38 results about "Ground resonance" patented technology

The invention discloses a flutter boundary prediction method in a wind tunnel flutter test under a turbulence excitation condition. At first, time domain response signals of typical measuring points in the process of the wind tunnel test are collected, trend removal and smooth pretreatment are carried out on the signals, self-correlation processing is conducted on processed turbulence response data, the processed turbulence response data serve as impulse responses of a system, band-pass filtering is carried out on the processed turbulence response data, and frequency selected by filtering is determined by dangerous modal frequency measured by a model ground resonance test; secondly, spectral analysis and modal parameter identification are carried out on the obtained and processed impulse responses with noise, and an MPM with high anti-noise performance is adopted to identify various order modal parameters; at last, frequency and damping ratios of various order modals are arranged according to different wind speeds when data are collected, and a mode that straight line fitting and polynomial fitting are combined is adopted to extrapolate a flutter boundary. According to the method, the flutter boundary under the turbulence excitation condition can be forecasted, a traditional flutter judgment method is improved, and improvements of accuracy and safety of the test are facilitated.

The invention discloses a bearing-less rotor ground resonance test method, and belongs to the technical field of bearing-less rotor test. The method comprises the following steps: 1, firstly constructing a bearing-less rotor simulation model and a testbed simulation model, and assembling the two models to carry out ground resonance calculation analysis; 2, verifying equipment in a test system, checking test pieces, debugging each system of the testbed, arranging test points, mounting a sensor, mounting an acceleration sensor at a propeller hub center to measure acceleration speeds in a rotation surface in two directions, mounting an acceleration sensor on a rotor shaft to measure an acceleration speed of the rotor shaft, and providing an overload monitor; 3, releasing an excitation signalby a signal source, transferring the excitation signal to an auto-bank unit via an excitation system, and driving paddles to carry out periodic variable-distance movement by the auto-bank unit; and 4,responding a time history by test points, acquired in the excitation process, on a rotor hub, a flexible beam and the testbed, and processing and analyzing a coupling modal frequency and modal damping of the bearing-less rotor and the testbed, so as to judge whether a ground resonance stability margin exists or not.

The invention relates to a nonlinear dynamic modeling method of an undercarriage wheel, belonging to a helicopter theoretical modeling technology. The nonlinear dynamics modeling method is characterized by comprising the following steps: carrying out vertical rigidity modeling, lateral rigidity modeling, course rigidity modeling and torsional rigidity modeling on the wheel according to structure parameters (inflation pressure, diameter and width) of the wheel, carrying out vertical damping modeling, lateral damping modeling, course damping modeling and torsional damping modeling on the wheel, solving an undermined coefficient in a model according to the static compression characteristic of the wheel and the dynamic rigidity damping test data by using a least square method, and acquiring an analytical expression capable of accurately simulating the nonlinear dynamic characteristic of the wheel in a time domain. The nonlinear dynamics modeling method disclosed by the invention can be used for dynamic characteristic analysis for the center of a propeller hub of a helicopter and ground resonance analysis, so that the accuracy of ground resonance stability analysis can be improved, great error caused to ground resonance analysis by linear processing is avoided and the helicopter weight cost caused by linear processing is avoided.

The invention belongs to the technical field of helicopter dynamic tests, and discloses a helicopter flight control system transmission characteristic test method. The method comprises the steps of respectively performing sweep frequency excitation on a rotor time domain total distance and a periodic variable pitch through a control system, operating an excitation signal through a hydraulic actuator cylinder, and measuring the rotor time domain total distance and the periodic variable pitch under excitation; and performing FFT on the time domain total distance and the periodic variable pitch of the hydraulic actuator cylinder and a paddle to obtain a control system total distance and periodic variable pitch frequency domain transfer function. The method can be used for providing data for establishing a helicopter flight control system mathematical model based on tests, and providing basic technical support for establishing a helicopter rotor and body coupling stability model considering a flight control system and stability comprehensive analysis. The method has a certain engineering reference value for design and analysis of ground resonance and air resonance of advanced helicopters in China, and has far-reaching significance for guaranteeing safe flight and stability improvement of modern advanced helicopters.

