67 results about "Aircraft fuel system" patented technology

An aircraft fuel system includes a fuel tank, a fuel pump configured to receive a fuel flow from the fuel tank via a fuel supply line. The fuel system further includes a tap fluidly coupled to an outlet of the fuel pump, at least one generator or permanent magnet machine fluidly connected to the fuel pump via the tap, and the tap is configured to deliver fuel to the at least one generator or permanent magnet machine. The fuel system further includes a gearbox mechanically connected to the at least one generator or permanent magnet machine.

The invention relates to an aircraft fuel system fault isolation and maintenance decision-making method based on a knowledge graph, and the method comprises the steps: employing an equipment tree analysis method to disassemble an aircraft fuel system into a plurality of maintenance units of different levels, and obtaining system knowledge data; obtaining fault knowledge data and maintenance knowledge data of each maintenance unit; constructing a fault isolation and maintenance knowledge graph according to the obtained system knowledge, fault knowledge and maintenance knowledge; and obtaining a to-be-matched fault early warning result output by the fault early warning system, searching a matched entity in the fault isolation and maintenance knowledge graph, and outputting a corresponding fault isolation and maintenance decision. According to the invention, accurate fault isolation suggestions and maintenance decisions can be quickly generated according to fault information, the maintenance work efficiency can be greatly improved, and accurate maintenance of equipment is realized.

A fuel system for a hydrogen fuelled aircraft propulsion system comprises a fuel tank configured to store hydrogen fuel, a main fuel conduit configured to provide fuel to a combustor of a gas turbine engine, a fuel heater comprising a catalytic combustor configured to catalytically combust a portion of the hydrogen fuel prior to delivery to the combustor and a heat exchanger configured to exchange heat between exhaust gases from the fuel heater and hydrogen fuel in the fuel conduit.

The utility model belongs to the field of aircraft fuel systems, and relates to an anti-negative overload oil supply device which comprises an oil tank, an oil delivery pipe connector is arranged at the top of the oil tank, an oil supply pipe connector is arranged on the side face of the oil tank, an oil suction pipe is arranged in the oil tank, and an oil outlet of the oil suction pipe is communicated with the oil supply pipe connector. Due to the fact that the three partition plates are distributed on the periphery of the oil suction port in the oil tank, air can be effectively prevented from entering the oil suction port. Therefore, the negative overload resistant small box can still supply oil continuously under the negative overload condition, and the negative overload resistant oil supply capacity is high.

The invention discloses an aircraft green inerting system with a water removal function. The aircraft green inerting system comprises an oil tank, a catalytic reactor, a water removal device and a control system. The system extracts oxygen and fuel oil steam in an oil tank through a first fan, the oxygen and the fuel oil steam are mixed with ram air, then the mixture is heated through an electric heater and burnt in a catalytic reactor to generate high-temperature inert gas, then the high-temperature inert gas is subjected to water removal through a first water separator, cooling through a heat exchanger, pressurization through a compressor and secondary water removal through a second water separator in sequence, and pressure reduction and temperature reduction are conducted through a cooling turbine; finally, low-temperature dry inert gas is formed and is reinjected into an oil tank through a fuel oil ejector, and the dual functions of inerting and water removal are achieved. The system monitors in real time through an oxygen concentration sensor, a temperature sensor and a flow sensor, dynamically adjusts the electric heating power and the rotating speed of a valve and a fan, and ensures safe and efficient operation. Catalytic combustion and high-pressure water removal technologies are combined, the oxygen concentration of the fuel tank and the water content of fuel are effectively reduced, and the device is energy-saving, environment-friendly, efficient in inerting and suitable for safety protection of various aircraft fuel systems.

