36 results about "Aviation" patented technology

The invention relates to a counter-flow flame combustion chamber which has a monocyclic cavity structure. The counter-flow flame combustion chamber is designed by adopting a concept of staged combustion, and combustion gas is supplied from a pre-combustion stage and a main combustion stage, wherein the pre-combustion stage adopts a diffusion combustion structural mode with stable whirlblast, and the main combustion stage adopts a combustion structural mode of combining premixing and pre-evaporation with counter-flow flame. The counter-flow flame combustion chamber mainly comprises a shunt diffuser, an external combustion chamber case shell, an internal combustion chamber case shell, a fuel nozzle, a head pre-combustion stage, a flame tube external wall main combustion stage, a flame tube internal wall main combustion stage, a flame tube external wall and a flame tube internal wall, wherein the head pre-combustion stage utilizes stable flame in a low-speed backflow area, which is generated by whirlblast air entering the combustion chamber from a pre-combustion stage whirlblast assembly; the main combustion stage comprises the flame tube internal wall main combustion stage and the flame tube external wall main combustion stage, and uniform oil-gas mixture jetted from the main combustion stage counters in a flame tube to form a counter jet flow which forms counter-flow flame under the ignition of pre-combustion flame. The invention has simple structure, and can effectively reduce disposal of pollutants besides ensuring the normal work condition of an aeroengine.

The invention discloses a device and a method for assisting airplanes to conduct airline extinguishment. The device is an onboard airline electronic system, comprising a navigation and combustion source sensor, a database system, a voice communication system, a data link, a human-computer interface, a display system, a data interface system and a comprehensive task processing system, wherein all systems are connected with a comprehensive task processing system. The device is utilized to conveniently formulate a flight plan on flying to the scene of fire, an extinguishment strategy flight plan and a round-trip water-drawing flight plan. Meanwhile, the device can monitor and alarm fire behavior display, extinguishment material spray suggestion, extinguishment material spray prediction and extinguishment flight tasks. The method and the device for implementing airline extinguishment researched by the invention can effectively assist pilots to conduct extinguishment, improve extinguishment success rate, save cost, and strengthen safety in the process of executing the tasks.

The invention relates to a periodic structure which is based on the local resonance band gap characteristic and has the low-frequency vibration suppression performance, and belongs to the technical field of aviation, aerospace and ship vibration noise control. The periodic structure comprises mass components, elastic components and a base body, and the mass components and the elastic components are periodically distributed. A periodic unit comprises metal structures, namely, the mass components. The side surface of each mass component is surrounded by a circle of rubber or silicone annular structure, namely, the corresponding elastic component. The side surface of the circumference of each mass component is a curved surface with a certain shape. Each elastic component is attached to the corresponding mass component and then surrounds the mass component to serve as a whole period to be arranged in the elastic structure base body. The elastic components are also attached to the base body. The periodic structure has the obvious band gap characteristic. The periodic structure based on the local resonance band gap characteristic has the good low-frequency vibration suppression performance.

A aerospace machine element product (1) to be managed comprises a plurality of assembled components (2) ((1) to (3)). Each of the plurality of components (2) ((1) to (3)) is manufactured through a material purchase process (S1), a forging process (S2), a heat treatment process (S3), a grinding process (S4), surface treatment process (S5), and inspection process (S6). The information of the processes (S1) to (S2) is recorded on an IC tag (4) together with the lot numbers of the processes. After the components (2) have been assembled in the aerospace machine element product (1), the recorded information regarding the components (2) on the IC tag (4) is recorded in a management computer system (11) so as to be related to a manufacturing number or a manufacturing lot number. Another IC tag (10) is attached to the aerospace machine element product (1) and the manufacturing number or the lot number is recorded on the IC tag (10). In addition, processing condition information about the forging, heat treatment, grinding, surface treatment processes is recorded on this IC tag (10).

The invention discloses an aero-engine turbine blade reliability evaluation method based on fracture mechanics. The method comprises the following steps: S1, establishing a simplified model of a bladesection by taking a turbine blade as a research object, calculating a stress intensity factor of the turbine blade containing an initial I-type crack at a blade root by applying a finite element method, substituting the stress intensity factor into a shape factor calculation formula, and determining a shape factor parameter; S2, establishing a turbine blade reliability model based on the generalized stress intensity interference model; and S3, establishing a probability density function of load and fracture toughness, establishing a probability life model and a reliability model for the bladein combination with a Paris formula, and solving the model to obtain the change condition of the reliability of the turbine blade along with the life. According to the method, initial defects existing in the blade are considered, on the basis of the initial defects, the crack propagation life of the I-shaped cracks on the surface of the blade under the cyclic load is established, and life prediction and reliability evaluation of the blade can be more accurate.

