Abstract: A bleed system including control circuitry and a variable jet pump. The control circuitry is configured to receive a signal indicative of a fluid parameter in the bleed system and cause the jet pump to alter a mixing ratio of a higher pressure gas and a lower pressure gas based on the signal. The jet pump is configured to combine the lower pressure gas and the higher pressure gas in the mixing ratio to generate a mixed gas. The jet pump is configured to supply the mixed gas to one or more gas loads in the bleed system. In examples, the control circuitry is configured to establish a system setpoint for the fluid parameter based on an operating status of the one or more gas loads.

Abstract: An aeroderivative gas turbine provided with a casing, a compressor including a rotor mounted on a generator shaft supported for rotation in the casing, a high pressure turbine arranged in the casing and with a rotor mounted on the generator shaft for co-rotation with the compressor rotor, a combustor, a power turbine arranged in the casing and including a rotor mounted on a turbine shaft to drive a load, wherein a thermal insulation coating is present to reduce heat dispersion through the casing.

Abstract: A gas turbine engine is provided, having a turbomachine and a rotor assembly driven by the turbomachine and operable at a first blade passing frequency (f1) greater than or equal to 2,500 hertz and less than or equal to 5,000 hertz during a high power operating condition; a heat exchanger positioned within an annular duct and extending substantially continuously along the circumferential direction, wherein an effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for a high power operating condition, and wherein the heat exchanger comprises a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the high power operating condition, the OARR equal to: sin ? ? ( 2 × ? × f 1 a 1 × L i ) 2 wherein a1 is equal to 13,200 inches per second during the high power operating condition.

Abstract: A method of operating a gas turbine engine comprising: extracting a flow of air from a compressor section of the gas turbine engine into a first conduit; flowing the extracted flow of air through the first conduit to a first location at a turbine section of the turbine section, wherein a second conduit is in fluid communication with the turbine section at a second location; flowing a heat transfer fluid to a first heat exchanger positioned in thermal communication with the flow of air through the first conduit, the heat transfer fluid in thermal communication with the extracted flow of air through the first conduit via the first heat exchanger; and modulating, via a flow control device, a portion of the flow of air extracted from the first conduit to the second conduit downstream of the first heat exchanger.

Abstract: In some examples, propulsion, electrical generation, and cooling system. The system comprises a gas turbine engine including a compressor and a bleed air outlet from the compressor, wherein the compressor is configured to compress a first fluid, wherein a portion of the compressed first fluid is directed out of the bleed air outlet to define bleed air from the compressor. The system also includes a turbo-generator including a combustor, wherein the combustor is configured to receive the bleed air from the compressor and combust a fuel with the bleed air, wherein the turbo-generator is configured to generate electrical energy via the combustion of the fuel by the combustor. The system also includes an air cycle cooling system configured to remove heat via an air cycle cooling process, wherein the air cycle cooling process is charged via the bleed air from the compressor. A compressor of the air cycle cooling system may be driven by a turbine of the turbo-generator or a turbine of the gas turbine engine.

Abstract: A turbofan engine includes a fan, a core turbine engine having one or more turbines and an exhaust section, and a hydrogen-exhaust gas heat exchanger in flow communication with the exhaust section and hydrogen flowing along a hydrogen supply line. The hydrogen-exhaust gas heat exchanger defines a load capacity factor determined by raising a product to a one-quarter power, the product being determined by multiplying a heat transfer surface area density associated with the hydrogen-exhaust gas heat exchanger by a process conductance parameter that relates characteristics of hydrogen, ambient air, and exhaust gas at takeoff, as well as a fan diameter of the fan and a number of turbine stages of the turbofan engine. The load capacity factor is between 4.37 and 28.65 for the fan diameter being between 0.5 and 3.5 meters and the heat transfer surface area density being between 500 m2/m3 and 10,000 m2/m3.

Abstract: A turbine engine having counter-rotating rotors comprising a first rotor, rotating in a first rotational direction, defining a first rotor set of blades axially spaced to define a gap, and a second rotor, rotating in a second rotational direction counter the first rotational direction. The second rotor further including a second set of blades received within the gap of the first rotor. A plurality of fluid passages is formed in the first rotor with an outlet facing the gap.