The present application belongs to the field of aircraft fuel system, and relates to a kind of integrated design light weight fuel transfer system and method of seaplane. Including Ⅰ, Ⅱ, Ⅲ, Ⅳ, Ⅴ, Ⅵ oil tank on the wing of the aircraft, pressure refueling connector, pressure refueling control valve, oil level controller, pressure refueling control panel, oil pump, low oil level signal. One oil pump integrates the flow of oil transfer system and the flow of emergency oil discharge system two working states, maximally reduces the number of pump;Pressure refueling pipeline, oil transfer pipeline, emergency oil discharge pipeline are integrated designed, reduce the number of pipeline;Thirdly, the low oil level signal of oil tank and the low oil level signal of oil quantity measuring system are shared designed during oil transfer and emergency oil discharge, the on-off control of oil transfer system and the on-off control of pressure refueling are shared designed, the on and off of oil transfer are controlled by pressure refueling control valve, reduce the setting of oil transfer cut-off valve.

The present invention relates to the technical field of aircraft fuel systems, and particularly to an ejector fuel transfer device for a multi-tank aircraft and its fuel supply method, which includes a supply fuel tank, a main fuel tank, a left front fuel tank, a left rear fuel tank, a right front fuel tank, a right rear fuel tank, three ejector pumps and a fuel booster pump; the ejector pumps are respectively arranged in the left rear fuel tank, the right rear fuel tank and the main fuel tank, and the fuel booster pump provides a dynamic pressure source for the ejector pumps to transfer fuel from the left front fuel tank, the left rear fuel tank, the right front fuel tank and the right rear fuel tank to the main fuel tank respectively; the inlet port of the fuel booster pump is split into a first supply oil pipe and a second supply oil pipe, and the ports of the first supply oil pipe and the second supply oil pipe in the supply fuel tank are arranged at different heights; the outlet port of the fuel booster pump is connected to the left engine and the right engine through a main supply oil pipe. Through the present fuel transfer device and its fuel supply method, the problems of high manufacturing cost, complex system setting and low reliability can be effectively solved.

An aircraft fuel system includes a first integrated boost stage having a boost pump configured to receive a fuel flow from a fuel tank via a boost supply line, at least one generator collocated with and operatively connected to the boost pump for driving the boost pump, and a tap fluidly connecting the boost pump and the at least one generator for selectively delivering an amount of the fuel flow to the at least one generator. The fuel system further includes a gearbox operatively connected to the at least one generator, and a main stage downstream of the boost stage and operatively connected to the gearbox. The main stage includes a main pump configured to receive the fuel flow from the boost pump along a main supply line, and a filter disposed along the main supply line upstream of the main pump for filtering the fuel flow.

The present application belongs to the field of aircraft fuel system, and relates to a four-engine aircraft fuel supply and fuel transfer system and method. The scheme combines the pressure fueling system and the fuel supply system without increasing the system development cost too much, can use the cross-feed fuel supply mode to eliminate the imbalance through the consumption of the engine, and can also use the transfer mode to transfer the fuel to other fuel tanks to quickly eliminate the imbalance. The fuel transfer / cross-feed can be automatically controlled by the fuel computer software, or manually controlled by the crew. The system can meet the fuel transfer demand and the cross-feed fuel supply demand, reduces the development and inspection and maintenance cost of the aircraft, and has outstanding economy and practicability.

This invention belongs to the field of testing technology, specifically relating to a dynamic shear test device and method for a flexible connection assembly of an aircraft fuel system. The test device includes a reciprocating assembly, a clamping assembly, and a loading assembly. A second push block in the reciprocating assembly is connected to a first locking block in the clamping assembly. The loading assembly is located on one side of the clamping assembly. The reciprocating assembly applies a shear load to the flexible connection assembly. The clamping assembly is responsible for clamping and fixing the test tube. The loading assembly injects pressure into the test tube. By connecting two test tubes using the flexible connection assembly, and controlling the pressurization module in the loading assembly to pressurize the test tube and the electric cylinder in the reciprocating assembly to apply thrust to the test tube, the internal working pressure and external shear load of the flexible connection assembly are tested. This provides standard equipment and methods for conducting life testing research on flexible connection assemblies, and also provides a reference for the selection of flexible connection assemblies.

