51 results about "Variable geometry" patented technology

The invention discloses a point cloud neural network aerodynamic heat rapid prediction method considering a variable geometric shape, and belongs to the field of aerodynamics, and the method comprises the steps: determining a research object, and constructing a data set; then, a three-dimensional point cloud neural network is adopted to extract global shape features, local shape features and incoming flow condition features of different variable geometric shapes under the same basic configuration to serve as point cloud neural network input parameters, and the same model is adopted to predict wall surface pressure and heat flux density; wherein the training data of the three-dimensional point cloud neural network adopts the wall surface heat flux density and pressure under the non-uniform wall temperature boundary condition. According to the method, efficient prediction of the wall surface pressure distribution and the heat flux density under the same basic configuration and under the non-uniform wall temperature boundary condition of different variable geometric shapes can be realized, and the prediction accuracy is improved.

The invention discloses a morphing wing device integrating variable sweepback and variable dihedral angle functions. The morphing wing device comprises two actuating mechanisms, the actuating mechanism comprises a metamorphic mechanism and a wing connecting assembly; the metamorphic mechanism comprises an input semi-ring and an output semi-ring, and each of the input semi-ring and the output semi-ring comprises a first short rod, a second short rod and a long rod; one joint in the input half ring is a coupling joint, the corresponding joint in the output half ring is a coupling joint, and the other four joints are revolute pairs; in the input semi-ring, the first short rod and the long rod are driven by a driving device to open and close; the wing connecting assembly comprises four hooke joints and two guide rods, the two hooke joints are arranged at joints between the first short rod and the second short rod respectively, and the other two hooke joints are arranged on the two long rods respectively; the guide rod is parallel to the longitudinal axis of the wing; fixed connecting blocks, sliding connecting blocks and sliding connecting grooves are fixedly arranged on the lower surfaces of the wings; the front guide rod penetrates through the fixed connecting block and the sliding connecting block; the sliding connection groove is located between the two hooke joints on the rear guide rod.

The application discloses a shaft-symmetrical rocket-based combined engine with a variable geometry throat, which comprises a combustion chamber and a square flow channel section in communication in an axial direction from front to back; the shell of the combustion chamber is surrounded by a bottom support bottom plate and a top plate, the bottom support bottom plate is horizontal, and the top plate is a semicircular arc shape which is arched upward; the square flow channel section is a cuboid cavity, and an upper variable structure geometry throat and a lower variable structure geometry throat which are consistent with the size of corresponding holes are embedded in the holes at the upper part and the lower part; the upper variable structure geometry throat and the lower variable structure geometry throat are both plate bodies in the shape of an inclined wedge, the tip of the inclined wedge faces the rear, and the inclined surface sides of the inclined wedge both face the cavity; the upper variable structure geometry throat and the lower variable structure geometry throat are used for changing the flow channel geometry throat area in the square flow channel section. The engine is matched with the flight state by reducing the throat area with the increase of the flight Mach number, the engine performance is improved, and the engine thrust vector adjustment can be realized.

The application provides a morphing aircraft fault detection model generation method and a fault detection method. The morphing aircraft fault detection model generation method comprises the following steps: obtaining a state control feedback matrix corresponding to each wingspan change rate based on an initial value of each wingspan change rate, an initial value of an aircraft state variable, and a preset flight linear change parameter model of the morphing aircraft; generating flight simulation data based on to-be-simulated data, wherein the to-be-simulated data comprises the initial value of the aircraft state variable, the wingspan change rate, and the state control feedback matrix, and each set of flight simulation data comprises flight state parameters at multiple time points; generating sample data based on the flight simulation data, wherein the sample data comprises the flight simulation data and a fault label corresponding to the flight simulation data; and obtaining the morphing aircraft fault detection model by training based on multiple sets of sample data. The application can effectively detect faults of the morphing aircraft.