The invention belongs to the technical field of aeroelasticity, and relates to a full-aircraft beam frame type reduction stiffness combination modeling method. A reduction method is used for obtaining a stiffness matrix of a connecting portion of a rear fuselage large opening and a wing body, and therefore the problem that accurate stiffness data of the complex structures are difficult to obtain through a previous computing method; moreover, a full-aircraft beam frame type reduction stiffness combination model is established in combination with a beam frame model for wings, a vertical tail, a horizontal tail and other components, as a result, accurate simulation of full-aircraft stiffness characteristics is realized, the computing scale is greatly reduced, computing efficiency is improved, model adjustment and parameter variation analysis after a full-aircraft ground resonance test are facilitated, the deficiency of the current modeling technology is remedied, and a novel path is opened up for aeroelastic simulation modeling.

The invention discloses a propeller hub center non-linear dynamic characteristic modeling method, which belongs to helicopter theoretical modeling technology. The propeller hub center non-linear dynamic characteristic modeling method is characterized by comprising the following steps: establishing an engine body coordinate system, and describing the motion of an engine body and undercarriages; carrying out the balance computation of single undercarriage, and establishing force and position relationship between airplane wheels and a buffer; carrying out the balance computation of the engine body on the undercarriage; and determining the non-linear constraint force of an undercarriage system to the motion of the engine body in a time domain, and inertial load acted to the engine body, establishing an engine body motion equation on a fixed coordinate system by adopting a d'alembert's principle, and simultaneously establishing the balance equation of each undercarriage and the motion equations of the engine body on the undercarriage to obtain a propeller hub center non-linear dynamic characteristic equation of the engine body on the undercarriage. The invention can be applied to helicopter propeller hub center non-linear dynamic characteristic analysis and ground resonance modeling and analysis and can improve the accuracy of the propeller hub center non-linear dynamic characteristic analysis and the ground resonance stability analysis.

The invention provides a multi-functional integrated circuit protection device applied to a 6-66kv power grid system. The multi-functional integrated circuit protection device comprises a composite circuit formed by nonlinear damping devices Rk1 and Rk2 and a nonlinear damping resistor RG3 having the characteristic of ultra-low impedance and an oversized energy capacity. According to the invention, the multi-functional integrated circuit protection device can effective damp and suppress the rising rate and amplitude of transient recovery voltage of a vacuum circuit breaker, and prevent the generation of over voltage. In the on-off of the circuit breaker, operating over voltage of oscillation generation due to mutual conversion of electric energy and magnetic energy is damped and suppressed to switch on-off of high short circuit capacity of the vacuum circuit breaker. Through the combined application of the nonlinear damping resistors and the circuit design of an artificial neutral point O, the multi-functional integrated circuit protection device has the functions of excellent lightning protection, damping and suppression of phase-phase and phase-ground resonance and operating over voltage, and especially excellent suppression of single-phase intermittent arc grounding over voltage, and serves to eliminate electric arc over voltage.

The invention discloses a helicopter rotor and fuselage coupling stability modeling method, and belongs to the helicopter dynamics modeling and analysis technology. The method includes: adopting a modal synthesis technology to perform comprehensive modeling on a rotor and fuselage dynamics coupling system, describing the motion of a fuselage and rotors in different coordinates, establishing structural dynamics finite element method models of isolated rotor blades and an airframe structure respectively, adopting an aerodynamic model to model aerodynamic force, and deducing a rotor / airframe coupling system kinetic equation by applying the Hamilton principle; establishing an automatic flight control system control model, and deriving an aerodynamic load matrix related to the pitch variable, thereby establishing a rotor wing and aircraft body coupling aeroelastic stability analysis model considering the flight control system; and finally, solving a characteristic value through a characteristic value method, and judging the stability of the coupling system through a characteristic value solution. The model can be used for calculating and analyzing the'ground resonance 'and'air resonance' stability of all advanced helicopters, and key technical support is provided for model design and development.