The invention discloses a design method for GH4169 alloy ring preforming blank with a tapered section, which solves the problems that the blank material with a rectangular section has low utilization ratio, the subsequent machining time is long, and the streamline structure of a ring is incomplete in the past, and on the like. The design method adopts the basic ideas that proper rolling proportion is selected according to the principle of the unchanging volume and the structural characteristics of the ring with the tapered section during the plastic deformation of metal, so that the structural sizes of parts of the preforming blank are obtained. The design method is simple, can effectively improve the utilization ratio of material, reduces the working hours for subsequent machining, can keep the complete inner streamline of the ring after machining is finished, and is beneficial to improving the performance of an aircraft engine and prolonging the service life.

The invention belongs to the field of aviation testing technology and relates to a performance testing method and testing hanger for a whole set of vibration isolator of an airplane APU (Auxiliary Power Unit). The testing method comprises the following steps: hanging the APU on the testing hanger; installing a testing sensor; and testing the performance of the whole set of vibration isolator. The testing hanger comprises a bottom frame made of hollow square steel, a rear vertical frame, a front vertical frame, inter-frame reinforcing ribs, a top longitudinal beam (4) and a support base fixed on the testing hanger. The performance testing method and the testing hanger can be used for testing the performance of the whole set of vibration isolator of the airplane APU in a simulated installing state and assessing the vibration isolating performance of the whole set of vibration isolator accurately, thereby laying down a basis for optimizing the installation structure of the APU.

The invention discloses a control method for a VIENNA rectifier of an aviation multi-electric engine. The control method comprises the following steps: establishing a stationary coordinate system of the VIENNA rectifier and obtaining a mathematical model under a two-phase rotating coordinate system by coordinate transformation; taking a DC side voltage as a control variable, and designing a sliding mode surface based on a global terminal sliding mode theory to design a voltage external loop controller; and taking the DC side voltage and a derivative thereof as input of fuzzy reasoning, and adjusting the sliding mode reaching law in real time according to the output sliding mode in a system state so as to control the DC side voltage. According to the control method disclosed by the invention, the robustness of the system can be improved by virtue of the global terminal sliding mode theory and an error is restrained to be zero within a definite time; the sliding mode reaching law is controlled in real time according to the experience of an expert by virtue of a fuzzy reasoning theory, so that not only is the quickness of system response improved, but also buffeting is effectively weakened.

The invention discloses a test method for testing a P-S-N curve of high-cycle fatigue of an aero-engine material by a small sample amount. The method comprises the steps that same probability distribution points of fatigue lives of same samples under different stress levels in respective parent distribution meet an expression of a formula (4) defined in the specification, and in the formula (4), nji is logarithmic life of the sample i under the jth stress level, mu j is a logarithmic life mean under the jth stress level, sigma j is a logarithmic life standard deviation under the jth stress level, and meanings are same when a subscript j is changed to k; the life sample points under the different stress levels have a one-to-one corresponding relationship; according to the corresponding relationship in the formula (4), the fatigue life on a stress level is equivalent to another stress level, so that sample information aggregation is realized; equivalent large samples are obtained througha small sample test; and according to equivalent large sample data, life distribution parameter statistics is performed. According to the method provided by the invention, based on a sample aggregation principle, fatigue life distribution can be obtained through a small amount of test data, and the relatively accurate P-S-N curve is obtained.

The invention belongs to the technical field of aerospace engineering, and relates to an engine test system and method, in particular to an engine test collecting-distributing type control system and method. The engine test collecting-distributing type control system comprises an upper computer based on the windows system, a plurality of lower computers based on the Linux system, an network switch for achieving interaction between the upper computer and the lower computers, a corresponding driving unit and a corresponding acquisition unit. The engine test collecting-distributing type control system guarantees accurate and real-time transmission of data through the methods that the upper computer and the lower computers adopt a TCP/IP communication protocol, the inventor defines fields of a protocol packet, and the protocol packet adopts IP addresses and a packet head which are bound for testing. By means of the engine test collecting-distributing type control system and method, the problems of high cost and difficult later maintenance and function extension and integration due to the fact that an engine test control system in the technical field of aerospace engineering needs to be imported are solved.