The application provides a deformable wing using a telescopic connecting rod mechanism, comprising: a plurality of frames; a plurality of connecting rod mechanisms, each of which is arranged on a frame, each of which comprises a telescopic first connecting rod, a non-telescopic second connecting rod and a connecting rod for connecting the first connecting rod and the second connecting rod; and a plurality of driving devices, each of which has a frame on the upper side and the lower side respectively, and each of which is connected with the first connecting rod in the two corresponding connecting rod mechanisms, so as to control the two first connecting rods to synchronously stretch and retract. The application can realize the length adjustment of the aircraft wing in the wing profile section direction, so as to change the wing profile area and the angle of attack, thereby enabling the aircraft to adjust the lift-drag ratio by controlling the movement of the four connecting rod mechanisms to adjust the wing area and the angle of attack when the lift needs to be adjusted. The application has the advantages of simple structure, stable control, high safety factor and the like.

The invention is a traction device for a vehicle, such as a wheel or a track, comprising one or more movable parts, referred to as drive parts, which can change position from a position enabling propulsion on the ground to a position ensuring propulsion in a fluid or on its surface. The main applications are land vehicles capable of traveling on roads as well as on uneven or muddy ground and of traveling on or in water, such as a motor vehicle, a boat capable of navigating shallow waters, a submarine, or an aircraft capable of landing on a beach or on the water.

The present invention includes a seat having an elastomeric seal and a core, the core having a groove disposed within the core forming a seal path, the seal path having a top centerline, a bottom centerline, an exposed edge, and a covered edge; the elastomeric seal is disposed within the groove; the elastomeric seal has a progressively varying seal protrusion height and protrusion width above the seat face and a progressively varying depth and width seal substructure below the seat face.

The invention belongs to the technical field of prediction and control of a supersonic air inlet channel global flow field of a ramjet engine, and particularly relates to a prediction method and a prediction system for a wide-speed-domain variable-geometry air inlet channel global flow field. The method comprises the steps that 1, wide-speed-domain variable-geometry air inlet channel flow field information is obtained; 2, based on the information obtained in the step 1, a variable geometry air inlet channel global flow field prediction data set is constructed; 3, a variable geometry air inlet channel wall surface sensor configuration scheme is constructed, and the pressure value of the variable geometry air inlet channel is collected; and 4, designing a global flow field prediction model structure of the air inlet channel based on the variable geometry air inlet channel acquired in the step 3, and realizing prediction of the global flow field of the wide-speed-domain variable geometry air inlet channel. The method is used for solving the problem that the accurate perception of the hypersonic flow field is in the future development of hypersonic velocity.

A shape forming tool for pre-carbonization shaping of a fibrous preform is provided, comprising a rounded preform and an inner die to fit within the preform. The rounded preform may be a cylindrical, conical, or frustoconical shape. The inner die may be of a non-uniform diameter and / or have complex features. The preform may shrink to conform to the inner die.

This invention discloses a method, apparatus, computer equipment, and computer-readable storage medium for correcting an adjustable guide vane clearance leakage loss model, relating to the field of variable geometry turbine adjustable guide vane technology. The method includes: constructing a database of multiple clearance flow fields for the blade tip / root clearance and structures containing a rotating shaft based on CFD experimental flow field data of a variable geometry turbine, obtaining independent leakage losses. A machine learning method, primarily using neural networks, is used to fit the corrected AMDCKO model, obtaining blade tip and root clearance loss models, which are then superimposed into a total clearance loss model without a rotating shaft. A rotation shaft correction coefficient KZ is introduced, and its diameter and position are used as variables to model the leakage loss at the tip of the adjustable guide vane containing the rotating shaft. Embedding this model into a one-dimensional performance prediction program can improve the accuracy and reliability of loss calculation, providing a scientific theoretical basis and technical support for the aerodynamic optimization design and parameter matching of variable geometry turbines.

The application discloses a variable wing for a fixed-wing aircraft, which comprises a rotating wing system arranged on both sides of the axis of the fuselage, the rotating wing system is installed on the fixed wing of the fuselage, a driving motor is installed on the fuselage, and the driving motor controls the operation of the rotating wing system through a transmission system. The variable wing for the fixed-wing aircraft can improve the flight performance of the fixed-wing aircraft, adapt to different flight environments, meet the task requirements of high speed, low speed and short take-off and landing, and realize the large-speed-range and full-time-period maneuvering flight of the aircraft.

The method can include determining (401) a mass flow rate W of working fluid circulating through the compressor stage, determining a control parameter value associated to the geometrical configuration of the variable geometry element based on the determined value of mass flow rate W; and changing (414) the geometrical configuration of the variable geometry element in accordance with the determined control parameter value.