A linking system for attaching a cross tube to a skid of a landing gear of a rotorcraft, the linking system comprising a hinge element, means for connecting the hinge element to the cross tube and means for connecting the hinge element to the skid, the hinge element in turn comprising a cross tube-sided part and a skid-sided part able to rotate with respect to one another, a torsion bar spring and a rotary damper element suitable for dissipating a ground resonance excitation energy when the rotation of the cross tube-sided part with respect to the skid-sided part occurs.

The invention discloses a design method for improving the elastic stability of an aerodynamic servo. The method comprises the steps: correcting a body elastic model and a transfer function through employing a whole-plane ground resonance test or whole-plane ground structure modal coupling test data design optimization technology, and then improving the elastic stability of the aerodynamic servo through employing a flight control law dynamic parameter adjustment technology; the aerodynamic force, the structural elasticity and the coupling relation between control systems can be well considered,so that the flight quality requirement, the stability margin requirement of the control systems and the pneumatic servo elastic stability requirement are met within the full envelope range, and the method is already applied and verified in the multi-control-surface unmanned aerial vehicle. The problem that the pneumatic servo elastic stability of a multi-control-surface unmanned aerial vehicle isdifficult to control is solved.

The invention relates to a fixing device of an aircraft control system, and belongs to the technical field of aeroelastic effect tests. The device comprises pedal fixing units (1), steering wheel fixing units (2), a basis (3) and support bar systems (4), wherein one end of each pedal fixing unit is connected with the basis, and a clamping end is arranged at the other end of each pedal fixing unit and is used for clamping a pedal connecting tube of an aircraft pedal; one end of each support bar system is connected to the basis, and the other end extends toward an aircraft steering wheel to form a hollow tube; each hollow tube is used for muff-coupling and fixing an end part of the corresponding steering wheel fixing unit; a second clamping end is arranged at the other end of each steering wheel fixing unit, and is used for clamping a steering wheel control handle of the aircraft steering wheel. Through the device, whole aircraft control system fixation can be realized in the process of an aircraft ground resonance test, each part is prevented from relative displacement in the fixation process, and measurement of each control surface rotating frequency in the resonance test is ensured.

A hydraulic damper assembly has a barrel-piston assembly and a spring assembly. The barrel-piston assembly has a barrel with a top portion having an open end, and a piston extending through the open end. The piston has a bottom portion that is located inside the barrel. The piston and the barrel also define an annular space in which is a barrel shoulder and a piston shoulder. The barrel shoulder radially extends into the annular space from the barrel top portion. The piston shoulder radially extends into the annular space from the piston bottom portion. The spring assembly is disposed in the annular space between the barrel shoulder and the piston shoulder. When the piston is extended from the barrel, the shoulders compress the spring assembly and, thereby, reduce the amount of force required to compress the piston which enables the damper to be charged to a higher nitrogen pressure. An object of the invention is to provide a damper assembly for aircraft landing gears that permit high nitrogen pressures to increase ground resonance clearance at high aircraft gross weight and not be fully extended at low aircraft gross weights to avoid ground resonance problems. Another object of the invention is to provide a cost effective damper assembly that is easily retrofitted into current designs of helicopter landing gears.

The invention belongs to a dynamic design technique of helicopters, relates to a dynamic modeling method of a multi-transmission rotor hub structure and aims to prevent ground resonance of a bearingless rotor and expand the air resonance boundary. The accurate modeling analysis and design means is provided in terms of rigidity configuration, displacement coordination for a flexible beam, a control lever, a damper and a sleeve; dynamic finite element modeling of a multi-transmission rotor hub structure is performed first, and degrees of freedom of non-dependent joints of component parts of the structure are expressed in joint degree-of-freedom linearity; according to the dependency and relativeness of the joint degrees of freedom of the component parts of the structure, the independent joints, the relative joints and the displacement coordination relation between the component parts of the structure and the connective joints are determined so as to establish a dynamic rigidity matrix of the structure. By the use of the dynamic modeling method, the key difficulties in the design and analysis of the complex multi-transmission rotor hub structure are broken through.

