57 results about "Aviation engineering" patented technology

The invention discloses a test system of an automatic flight control system, and belongs to the aviation engineering test field. The test system is characterized by comprising an automatic flight control signal transmission connecting and disconnecting device for detecting crosslinking signals of an automatic flight control computer, an airborne data bus coupler, an automatic flight control system test interface device, a throttle implementation mechanism, a return actuator, an emulator of a master flight control computer, an air data system emulator, an inertial navigation system emulator, a radio altimeter emulator, a flight management system emulator, a flight state display emulator, an engine parameter system emulator, and test application software. Emulation of crosslinking airborne equipment in an automatic flight control system test can be realized, and a test on the automatic flight control system can be carried out.

The invention belongs to the field of aircraft engineering, and provides a nose-wheel steering system of an aircraft. According to the nose-wheel steering system, a steering control box is connected with an instruction sensor, a combined hydraulic valve and a feedback sensor through cables; one end of the instruction sensor is fixedly arranged below a floor of a cockpit of the aircraft by using bolts, and the other end of the instruction sensor is hinged to a foot pedal through bolts; the combined hydraulic valve is fixedly arranged on a bulkhead plate of a nose landing gear through bolts; the feedback sensor is arranged on a steering actuator; one end of the feedback sensor is fixedly arranged on a steering actuator shell through a hoop, and the other end of the feedback sensor is connected with a piston rod of the steering actuator through a clamping piece; the steering actuator is arranged on the nose landing gear of the aircraft; a piston rod of the steering actuator is hinged to a rotary sleeve of the nose landing gear through a joint bearing and a bolt. The nose-wheel steering system of the aircraft is used for steering at a low-speed stage and swing prevention at a high-speed stage during sliding of the aircraft, and has good operation flexibility and anti-swing stability.

The invention discloses a test system of a high-lift control system and a method thereof and belongs to the field of an aviation engineering test. The test system is characterized by comprises a high-lift signal transmission interruption and connection apparatus, a high-lift control system test interface device, a flap and edge slat super control panel, a slotted flap super control handle, a slotted flap PDU (power driving unit), a slotted flap anti-retracting brake apparatus, a simulator of a slotted flap rudder surface position sensor, an avionics system simulator and test application software. Simulation of crosslinking equipment in a test of the high-lift control system is realized, and the high-lift control system is tested.

The invention pertains to aeronautical engineering and consists of an airborne platform that can be built to large sizes without requiring a rigid structure of comparable dimensions and which uses both buoyancy and the aerodynamic Magnus effect for lift. The aerodynamic lift is generated in lifting bodies (1), which also contain buoyant gas. The lifting bodies (1) are stacked in a column, at the bottom of which there is a structural anchoring module (2) which also contains buoyant gas. The lifting bodies (1) and anchoring modules (2) are connected by slender structural elements which, when taken together as a whole form a non-rigid assembly. The platform may be tethered or configured as an aircraft, for which purpose other features may be added, such as a propulsion system (11), a crew gondola (6), cables to (7) and from (8) a swivel (12) and a payload (10) connected to said cables.

The invention relates to an aircraft control stick force and displacement measuring device and a measuring method thereof, belonging to the field of aeronautical engineering experiments and tests. The invention is characterized in that the measuring device consists of a control handle, a force signal detecting unit, a displacement signal detecting unit and a mounting and connecting unit; the controlling handle, the force signal detecting unit, the displacement signal detecting unit and the mounting and connecting unit are combined and integrated together into a 1-shaped structure in space; the force signal detecting unit and the displacement signal detecting unit are distributed at an upper side and a lower side of the mounting and connecting unit; the measuring device is connected with the control stick through the mounting and connecting unit; force is applied to the control stick by the controlling handle and the control stick is controlled to move; and the magnitudes of the controlling force and displacement are measured by the force signal detecting unit and the displacement signal detecting unit. The aircraft control stick force and displacement measuring device has the characteristics of compact and simple structure, small space dimension, low cost, reliability in work, high testing accuracy and convenience for mounting, using and adjusting, and is suitable for measurement of the controlling force and the displacements of various kinds of aircraft control sticks.