The invention discloses a variable-shape aircraft backstepping control method, and belongs to the field of aircraft control. The implementation method comprises the following steps: establishing a variable shape aircraft system model; and setting a high-order fixed time disturbance observer. And estimating the state and disturbance of the variable shape aircraft system model. And establishing a controller comprising an attitude angle control loop and an attitude angular velocity control loop, and obtaining a control instruction based on an estimation result. And substituting the control instruction into the shape-variable aircraft model to obtain the attitude change of the aircraft. According to the high-order fixed time disturbance observer, auxiliary variables are introduced to carry out high-order differential expansion on disturbance estimation, so that the estimation precision of uncertain disturbance is improved; under the condition of not depending on the initial state of the system, it is ensured that the tracking error converges to any small field of zero within fixed time; according to the backstepping control method based on the fixed-time high-order disturbance observer, effective observation and estimation of multi-source uncertain time-varying disturbance existing in a control system can be achieved.

This invention relates to the field of multibody system dynamics, specifically a dynamic modeling method for simulating the flight motion of a morphing aircraft with an arbitrary number of rotating components. The method first abstracts the morphing aircraft as a tree-like multibody system containing virtual rigid bodies and virtual hinges. Rigid bodies and hinges are uniformly numbered using a depth-first traversal search, with the relative rotation angles between rigid bodies used as generalized coordinates. Then, a path matrix is ​​used to describe the system topology, and the center-of-mass velocity, center-of-mass acceleration, angular velocity, and angular acceleration of each rigid body are derived using the generalized coordinates. Next, based on the above kinematic parameters, the deflection velocity, deflection angular velocity, generalized active force, and generalized inertial force are calculated and substituted into the Kane equations to establish the multibody system dynamic equations. Finally, the dynamic equations are solved, and the motion state is updated according to a set time step, repeating the process until the simulation of the set duration is completed. This invention has strong versatility and good scalability, and is applicable to the flight motion simulation of morphing aircraft with an arbitrary number of rotating components.

A compressor system (400) for a heating, ventilation, air conditioning, and refrigeration (HVAC&R) system (10) includes a low stage compressor (402) having a first pre-rotation vane (PRV) (406) and a first variable geometry diffuser (VGD) (408) configured to operate to regulate a capacity of the low stage compressor (402), wherein a first operational linkage for the low stage compressor (402) causes a first position of the first VGD (406) to be dependent upon a second position of the first PRV (408). The compressor system (400) also includes a high stage compressor (404) having a second PRV (410) and a second VGD (412) configured to operate to regulate a capacity of the high stage compressor (404), wherein a second operational linkage for the high stage compressor (404) causes a third position of the second PRV (410) to be dependent upon a fourth position of the second VGD (412).

The present invention includes a seat having an elastomeric seal and a core, the core having a groove disposed within the core forming a seal path, the seal path having a top centerline, a bottom centerline, an exposed edge, and a covered edge; the elastomeric seal is disposed within the groove; the elastomeric seal has a progressively varying seal protrusion height and protrusion width above the seat face and a progressively varying depth and width seal substructure below the seat face.

The utility model relates to a ship model structure transformation device of a morphing small water plane twin hull, and aims at solving the problem that the fixed small water plane twin hull ship model cannot be morphed and the tow point height cannot be artificially adjusted. The ship model structure transformation assembly is installed on the main hull, the tow point height adjusting assembly is installed on the ship model structure transformation assembly, the longitudinal support pipe group is fixedly installed on the main hull through a plurality of load-bearing columns and a plurality of square pipe seats, the vertical support pipe group is fixedly installed in the middle of the longitudinal support pipe group through a plurality of semi-detachable T-shaped three-way connectors, and the transverse support pipe group is fixedly installed at the end of the longitudinal support pipe group through a plurality of semi-detachable T-shaped three-way connectors. The utility model is used in the water surface navigation technical field.

