular to a multi-objective optimization method for a pod-shaped composite space thin-wall deployable structure. The method comprises the following steps: selecting an optimal sample point for uniformizing the space deployable structure in an interval range of structure parameters; constructing a finite element analysis model of the pod-shaped composite material space thin-wall deployable structure; the method comprises the following steps: by taking a damage failure criterion as a failure mode judgment basis of fibers and a matrix of a pod-shaped composite material, describing an evolution path of material rigidity in a damage development process by utilizing a bilinear rigidity degradation model after failure occurs, and constructing a random forest model based on adaptive particle swarm optimization as an agent model; and a Pareto leading edge optimal solution is obtained by calling the proxy model and is used as an optimal pod rod structure. In this way, the axial bearing capacity and folding performance of the unfolded structure are comprehensively improved, and the optimized structure can show more excellent mechanical performance and adaptability to different scenes in a complex space environment.

The invention belongs to the technical field of space energy, and particularly relates to a pod rod driving type solar wing capable of being wound, unfolded and folded. Comprising a solar blanket winding shaft assembly, unfolding driving devices, a composite material deployable structure, a perovskite solar blanket and a base assembly, the two ends of the perovskite solar blanket in the length direction are connected with the solar blanket winding shaft assembly and the base assembly respectively, and the two ends of the solar blanket winding shaft assembly are connected with the two unfolding driving devices respectively; the two groups of unfolding driving devices are connected with the base assembly through two composite material deployable structures located on two sides of the perovskite solar blanket in the width direction, and the unfolding driving devices are used for driving the perovskite solar blanket and the composite material deployable structures to be synchronously unfolded and wound; and the composite material deployable structure provides support for the unfolded perovskite solar blanket. Compact storage and efficient unfolding of the solar wing are achieved, the solar wing can be unfolded and rolled repeatedly, and the solar wing is suitable for complex and diversified aerospace mission requirements and has both flexibility and practicability.

A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.

The invention discloses a sectional driving method of a Miura paper folding derived space deployable structure, which comprises the following steps of: setting a corresponding driving rule, namely applying driving to a hinge component at the central position of a system or a plurality of hinge components in the same folding and unfolding direction by adopting a sine type driving function and a multi-point central driving mode; for the folding process, according to a deduced segment driving equation, the position of a driving segment is solved through a Newton-Raphson method, the whole folding process is divided into two driving stages, in the first stage, driving is applied to a connecting hinge in the longitudinal folding and unfolding direction, and after the connecting hinge rotates to the optimal driving segment position, the second stage is started; in the second stage, driving is applied to a connecting hinge in the transverse folding and unfolding direction, and the folding process is completed; for the unfolding process, the driving application sequence is opposite to that of the unfolding process. The invention provides a special driving path of the Miura paper folding derived deployable structure, under the condition of the same deploying and retracting time, the more stable and complete deploying and retracting process is realized, and the reliability and stability of the structure are improved.

This invention belongs to the field of aerospace composite material manufacturing technology, specifically relating to a method and equipment for forming thermoplastic composite strain energy rods. First, prepreg is laid up to form a preform, which is then pressed by a flatbed pre-pressing device to form a preform sheet intermediate. Subsequently, the preform intermediate is fed into a hot-pressing module with a central heating and two-end cooling gradient temperature field. It undergoes melting and molding in a high-temperature zone and controlled cooling in a low-temperature zone to achieve final shaping. Finally, the formed rod is cooled and shaped, then pulled out at a uniform speed by a traction device and coiled and collected by a winding device, achieving continuous production. The hot-pressing module of the supporting hot-pressing equipment is equipped with a zoned heating and integrated cooling system, which can precisely construct a gradient temperature field and control the cooling rate. This invention, by using gradient temperature control to precisely regulate the material's crystallization behavior and release internal stress, effectively ensures the straightness, dimensional stability, and consistent deployment performance of the strain energy rod, perfectly meeting the manufacturing requirements of strain energy rods for large deployable structures in spacecraft.

The present invention discloses a simulation analysis method for an automatic retraction structure of a double elastic rod. The elastic rod is an arc-shaped shell structure with a certain curvature and is often used as the main supporting and load-bearing component of a large spacecraft structure. The elastic rod undergoes a winding and retraction process under the action of an external load. After the locking device is opened, the elastic rod can gradually rebound and unfold by releasing the strain energy generated by bending. The principle is simple, and it is a space deployable mechanism that has been widely used. This method uses the ABAQUS finite element simulation software to assemble according to the actual spatial position of the deployable structure of the elastic rod, considers the interaction between various parts of the structure and the actual contact situation, applies a load to realize the working winding process, and selects the explicit dynamics algorithm for solution in the analysis step. A finite element model of the double elastic rod winding is established using the S4R shell element and mechanical simulation analysis is carried out. The present invention can provide a reference for the cross-section design of the winding elastic rod and the design of the system working prototype.