The invention discloses an airplane steering wheel multi-axis force measurement component and a measurement method thereof, which belong to the aeronautical engineering test detection field. The measurement component consists of a control handle, a horizontal force signal detection unit, a vertical force signal detection unit and an installation connection unit, which are combined horizontally in turn. The force signal detection units and the control handle are respectively arranged at the two sides of the installation connection unit which serves as the center. The control handle is arranged at the outmost end and the two ends of the force signal detection units are respectively connected with the installation connection unit and the control handle. The measurement component is connected with the steering wheel through the installation connection unit. Control force is exerted onto the steering wheel through the control handle and controls the movement of the steering wheel. The force signal detection units measure the control force. The invention has the advantages of compact and simple structure, reliable work, high measurement precision and convenient installation, use and adjustment, accords with the actual control situation of the airplane steering wheel and is applicable to the measurement of the control force of various airplane steering wheels.

The invention relates to a method for producing a processing product made by aluminum alloy or an integral multifunctional structural element made by the aluminum alloy. The method comprises the step of thermal process, being characterized in that the method also comprises a step of processing cold plasticity deformation after the thermal process. The structural element imposed by the difference between the two regions at least at least 2% of the average generalized plastic deformation, at least 3% of the difference between preferred. Preparation of this invention can be used for aeronautical engineering in particular, the structure of components, components in the space above with the use of variable nature, and at the same time with the existing components have the same shape characteristics. The present invention method and has a controlled economy, it did not temper the major components to enable the use of the nature of change.

The invention relates to aeronautical engineering, in particular to methods of flight due to creating of forces on air propellers, namely using thrust vectoring of direction and amount of force created by air propellers of opposite rotation with the axis, mainly in the direction of flight, in the expanded range of speeds, from 50 m / s to high near-sonic speed of flight. The invention may be applied for horizontal flight and maneuvering in flight on vertical take-off aircrafts using rotation of main rotors from the vertical stand of rotor axes during the take-off to almost horizontal position of rotor axes in horizontal flight, thus both the rotation of rotors and the change of position of an aircraft together with its rotors may be applied, and it may be also used in the horizontal take-off aircrafts with almost horizontal position of axes during the horizontal flight, including the planes with airscrew propellers.

The invention relates to an inflation fatigue test loading device, aiming at solving the technical problems in the technical field of existing aeronautical engineering that fatigue damages are easy to generate by a structure influenced by inner and outer pressure difference. The inflation fatigue test loading device comprises a filter, reducing valve and lubricator (FRL) part, a two-position three-way solenoid valve, a three-position five-way solenoid valve, a pipeline throttle valve, a three-position five-way solenoid valve, a safety valve and a ball valve; the two-position three-way solenoid valve is connected with the filter, regulator and lubricator part through a pipeline; the three-position five-way solenoid valve is connected with the two-position three-way solenoid valve through a pipeline; the ball valve is connected with the three-position five-way solenoid valve through a pipeline; the safety valve is mounted on a front end pipeline of the ball valve; an input end of the pipeline throttle valve is connected with a pipeline at the front end of the three-position five-way solenoid valve; an output end of the pipeline throttle valve is connected with the input end of the three-position five-way solenoid valve; the output end of the three-position five-way solenoid valve is connected with a pipeline at the rear end of the three-position five-way solenoid valve. The inflation fatigue test loading device is widely applied to the technical field of the aeronautical engineering.

The invention discloses on-line detection equipment and an on-line detection method for a fly-by-wire flight control system, and belongs to the field of aeronautical engineering testing. The on-line detection equipment is characterized by consisting of a signal transmission and disconnection device and flight testing interface equipment, wherein the signal transmission disconnection device is used for completing transmission and disconnection of related crosslinking signals of a fly-by-wire flight control computer; the signal transmission disconnection device is of a modular combined structure and consists of a plurality of unit modules which are relatively independent in structure and relatively complete in function; each independent unit module can complete transmission and disconnection of similar or approximate signals; the flight testing interface equipment consists of an operation controller, a flight testing data bus interface plate, an onboard data bus interface plate, an analog quantity signal interface plate, a switching value signal interface plate, a digital value signal interface plate, a tester bus or a computer bus and detection application software, and is used for extracting, analyzing and displaying and output parameter information of the fly-by-wire flight control computer in various tests on the fly-by-wire flight control computer or the fly-by-wire flight control system.