Abstract: A method for controlling the clearance between the blade tips of a high-pressure turbine of a gas turbine aircraft engine and a turbine shroud, including the controlling of a valve delivering a stream of air to the turbine shroud, this method further including the following steps: the detection of a transient acceleration phase of the engine; the receiving of an item of data representative of the gas temperature at the outlet of the combustion chamber of the engine; a valve opening command, to deliver the air stream to the turbine shroud or to increase the flow rate of the delivered air stream, if the transient acceleration phase is detected and if the gas temperature at the outlet of the combustion chamber is greater than a first temperature threshold corresponding to a degraded clearance characteristic of an aged engine, this threshold being less than an operating limit temperature of the engine.

Abstract: A cooling system for a gas turbine engine may comprise a plenum extending circumferentially around an outer engine case structure. The plenum may comprise a supply conduit and a return conduit. The supply conduit and the return conduit may be in fluid communication with a heat exchanger. The heat exchanger may be disposed between the outer engine case structure and an inner engine case structure. The plenum may be configured to provide enhance heat transfer for the cooling system.

Abstract: The invention relates to an assembly for a turbomachine (1) extending about an axis (X) comprising: —an inner annular platform (13) and an outer annular platform (14) delimiting a flow channel (3) for a flow flowing from upstream to downstream, and —at least one fixed blade (12) extending radially between the inner annular platform (13) and the outer annular platform (14), said fixed blade (12) being profiled with a leading edge (15), characterized in that the inner annular platform (13) and/or the outer annular platform (14) comprises, upstream of the leading edge (15) of the fixed blade (12), a furrow (21) for channeling the fluid flowing in the channel (3), having a segment forming an upstream limit (22) and a segment forming a downstream limit (23), the length of the upstream limit (22) being greater than the length of the downstream limit (23).

Abstract: A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

Abstract: A loading/unloading method for a gas turbine system is disclosed. The gas turbine system includes a combustion section featuring a primary combustion stage with a first plurality of fuel nozzles and a downstream, secondary combustion stage with a second plurality of fuel nozzles. For loading, the method progresses through each of a plurality of progressive combustion modes that sequentially activate a higher number of at least one of the first or second plurality of fuel nozzles; and for unloading, the method progresses through each of a plurality of progressive combustion modes that sequentially activate a lower number of at least one of the first or second plurality of fuel nozzles. During each combustion mode, regardless of whether loading or unloading, a primary combustion stage exit temperature of a combustion gas flow is controlled to be within a predefined target range corresponding to the respective combustion mode.

Abstract: A turbine engine includes a duct defining an annular passage, at least two heat exchangers arranged within the annular passage and spaced circumferentially apart, a passage between the at least two heat exchangers, and a forward flow control device operable for controlling airflow through the passages.

Abstract: A gas turbine engine component assembly comprising: a first component having a first surface and a second surface opposite the first surface, wherein the first component includes a cooling hole extending from the second surface to the first surface; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween; and a lateral flow injection feature integrally formed in the first component and fluidly connecting a flow path located proximate to the second surface of first component to the cooling channel, the lateral flow injection feature being configured to direct airflow from the airflow path through a passageway and into the cooling channel at least partially in a lateral direction parallel to the second surface of the second component such that a cross flow is generated in the cooling channel.

Abstract: A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core.

Abstract: A stator vane is configured to guide a flow of a fluid flowing in an axial direction. The stator vane includes a stator vane body which extends in a radial direction with respect to an axis and includes a radially outer end portion which is supported by a casing and a stator vane inner circumferential surface which is an end surface facing an inside in the radial direction and faces an outer peripheral surface of a rotary shaft via a clearance, and an internal channel which is defined inside the stator vane body and includes a first end which is open on the stator vane inner circumferential surface and a second end which is connected to a pressure source having a pressure different from a pressure around the stator vane body.

Abstract: An aircraft propulsion system, including a turbojet and a heat exchanger system which includes a main heat exchanger, a hot air supply pipe, a transfer pipe transferring hot air to an air management system, a main supply pipe supplying cold air from the fan duct, an evacuation pipe expelling air to the outside, a sub heat exchanger with a high pressure pipe going therethrough, a sub supply pipe supplying cold air and including a sub regulating valve, a sub evacuation pipe expelling air, a temperature sensor, and a controller controlling the system according to the temperature measured The sub regulating valve comprises a door articulated between closed and open positions, a return spring constraining the door in the open position, and a realizing system, controlled by the controller, mobile between a blocking position and a realizing position in which the door is released to move to the open position.