The invention belongs to the technical field of aero-engine design, and particularly relates to an interpolation method suitable for an aero-engine real-time model. The method comprises the followingsteps: firstly, recording an interpolation algorithm of a current rotating speed position, a current pressure ratio position and a current adjustable blade angle position and the slope; performing first interpolation according to the current rotating speed; obtaining four discrete points, carrying out second interpolation according to the current adjustable blade angle; obtaining two discrete points, and finally, carrying out third interpolation according to the current voltage ratio to obtain the flow and the efficiency of the current working point; recording the interpolation position and the slope of the previous period in the difference calculation process of the next working point, judging whether the position value and the slope value need to be updated or not. Repeated circulation caused by searching is reduced, the interpolation efficiency in a real-time model is improved, and the real-time model running speed is increased.

The invention discloses an adaptive multiple input multiple output method and system for aviation communication. The adaptive multiple input multiple output method for the aviation communication comprises the following steps of: 1, establishing a three-dimensional multiple input multiple output channel model; 2, computing optimal channel capacity; 3, acquiring a distance between an airplane and a ground antenna array and a direction angle of the airplane relative to the ground antenna array; 4, computing a distance dr between every two adjacent antennae in a virtual receiving antenna array on an airplane body according to the optimal channel capacity, the distance and the direction angle by using an adaptive minimum mean square error interpolation algorithm; and 5, computing a distance dt between every two adjacent antennae in a virtual transmitting antenna array on the airplane body according to the optimal channel capacity, the distance and the direction angle by using an adaptive delay algorithm. According to the adaptive multiple input multiple output method and system disclosed by the embodiment of the invention for the aviation communication, maximum system capacity can be provided according to the position of the airplane in real time.

The invention discloses a recovery method of an alkyl-halide fire extinguishing agent in a fire extinguishing bottle for aviation, which belongs to the technical range of recovery of fire extinguishing agents. The recovery method comprises the following steps of: firstly, establishing a fire extinguishing agent recovery system which enables a fire extinguishing agent to be liquefied by reducing the temperature and increasing the pressure simultaneously; setting the temperature of a refrigeration house to be -40 DEG C and the volume of a pressure storage tank to be 35L according to practical needs; placing the pressure storage tank in the small refrigeration house or refrigerator, placing the fire extinguishing bottle outside the small refrigeration house or refrigerator; connecting the pressure storage tank to the fire extinguishing bottle and the fire extinguishing agent storage tank respectively; opening an inlet shutoff valve to release the fire extinguishing agent in the fire extinguishing bottle to the pressure storage tank, wherein the pressure storage tank is higher than the fire extinguishing agent storage tank, and the liquid fire extinguishing agent in the pressure storage tank flows towards the fire extinguishing agent storage tank under the action of gravity; and automatically recovering according to specific pressure. The recovery method diclosed by the invention has the advantages that the fire extinguishing agent discharged into the atmosphere is obviously reduced, so that the environment is protected, and the purposes of reducing the waste of the fire extinguishing agent and improving the recovery efficiency of the fire extinguishing agent are fulfilled.

The invention relates to a manufacturing method for a high-speed permanent magnetic rotor. By performing the processes of assembling, polishing, dynamic balance and the like, and setting reasonable parameters, it can be ensured that magnetic poles of the motor rotor are firmly bonded on a rotor iron core at a high speed of 40,000 rpm. The permanent magnetic rotor manufactured with the method can work under a condition that the rotational speed is 50,000 rpm for a long term, is successfully applied to power motors of digital electronic control systems of aircraftengines, and meets operating requirements of high rotational speeds of aviation motors.

Provided in the invention is a steady-state-temperature-field-included simulation test bench of an aero-engine combustor. The simulation test bench comprises a box body, a combustion chamber simulation device, a test device, an air supply system, a ventilation system, a cooling system and a control system. The ventilation system includes an air suction type exhaust hood installed inside the lowerbox body and a wall-mounted industrial ventilation fan installed at the top of the box body. The air supply system and the cooling system are installed inside the lower box body. The control system includes a controller installed inside the lower box body and a human-machine interaction device installed on the outer side of the box body; and the controller is connected with the wall-mounted industrial ventilation fan, the human-machine interaction device, the combustion chamber simulation device, the test device, the air supply system, and the cooling system respectively. The simulation test bench can be used for simulating the actual working condition of the aero-engine combustion chamber; and the enabling performance of the thermal barrier coating material is studied preliminarily beforerun trial examination of the ground track. The steady-state-temperature-field-included simulation test bench has advantages of low cost, short cycle and low risk and the like.