The invention discloses a testing system and method for an automatic flight control system of an airplane, and belongs to the field of aviation engineering tests. The testing system is composed of an automatic flight control signal transmission disconnection device, automatic flight control system test interface equipment, an airborne cross-linking data bus coupler, a flight simulation system, an airborne cross-linking equipment simulator, a field visualization operation guide unit and simulation test application software. Simulation of airplane motion characteristics in an automatic flight control system test and simulation of cross-linked airborne equipment are carried out by taking an automatic flight control device, an automatic flight control computer, a flight control system bus coupler, a fly-by-wire flight control computer and a fly-by-wire flight actuation system as objects; and operation guidance in the test of the automatic flight control system is completed in a visual form, so that standardized operation of the test is realized, and the test efficiency of the test is improved.

The invention belongs to the technical field of aeronautical engineering and particularly relates to a dynamics simulation model of an aircraft arrestment landing process. The method is characterized in that a many-body dynamics theory is combined, a complete aircraft arrestment landing dynamics mathematic model is established, and arrestment landing use conditions of various typical landing attitudes are comprehensively considered, so that a complete aircraft arrestment landing dynamic characteristic analysis method of the system is established. The method has the beneficial effects that by virtue of dynamics simulation calculation, the aircraft landing tailhook load and the load impacting an aircraft body are obtained, reference is provided for the aircraft arrestment characteristic analysis, and meanwhile a basis is provided for designing a tailhook; the modeling time can be saved, the manufacture and test expenses of a test piece are reduced, the working efficiency is improved, and the method has important engineering application value.

The invention discloses a method for generating an aircraft operating characteristic curve, and belongs to the technical field of aeronautical engineering testing. The method comprises a first step ofperforming data smoothing or digital filtering processing on an original recorded time series operating force data set and a corresponding operating displacement measurement data set; a second step of extracting slope varied data segment points in the measurement data series; a third step of segmenting measurement data according to the data segment points extracted in the second step, fitting themeasurement data of each segment by using a least-squares method to obtain the slope kj and intercept bj of the segment of straight line, and using Lj(Kj,bj) for denoting the j-th straight line; anda fourth step of calculating a point of intersection Rj(xj,fj) of the adjacent two straight lines Lj(kj,bj) and Lj-1(kj-1,bj-1) until the calculation of the points of intersection of all the straightlines is completed and the points of intersection of all the straight lines are obtained, and connecting the points of intersection of the straight lines in sequence to obtain an operating characteristic curve composed of the plurality of straight lines. The method of the invention is not subject to the subjective will of the test personnel, and the fitting effect is good.

The invention discloses a springback type angle displacement measurement device, and belongs to the field of aeronautical engineering test detection. The device is characterized by comprising an angle displacement detecting unit, a springback reset unit, a transmission connecting unit and an installation locating unit, wherein a thrust end of the springback reset unit is fixed at a bearing fixed end of the angle displacement detecting unit; the other thrust end is fixed at a rocker arm of the transmission connecting unit; a tail end of the rocker arm of the transmission connecting unit is fixed on a rotary shaft of the angle displacement detecting unit; an end of the other rocker arm is in contact with an upper surface of the installation locating unit; and the control surface drift angle is measured in a way that the transmission connecting unit follows deflection of the airplane control surface and drives the rotary shaft of the angle displacement detecting unit to deflect. The device has the characteristics of compact and simple structure, small space size, low cost, reliability in operation, high test precision, accordance with the actual motion situation of the airplane control surface, and convenience in installation, usage and adjustment, and is suitable for measurement of drift angles of various airplane control surfaces.