The invention relates to a fuel tank suitable for a ground simulation test of an aircraft fuel system and a fixing frame of the fuel tank, the fuel tank comprises four groups of fuel tank test pieces with the same structure and capacity as the fuel tank on an aircraft, and the fuel tank comprises a first fuselage fuel tank test piece, a second fuselage fuel tank test piece and two groups of left and right flank fuel tank test pieces, the four groups of fuel tank test pieces and the test equipment are connected through a test pipeline and a communication line; the fixing frame is of a truss structure, a first fuselage fuel tank test piece installation area and a second fuselage fuel tank test piece installation area are arranged in the truss structure, left and right side wing fuel tank test piece installation areas are symmetrically arranged on the two sides of the truss structure, and connecting flanges fixedly assembled with the test bed outer frame are arranged at the front end and the rear end of the truss structure. And the fixed frame is driven by the test bed outer frame to realize pitching and overturning actions. According to the method, the test cost is reduced, the verification period is shortened, the performance parameters in the limit state can be obtained, and the reliability of the test conclusion is ensured.

An aircraft fuel system includes a first fuel storage tank arranged to store a first fuel, a fuel oxygen reduction unit arranged to generate a deoxygenated fuel from the first fuel, and a second fuel storage tank arranged to store the deoxygenated fuel generated by the fuel oxygen reduction unit, and being arranged to supply the deoxygenated fuel to an engine.

The invention discloses a split type centrifugal pump convenient to maintain, and belongs to the field of aircraft fuel systems, the split type centrifugal pump convenient to maintain comprises a pump body, a main cavity, an auxiliary cavity and a first fuel runner are arranged in the pump body, the main cavity and the auxiliary cavity are communicated through the first fuel runner, a pump core is arranged in the main cavity, and a sliding valve assembly is arranged in the auxiliary cavity; the pump core comprises a shell, a motor, a rotating shaft and an impeller, the motor drives the impeller to rotate through the rotating shaft, an oil inlet hole and a second fuel oil flow channel are formed in the shell, the impeller and the oil inlet hole are oppositely arranged, the second fuel oil flow channel is communicated with the first fuel oil flow channel, and the oil inlet hole is communicated with the second fuel oil flow channel; the side wall of the auxiliary cavity is provided with an oil inlet used for being communicated with an oil tank, the sliding valve assembly is used for cutting off a channel between the oil inlet and the first fuel oil flow channel in the overhauling process and conducting the channel between the oil inlet and the first fuel oil flow channel in the oil pumping process, and therefore the centrifugal pump can be more convenient to disassemble and maintain.

The present application belongs to the field of aircraft fuel system processing, and relates to a cutting device for polyurethane foam shaping and a use method thereof. The cutting device is used for cutting polyurethane foam by using the shape of a nichrome wire model at high temperature. The nichrome wire model with different shapes is placed in a main body, is lifted to slide along a sliding groove at the top of the main body to one side of the main body, and is rotated to be horizontally placed by a motor. A motor is arranged at the lower part of the side of the main body, and an operation table parallel to the nichrome wire model is fixed above the output shaft of the motor. The polyurethane foam is placed on the operation table, and is controlled to move upwards until the polyurethane foam is attached to the nichrome wire model, so that the cutting of the polyurethane foam is completed. The cutting device uses the nichrome wire model to replace a cutter, can realize one-piece shaping, has a smooth cutting surface, and can prevent the generation of excess scraps by using the electric heating mode, so as to avoid the hidden danger of excess substances. The cutting efficiency and cutting quality of the cutting device are greatly improved, and great convenience is brought to the later inspection.

An aircraft fuel system includes a first fuel storage tank arranged to store a first fuel, a second fuel storage tank arranged to store a reduced oxygen fuel, and a fuel tank inerting system arranged to provide a reduced oxygen gas through a fuel contained within the second fuel storage tank to generate the reduced oxygen fuel within the second fuel storage tank. The fuel tank inerting system may include a fuel tank sparging system to inject the reduced oxygen gas directly into the fuel contained within the second fuel storage tank to generate the reduced oxygen fuel.

