42 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.

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 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.

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.

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.

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.

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.

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.

This invention relates to the field of space deployable structure technology, and more particularly to a deployable light-shielding structure based on shape memory materials and its shaping and deployment method. The structure comprises an outer frame and an inner light-shielding cover. The outer frame consists of a top frame, a bottom frame, and multiple shape memory L-shaped beams connecting the two, forming a frustum-shaped boundary after unfolding. The number of inner light-shielding covers matches the sides of the frame; each includes a fan-shaped shape memory skeleton with concentric annular curved lines and a light-shielding cloth. The skeleton can be folded along the curved lines into a radially corrugated laminated ring shell to achieve ultra-high axial compression, and after heating, it can be deterministically unfolded into a planar fan-shaped ring to fill the boundary. This invention adopts a two-stage unfolding method of "boundary first, then filling." This structure integrates the geometric determinism, high folding-to-unfold ratio, and high energy storage characteristics of curved origami with the active controllability of shape memory materials, possessing advantages such as lightweight, high reliability, high precision, smooth and impact-free unfolding, and customizable structure.

Underwater Deployable Structures (UDSs) capable of achieving inflation via mechanical pumps, moisture-activated expanding foams, water swelling material, or a hybrid of a mechanical pump and water swelling material. Moisture expanding foams begin as polyurethane-based resins of low viscosity, and react in the presence of water to become solid foam. In their final forms, the foams exist as open- or closed-cell, and vary in strength, elasticity, and rigidity. Water swelling material is also disclosed wherein the water swelling material is capable of achieving expansion numerous (e.g., up to about 250) times its initial volume when in contact with water and can shrink back to its initial volume upon dehydration, making the UDSs reusable. A mechanical pump may optionally be used to assist in initial inflation until the water swelling material reaches full expansion.

A space unfolding and folding device with a shape self-sensing function belongs to the field of space unfolding structures and comprises pod rod structure unfolding and folding equipment and shape sensing equipment. The pod rod structure unfolding and folding equipment comprises an outer frame, a motor, a guide mechanism and a pod rod structure; the outer frame is of a cuboid structure, and the motor is fixed to the middle of the inner wall of the frame. One end of the pod rod structure is fixedly connected with the motor, and the other part is wound on the motor; the other end of the pod rod structure penetrates through the guide mechanism and extends out of the outer frame; the shape sensing equipment is attached to the outer surface of the pod rod structure; data collected by the shape sensing device is transmitted to the upper computer, and the upper computer analyzes the data and then outputs real-time shape data of the structure. The guide mechanism is a guide sleeve with a linear bearing and is installed at the position of an outer outlet of the pod rod structure to restrain the unfolding and folding movement direction of the pod rod structure. According to the invention, the shape sensing equipment and the space deployable structure are integrated, so that accurate measurement of the shape in the service process of the space structure is realized.

The invention belongs to the technical field of spaceflight equipment, and particularly relates to a spatial large-scale secondary variable-dimension deployable light shield. According to the technical scheme, the device comprises a central hexagonal prism frame, circumferential unfolding mechanisms, axial unfolding mechanisms, correcting devices, circumferential tensioning cables and a cover film, the six circumferential unfolding mechanisms are evenly distributed around the central hexagonal prism frame, and each circumferential unfolding mechanism is connected with one axial unfolding mechanism through one correcting device; the six circumferential unfolding mechanisms are connected through a plurality of circumferential tensioning cables, the six axial unfolding mechanisms are connected through a plurality of circumferential tensioning cables, and the cover film is arranged outside the six circumferential unfolding mechanisms and the six axial unfolding mechanisms. Secondary variable-dimension unfolding of the whole body is achieved through the sequence that two-dimensional unfolding is conducted through the circumferential unfolding mechanism and then one-dimensional unfolding is conducted through the axial unfolding mechanism. According to the invention, the development requirements of large-scale and high-resolution optical equipment, stray light resistance and thermal deformation prevention of an optical system, light weight of a deployable structure and high storage rate can be met.

The application discloses a winding rib type electrostatic forming film reflecting surface antenna device, which comprises a center wheel cylinder, the center wheel cylinder is a prism, and a plurality of elastic winding ribs are fixedly connected to the side surface of the center wheel cylinder; a flexible cable is arranged between two adjacent elastic winding ribs; a base cable net electrode is laid above the elastic winding rib and the flexible cable; the base cable net electrode is composed of triangular meshes; a film reflecting surface is arranged above the base cable net electrode; the film reflecting surface is in a parabolic shape; the base cable net electrode and the film reflecting surface are connected with an electrostatic power supply through a cable; and the film reflecting surface is connected with the center wheel cylinder. The application adopts an elastic winding rib type deployable structure as a support structure of the whole antenna system. The winding rib type structure can be folded and unfolded, and only occupies a small space when being folded, and the folding volume is significantly reduced.