The invention relates to a chromizing and aluminizing process for an aero-engine blade, and belongs to the technical field of aviation equipment. The blade obtained through the method of sand blowing,cleaning, blending, stirring, protecting, loading to a tray, loading to a rack, loading to a can, exhausting, heating, furnace charging, chromizing and aluminizing, cooling, taking out of the can, cleaning, polishing and brushing, citric acid washing, cleaning, blow-drying, inspecting, diffusing and final inspecting has the advantages that the surface state of a permeated layer is uniform and consistent, and the depth of the permeated layer is uniform and consistent.

Optimizing an "improper" combination, a combination consisting of an even plate or shell and a conceptually irregular set of reinforcing stiffeners, enables to attain said combination's hitherto unknown unique synergetic attribute, the attribute of total disappearance of local buckling phenomena of said combination when compressing thereof up to destruction of structural material(s) thereof, and therefore enables to utilize strength potentialities of the material per se utterly, just us utterly as when stretching thereof, and besides irrespective of plates/shells shape and size, irrespective of structural materials' kind, volume and weight, and irrespective of operational temperatures and type of external load distribution, etc. Accordingly, given equal other conditions, cardinal weight reduction or bearing capacity's raise of thin-walled structure's bearing skin is attainable that can be utilized for any aerospace thin-walled structure, very extensively in naval engineering, and in civil engineering sometimes also; besides a reduction of expenditures, more efficient utilization of energy resources and hence enhancement of the environment of mankind is provided too. I call said synergic attribute "Superstability".

The invention discloses a multi-rotor aircraft with an electric wire supplying power and application of the multi-rotor aircraft, belongs to the technical field of aviation, and particularly relates to a combined-type aircraft. The combined-type aircraft comprises a fixed wing aircraft, the multi-rotor aircraft and a ground power supply device; the structure of the fixed wing aircraft is similar to that of a regular fixed wing aircraft, and the fixed wing aircraft is further provided with a connecting end; the multi-rotor aircraft comprises a machine body, a connecting device, a radar detecting device and a power supply distributor; the ground power supply device comprises a controller, an electric wire collecting and putting implement and the electric wire; one end of the electric wire is fixedly connected to the power supply distributor, the other end of the electric wire is connected to a power source and is twined on the electric wire collecting and putting implement, and the electric wire transmits power supply for the multi-rotor aircraft; the multi-rotor aircraft flies in the sky with the electric wire, and the connecting device joints the connecting end so as to make the multi-rotor aircraft and the fixed wing aircraft connected and combined to load or carry the fixed wing aircraft. According to the multi-rotor aircraft with the electric wire supplying power and the application of the multi-rotor aircraft, the power source is provided for the multi-rotor aircraft through the electric wire, the combination of the multi-rotor aircraft and the fixed wing aircraft is utilized, and a new scheme is provided for the taking off and landing of the fixed wing aircraft.

The invention belongs to the technical field of aviation, and relates to a heating device special for an aircraft oil tank. Currently, most of existing heating devices are each of an insertion type structure, as aviation lubrication oil belongs to combustible liquid, the existing heating devices make direct contact with the lubrication oil, and fire risks are caused. The heating device isolates aheating assembly and the lubrication oil and then conducts heating through the thermal radiation principle. Heating is conducted on the outer side of the oil tank. The heating device comprises a controller and a shell component installed on the outer surface of the oil tank; a plurality of heating pieces are arranged on the outer surface of the shell component; the shell component is made of aluminum alloy materials; and a platinum resistor and a temperature relay can effectively ensure the effect that the heating device always keeps working within the safe temperature range. Efficient transferring of heat energy of the heating device is ensured, the heating device can be independently used or multiple heating devices are used at the same time according to the heating requirements, the structure is simple, dismounting is convenient, and interchangeability is good.