Abstract: The present application provides a method of optimizing fan usage in a cooling water system having a number of fans with a heat exchanger to cool a cooling fluid for use with a number of gas turbine subsystems. The method may include the steps of running all of the fans at base load, calculating a heat transfer capability of each fan at base load, calculating a temperature difference between an actual temperature and a target temperature of the cooling fluid, selecting a minimum target temperature of the cooling fluid, calculating a target thermal energy of the cooling fluid for the minimum target temperature, calculating a number of the fans to be turned on or off by dividing the target thermal energy with the heat transfer capability of each fan, and turn on or off the calculated number of fans in a predetermined manner with an objective of balancing the running hours of each fan.

Abstract: A transition piece of an embodiment leads a combustion gas generated by a combustor liner to a turbine part in a gas turbine facility. An outlet portion from which the combustion gas flows out to the turbine part in the transition piece includes: an inner peripheral wall located inside in a radial direction of the turbine part; an outer peripheral wall located further outside than the inner peripheral wall in the radial direction; and support struts provided between the inner peripheral wall and the outer peripheral wall. In the inner peripheral wall, first insertion grooves are formed. In the outer peripheral wall, second insertion grooves are formed. In the support struts, first end portions located inside in the radial direction are inserted into and fixed in the first insertion grooves, and second end portions located outside in the radial direction are inserted into and fixed in the second insertion grooves.

Abstract: A bleed air control system and method include first and second flow paths, first and second flow control valves, and a controller. The first flow path is configured to carry a first bleed air flow and includes a heat transfer device configured to transfer heat to or from the first bleed air flow. The second flow path is configured to carry a second bleed air flow. The first and second flow paths merge downstream to provide a combined bleed air flow. The first flow control valve is positioned to control the first bleed air flow, and the second flow control valve is positioned to control the second bleed air flow. The controller is configured to control the first and second flow control valves to control a temperature and flow of the combined bleed air flow.

Abstract: A gas turbine engine including core engine is provided. Air may enter the core engine through an inlet and travel through and engine air flowpath extending through the core engine, e.g., generally along an axial direction of the gas turbine engine. The gas turbine engine additionally includes a cooling air flowpath extending outwardly generally along the radial direction of the gas turbine engine. The cooling air flowpath extends between an inlet in flow communication with engine air flowpath and an outlet defined by an opening in an outer casing of the core engine. Moreover, the gas turbine engine includes a heat exchanger positioned at least partially within the outer casing the core engine with the cooling air flowpath extending over or through the heat exchanger.

Abstract: A method of controlling turbine tip clearance includes measuring turbine speed; measuring turbine temperature; measuring parameters indicative of current operating conditions; determining limits for the turbine speed and turbine temperature; calculating target tip clearance from the turbine speed, turbine temperature and parameters, to optimise turbine efficiency within the turbine speed and turbine temperature limits; and controlling turbine tip clearance apparatus to the calculated target tip clearance.

Abstract: A gas turbine engine includes a first tap connected to a compressor section to deliver air at a first pressure. A heat exchanger is downstream of the first tap. A cooling air valve selectively blocks flow of cooling air across the heat exchanger. A cooling compressor is downstream of the heat exchanger and pressurizes the air from the first tap to a greater second pressure. A shut off valve selectively stops flow of the air between the heat exchanger and the cooling compressor. A controller controls the cooling air valve, the shut off valve, and the cooling compressor such that the flow of the air is stopped between the heat exchanger and the cooling compressor only after the controller has stopped the cooling compressor. A monitoring system communicates with the controller and includes a pressure sensor and a temperature sensor downstream of the cooling compressor.

Abstract: A turbine injection system for a gas turbine engine includes a first end operable to receive air from a heat exchanger, a second end operable to distribute mixed cooling air to a turbine stage, an opening downstream of said first end and a mixing plenum downstream of said first end and said opening. The opening provides a direct fluid pathway into said turbine injection system.