The invention belongs to the field of software safety engineering of aeronautic equipment complex software engineering, and discloses a risk indicator indexing matrix-based aeronautic equipment software importance grading method. Through combining characteristics of domestic aeronautic engineering and considering operability of practical engineering application, the risk indicator indexing matrix-based aeronautic equipment software importance grading method is put forward; and according to the method, two factors in are considered when system risk analysis is carried out: a system risk severity grade and a system risk software control grade. Risk importance is considered from two aspects: influence on system safety (casualties, equipment damage and environment destroy), and influence on mission functions (flight missions and operational missions), so as to provide basis for aeronautic equipment mission system software importance grading. The factor of risk occurrence possibility whichis difficult to calculate or assessment in engineering is not considered, so that the practicability and operability of engineering application are improved.

The invention relates to a testing system and method for an aircraft management system, and belongs to the field of aviation engineering test. The testing system comprises interface test equipment 1, a bus coupler 5, bus monitoring equipment 6, a bus analyzer 7, airborne crosslinking equipment 8, a network interface coupler 9, airplane model simulation equipment 10, a cabin control system 11 and visualization equipment 12. The testing system takes an aircraft management system computer 2, a flight control system 3 and an electromechanical control system 4 as test objects, simulates node communication, mode switching, redundancy management and control strategy related functions in an aircraft management system test, guides various operations in the test through visual equipment, realizes standardized operation of the test, and improves the test efficiency.

The invention relates to solid-state laser gyroscopes for measuring rotational speeds or angular positions which are used, in particular for aircraft engineering. A gyroscope performances depend on the temperature stability of the scale factor thereof (S) which is 4A / L, wherein L and A are the optical length and the area of the laser cavity, respectively and is a laser emission mean wavelength without the Sagnac effect. Conventionally, for gas lasers, each parameter of the scale factor is selected in such a way that it is temperature independent. The solid state lasers which are very different from the gas lasers are not treatable in the same manner. Said invention offers the laser cavity optical architecture based on the scale factor global conservation, wherein each parameter is temperature variable. An architecture which makes it possible to avoid optical mode hopes and several digital examples proving that inventive principles are applicable to standard materials are also disclosed.

The invention discloses a motion-linked cabin layout converting device for airplane simulator, belonging to the technical filed of aeronautical engineering, comprising a control mechanism A, a control mechanism B, a connection mechanism A, a connection mechanism B, a rocker arm A, a rocker arm B, a rocker arm C, a rocker arm D, a rocker arm spout A, a rocker arm spout B, a rocker arm spout C, a rocker arm spout D, a longer connection rod A, a longer connection rod B, a shorter connection rod A, a shorter connection B and a clamping device. The rocker arm A and the rocker arm B are vertically installed on a right base / back base simulator support; the rocker arm C and the rocker arm D are vertically installed on a left base / front base simulator support. The invention provides a motion-linked cabin layout converting device for airplane simulator, aiming at the fixed tandem or parallel cabins of the existing traditional helicopter simulator control cabin, which not only can save large amount of time and manpower and resources cost, but also perform man-machine interaction experiment research and adjustment according to actual situation.

The invention discloses a boarding gate operating mechanism reliability test automatic control method and device, and belongs to the field of aviation engineering technology. The method comprises the steps: firstly employing the movement of articulation points on a boarding gate to replace the overall movement of the boarding gate; secondly calculating the moving tracks of the articulation points; thirdly giving the movement instruction for all articulation points at the same moment according to the moving tracks of the articulation points; fourthly starting a driving device to drive the movement of all articulation points; finally carrying out the above process repeatedly according to the time sequence, and completing the operation of controlling the boarding gate to open and close till a boarding gate operating mechanism reliability test is completed. The articulation points comprise a handle shaft, a hinge movable shaft, and a hinge fixed shaft. A driving device for driving the handle shaft makes synchronous movement with a device for driving other articulation points. The invention provides a matched device according to the above method, and the device employs automatic control to solve a problem of repeated manual operation of a testing worker.