A method of health monitoring aircraft fuel system structures can include monitoring guided waves emitted by at least one guided wave emitter mounted to an inner hydrogen tank using at least one guided wave detector mounted to the inner hydrogen tank to monitor for a change versus a first baseline. The method can include monitoring guided waves emitted by at least one guided wave emitter mounted to an outer hydrogen tank surrounding the inner hydrogen tank using at least one guided wave detector mounted to the outer tank to monitor for a change versus a second baseline. The method includes outputting an alert to at least one of aircraft avionics and / or an onboard integrated Vehicle Health Management system (IVHM) if a change versus either or both of the first or second baselines is detected.

During a method, an aircraft powerplant is provided. The aircraft powerplant includes an injector and a fuel system configured to deliver fuel to the injector. The fuel system includes a first tank, a first circuit and a second circuit arranged in parallel with the first circuit between the first tank and the injector. The second circuit includes a second tank. The fuel is directed from the first tank to the injector during a first mode through the first circuit and independent of the second circuit. The fuel is directed from the second tank to the injector during a second mode through the second circuit and independent of the first circuit.

Disclosed is a double-wall pipe 10 for an aircraft fuel system. The double-wall pipe 10 comprises an inner wall 11 defining a first open end, a second open end, and an inner region 14 connecting the f

The present application belongs to the technical field of aircraft fuel system, and relates to an automatic device for jet cleaning of an aircraft fuel system and a use method thereof. The present application adopts a form of a stepping motor matched with a rudder to realize rotation of a three-degree-of-freedom mechanism, achieve accurate positioning in space, adopt program setting of a fixed jet route, drive and control a three-degree-of-freedom steering mechanism to perform omnidirectional jet cleaning according to a specified route, can effectively remove excess materials in the fuel tank, simultaneously increase a fuel cooling system device, realize real-time monitoring of fuel temperature by a built-in temperature sensor, and automatically realize a fuel cooling function, provide suitable temperature fuel for the jet device, avoid the influence of high temperature on each system in the fuel tank, and realize automatic and intelligent cleaning. The automatic cleaning device has a simple operation process, complete cleaning effect, improved cleaning efficiency, reduced production cost, and improved economic benefit of the company.

The application belongs to the technical field of aircraft fuel system, and particularly relates to a nitrogen-rich gas flow distribution device for aircraft fuel tank inerting. The device comprises: an air separation device for generating nitrogen-rich gas; a flute-shaped pipe, a fuel tank pressure sensor and a fuel tank oil quantity sensor arranged in the fuel tank; one end of an inerting pipeline is connected with the air separation device, and the other end is connected with the flute-shaped pipe through the fuel tank; a one-way valve, a flow regulating ball valve, an oxygen concentration sensor and a flow sensor are sequentially arranged in the inerting pipeline from upstream to downstream; an upper computer is used for collecting wheel load signals; a controller is used for judging the current flight phase of the aircraft according to the fuel tank pressure and the wheel load signals, and controlling the opening degree of the flow regulating ball valve according to the current flight phase of the aircraft, the fuel tank pressure and the fuel tank oil quantity, so that the nitrogen-rich gas corresponding to the flow is filled into the fuel tank. The application can realize the monitoring of multiple parameters and timely adjust the flow of the nitrogen-rich gas entering the fuel tank.

The invention relates to the field of aircraft fuel oil, in particular to a fault analysis method for a multivariable complex system, which comprises the following steps of: 1, reading all test variable data of a ground simulation test of an aircraft fuel oil system; step 2, inputting fault variables in all test variable data of the ground simulation test of the aircraft fuel system, and determining a correlation coefficient threshold; step 3, performing correlation analysis on each fault variable; according to the method, aircraft fuel system ground simulation test data serves as a sample, aircraft fuel system ground simulation system fault and correlation analysis is achieved around fault variable data, correlation coefficients among fault variables are obtained through calculation, and fault detection is emphasized on strong correlation fault variables.