The invention discloses a method for acquiring a kinetic parameter of a helicopter hub center. The method for acquiring the kinetic parameter of the helicopter hub center comprises the following steps: step 1, establishing a helicopter finite element model including a helicopter rotor and an airframe; step 2, setting a calculation frequency bandwidth, performing modal analysis on the helicopter finite element model, and acquiring frequency response data of the overall mode; step 3, performing kinetic characteristic response simulation calculation on the hub center to obtain frequency response data of the hub center, and drawing an amplitude-frequency response curve and a phase-frequency response curve based on the frequency response data of the hub center and the frequency response data of the overall mode; and step 4, acquiring the kinetic parameter of the helicopter hub center. The method for acquiring the kinetic parameter of the helicopter hub center can provide a parameter for preliminary design phase analysis of helicopter ground resonance.

An aspect provides a skid landing gear assembly including a cross member disposed between a first skid tube and a second skid tube, and a tension cable member configured to be disposed in the cross member; wherein the tension cable member resists outward deflection of at least a portion of the skid landing gear assembly when a downward load is applied to the skid landing gear assembly. In other aspects, there are methods of avoiding ground resonance in a skid landing gear assembly, methods of assembling a skid landing gear assembly, methods of operating a helicopter with a skid landing gear assembly, and methods of improving the service life of a skid landing gear assembly.

The invention belongs to the field of wind tunnel tests, and particularly relates to a design parameter control method for transonic flutter wind tunnel model processing. The method comprises the steps that a metal beam frame is adopted for manufacturing an airplane force bearing component, a first flutter wind tunnel model is generated, and a first ground resonance test is conducted on the first flutter wind tunnel model; a balancing weight is adhered to the first flutter wind tunnel model to generate a second flutter wind tunnel model, and a second ground resonance test is performed on the second flutter wind tunnel model; accessories, embedded sensors and pasted foam fillers are mounted on the second flutter wind tunnel model; a composite material skin test piece is obtained, and a material tensile test and a composite material skin test piece resonance test are performed on the composite material skin test piece; a composite material skin is laid in the mold, the second flutter wind tunnel model is put into the composite material skin for medium-temperature curing to generate a third flutter wind tunnel model, and a third ground resonance test is performed on the third flutter wind tunnel model.

The invention discloses an airborne IMU high-precision reference standard acquisition method based on wing deformation. The method comprises the following steps that a finite element model is established, a ground resonance experiment is conducted, then the layout of FBG sensors is designed, space coordinates of all points on a wing are determined according to measured data, and finally a reference standard is determined. According to the invention, the application of providing the reference standard for the airborne IMU based on the FEM technology and the FBG sensing technology is realized, the reference standard of the IMU can be quickly determined, and the universality is high.

A longitudinal helicopter rotor and fuselage coupling stability modeling method belongs to the helicopter dynamics modeling and analysis technology, adopts a fully-articulated rigid blade model, and comprises the following steps: firstly, establishing coordinate systems of all systems and relationships among the coordinate systems, establishing a fuselage motion model, a rotor motion model and anaerodynamic model; after rotor coupling response solving and multi-blade coordinate transformation, combining a rotor body undercarriage blade model mass damping stiffness matrix, establishing a tandem double-rotor helicopter ground resonance analysis model, calculating ground resonance by adopting a characteristic value method, and judging ground resonance stability through a characteristic valuereal part. The model can be used for calculating and analyzing the ground resonance stability of the tandem double-rotor helicopter, and key technical support is provided for model design and modification development.

The invention discloses a ground resonance test method for a bearingless rotor, belonging to the technical field of bearingless rotor tests. Including: step 1, first constructing the simulation model of the bearingless rotor and the test bench, and after assembling the two models, performing calculation and analysis of the ground resonance; Commissioning; arrangement of measuring points and installation of sensors; an acceleration sensor is installed in the center of the propeller hub to measure the acceleration in two directions in the rotating plane; an acceleration sensor is installed on the rotor shaft to measure the acceleration of the rotor shaft and provide overload monitoring; steps 3. The signal source releases the excitation signal, and the excitation signal is transmitted to the automatic tilter through the excitation system, and the automatic tilter drives the blades to perform periodic variable pitch motion; step 4, during the excitation process, collect the rotor hub, the flexible beam and the test bench. The response time history of the measuring point is used to process and analyze the coupling modal frequency and modal damping of the bearingless rotor and the test bench, and judge whether there is a ground resonance stability margin.