The application provides a deployable speed on-orbit adjustable damping deployable locking joint and a spacecraft, and relates to the field of spacecraft structures and mechanisms, and comprises a first double-wing driving constant-torque deployable hinge, a rotating damping mechanism and a second double-wing driving constant-torque deployable hinge, wherein the first double-wing driving constant-torque deployable hinge, the rotating damping mechanism and the second double-wing driving constant-torque deployable hinge are connected to a deployable structure respectively, the spherical rotating pair of the first double-wing driving constant-torque deployable hinge and the second double-wing driving constant-torque deployable hinge forms a deployable locking joint rotating shaft, and the deployable locking joint rotating shaft is coaxial with the rotating shaft of the rotating damping mechanism. The deployable speed on-orbit adjustable damping deployable locking joint has the beneficial effects of system simplification, deployable speed on-orbit adjustability, small locking impact and the like, and has better application adaptability and expansibility.

Deployable structures are described, in particular linearly-deployable structures, such as masts or booms. The masts may be stowed for transport and then deployed at their destination in space or on earth. A deployment system includes a storage reel storing a stowed elongate band. A drive mechanism biases and guides the band helically out of the storage reel to form an elongated mast. Adjacent edges of the deployed band may secure together using openings and corresponding protrusions, such as rivets. A welding system may use a rotating welder to weld adjacent edges of the band as it deploys. The band may be formed of multiple band segments attached together by connectors such as doublers. Protrusions such as rivets or other fasteners may attach the connectors to opposing sides of the band segments. A cylindrical space habitat or other macrostructure may be formed using multiple deployable masts that connect large rings.

The invention provides a foldable antenna and solar wing unfolding device based on a scissor hinge structure. The foldable antenna and solar wing unfolding device comprises a scissor hinge unfolding structure, a foldable antenna panel, a star fixing end and a solar wing, the scissor hinge unfolding structure is a continuous structure formed by repeating the same units, and the number of the units can be adjusted according to the actual length requirement of the antenna; the foldable antenna panel is composed of a plurality of daughter boards and is connected with the scissor hinge unfolding structure; the star body fixing end is connected with the scissor hinge unfolding structure and the foldable antenna panel; the solar wing is connected with the scissor hinge unfolding structure; based on the expandable structure of the scissor hinge, the antenna plate and the solar wing are integrated, and the design that multiple functions are achieved through a single structure is achieved.

The invention relates to the technical field of discharge defect early warning, in particular to a major discharge defect early warning device which comprises a transverse base, and an electric control type assembly housing is movably assembled on the lower surface of the transverse base. According to the early warning device applied to the major discharge defect, the design of an openable structure is adopted, so that the device can be quickly assembled and disassembled on a cable, and the operation is very simple and convenient; the optical detection module is mounted on the inner wall of the electric control type assembly housing through the embedded sliding seat, an annular guide rail can be formed when the electric control type assembly housing is turned over downwards to be closed, then the embedded sliding seat is controlled to rotate by utilizing an electromagnetic driving principle, surrounding type optical detection on a power supply is facilitated, and the detection efficiency is improved. The detection range is wider; the end electric drive wheel set and the electric control spiral lifting module matched with the overturning type lateral adjusting arm are assembled on the transverse base, so that the equipment can conveniently operate in a narrow channel or high altitude, and the application range is wider.

This invention relates to an auxiliary folding device for a push-plate type flexible deployable structure, comprising an upper push-plate assembly, a lower push-plate assembly, and a straightening and limiting assembly. The upper push-plate assembly is located on the upper center of the lower box plate of the flexible deployable structure, and the lower push-plate assembly is located on the lower center of the lower box plate. The straightening and limiting assembly is located on both sides of the upper side of the lower box plate, on either side of the upper push-plate assembly. The lower push-plate is used to lift the solar panel in the folding direction, the upper push-plate is used to store the folded solar panel, and the straightening and limiting assembly is used to ensure the straightening and limiting of the solar panel during the folding process, preventing overall deformation or misalignment of the solar panel that would hinder its folding. This invention provides an efficient, convenient, versatile, and environmentally friendly auxiliary folding and retraction device for flexible deployable structures, solving the technical bottlenecks and operational risks of low efficiency and insufficient safety in existing flexible deployable structure folding technologies.

