159 results about "Bleed air" patented technology

A gas turbine engine is provided that includes a fan section, a compressor section, a combustion section, a turbine section, an annular fan bypass air path, a bleed air structure, an annular compartment, and an anti-icing system. The compressor section has a low and high pressure compressor sections. The fan bypass air path is defined by inner and outer bypass wall structures. An outer radial boundary structure defines an outer radial boundary of a core gas path within the low pressure compressor section. The bleed air structure is disposable in an open state or in a closed state. The annular compartment is defined by the inner bypass wall structure, the outer radial boundary structure, and a firewall. The anti-icing system is disposable in an activated state or in a deactivated state. In the activated state the anti-icing system is configured to mitigate ice accumulation within the annular compartment.

The invention relates to the technical field of airborne heat management systems, in particular to an airborne low-compensation heat management system, an aircraft and a system control method. The heat management system comprises an air compression assembly, a refrigeration assembly, a heating assembly and an air supply assembly. The air compression assembly comprises an air compressor and an air inlet unit; the air inlet unit communicates with an inlet of the air compressor. The refrigeration assembly comprises a first cooling turbine, a second cooling turbine, an equipment radiator, a refrigeration pipeline unit and a refrigeration valve unit; the heating assembly comprises a first air entraining pipeline and a temperature adjusting valve. The first air entraining pipeline is used for communicating with an air entraining opening of the engine assembly. The temperature adjusting valve is communicated with the first air entraining pipeline; the air supply assembly comprises an air supply pipeline; an outlet of the first cooling turbine is communicated with an air supply pipeline through a flow limiting valve; and the first air entraining pipeline is communicated with the air supply pipeline. In this way, the problems that the controllability is low after a refrigerating circuit and a warm air supply pipeline of airborne equipment intersect, and too much engine bleed air needs to be consumed for cabin temperature adjustment are solved.

The invention relates to the technical field of sealing performance detection, and discloses a detection system and a detection method for detecting the two-way leakage rate of a spherical sealing valve of an aircraft bleed air system. The system comprises an air inlet device, a sealing device and a detection device, the gas inlet device provides clean test gas with controllable pressure; the sealing device is used for fixing the spherical sealing valve and simulating two sealing states of the spherical sealing valve in the high-pressure stage regulator to form two independent test gas paths; the detection device is connected with an outlet of the sealing device, is used for quantitatively or qualitatively detecting leaked gas and comprises a flow monitoring unit and a pressure difference monitoring unit; the sealing device can respectively test the sealing performance of the ceramic ball and the lift valve seat and the sealing performance of the ceramic ball and the valve mounting seat by replacing or adjusting the limiting structure; according to the invention, component-level bidirectional sealing performance detection is realized, the test precision and the maintenance efficiency are improved, defects can be identified before assembly, and the rework rate is reduced.

The application relates to the technical field of sealing performance detection, and discloses a detection system and a detection method for detecting bidirectional leakage of a spherical sealing valve of an aircraft bleed air system; the system comprises an air inlet device, a sealing device and a detection device; the air inlet device provides pressure-controllable and clean test gas; the sealing device is used for fixing the spherical sealing valve and simulating two sealing states of the spherical sealing valve in a high-pressure stage regulator, thereby forming two independent test gas paths; the detection device is connected to the outlet of the sealing device and is used for quantitatively or qualitatively detecting leakage gas and comprises a flow monitoring unit and a differential pressure monitoring unit; the sealing device can test the sealing performance of a ceramic ball and a poppet valve seat and the sealing performance of the ceramic ball and a valve mounting seat by replacing or adjusting a limiting structure; the application realizes component-level bidirectional sealing performance detection, improves test precision and maintenance efficiency, can identify defects before assembly, and reduces the rework rate.

A turbine engine includes an engine core, an engine drive shaft, an accessory gearbox (AGB), an air turbine starter (ATS), and a bleed air circuit. The engine core has a fan section, a compressor section, a combustion section, and a turbine section arranged in series. The engine drive shaft is operatively connected to the fan section, the compressor section, and the turbine section. The AGB and the ATS are operatively connected to the engine drive shaft.