A dynamic driving optimization method for deployable serial-parallel mechanisms in space is proposed, belonging to the field of spacecraft structural optimization design. Addressing the challenges of multi-scale high-fidelity modeling, integration of transient dynamic simulation and optimization design, lack of multi-objective collaborative optimization mechanisms in driving strategies, and insufficient model parameter calibration and robustness in existing deployable structures, this method constructs a high-fidelity dynamic model incorporating coil spring drive, rope linkage, and locking constraints. Parameter calibration is performed using experimental data, and a multi-constraint, multi-objective optimization model is established. The driving parameters are externalized, and the optimization algorithm is linked with the dynamic solver via an interface. This automatically performs parameter updates, transient simulations, and performance evaluations, iteratively obtaining the optimal set of driving parameters. Performance trade-off analysis and verification are conducted based on the optimization results. This method is applicable to the deployment driving design and performance optimization of large-scale deployable space structures.

The invention discloses a construction method of a two-dimensional three-way woven composite material in-plane mechanical isotropy size effect prediction model, and the method discusses the size effect of the tensile property of a TWF system in different loading directions. An elastic mechanics theory and analytic geometry are combined, a new mechanical model is established, and the relation among the sample size, the tensile modulus and the loading angle is quantitatively described. The model clarifies how the local fiber bundle constraint and the reinforcement effect of the woven structure influence the macromechanical behavior, theoretical derivation reveals the size effect of the three-stage mechanical response, and two key aspect ratio thresholds are delimited. The model constructed by the method can overcome the limitation that a typical volume element analysis method neglects a boundary effect, and a theoretical tool is provided for material and size coordination design of spaceflight deployable structures such as space deployable antennas.

An antenna array has a plurality of square or rectangular antenna assemblies. Each assembly includes a first antenna assembly surface with a solar cell and a second antenna assembly with one or more antenna elements. The antenna assemblies are interconnected without gaps therebetween to form a first contiguous array surface comprised of the first antenna assembly surfaces and a second contiguous array surface comprised of the second antenna assembly surfaces. The antenna assemblies are connected together by mechanically stored-energy connectors, such as spring tape, that self-deploy the array in space without the use of electric energy.

The invention provides a spacecraft rope net structure system based on spinning stability and a design optimization method, and relates to the technical field of spacecraft large flexible space deployable structures. According to the method, a spacecraft ultra-large space deployable parabolic antenna serves as a research object, dynamic stable balance is achieved through centrifugal force generated by spinning of an antenna structure system, parametric modeling is achieved for design parameters of the antenna structure system, then optimization design is conducted on the structure design parameters subjected to parameterization processing, and the design precision of the antenna structure system is improved. And obtaining key factors influencing reliable unfolding and spinning stable balance of the antenna. Based on the obtained physical quantity characteristics of the key factors, the probability distribution criterion of the key factors is determined, random targeting analysis is carried out, the weight of the key factors on the dynamic stability influence is evaluated, and the optimization design problem of the spinning antenna structure system and the design problem that ground effective test verification cannot be carried out are solved.

The present application relates to the field of structural health detection, and particularly relates to a device abnormality detection method for wind-induced deformation measurement of large deployable structures, comprising the following steps: obtaining time series response data of a healthy structure under wind load excitation to form baseline reference data; and normalizing the data; combining empirical mode decomposition and Hurst index analysis to carry out signal denoising; establishing a Hankel matrix of each channel data based on the denoised data; calculating the orthogonal projection vector of the matrix; obtaining a state matrix and an output matrix of the large deployable structure through matrix decomposition; and calculating a reference mode matrix of the structure in a healthy state; in the service process of the structure, data acquisition, preprocessing and signal denoising are carried out in the same way as the above steps to obtain a current mode matrix corresponding to the service state; the distance between the current mode matrix and the reference mode matrix is calculated; the damage position of the deployable structure is determined by comparing the distance changes of each order; and the abnormality detection of the structure is realized.

A deployable structure includes a reconfigurable member formed from a plurality of elements defining a first end, a second end that is opposite the first end, a first side extending between the first end and the second end, and a second side, opposite the first side, extending between the first end and the second end. A first inflatable rigidizing member extends along the first side across each of the plurality of elements and a second inflatable rigidizing member extends along the second side across each of the plurality of elements.

This invention discloses a dual-motor driven bidirectional deployable truss structure. The dual-motor driven bidirectional deployable truss includes two motors, with multiple bidirectional deployable structural units arranged between the two motors, connected in series to form a closed loop. Each motor is connected to one bidirectional deployable structural unit via a winding mechanism, and the two motors are respectively installed at both ends of the deployable truss. This truss can effectively improve the stiffness and deployment-to-retraction ratio of the truss structure. Furthermore, the dual-motor drive, installed at both ends of the deployable truss, can effectively increase the fundamental frequency of ultra-large aperture antennas and enhance the stability of the entire structure.