The invention belongs to the technical field of aeronautical engineering, and particularly relates to an air and hydraulic pressure system with a power-controlled unmanned aerial vehicle launcher. The air and hydraulic pressure system comprises an oil tank, a hydraulic pump, a one-way valve assembly, an energy accumulator and a motor assembly, all of which are sequentially connected. The motor assembly is connected back to the oil tank, and a loop is formed. According to the air and hydraulic pressure system, the dynamic responsiveness of a launching device is increased through a direct drive roller structural form, and the beneficial effects of being sufficient in power and high in speed are achieved. Compared with current unmanned aerial vehicle air and hydraulic pressure launchers, no pulley speed increase system is adopted, so that the mechanical structure of the launcher is simple. Meanwhile, the single hydraulic loop is adopted in the air and hydraulic pressure system, redundant design is carried out on key control elements, effective integration of the air and hydraulic pressure system is achieved, and the reliability of the air and hydraulic pressure system is improved.

The invention belongs to the technical field of aviation engineering, and in particular relates to an airport power-assisted take-off and landing device dragged by a superconducting outer rotor motor,which comprises a transverse pull rope, a fixed pulley assembly symmetrically arranged at two sides of an airport runway and a winding drum assembly symmetrically arranged at the two sides of the airport runway, wherein the transverse pull rope sequentially bypasses the winding drum assembly, the fixed pulley assembly at the same side, the fixed pulley assembly at an opposite side and the windingdrum assembly at the opposite side; the transverse pull rope spans the airport runway; the winding drum assembly is driven by the superconducting outer rotor motor; during take-off, a towed vehicle pulls the transverse pull rope and hangs into an aircraft lanyard mechanism, and an aircraft engine and the superconducting outer rotor motor cooperate to start power-assisted take-off, thereby increasing stability of the aircraft, shortening an aircraft taxiing distance, reducing an airport runway length, reducing fuel consumption, and reducing environmental pollution near the airport; when landing, the transverse pull rope is hung through a tail hang mechanism of the aircraft, the superconducting outer rotor motor is dragged to generate electricity, the aircraft slows down and land, and kinetic energy of the aircraft is recovered into electric energy, which is more environmentally friendly.

The invention relates to a simulation platform and a test method for an airspeed anomaly auxiliary function of a large aircraft, and belongs to the field of aviation engineering test. The system comprises an information display device 1, an interface test device 3, a bus interface coupler 4, a cockpit operating system 5, a pilot interface unit 6, a bus monitoring device 7, a data analyzer 8, a signal conversion device 9, an aircraft model simulation device 10, an airspeed parameter injection device 11 and a visualization device 12.

The invention belongs to the field of high polymer materials, and particularly relates to low-volatile organic compound (VOC) high-strength high-toughness nylon and a preparation method thereof. The low-VOC high-strength high-toughness nylon is prepared by treating low-flexibilizer-content nylon by a supercritical fluid technology, wherein the low-flexibilizer-content nylon component comprises thefollowing components: 90-100 parts by weight of nylon, 1-10 parts by weight of a flexibilizer and 0.1-1 part by weight of an assistant. The invention overcomes the defects that the addition amount ofa flexibilizer in toughened nylon in the prior art is too high, then the mechanical property of the toughened nylon is greatly reduced, meanwhile, cost is increased along with the increase of the addition amount of the flexibilizer, and VOC (volatile organic compounds) also can be dissipated to cause large danger to human bodies. The mechanical property of the original nylon can be effectively maintained under the condition that the content of the flexibilizer in the toughened nylon is reduced, an ideal balance effect between toughness and rigidity is achieved, and meanwhile the advantages ofeffectively reducing cost and reducing the dissipation amount of VOC are also achieved. The low-VOC high-strength high-toughness nylon can be mainly applied to the fields of food, automobile engineering, aviation engineering, buildings and the like.

The invention discloses a novel landing-gear-drop-test driving rotating device and a test method, belongs to the technical field of aeronautical engineering, and relates to a driving rotating mechanism in a landing-gear drop test. The novel landing-gear-drop-test driving rotating device is mainly composed of a friction wheel, a motor, lifting threaded rods, supports, guiding rails, an actuator cylinder, an adjusting air valve, a motor rotating speed controller and the like. The device is used for providing the pre-rotating speed for airplane wheels in the landing-gear drop test; according to the actual requirement, the motor and the type of a driving rotating wheel can be replaced, the height of the motor is adjusted, the rotating speed of the airplane wheels is controlled, and the test shows that the novel landing-gear-drop-test driving rotating device is stable and reliable.