The application discloses a sliding cable driven crescent-shaped deployable cable net structure in the technical field of deployable structures and cable net structures, which comprises an inner ring end fixed node, an inner ring free node, an outer ring fixed node and three cables; the inner ring end fixed node and the inner ring free node are located on a first circular arc line; the outer ring fixed nodes are uniformly fixed on second and third circular arc lines and are symmetric about the plane where the first circular arc line is located and the normal plane of the line connecting the inner ring end nodes; a first cable passes through the nodes on the first circular arc line in sequence, a second cable alternately passes through the nodes on the second circular arc line and the nodes on the first circular arc line, and a third cable alternately passes through the nodes on the third circular arc line and the nodes on the first circular arc line; the lengths of the three cables are adjustable, the folding and unfolding of the cable net structure are realized, different building roof opening requirements are met, the structure composition and control mode are simple, and the structure has a high storage ratio.

The invention discloses a spacecraft heat balance dynamic simulation evaluation method, and particularly relates to the technical field of spacecraft heat balance. Constructing a spacecraft kinematics model with an expandable structure and a thermal network model; establishing a thermal-structure-motion multi-field coupling interface, and realizing dynamic driving of a structure deployment state on thermal connection parameters; in the joint simulation process, the thermal coupling matrix is updated in real time, so that the thermal connection state continuously changes along with structure deployment; key thermal node temperature sequence data are recorded, thermal response characteristics of the key thermal node temperature sequence data in the whole deployment process are evaluated, and the problems of temperature anomaly and thermal instability are identified; reversely positioning a potential thermal control defect area and a deployment strategy risk point by combining the deployment state and the thermal connection state change, and outputting a thermal response map and a design optimization suggestion; according to the method, the accuracy and responsiveness of thermal simulation in the structure deployment process are improved, thermal connection dynamic modeling and risk diagnosis in the deployment process are achieved, and the method is suitable for thermal control design and thermal safety evaluation of the spacecraft with the deployable structure.

The invention discloses a flexible winding structure unfolding dynamics modeling analysis method based on multiple scales. The method comprises the following steps: S1, establishing geometric models of a spacecraft flexible winding structure and a heterogeneous flexible buckle structure; s2, based on the parameterized geometric model established in S1, establishing a kinetic model of the heterogeneous flexible buckle structure; s3, acquiring a change curve of load output along with the influence of the heterogeneous flexible buckle parameters; s4, establishing an equivalent mechanical model of the heterogeneous flexible buckle structure; s5, on the basis of the equivalent mechanical model established in the step S4, the elastic thin-wall rods and the flexible thin-film structure are combined to be assembled to form an overall dynamic analysis model of the flexible deployable structure; s6, performing influence evaluation on the key factor parameters of the heterogeneous flexible buckle structure; and S7, on the basis of an influence evaluation result, comparing the influence evaluation result with a specified index parameter. The method is applied to dynamic modeling and analysis of the large flexible space deployable structure of the spacecraft, and design and development of a large flexible space deployable structure system are effectively guided.

The application provides a method for predicting the whole satellite dynamics characteristics of a deployable structure during on-orbit deployment, which comprises the following steps: establishing a flexible dynamics equation of the whole satellite and a flexible multi-body dynamics equation set of the deployment process of the deployable structure; performing one-step numerical integration on the flexible multi-body dynamics equation set; replacing the hinges of the deployable structure which are deployed to the position or reach the preset jam with six-degree-of-freedom force element connections according to the integration results, locking the jammed hinges and the corresponding connection components, and establishing a multi-body dynamics structure dynamics model of the mixed semi-locking and semi-deployment; eliminating the rigid body mode through modal analysis to obtain a deployable structure dynamics model; and obtaining a semi-locking and semi-deployment dynamics model of the satellite during on-orbit deployment based on the deployable structure dynamics model and the flexible dynamics equation of the whole satellite; thus, the application can solve the whole satellite attitude stabilization control input and deployment failure disposal whole satellite model disposal problems caused by the locking during the on-orbit deployment of the satellite-borne deployable structure.