This application belongs to the field of aero-engines, and specifically relates to a bleed air device for a combustion chamber test piece, comprising: a disc-shaped cover plate, the cover plate including an outer disc surface and an inner disc surface, an annular gas collecting ring installed on the outer edge of the outer disc surface, the gas collecting ring having an annular cavity; the inner disc surface and the bleed air chamber cover plate with a circular groove forming a cylindrical cavity; two channels communicating with the annular cavity through through holes at the same position on the cover plate and the gas collecting ring, the annular cavity communicating with the cylindrical cavity through multiple circumferentially evenly distributed guide bends; the cylindrical cavity communicating with a gas outlet pipe connected to the inner side of the bleed air chamber cover plate; the bleed air device of this application can better simulate the bleed air under the actual engine environment of the combustion chamber test piece, and the number of through holes that can be adjusted by the rotatable gas collecting ring corresponds to different combustion chamber test piece environments.

A turbofan engine and a bleed air door arrangement for reducing the effects of rain on engine performance. The bleed air door arrangement includes a bleed air window, a valve plate, and a grid; the bleed air window is configured in an intermediate casing between an outer and an inner duct, through the intermediate casing; the valve plate extends upstream from a downstream edge of the bleed air window into the inner duct to open the bleed air window; the grid extends from the valve plate into the inner duct to break up water droplets in the inner duct.

This application belongs to the field of aero-engine technology, and specifically relates to a method and apparatus for simulating aircraft bleed air during sustained test of a turbofan engine, wherein the inlet temperature T1 and inlet pressure P1 of the engine are preset to a range; Step S2: Calculate the aircraft bleed air volume W fy The range, the total bleed air temperature T of the aircraft fy The range and total bleed air pressure P fy Range; Step S3: Calculate the cold diameter D of the bleed air nozzle. fy Range; Step S4: Select the cold-state diameter D of the bleed air nozzle fy An aircraft bleed air volume verification test was conducted to obtain the bleed air volume W for the test. fyr When the induced draft volume W in the experiment fyr The requirements are met, and the cold-state diameter D of the expiratory air nozzle is adopted. fy When the induced draft volume W in the experiment fyr If the requirements are not met, return to step S4; ensure that the calculation error of the aircraft bleed air volume is within the required range, and obtain the actual bleed air volume of the engine in various states to the maximum extent through test program design. This test method has the advantages of simple test method, low test cost and high replacement efficiency.

Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

The invention relates to the technical field of engine combustion chambers, and discloses a transition section flame stabilizer and an aero-engine with the same. The transition section flame stabilizer comprises an overflowing assembly, a fuel gas concave cavity is formed in the side, facing an inner cavity of the transition section, of the overflowing assembly, a main gas guiding flow channel is formed in the overflowing assembly, a first gas inlet and a second gas inlet are formed in the area, close to the inlet end, of the main gas guiding flow channel, and a plurality of oil outlet holes are formed in the side, facing the fuel gas concave cavity, of the oil supply assembly; and a supporting plate air guiding channel is formed in the radial supporting plate assembly, and the inlet end of the supporting plate air guiding channel and the inlet end of the main air guiding flow channel face the same direction. By means of the split design of the main air guide flow channel for inlet air flow of the transition section, most of the air flow keeps flowing in an original path, only a small part of the air flow participates in combustion organization, meanwhile, the radial supporting plate assembly does not completely penetrate through the transition section, effective mixing of high-temperature fuel gas and cold air after combustion can be achieved under the condition that an additional mixing structure is not added, and the combustion efficiency is improved. The transition section outlet temperature field stability is improved.

A fuel system for a gas turbine engine includes a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit and at least first and second heat exchangers. The first heat exchanger is configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and the second heat exchanger is provided upstream in hydrogen flow of the first heat exchanger and is configured to transfer heat from a further heat exchange fluid to hydrogen fuel in the main fuel conduit. In an embodiment, the further heat exchange fluid is compressor bleed air bled from a core compressor of the gas turbine engine.

The utility model discloses an aircraft bleed air auxiliary test device, and relates to the technical field of aviation aircraft valve test and maintenance, the aircraft bleed air auxiliary test device comprises a test cabin and a test bench, the middle part of the upper surface of the test bench is provided with an aircraft valve through a tool module, and two ends of the aircraft valve are respectively provided with an air inlet adapter and an air outlet adapter; a first valve connector is arranged at the other end of the aircraft valve, and a second valve connector is further arranged at the position, close to the middle, of the aircraft valve. According to the design, the rapid clamping and replacing efficiency of the aircraft valve can be improved, various valves are connected in series / in parallel, different valves can be opened and closed according to different testing requirements, switching of testing pipelines is achieved, the testing efficiency is improved, the testing cost is reduced, and the testing efficiency is improved. The multi-purpose machine is achieved, full-intelligent data collection is adopted, all parameters such as pressure and flow can be fed back to an external operation computer in real time, and real-time checking is facilitated.