The invention discloses a flight control system onboard test platform and a method thereof, and belongs to the field of aviation engineering tests. The system is characterized by comprising a fly-by-wire flight control signal disconnection unit, an automatic flight control signal disconnection unit and a high lift control signal disconnection unit which are used for completing airborne signal detection of a fly-by-wire flight control subsystem, an automatic flight control subsystem and a high lift control subsystem, a fly-by-wire flight control system FTI, an automatic flight control system FTI and a high lift control system FTI which are used for completing airborne parameter extraction, storage and analysis of all subsystems, a system open-loop control unit and a frequency response analysis system which are used for completing stability storage and modal coupling tests of the flight control systems; and a flight simulation system, an airborne crosslinking equipment simulation system, a test data processing unit and a comprehensive control management system which are used for completing closed-loop tests of the flight control systems.

The invention relates to the field of aeronautical engineering dynamics, and discloses a modular closed low-speed direct-current wind tunnel. The wind tunnel comprises a tunnel body structure and a power system, and is characterized in that the tunnel body structure comprises a tunnel body bracket, and the tunnel body bracket is sequentially provided with an air inlet section, a test section and an expansion section; the power system comprises a draught fan, a movable sliding rail and a power support; the tunnel body structure is in butt joint with the power system through the movable slidingrail; fan connecting lugs are welded to the two sides of the outer wall of the fan correspondingly and connected with the movable sliding rail through bolts; a plurality of through holes are formed inthe upper surface and the lower surface of the movable sliding rail respectively, and the movable sliding rail is connected with the power support. The whole structure has the characteristic of modularization, the performance is reliable and stable, the use and operation are very convenient and fast, the practicability and adaptability are high, the hole body structure and the power system both adopt independent movable brackets, and the weight is low; and the tunnel wall adopts a streamline design meeting aerodynamic requirements, so that the whole operation is more energy-saving and environment-friendly.

The weaving loom of the invention is of the kind comprising a warp thread supply system (1), a frame type warp thread sequencing system (2), a reed (9) and a weft thread insertion system (10). The sequencing system (2, 25, 40) comprises means for driving side-by-side frames (27, 28; 41‑43). Application to the production of woven preforms in a single piece for a composite component, notably for aeronautical engineering.

The invention belongs to the technical field of aeronautical engineering, and particularly relates to a method for calculating longitudinal damping force generated by upward raising of a tail hook after cable hooking in the airplane cable hooking process. The method is characterized in that through studying the moving characteristics of the tail hook after the airplane cable hooking, using the airplane general parameters and analyzing the dynamical characteristics when the tail hook collides with the pavement, a tail hook moving dynamical equation is built, and a tail hook longitudinal damping force analysis method is obtained. The method is favorable for engineering personnel to control through parameters, so that the expected longitudinal damping force can be realized.

The invention belongs to the technical field of aviation engineering, and particularly discloses a take-off device for an ornithopter. The take-off device comprises a transverse cable, two pulley components, and two drum components driven by outer rotor motors; the two pulley components and the two drum components are arranged in symmetry relative an airport runway; the transverse cable is sequentially wound around one drum component, the fixed pulley component on the same side, the fixed pulley component on the opposite side and the other drum component on the opposite side. The transverse cable crosses the airport runway, and the drum components are driven by the outer rotor motors. During take-off, the transverse cable is pulled by a towing vehicle to be hung on a padlock mechanism at the front end of the ornithopter, an ornithopter engine and the outer rotor motors are cooperatively started to assist the take-off, and the outer rotor motors help provide the initial speed of the ornithopter, so that the stability of the ornithopter is enhanced, the sliding distance of the ornithopter is shortened, and an ornithopter runway is shortened; by the design of the pulley components, the take-off direction can be selected according to the actual wind direction; through electric power assisted dragging, the fuel consumption is reduced, and the surrounding environmental pollution of an airport is reduced.