The invention discloses a device for ground resonance test of a coaxial dual-rotor helicopter, and relates to the field of aircraft test devices. The device comprises upper rotor propeller blades (1), an upper rotor propeller boss (2), an upper rotor boss shaft (3), an upper rotor driving gear (4), an electric motor (5), a box body (6), a spring (7), a damper (8), a supporting plate (9), a cardan joint (10), a chassis (11), brackets (12), bearings (15), pins (16), a counter weight (17), a lower rotor boss shaft (18), a lower rotor driving gear (19), an electric motor bevel gear (20), a lower rotor propeller boss (22) and lower rotor propeller blades (23). The device disclosed by the invention can simulate the resonance effect of the coaxial dual-rotor helicopter, and a test device for researching the resonance of the coaxial dual-rotor helicopter is provided.

The invention discloses an airfoil ground resonance test wind power loading device and relates to the technical field of aircraft aeroelasticity. The airfoil ground resonance test wind power loading device comprises a pneumatic load flow dividing dredging structure (2) and a pneumatic loading mechanism (3) used for applying a pneumatic load to the pneumatic load flow dividing dredging structure (2). The pneumatic load flow dividing dredging structure (2) is arranged to be of a cone cylinder structure; the small end of the cone cylinder structure is provided with a flow baffle plate (21) provided with flow baffle holes (211); and the pneumatic load flow dividing dredging structure (2) is mounted in the pneumatic pressure center position of an airfoil (1) through the flow baffle plate (21). The airfoil ground resonance test wind power loading device has the advantages that the airfoil ground resonance test wind power loading device can adjust the loading position, the loading angle and the magnitude of the applied load along with deformation of the airfoil, thus deformation of various typical states of an airplane is advantageously simulated, the operation flow is simple, and the test efficiency and the reliability of results can be greatly improved.

The invention provides a real-time measurement device of the deformation of a test piece free end of a ground resonance test, relating to the technical field of aircraft aeroelasticity. The device comprises a measurement base which is connected to the free end of a test piece and can rotate freely around an x-axis and a y-axis, and the degree of freedom of rotation around a z-axis is constrained. The device also comprises a measuring instrument which is installed on the measurement base, and a displacement ruler which is fixed arranged and is provided with displacement scales in z-axis and x-axis directions. The measuring instrument can be maintained at a horizontal state on the measurement base and indicate the displacement scales on the displacement ruler, and x, y and z are in a customized three-axis coordinate system. The device has the advantages that an angle of measurement can be adjusted automatically based on a self gravity effect with the deformation of the test piece, the test precision is ensured, the device is convenient and fast to use, the processing cost is low, the test efficiency can be greatly improved, and the reliability of a test result is effectively ensured.

The present invention relates to a layout method for a shimmy damper and belongs to the field of helicopter kinetics physical design, comprising: performing dynamics modeling analysis on a layout scheme of the yaw damper; taking a hub center as the origin of a coordinate system and an extending direction of a blade K as X axis; establishing a coordinate system describing asymmetry in-blade layout of the blade K and a blade K+1 of the yaw damper; establishing a yaw damper-to-blade damping torque formula in the new scheme: a collected damping torque formula and a periodic damping torque formula; and according to a stiffness damping of yaw damper that needs to be satisfied, obtaining an optimal geometrical parameter. The present invention is applicable to blade yaw damper design of a machine under development or other helicopter models, is capable of meeting the requirements of eliminating both ground resonation and torsional vibration instability, and can improve the rotor yaw vibration damping and stiffness.

A method for analyzing ground resonance of a helicopter in a taxiing or taxiing condition include obtaining a lateral stiffness coefficient of a tire of a helicopter, an equivalent damping of a damper, and a method for analyzing ground resonance of a helicopter in a taxiing or taxiing condition. The effective mass, stiffness and damping of the damper are obtained according to the vibration frequency of the rolling mode. The equivalent damping, effective mass, effective stiffness and effective damping of the damper are substituted into the rotor. Coupling equations of airframe system are derived to obtain the ground resonance characteristics in taxiing or taxiing state.