The application provides a compressor tester provided with an axial flow turbine and relates to the technical field of compressor testing. The tester is provided with an axial flow turbine between the exhaust end of a compressor cylinder and the air inlet end of an exhaust cylinder. The axial flow turbine comprises turbine moving blades and turbine adjustable guide vanes. The turbine moving blades are arranged on the extension section of a rotor and are used for rotating and working under the action of high-pressure gas to reduce the exhaust temperature and generate reverse axial force. The turbine adjustable guide vanes are arranged on the extension section of the compressor cylinder and are used for adjusting the airflow direction and flow rate entering the turbine moving blades. The tester is also provided with a bleed air system and an automatic adjusting and control system. The application utilizes the working of the axial flow turbine under the action of high-pressure exhaust to reduce the exhaust temperature, reduce thermal stress, and generate reverse axial force opposite to the axial force direction of the rotor. The active and accurate control of the axial force is realized through the coordinated adjustment of the adjustable guide vanes and the bleed air system, the equipment structure is simplified, and the energy consumption and maintenance cost are reduced.

The invention relates to the technical field of equipment parameter monitoring, in particular to a gas turbine combustion chamber outlet temperature real-time evaluation method and device and a gas turbine. The method comprises the steps that atmospheric environment parameters, fuel supply parameters, exhaust parameters and unit operation parameters are monitored in real time; on the basis of the first law of thermodynamics, the atmospheric environment parameters and the fuel supply parameters serve as input energy items, the exhaust parameters and the unit operation parameters serve as output energy items, a first evaluation model is constructed, and the air flow of the inlet of the air compressor is predicted in real time through the first evaluation model; based on the compressor inlet air flow, the bleed air proportionality coefficient and the pressure loss coefficient, the combustion chamber outlet pressure is predicted through a second evaluation model; and a third evaluation model is constructed by combining the critical flow state maintained by the throat of the turbine first-stage guide blade, and the combustion chamber outlet pressure is input into the third evaluation model to calculate the combustion chamber outlet temperature. Online evaluation of the outlet temperature of the combustion chamber of the gas turbine can be realized, and the monitoring efficiency is improved.

Systems, apparatus, articles of manufacture, and methods to provide for cross-diffuser bleed are provided herein. An example gas turbine engine includes a frame defining a cavity in a forward side of a diffuser of the gas turbine engine; a compressor including the diffuser, the diffuser defining a primary flow path to provide air flow to a combustor and including at least one conduit, the at least one conduit fluidly coupled to the cavity; and a downstream sink fluidly coupled to the cavity via the at least one conduit defining at least a portion of a bleed air path.

An aircraft, an aircraft bleed air system, and a method of providing thrust for an aircraft. The bleed air system includes a bleed air turbine configured to receive at least a first portion of bleed air from a jet engine and to deliver the at least first portion of bleed air to a cabin, a cabin bleed air compressor operably coupled to the bleed air turbine and configured to receive cabin bleed air from the cabin, and a cabin bleed air nozzle configured to receive cabin bleed air from the cabin bleed air compressor. The aircraft includes a fuselage, a wing, a jet engine, and the bleed air system. The method includes the steps of expanding at least a first portion of bleed air from a jet engine to drive a cabin bleed air compressor, compressing cabin bleed air from a cabin with the cabin bleed air compressor, and ejecting the cabin bleed air into an environment.

Propulsion system (1) for an aircraft, comprising - at least one gas turbine (2), - at least one fluid system (4) for providing a heated fluid (6), - comprising an external auxiliary power unit (40) - wherein the at least one gas turbine (2) and the at least one fluid system (4) are coupled and arranged in such a fluid-conducting manner that at least one section (30, 34, 36) of the gas turbine (2) is supplied with the heated fluid (6), - wherein the fluid system (4) comprises a combustion unit (12) designed separately from the auxiliary engine (40), which is configured to produce the heated fluid (6) by burning fuel, - characterized in that a fluid supplied to the combustion unit (12) is bleed air from the auxiliary engine (40).

