239 results about "Bleed air" patented technology

Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.

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.

A swirl recovery vane (SRV) includes a trunnion configured to pivotally connect to an exterior surface of an inner platform of an SRV structure system. The SRV includes a body extending radially from the trunnion to an unshrouded distal vane tip of the vane. The SRV includes a heat exchanger that includes a gas inlet configured to couple a bleed air valve of a compressor section of a gas generator to a gas channel. The heat exchanger includes the gas channel adapted to: enable gas to flow into the gas inlet as cool gas and out through a gas outlet as warm gas; and extend into the body of the vane to enable heat transfer from a heat source within the vane to the cool gas. The heat exchanger includes the gas outlet configured to output the warm gas from the channel to an ambient environment outside of the vane.

An environmental control system pack for an aircraft that includes a ram air circuit and a bleed flow circuit. The ram air circuit receives air from an atmosphere on an exterior of the aircraft and expels the air through an outlet back into the atmosphere. The bleed flow circuit receives bleed air and cools the air prior to the bleed air being delivered to an interior space. The environmental control system pack uses energy from the bleed air to drive a fan in the ram air circuit. This causes an increase in a compression power of the air in the ram air circuit to increase a thrust of the air exiting the ram air circuit through the outlet.

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.

The invention relates to the technical field of aero-engines, and discloses a combustion chamber with an adjustable air entraining structure, the combustion chamber comprises a cartridge receiver structure, the air entraining structure and an adjusting structure, the cartridge receiver structure comprises a first cartridge receiver and a second cartridge receiver, the first cartridge receiver and the second cartridge receiver are connected with each other and jointly define a mounting cavity, and the first cartridge receiver is provided with a first communication part and a second communication part; the air entraining structure is provided with a communicating cavity, the communicating cavity is provided with a first opening, a second opening and a third opening, the first opening communicates with the first communicating part, the second opening communicates with the second communicating part, and the third opening communicates with the engine; the adjusting structure comprises a first adjusting assembly and a second adjusting assembly, the first adjusting assembly covers the first communicating part and is arranged on the first casing, and the second adjusting assembly covers the second communicating part and is arranged on the first casing; ignition delay and flameout of the auxiliary power device in a small state are avoided, low-frequency excitation of the flame tube in a large state is avoided, and performance degradation of the combustion chamber in an intermediate value state is avoided.

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.

The invention belongs to the technical field of aircraft fuel tank inerting, and particularly relates to a method for calculating the shedding rate of a catalyst coating of an ozone converter in single flight. Comprising the following steps: S1, acquiring the mass M of a catalyst coating of the ozone converter; s2, acquiring flight profile data, wherein the flight profile data comprises a relationship H (t) between the flight height H and time t of the aircraft, time h required by one flight cycle of the aircraft, and time t1 and t2 for starting working and closing of an inerting system; s3, the relation m (H) between the mass m of dust contained in unit volume of entrained air in the entrained air of the inerting system and the flight height H is obtained; s4, the air entraining flow Q of the inerting system is obtained; s5, obtaining a guarantee value F and a dust holding capacity M1 of the minimum replacement interval of the filter; and S6, calculating the single flight shedding rate mu of the catalyst coating of the ozone converter.

A gas turbine engine includes a steam generating system, and a combustion section having a first combustor and a second combustor in fluid communication with a secondary combustion zone of the first combustor. The second combustor is operable to receive compressor bleed air and fuel, along with steam from the steam generating system. During a non-idle operating state and a power-augmentation operating state, when the steam system is operable, steam is provided to the second combustor and then to the secondary combustion zone of the first combustor. When the steam generating system is inoperable, the fuel and bleed air are provided to the second combustor and combusted, and combustion products are provided to the first combustor. When the steam generating system is partially operable, the steam, the bleed air, and the fuel are provided to the second combustor, and combustion products are provided to the first combustor.

A fluid thermal control system for a gas turbine engine is provided. The gas turbine engine has a compressor section and an engine fluid source that produces a flow of engine fluid. The fluid thermal control system includes an airflow inlet, an airflow outlet, a heat exchanger, an airflow inlet valve, and a bleed air valve. The heat exchanger permits heat transfer between a heat exchanger airflow and the flow of engine fluid. The airflow inlet valve permits the airflow from the source of airflow to engage with the heat exchanger as at least part of the heat exchanger airflow. The bleed air valve is controllable between closed and open configurations. The bleed air valve is in communication with the compressor section and is configured to receive compressor bleed air. In the open configuration, the bleed air valve allows compressor bleed air to engage with the heat exchanger.