The invention relates to the technical field of power grid maintenance tools, in particular to a power distribution 10 kilovolt safety insulation baffle. Comprising a fixing belt, the fixing belt is fixedly connected with a plurality of fixing shells which are evenly distributed, the fixing shells are fixedly connected with supporting rods, the supporting rods are jointly and fixedly connected with flexible cloth, and the flexible cloth is used for forming an annular insulation barrier surrounding the telegraph pole when the supporting rods are unfolded. And the fixed belt is fixedly connected with a detection module. According to the scheme, through the flexible cloth of the flexible deployable structure and the supporting rods of the umbrella rib type supporting mechanism, a complete annular insulating barrier is formed after unfolding, all-directional protection is ensured, the overall weight is greatly reduced, the storage size is reduced, complex tools are not needed during installation, and fixing and closing can be rapidly completed by a single person; the high-altitude operation efficiency and safety are remarkably improved, and the humidity, the temperature and the wind speed of the external environment can be monitored in real time through the detection module so that the safety of the working environment can be guaranteed.

The invention relates to the field of folding structures, and particularly discloses an easy-to-unfold structure, a folding leg frame, a storage box frame and a table, the easy-to-unfold structure comprises a frame component, a supporting rod rotatably installed on the frame component and an elastic piece arranged between the supporting rod and the frame component, the frame component is provided with an inserting part, the inserting part is provided with an inserting opening, and the folding leg frame and the storage box frame are connected through the inserting opening. One end of the supporting rod is an insertion end, and the insertion end is suitable for being inserted into the insertion port of the insertion part; a supporting guide part is arranged on the side face of the inserting part, and a sliding guide part corresponding to the supporting guide part is arranged at the inserting end. In the conversion process of the supporting rod from the folded state to the unfolded state, the insertion end of the supporting rod and the insertion part of the frame component do not need to be in manual butt joint, under the pulling force effect of the elastic piece, through sliding fit of the sliding guide part and the supporting guide part, after the supporting rod is unfolded, the insertion end of the supporting rod can be automatically aligned with the insertion part; the supporting rod and the frame component are convenient and efficient to assemble.

The invention discloses a modular directional extensible structure based on temperature control and a use method thereof. The modular directional extensible structure comprises a top plate; the plurality of single-layer deformation structures are connected with the top plate after being overlapped; the single-layer deformation structure comprises two supporting frames which are tightly attached to each other and are arranged in a straight line. A first arch structure mounted on one of the support frames; the second arched structure is mounted on the other supporting frame; and the supporting rod is arranged on the first arch structure and the second arch structure. The bi-material arch serves as a core deformation area and is composed of an elastomer and a shape memory polymer. The material combination endows the extensible structure with the capacity of achieving shape conversion and space steering according to temperature changes, and possibility is provided for multifunctional operation of the extensible structure in a complex environment.

A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.

The invention provides a design method for an expandable structure of a paper-cut metamaterial, and belongs to the technical field of paper-cut metamaterials. The specific design steps comprise: S1, determining basic information of the paper-cut metamaterial, wherein the basic information comprises a closed configuration, an unfolded configuration, shapes of paper-cut units, and a connection mode between the paper-cut units; s2, a closed contour is determined based on the closed configuration, an unfolded contour is determined based on the unfolded configuration, paper-cut units are drawn in the unfolded contour, the multiple paper-cut units are arranged according to a preset mode to form a paper-cut pattern, and the area of the closed contour is the same as that of the unfolded contour; s3, constructing area-preserving transformation mapping from the unfolded contour to the closed contour so as to map the paper-cut pattern in the unfolded contour into the closed contour; and S4, based on a connection mode between the paper-cut units, adding a connection design between the paper-cut units on the paper-cut pattern in the closed contour so as to determine a to-be-cut side line in the paper-cut pattern. The design method provided by the invention is simple, rapid and convenient to operate.

The present invention belongs to the technical field of structural member strain monitoring, and relates to a layout optimization method for deformation monitoring of large deployable structures, including the following steps: establishing a finite element model for the large deployable structure; carrying out theoretical modal analysis of the large deployable structure for the finite element model, uniformly setting dense sensing measurement points on the surface of the structure, and constructing an error compensation model according to the strain transfer efficiency of the fiber optic sensor; obtaining the corresponding mass matrix [M] and stiffness matrix [K] at all measurement points, and obtaining the modal vibration modes of each measurement point and the vibration mode matrix of the entire system; calculating the orthogonality of each order of modes of each multi-degree-of-freedom structure according to the vibration mode matrix; realizing the update of the modal orthogonality fitness function through selection, crossover, and mutation steps to optimize the position of the sensor measurement points; and optimizing the position of the sensing measurement points. After determining the measurement point positions, a fitness function is constructed with the goal of minimizing the spacing between all sensing measurement points to obtain the final serial connection order of the fiber optic measurement points. The method of the present invention has less operation time.