The application provides an aero-engine bleed air component test system, which comprises a bleed air flow path, an ambient air flow path, a bleed air online analysis device and an ambient air online analysis device. The bleed air flow path is connected with the bleed air online analysis device and used for analyzing the components of the engine bleed air. The ambient air flow path is connected with the ambient air online analysis device and used for analyzing the components of the ambient air. The test system further comprises a switching flow path, which is connected with the bleed air flow path and the ambient air flow path. In a first test state, the switching flow path is closed, the engine bleed air in the bleed air flow path enters the bleed air online analysis device, and the ambient air in the ambient air flow path enters the ambient air online analysis device. In a second test state, the switching flow path is opened, and the ambient air in the ambient air flow path enters the bleed air online analysis device.

A system includes a fan exit guide vane located aft of a rotor fan. The fan exit guide vane defines a first channel for a fluid from a thermal management system and a second channel therein for receiving bleed air from a low-pressure compressor. An inlet connects to a first end of the first channel to receive the fluid. An outlet connected to a second end of the first channel provides a cooled fluid. In a first mode, the fluid passing through the first channel is cooled by air from the rotor fan passing over exit guide vane and is further cooled by the bleed air passing through the second channel. In a second mode, the hot fluid passing through the first channel is cooled only by air from the rotor fan passing over the exit guide vane to provide the cooled fluid and bleed air is not provided.

This invention relates to an air preparation system and an aircraft including such a system. The air preparation system includes an engine bleed air branch, a cabin exhaust branch, a compressor, a one-way valve, and a main heat exchanger. The engine bleed air branch draws air from the bleed air manifold to supply the compressor or the one-way valve, and is equipped with an air preparation system pressure regulating valve. The cabin exhaust branch draws air from the cabin exhaust area to supply the compressor or the one-way valve, and is equipped with a cabin exhaust supply valve. The one-way valve is arranged in parallel with the compressor. The main heat exchanger is located downstream of the compressor and the one-way valve and is in fluid communication with them. The main heat exchanger is used to regulate the temperature of the air from the compressor or the one-way valve before supplying it to the inerting system. According to the above technical solution, this invention can achieve the following beneficial technical effects: reducing engine bleed air consumption, thereby reducing aircraft fuel consumption.

An environmental control system bleed cooling system including bleed air from the environmental control system bleed air system; a heat exchanger fluidly coupled between the environmental control system bleed air system and a gas turbine engine; and ambient cooling air fluidly coupled with the heat exchanger, wherein the ambient cooling air is configured to remove thermal energy from the bleed air flowing through the heat exchanger.

A gas turbine engine for an aircraft includes an engine core with a turbine, combustor, compressor, and core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the core and driven by the shaft. The engine includes a nacelle surrounding the fan and core and defining a bypass duct radially outside of the core. The bypass ratio, defined as the ratio of mass flow rate through the bypass duct to mass flow rate through the core at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators. At least one of the actuators is configured to actuate a bleed air valve. The engine includes a fuel supply system that supplies fuel for combustion in the combustor, and supplies fuel to fueldraulically drive the at least one actuator configured to actuate the bleed air valve.

The utility model discloses a turbocharger impeller accelerating device realizes turbocharger impeller bleed air acceleration. Including the front section impeller cover shell, the rear section impeller cover shell, the front section impeller cover shell, rear section impeller cover shell all have the inner race, and the inner race between the front section impeller cover shell, rear section impeller cover shell is connected seal, and the rear section impeller cover shell has the outer race, and the outer edge of front section impeller cover shell, the outer race of rear section impeller cover shell is correspondingly equipped with the connecting flange, and the connecting flange on the front section impeller cover shell, rear section impeller cover shell is through fastener solid -state link, makes the front section impeller cover shell, rear section impeller cover shell and is connected as a whole, and the outer race of rear section impeller cover shell forms the gas passage with the inner race of front section impeller cover shell, rear section impeller cover shell, and the rear end of rear section impeller cover shell is provided with the air inlet hole of gas passage, and the front end of front section impeller cover shell is equipped with a plurality of injection holes, and injection hole is respectively communicated with the gas passage, the inner hole of front section impeller cover shell.