The invention belongs to the technical field of aircraft refrigerating systems, and particularly relates to an airborne auto-cascade refrigerating system control method and system. The cold source temperature T1 at the fuel supply position of the aircraft fuel system and the heat sink temperature T2 at the liquid return position of the airborne equipment needing temperature control are collected, and when T2-T1 is larger than T02, the airborne auto-cascade refrigeration system operates in an auto-cascade mode; and when T2-T1 is greater than or equal to T01 and less than or equal to T02, acquiring gravity overload or attitude change data of the aircraft, acquiring the temperature T3 and the pressure P3 at the inlet of the gas-liquid separator, constructing an agent model representing the mapping relation of the two-phase inlet state, the gravity condition and the separation efficiency of the gas-liquid separator, and calculating the gas-liquid separation efficiency of the gas-liquid separator through the agent model. The separation efficiency is compared with a preset separation efficiency threshold value, and the airborne auto-cascade refrigeration system is controlled to operate according to an auto-cascade mode or a single-stage compression mode according to a comparison result; when T2-T1 is smaller than T01, the airborne auto-cascade refrigeration system operates in a single-stage compression mode, T01 is the lower threshold value of the difference value between the heat sink temperature T2 and the cold source temperature T1, and T02 is the upper threshold value of the difference value between the heat sink temperature T2 and the cold source temperature T1. According to the invention, efficient and reliable operation of the airborne refrigeration cycle system can be realized in a complex dynamic environment and a wide temperature range working condition.

The invention belongs to the field of ground full-size simulation tests of aircraft fuel systems, and particularly relates to an aircraft pressure refueling and ground pressure refueling rapid switching control system and method. The pressure refueling control system and the ground pressure refueling control system share test resources to the maximum extent, the pressure at the inlets of the pressure refueling connector and the ground pressure refueling connector is accurately controlled, the control systems are rapidly switched, the control resources are rapidly and conveniently configured, the test efficiency is greatly improved, and the test resources are saved.

The invention belongs to the technical field of aircraft fuel systems, and discloses a high-cavitation-resistance self-suction centrifugal pump and an integrated impeller thereof, a hub of the impeller is arranged to be a cone in the axial direction, two long spiral blades are arranged on the conical surface of the hub, and the long spiral blades are coiled from the axial front end of the hub to the axial tail end of the hub; the two short spiral blades are respectively arranged on the hub at intervals of 90 degrees and-90 degrees of the long spiral blades, and the two short spiral blades are coiled from the axial tail end of the hub to the axial middle section of the hub; a circle of balance holes are formed in the axial middle section of the hub corresponding to the front end parts of the short spiral blades; a center hole is axially formed in the center of the hub, and a through hole communicated with the center hole is radially formed in the front end of the hub. The rotary plate valve is arranged at the front end of the pump shell of the fuel pump, so that the pump core is replaced without discharging oil when the pump core is disassembled; and the inducer and the centrifugal impeller are integrally designed, so that the axial size is reduced, and a balance hole is formed in the centrifugal impeller, so that the axial force is reduced.

The invention discloses an aircraft fuel quantity measuring method, device and system and an aircraft, and belongs to the technical field of aircraft fuel systems, and the method comprises the following steps: obtaining first measurement data and second measurement data; substituting the fuel oil temperature and the fuel oil environment pressure into the fitting relation to obtain a fuel oil density calculation value; comparing the fuel density measurement value with a preset density change threshold value, and taking the fuel density measurement value or the fuel density calculation value as a fuel density input value according to the density change threshold value; obtaining a fuel volume input value according to the first measurement data or the second measurement data; and obtaining the fuel quantity according to the fuel density input value and the fuel volume input value. The fuel quantity in the aircraft fuel tank is measured through the optical fiber sensing system, the density is measured through the pressure difference mode and the fitting calculation mode, fuel density double-source redundant backup is achieved, and the reliability of measurement and settlement is improved.