An assembly for an aircraft propulsion system includes an open propulsor rotor (38) having variable geometry propulsor blades (78) and an open guide vane / stator assembly (66) having a plurality of guide vanes (stator blades) (100) in a fixed position relative to the rotor (38) and assembly. The geometry (e.g., pitch, camber, etc.) of the guide vanes (100) can be manually adjusted by ground personnel but are incapable of adjustment during a flight mission (a flight). The open propulsor rotor (38) is configured to rotate about an axis, and the guide vane assembly (66) includes the plurality of guide vanes (100) arranged circumferentially about the axis and is disposed axially next to and downstream of the propulsor rotor (38).

A road vehicle having a roll axis, a pitch axis and a yaw axis and comprising a rear aileron; characterized in that the rear aileron comprises two airfoils coupled to each other in a rotary manner around a pivoting pin with an axis parallel to the yaw axis so that the aileron can assume a variable geometry between an opening configuration, in which the airfoils are in series to each other along an axis parallel to the pitch axis, and a closing configuration, in which the airfoils are in parallel to each other along an axis parallel to the roll axis, going through a plurality of semi-opening configurations, in which the airfoils define a non-zero negative sweep angle toward the front portion of the vehicle.

The present application belongs to the field of aviation power technology, and particularly relates to a variable-geometry ducted fan. The variable-geometry ducted fan comprises a variable-pitch fan, a tail nozzle and a driving mechanism, the variable-pitch fan is arranged on a casing, the variable-pitch fan is connected with the driving mechanism, and the tail nozzle is arranged at the rear end of the casing. The beneficial effects are that the flow size and the outlet jet direction of the fan are changed through the variable-geometry method, and the relationship between the fan rotating speed, the flow, the pressure ratio and other parameters is adjusted, so that the fan can realize different pressure ratio / flow characteristics, thereby better matching the different requirements of the aircraft on the power system under the conditions of take-off, cruising and flat flight.

The invention belongs to the field of aero-engine design, and particularly relates to a variable geometry two-stage low-pressure turbine aerodynamic design method which comprises the steps that a typical state point with the maximum flow and a typical state point with the minimum flow are selected respectively, through-flow design and turbine one-dimensional aerodynamic design are conducted, and a one-dimensional design result is obtained; determining turbine element blade profiles according to a one-dimensional design result, and establishing adjustable blade configurations of each stage of turbine through three-dimensional modeling; and establishing and operating a three-dimensional model of each stage of turbine, collecting turbine working parameters of each stage of turbine at typical state points of maximum flow and minimum flow, and carrying out turbine characteristic calculation of the turbine at different rotating speeds along with the change of the expansion ratio to obtain the turbine working parameters of each stage of turbine. Compared with a single-stage variable geometry turbine, the efficiency of the turbine is improved under the limited geometric size, the efficiency of the turbine is improved by 3-4% at the typical state point with the minimum flow compared with a single-stage variable geometry turbine, the variable geometry low-pressure turbine design with the wide flow adjusting range and high efficiency is achieved, and the overall performance of an engine is improved.

A system and a method for configuring at least one variable geometry mechanism (VGM) of an aircraft engine are provided. Pass-off testing data for the aircraft engine is obtained, the pass-off testing data indicative of an actual value of at least one operating parameter of the aircraft engine. Based on the pass-off testing data, at least one trim value to be used to adjust a setting of the at least one VGM to bring the actual value of the at least one operating parameter towards a target value is determined, a running line of the aircraft engine being substantially constant when the actual value of the at least one operating parameter is at the target value. The setting of the at least one VGM is adjusted, during pass-off testing of the aircraft engine, using the at least one trim value.

The application belongs to the technical field of aero-engines, and particularly relates to a variable-geometry diffuser adjustable device, which comprises a diffuser, a casing, a flame tube, an adjustable structure fixing disc, an adjustable structure rotating disc and a driving mechanism. The casing forms an annular space for mounting the flame tube. The diffuser is mounted at the front end of the casing to pass air into the annular space. The adjustable structure fixing disc and the adjustable structure rotating disc are mounted between the diffuser and the flame tube to adjust the air flow between the diffuser and the flame tube. The adjustable structure fixing disc is fixed on the casing. The adjustable structure fixing disc is provided with circumferentially distributed first air inlet holes. The adjustable structure rotating disc is rotatably and concentrically mounted on the end face of the adjustable structure fixing disc. The driving mechanism drives the adjustable structure rotating disc to rotate relative to the adjustable structure fixing disc. The adjustable structure rotating disc is provided with second air inlet holes. The angle between the adjustable structure fixing disc and the adjustable structure rotating disc is changed, and then the communication area of the first air inlet holes and the second air inlet holes is changed.