The invention discloses a coaxial twin-rotor helicopter ground resonance modeling method, which belongs to the technical field of helicopter ground resonance modeling. Including step 1, first establish the coordinate system of each system and establish the transformation relationship between each coordinate system, describe the motion of the body and rotor blades in the established coordinate system; step 2, determine the restraint force of the landing gear on the motion of the body, and then Establish the body motion equation; step 3, mirror the lateral, roll and yaw degrees of freedom of the nonlinear motion equation of the upper rotor, and replace the distance between the hub center of the upper rotor and the center of gravity of the body with the hub center of the lower rotor The distance from the center of gravity of the fuselage to obtain the nonlinear motion equation of the lower rotor; step 4, solve the rotor motion balance point, and linearize the rotor nonlinear motion equation at the rotor motion balance point; step 5, perform multi-blade coordinate transformation, and The linearized rotor motion equation and body motion equation are assembled to obtain the ground resonance analysis model of the coaxial dual-rotor helicopter.

The invention provides a connector with a grounding shielding structure. The connector comprises an insulating casing, wherein the insulating casing is provided therein with a conductive shielding member and a conductive plastic plate; the bottom of the insulating casing is connected with an insulating base; the insulating casing is further provided therein with a terminal assembly; the conductiveshielding member comprises a conductive shielding plate; the conductive shielding plate is provided with an upper receding through hole; a conductive elastic piece is further fixed to the conductiveshielding plate; the conductive plastic member is provided with a lower receding through hole; the projection of the upper receding through hole on the conductive plastic plate coincides with the lower receding through hole; the insulating base is provided with a plate mounting groove; the bottom of the plate mounting groove is provided with a lower stopping hole; the projection of the lower stopping hole on the conductive plastic plate is in the lower receding through hole; the terminal assembly includes a signal terminal and a ground terminal; the signal terminal and the ground terminal arefixedly connected with the lower stopping hole; and the conductive elastic piece is in contact with the ground terminal. The connector can reduce the grounding resonance and can absorb and eliminate the possible resonance energy.

The invention discloses an airfoil ground resonance test wind power loading device and relates to the technical field of aircraft aeroelasticity. The airfoil ground resonance test wind power loading device comprises a pneumatic load flow dividing dredging structure (2) and a pneumatic loading mechanism (3) used for applying a pneumatic load to the pneumatic load flow dividing dredging structure (2). The pneumatic load flow dividing dredging structure (2) is arranged to be of a cone cylinder structure; the small end of the cone cylinder structure is provided with a flow baffle plate (21) provided with flow baffle holes (211); and the pneumatic load flow dividing dredging structure (2) is mounted in the pneumatic pressure center position of an airfoil (1) through the flow baffle plate (21). The airfoil ground resonance test wind power loading device has the advantages that the airfoil ground resonance test wind power loading device can adjust the loading position, the loading angle and the magnitude of the applied load along with deformation of the airfoil, thus deformation of various typical states of an airplane is advantageously simulated, the operation flow is simple, and the test efficiency and the reliability of results can be greatly improved.

The invention relates to a fixing device of an aircraft control system, and belongs to the technical field of aeroelastic effect tests. The device comprises pedal fixing units (1), steering wheel fixing units (2), a basis (3) and support bar systems (4), wherein one end of each pedal fixing unit is connected with the basis, and a clamping end is arranged at the other end of each pedal fixing unit and is used for clamping a pedal connecting tube of an aircraft pedal; one end of each support bar system is connected to the basis, and the other end extends toward an aircraft steering wheel to form a hollow tube; each hollow tube is used for muff-coupling and fixing an end part of the corresponding steering wheel fixing unit; a second clamping end is arranged at the other end of each steering wheel fixing unit, and is used for clamping a steering wheel control handle of the aircraft steering wheel. Through the device, whole aircraft control system fixation can be realized in the process of an aircraft ground resonance test, each part is prevented from relative displacement in the fixation process, and measurement of each control surface rotating frequency in the resonance test is ensured.