The utility model discloses a full dynamic pressure type air refrigeration system test system, which belongs to the technical field of air refrigeration test systems and comprises a rack, a centrifugal compressor and a turbo expander are arranged on the rack, and the exhaust end of the centrifugal compressor is communicated with the air inlet end of the turbo expander through a first connecting pipeline. The first connecting pipeline is provided with a pressure regulating valve, the air inlet end of the centrifugal compressor is provided with a compressor air inlet pipeline, and the exhaust end of the turbo expander is detachably connected with a turbo expander exhaust pipeline; first pressure detection elements and first temperature detection elements are arranged on the compressor air inlet pipeline and the first connecting pipeline close to the exhaust end of the centrifugal compressor, and second pressure detection elements and second temperature detection elements are arranged on the turbo expander exhaust pipeline and the first connecting pipeline close to the air inlet end of the turbo expander. According to the utility model, the matching performance test of the all-dynamic-pressure centrifugal compressor and the all-dynamic-pressure turbo-expander in different combinations can be realized; and matching performance tests under different working conditions can be carried out.

A compressed air supply system of an aircraft includes: a bleed air passage through which bleed air from a gas turbine engine flows; an auxiliary compressor including a compressor inlet and a compressor outlet, the compressor inlet being fluidly connected to the bleed air passage; a supply passage which is fluidly connected to the compressor outlet and through which compressed air from the auxiliary compressor is supplied to an air system of the aircraft; a driver that rotates the auxiliary compressor; a sensor that detects pressure or temperature of the supply passage; and circuitry that controls an output rotational frequency of the driver in accordance with a detection signal of the sensor.

The application relates to the field of general structure design of an aero-engine, and discloses an aero-engine duct airflow distribution mechanism and a control method thereof. The bleed air structure and the flow path control structure are integrated in the rotating casing and the support plate structure of the inner and outer ducts. In the turbofan mode, the leading edge and the trailing edge of the corresponding support plate of the outer duct and the inner duct are matched, so that the duct flow is smooth. In the turbojet mode, the inner and outer duct flow passages are blocked, and the outer duct airflow is guided into the inner duct. The application can support the demand that the variable flow path aero-engine works in the turbofan engine mode at low flight speed and works in the turbojet mode at high flight speed, widen the flight speed domain of the conventional turbofan engine, realize the consideration of high economy at low speed and large thrust at high speed, and effectively reduce the structural complexity of the engine.

The present application relates to the technical field of aviation system, more particularly to an aviation engine integrated with an inerting system and an aviation aircraft. It solves the problem that the traditional on-board nitrogen generation system technology needs to use the engine bleed air, thereby affecting the engine performance. The aviation engine comprises a compressor, a rotating shaft, a gas turbine, a power turbine, a generator, a compressor air splitter, an anode tail gas splitter, a combustion chamber, an anode tail gas blower, a reforming reactor, a fuel inerting system, a fuel pump, an electric propulsion system and a cathode tail gas splitter, wherein the fuel inerting system comprises a solid oxide fuel cell, a reformer heat exchange channel, a first heat exchanger, a second heat exchanger and an oil tank. It reduces the influence of the inerting system on the weight of the aircraft, avoids the additional demand of the nitrogen generation system on the engine bleed air, and also meets the demand for electric energy under the trend of aircraft electrification, thereby ensuring the normal working performance of the engine.

A system for bleeding air from a compressor of an aircraft engine, the system comprises a bleed off valve and a plenum for receiving bleed air from the compressor via the bleed off valve. The plenum has a first outlet for discharging the bleed air from the plenum into the atmosphere, and a second outlet fluidly connected to the compressor for reinjecting a portion of the bleed air into an air intake portion of the compressor. A backflow valve is provided for selectively closing the first outlet of the plenum. The backflow valve has an actuator operable to displace a closure from a closed position to an open position upon a fluid pressure inside the plenum reaching a pressure threshold.

The clearance control system and the air entraining device are used for meeting different cooling air requirements of a high-pressure turbine active clearance control system and a low-pressure turbine active clearance control system. The air entraining device is used for entraining air from the outer duct and comprises an air entraining pipe and a partition plate; the air entraining pipe comprises a wall body, and an air entraining channel is enclosed by the wall body; the partition plate is located in the air entraining pipe and evenly divides the air entraining channel into a low-pressure air entraining channel and a high-pressure air entraining channel, the low-pressure air entraining channel is used for entraining air to a low-pressure turbine case, and the high-pressure air entraining channel is used for entraining air to a high-pressure turbine case; the partition plate can rotate so that the circulation sectional area of an inlet of the low-pressure air entraining channel can be different from the circulation sectional area of an inlet of the high-pressure air entraining channel.