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.

The application provides an aircraft bleed air online adjusting device, and belongs to the technical field of aero-engines. The device comprises a measurement control system, an engine, a bleed air main pipe, a bleed air flow measurement and control device, a cabin simulator and a cabin pressure control valve. The bleed air main pipe is provided with a first gas temperature and pressure measurement point. The bleed air flow measurement and control device is provided with a second gas temperature and pressure measurement point. The cabin simulator is provided with a cabin pressure measurement point. The bleed air flow measurement and control device is used for online measurement and control of the bleed air flow of the engine. The first gas temperature and pressure measurement point, the second gas temperature and pressure measurement point and the cabin pressure measurement point are used for measurement of along-the-way gas parameters. The cabin simulator is used for simulation of the volume and pressure of the aircraft cabin. The cabin pressure control valve is used for online adjustment of the pressure of the cabin simulator. Through the processing scheme, the engine bleed air flow test and control are adjusted, the bleed air is consistent with the flight environment, and the adjustment and measurement accuracy are improved.

The application provides a radial diffuser blade design method and a radial diffuser. The method comprises the following steps: obtaining each order natural frequency corresponding to a diffuser blade; obtaining the relationship between a gap length and each order natural frequency; determining a minimum gap length based on an excitation frequency, the relationship between the gap length and each order natural frequency; and performing a cutting angle operation on the diffuser design blade according to the minimum gap length to complete the design. The frequency-controllable radial diffuser blade can effectively avoid the problem that the airflow of the upstream rotor resonates with the downstream diffuser blade; and on the basis of considering the constraint state change of the contact surface between the radial diffuser blade and the bleed air casing, the excitation frequency of the upstream rotor is combined to effectively avoid the high-cycle fatigue failure of the radial diffuser blade caused by the uncertain blade tip constraint state, and the service life of the diffuser blade is ensured.

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.

To provide an environmental control system pack for an aircraft.SOLUTION: An environmental control system pack 20 for an aircraft includes a ram air circuit 40 and a bleed flow circuit 30. The ram air circuit receives air from an atmosphere on an exterior of the aircraft and expels the air through an outlet back into the atmosphere. The bleed flow circuit receives bleed air and cools the air prior to the bleed air being delivered to an interior space. The environmental control system pack uses energy from the bleed air to drive a fan in the ram air circuit. This causes an increase in a compression power of the air in the ram air circuit to increase a thrust of the air exiting the ram air circuit through the outlet.SELECTED DRAWING: Figure 4

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.

An aircraft heat exchanger arranged longitudinally and including a casing with an inner chamber configured so that coolant flows longitudinally from an inlet to an outlet, a plurality of laterally spaced longitudinally extending inner chamber plates, and a plurality of channels defined between contiguous plates. Each plate includes a leading edge oriented towards the inlet and configured to divert coolant towards the channels. A plurality of the plates include an inner hollow area configured to conduct a flow of hot bleed air therethrough. Leading edges of a first group of plates are arranged in a stepped pattern. A gap defined between the first group of plates and a first casing lateral wall establishes a fluid coolant communication through the first group of plates between the casing inlet and outlet. The lateral distance between each leading edge of the first group of plates and the first lateral wall decreases longitudinally.

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.

A gas turbine engine 10 for an aircraft comprises an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor. The engine c

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.

The invention discloses an air purification device, which can avoid engine bleed air failure caused by cracking of an inner core of an ozone converter while ensuring the ozone conversion efficiency. The air purification device is installed between an air source air entraining pipeline of the airplane and an inlet pipeline of an air conditioner assembly, the air purification device comprises a first pipeline and a second pipeline which are arranged in parallel, one ends of the first pipeline and the second pipeline are connected with the air source air entraining pipeline, and the other ends of the first pipeline and the second pipeline are connected with the inlet pipeline of the air conditioner assembly; the ozone converter is arranged on the first pipeline and is used for converting ozone in the entrained air; the filter is arranged on the second pipeline and is used for filtering pollutants in the entrained air; the three-way valve is connected between one end of the first pipeline and one end of the second pipeline and the air source air entraining pipeline and is configured to be switched according to the external environment of the airplane, so that the air source air entraining pipeline is communicated with the first pipeline or the air source air entraining pipeline is communicated with the second pipeline.