A robotic surface cleaning device is described. The robotic surface cleaning device can be configured to traverse a floor surface in a room area to disinfect various surfaces, as well as disinfect the air in the room. The robotic surface cleaning device may include a maneuvering base. The robotic surface cleaning device may also include a support tower attached to the maneuvering base and at least one deployable structure attached to the support tower, wherein the at least one deployable structure is capable of extending and retracting to pass over surfaces in the room area. The at least one deployable structure may include one or more ultraviolet (UV) light sources configured to disinfect surfaces in the room area.The support tower may include at least one air intake vent configured to draw in room air, at least one air sanitizing unit configured with at least one array of UV light-emitting diodes (LEDs), and at least one exhaust vent configured to emit sanitized air.

Methods for constructing and / or deploying / retracting structures for space deployable applications are provided. Deployable structures may include retractable, inflatable and / or self-deploying structural tubes as part or all of the infrastructure of the system for use in space deployable applications. The structural tubes herein may be rolled and / or bent and may be configured to transition from a stored state to a deployed state. The deployed state may be an elongated state to be used as a structural support.

The invention discloses a fast deployable structure utilizing continuous jump buckling and a using method thereof. The extensible structure comprises a top plate and a bottom plate; the deformation mechanism is arranged between the top plate and the bottom plate; the deformation mechanism comprises a supporting center; the plurality of supporting rods are arranged on the supporting center, and the plurality of supporting rods are arranged on the same plane; and the spring steel sheets are arranged at the outer ends of the two adjacent supporting rods. The spring steel sheets are used and combined with the angle-adjustable supporting structure, so that after the structure is unfolded, the height of the structure is increased sufficiently (the height of the unfolded structure is 2.74 times that of the unfolded structure), and the overall stability and the bearing capacity of the structure are effectively guaranteed. Due to the high elasticity and fatigue resistance of the spring steel, the structure is not prone to fatigue damage in the multiple unfolding and shrinking processes, and the service life is long.

The application discloses a bidirectional expansion thick plate paper-cut array deployable structure, which meets folding conditions and does not generate motion interference, and comprises a first thick plate paper-cut array and a scissors driving mechanism matched with the first thick plate paper-cut array to realize unfolding and folding of the first thick plate paper-cut array; the first thick plate paper-cut array is composed of eight first-type basic folding and unfolding units and 52 plate surfaces connected through hinge hinges; the hinge hinges are bolted with the plates; eight plate surfaces form a first-type basic folding and unfolding unit; the first-type basic folding and unfolding unit is networked and constructed into an array through longitudinal and transverse expansion. The thick plate paper-cut array is coupled with the scissors driving mechanism to realize stable unfolding of the thick plate paper-cut array, and solves the problems of small folding and unfolding ratio, uneven working surface, limited working area and single unfolding form of the existing deployable structure.

A simulation device for microgravity deployment experiments on the ground is mainly composed of an adjustable leveling structure assembly, an adjustable anti-creep rope assembly, a magnetic levitation de-rotation structure assembly, an adjustable counterweight assembly, and a safety protection rope lock. This device significantly reduces resistance during operation and improves redundancy safety and reliability during microgravity deployment experiments. The invention features simple structure, low rotational resistance, high suspension compensation accuracy, strong versatility and design flexibility, high redundancy safety protection, and good scalability, making it suitable for deployment experiments of large-scale space deployable structures such as space deployable antennas and flexible deployable cabins.

The application relates to a thermal-mechanical coupling dynamic modeling method for a gap-hinged metamorphic spacecraft antenna, and relates to the technical field of spacecraft deployable structure dynamic modeling. The existing model cannot accurately and comprehensively obtain the time history response of the antenna during the antenna deployment process.According to a constructed temperature field interpolation model, a transient heat conduction equation, a temperature-related material parameter model, a gap motion pair contact criterion model, a normal contact collision force model, a tangential friction and stick-slip conversion model, a temperature-related mass matrix and stiffness matrix or generalized damping matrix, a thermal-structure-gap contact coupling dynamic model of a gap-hinged metamorphic spacecraft antenna under extreme thermal environment is formed, the model is solved by using an HHT-alpha implicit integration and Newton-Raphson iteration algorithm, and the time history response of the antenna during the antenna deployment process is obtained.The application is used for establishing a thermal-mechanical coupling dynamic model.