The present invention relates to methods and systems for managing engines in a low power state. Methods and systems for operating an aircraft having two or more engines are described. One method includes operating the two or more engines of the aircraft in an asymmetric operating regime, wherein a first one of the engines is in an active mode to provide power to the aircraft and a second one of the engines is in a standby mode to provide substantially no power to the aircraft; managing the first engine in the active mode using first management logic; and managing the second engine in the standby mode using second management logic, the second management logic based on a target compressor speed adjusted using a trim value and a variable geometry mechanism (VGM) setting, the trim value dependent on at least one parameter of the second engine in the standby mode.

The application belongs to the field of aero-engine design, and particularly relates to a variable-geometry two-stage low-pressure turbine aerodynamic design method, a flow maximum and a flow minimum typical state point are respectively selected to perform through-flow design and turbine one-dimensional aerodynamic design, and one-dimensional design results are obtained; turbine base element blade profiles are determined according to the one-dimensional design results, and three-dimensional modeling is respectively performed to establish adjustable blade configurations of each stage of turbine; three-dimensional models of each stage of turbine are established and run, turbine working parameters of each stage of turbine at the flow maximum and flow minimum typical state points are collected, turbine characteristic calculation of the turbine under different rotating speeds with changes in expansion ratio is performed, and turbine working parameters of each stage of turbine are obtained. The efficiency of the turbine is improved under limited geometric dimensions, compared with a single-stage variable-geometry turbine, the turbine efficiency at the flow minimum typical state point is improved by 3-4%, wide flow regulation range and high-efficiency variable-geometry low-pressure turbine design are achieved, and the overall performance of the engine is improved.

The invention provides a variable geometry non-axisymmetric elastic device for actively controlling a turbine blade tip clearance. The variable geometry non-axisymmetric elastic device is composed of an elastic retaining ring, a micro motor, a lead screw, a clearance sensor and a displacement plate. The execution unit and the controller jointly form a control part, the elastic retaining rings are evenly arranged on the inner wall of the air cylinder, and the control part is connected with the movable ends of the retaining rings and used for controlling the displacement of the movable ends of the retaining rings in the radial direction so as to adjust the turbine blade tip gap size under different working conditions, and therefore the turbine blade tip gap size is always close to the optimal blade tip gap size. The deformation of the elastic retaining ring enables the whole annular surface formed by the elastic retaining ring and the inner wall of the air cylinder to present different non-axisymmetric space curved surface shapes. The variable geometry non-axisymmetric elastic retaining ring for actively controlling the turbine blade tip clearance has the following advantages of being high in response speed, simple in structure, convenient to install, low in cost, light in weight, high in reliability, easy to maintain, long in service life of components and the like.

The present invention discloses a method for designing a variable geometry asymmetric nozzle coupled with a rocket thrust chamber, including a variable geometry three-dimensional asymmetric nozzle designed based on streamline tracing technology; a three-dimensional model of the rocket thrust chamber designed based on the characteristic line method; and a coupling method design for the rocket thrust chamber and the three-dimensional asymmetric nozzle. The present invention installs three rocket thrust chambers on the upper expansion surface of the three-dimensional asymmetric nozzle. The high-energy jet provided by the rocket thrust chambers improves the thrust performance of the three-dimensional asymmetric nozzle at low Mach numbers. The effective combination of the rocket thrust chambers and the asymmetric nozzle ensures high-performance operation of the asymmetric nozzle over a wide speed range, better meeting the use requirements of engines and aircraft.

A filtration system comprises a housing, a cover, and a filter element having a variable geometry sealing member. The housing comprises a sealing surface disposed at a housing first end. The variable geometry sealing member is disposed at the sealing surface. A shape of the variable geometry sealing member is defined by a first curve having a first radius and a second curve having a second radius. The first curve and the second curve are non-concentric such that a thickness of the variable geometry sealing member is non-uniform. The cover is coupled to the housing such that a surface of the cover compresses the variable geometry sealing member against the sealing surface.