The invention belongs to the field of aircraft engine design, and relates to a design method for comprehensive suppression of infrared and electromagnetic signal features of an aero-engine. The method comprises the following steps: acquiring an infrared signal radiation intensity target value and an electromagnetic signal RCS target value; determining the average temperature T1, meeting the infrared signal radiation intensity target value, of the visible area of the rear cavity of the engine; determining the cooling flow W1 of the visible part of the rear cavity of the engine according to the average temperature T1; determining an engine cooling channel structure and an opening cooling form which meet the temperature reduction requirement; by electromagnetic simulation calculation, determining angle combinations of multiple groups of typical parts of the engine meeting the requirement that the electromagnetic signals are not higher than the RCS target value of the electromagnetic signals; and by iterative optimization, determining the angle combination of the typical parts of the engine with the minimum electromagnetic signal under the condition of engine thrust. According to the invention, the problem of comprehensive suppression of the infrared and electromagnetic signal characteristics of the engine is solved, so that the infrared signals of the engine are lower and the electromagnetic signal characteristics are smaller.

Based on relation of conductivity and rigidity of CZ state 2Al2 aluminum alloy decided by its solvus temperature, cooling rate, and quenching natural aging time, the method detects conductivity and rigidity of the aluminum alloy after natural aging time, and derives solvus temperature of the aluminum alloy in reverse to determine whether overheating exists in the aluminum alloy. Features are: accurate, handy, quick, and nondestructive. The invention can check over whether overheating exists, and also can carry out preliminary filtration for further detection by metallography method. The invention is applicable to nonferrous metals processing, motor, and aviation etc. industries of using 2Al2 aluminum alloy.

The invention relates to an aircraft engine test site digital bus control device. The aircraft engine test site digital bus control device comprises a control system, a communication interface, an output interface and a power supply interface, and the control system comprises a main control circuit board, a central processing unit, an output module, a power supply module, a communication module and a relay module; the power supply interface is connected with the main control circuit board; the communication interface is connected with an engine test bed; the central processing unit is connected with the relay module which is connected with the output interface, and the output interface is connected with engine test bed site test equipment through a cable. The aircraft engine test site digital bus control device can monitor the status of an equipment site in real time and has good performance of safety, stability and anti-interference, intelligence of the control device is achieved better, and the requirement for aircraft engine test equipment site control is met.

The invention provides a lubricating oil sealing structure and an aircraft engine. The lubricating oil sealing structure comprises a grate tooth ring, a sealing shell and two graphite rings, wherein arotating shaft is sleeved with a grate tooth ring, the grate tooth ring is provided with an outer ring surface, and two groups of grate teeth are axially arranged at intervals front and back around the outer ring surface; the sealing shell is arranged around the periphery of the grate tooth ring and is provided with an axial positioning assembly, and the sealing shell is provided with a pre-rotating blade runner; the two graphite rings are respectively sleeved between the grate tooth ring and the sealing shell, are axially arranged front and back in a spaced mode, and respectively correspondto the two groups of grate teeth in a clearance fit mode, the two graphite rings are axially positioned through the axial positioning assembly, and the two graphite rings are of a floating ring structure; and the lubricating oil sealing structure is configured so that external air flow sequentially flows through the pre-rotating blade runner, the gap between the two graphite rings, the gap betweenthe two groups of grate teeth and the gap between the two graphite rings and the two groups of grate teeth and then respectively flows to the spaces on the front side and the rear side of the sealingshell.

The invention belongs to the field of aviation permanent magnet synchronous motor control systems, and discloses an AD2S1210-based aviation permanent magnet motor rotor position angle redundancy control system and method. The space vector control FOC algorithm unit is used for generating six paths of PWM (Pulse Width Modulation) waves to drive the three-phase bridge type inversion module to be switched on and switched off by collecting armature current, rotor position angle and rotating speed information of the permanent magnet synchronous motor and adopting a rotating speed and current double-closed-loop control strategy, so that the permanent magnet synchronous motor is controlled to reach a target rotating speed or torque; the AD2S1210 rotary transformer decoding chip is used for collecting rotor position angle information and rotating speed information of a rotary transformer coaxially installed on the permanent magnet synchronous motor. The rotor position angle information and the rotating speed information are decoded and then sent to the position angle redundancy control algorithm unit; and the position angle redundancy control algorithm unit is used for judging the received rotor position angle information, outputting a correct position angle signal to the space vector control FOC algorithm unit, converting the received rotating speed information and outputting the converted rotating speed information to the space vector control FOC algorithm unit.