The present invention relates to an aircraft fuel system quick-release connector assembly and an aircraft. The aircraft fuel system quick-release connector assembly includes a male component on the aircraft fuel system side and a female component on the engine side. The male component on the aircraft fuel system side includes an externally threaded portion at one end thereof, a flange located behind the externally threaded portion, and a boss located on the flange. The female component on the engine side includes an inner slidable sleeve, an intermediate sleeve, a second internally threaded sleeve, and an outermost slidable sleeve, which are arranged in sequence from the inside to the outside. The internally threaded portion of the second internally threaded sleeve is used to cooperate with the externally threaded portion to form a threaded connection, and the outermost slidable sleeve is provided with a slot for forming a snap fit with the boss. The female component on the engine side also includes a first return spring and a second return spring. The first return spring is fixed to the second internally threaded sleeve and the outermost slidable sleeve, and the second return spring is fixed to the intermediate sleeve and the inner slidable sleeve.

The application provides an airplane fuel system semi-physical simulation test system and method, and relates to the technical field of airplane fuel system simulation test, which comprises: a host computer, which is used for simulating an airplane avionics system, issuing and receiving ARINC429 signals and hard-wired signals; a real-time simulation machine, which is in communication connection with the host computer, is used for running a fuel system controlled object model, simulating a real airplane fuel system, and calculating simulated fuel tank state simulation signals and sensor simulation signals in the fuel system in real time; a slave computer, which is in communication connection with the host computer and the real-time simulation machine, and comprises a reflection memory board, a communication board and an analog board; and a fuel controller, which is in electrical connection with the communication board and the analog board of the slave computer. The application improves the safety of the airplane fuel system research and development test process, reduces the test bench construction cost, and improves the software and hardware research and development test iteration efficiency.

The invention discloses a jet pump with one-way and oil drainage functions at an outlet, and belongs to the field of aircraft fuel systems, the jet pump comprises a nozzle, a main box body, a jet pipe and a jet seat, the nozzle and the jet pipe are connected to the front side and the rear side of the main box body respectively, the jet seat is connected to the front end of the jet pipe, a suction inlet is formed in the upper side of the main box body, and an oil drainage port is formed in the lower side of the main box body. The nozzle, the injection pipe, the suction inlet and the oil outlet are all communicated with a cavity in the main box body; a turning plate is arranged on the front end face of the injection pipe and rotationally connected with the injection pipe, an injection hole is formed in the front end of the injection pipe, and the area of the turning plate is larger than that of the injection hole so as to seal the injection hole. The suction inlet is used for being communicated with an oil delivery tank, and an oil drain valve is arranged in the oil drain outlet, so that oil can be prevented from flowing back, and normal work of an engine is not affected.

This invention discloses a flight test method for unusable fuel quantities in a civil propeller-driven single-engine aircraft, comprising: determining safety measures for flight tests with unusable fuel quantities; capturing key discrimination parameters and acquiring safety data during ground tests; conducting safety risk analysis on flight tests with unusable fuel quantities and establishing corresponding risk mitigation measures based on the risk analysis conclusions; formulating flight test subjects for unusable fuel quantities based on the geometry of the test fuel tank and the location of the fuel inlet; conducting flight tests with unusable fuel quantities based on the formulated flight test subjects, with only one flight test subject conducted in each flight test; measuring the unusable fuel quantity in the aircraft's test fuel tank after each flight test, and obtaining the most unfavorable fuel quantity during the flight test based on the unusable fuel quantity of each flight test subject. The technical solution provided by this invention solves the airworthiness certification problem of flight test subjects for unusable fuel quantities in the fuel system of civil propeller-driven single-engine aircraft.

The invention provides a vacuum insulated conduit assembly, a hydrogen fuel system and an aircraft fuel system. The vacuum insulated conduit assembly includes a first conduit section and a second conduit section. The first conduit section has a first outer wall and a first inner wall within the first outer wall. The first outer wall and the first inner wall are concentric and spaced apart along a common central axis to define a first space. The second conduit section has a second outer wall and a second inner wall within the second outer wall. The second outer wall and the second inner wall are concentric and spaced apart along a common central axis to define a second space. The first inner wall includes a first inner connector and the second inner wall includes a second inner connector. The vacuum insulated conduit assembly has a shoulder and a sleeve portion supported on and sliding along the shoulder that selectively provides access to an area where the first and second inner connectors are located. The vacuum insulated conduit assembly also has a sealing element connected to the shoulder configured to provide a seal between the shoulder and the sleeve portion.