Abstract: An ejector is presented. The ejector includes a primary fluid inlet to receive a primary fluid. The ejector further includes a secondary fluid inlet to receive a secondary fluid. Furthermore, the ejector includes a nozzle fluidly coupled to the primary fluid inlet and the secondary fluid inlet. The nozzle includes a secondary pilot inlet to receive at least a portion of the secondary fluid from the secondary fluid inlet, and a nozzle outlet including a plurality of primary openings for discharging the primary fluid and a secondary opening for discharging the secondary fluid. A turbo-machine having the ejector is also presented.

Abstract: A method of modulating cooling flow to an engine component based on a health of the component is provided. The method includes determining a cooling flow requirement of the engine component for each of a plurality of operating conditions and channeling the determined required flow to the engine component during each respective operating condition of the plurality of operating conditions. The method also includes assessing a health of the engine component. The method further includes modifying the determined cooling flow requirement based on the assessed health of the engine component, and supplying the modified cooling flow requirement to the engine component during each subsequent respective operating condition of the plurality of operating conditions.

Abstract: A jet engine with a chamber which is delimited by a housing and inside of which a rotatable appliance that can be impinged by hydraulic fluid is arranged, wherein hydraulic fluid and air introduced through the housing into the chamber can be discharged from the chamber. In the area through which air and hydraulic fluid can be discharged from the chamber, the housing is embodied with a separating appliance in the area of which air and hydraulic fluid can be separated from each other, and/or a deflector appliance, by means of which air, hydraulic fluid and/or an air-oil mixture in the chamber can be guided in the direction of the area in a targeted manner, is provided upstream of the area of the housing.

Abstract: Disclosed herein is a system for cooling a gas turbine. The system for cooling a gas turbine cools a turbine disk unit by individually supplying cooling air to each of a plurality of turbine disks.

Abstract: In an embodiment, a method includes flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path; evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; and modulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation.

Abstract: A gas turbine engine comprises a compressor section, a combustor, and a turbine section. The turbine section includes a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap taps air having been compressed by the compressor, the tapped air being passed through a heat exchanger. A vane section has vanes downstream of the combustor, but upstream of the first stage blade, and the air downstream of the heat exchanger passes radially inwardly of the combustor, along an axial length of the combustor, and then radially outwardly through a hollow chamber in the vanes, and then across the blade outer air seal, to cool the blade outer air seal.

Abstract: A heat exchanger assembly for a gas turbine engine that includes an outer engine case. The heat exchanger assembly includes at least one cooling channel, the at least one cooling channel is configured to receive a flow of fluid to be cooled. At least one first coolant flow duct that is configured to receive a flow of a first coolant, wherein the at least one cooling channel is disposed between a first inlet and a first outlet. The heat exchanger assembly further include at least one second coolant flow duct that is configured to receive a flow of a second coolant, wherein the at least one cooling channel is disposed between a second inlet and a second outlet.

Abstract: A method of controlling a rotor tip clearance in a gas turbine engine (10). The method comprises determining an engine or component remaining useful life Tr, and controlling a tip clearance control arrangement (38) to maintain a rotor tip clearance (36) at a target tip clearance Dtarget. The target tip clearance Dtarget is determined in accordance with a function of remaining engine life Tr.

Abstract: An airflow control system control system for a gas turbine system according to an embodiment includes: an airflow generation system including a plurality of air moving systems for selective attachment to a rotatable shaft of a gas turbine system, the airflow generation system drawing in an excess flow of air through an air intake section; and a mixing area for receiving an exhaust gas stream of the gas turbine system; the airflow generation system: directing a first portion and a second portion of the excess flow of air generated by the airflow generation system into the mixing area to reduce a temperature of the exhaust gas stream; and directing a third portion of the excess flow of air generated by the airflow generation system into a discharge chamber of a compressor component of the gas turbine system.

Abstract: A system includes a foreign object damage (FOD) screen configured to be disposed upstream of an air intake of a gas turbine engine and to keep debris from entering the air intake. The FOD screen is configured to extend across a fluid flow path extending through the air intake into the gas turbine engine. The FOD screen includes a flexible, woven fabric made of a non-metal material and configured to absorb and dissipate energy from the debris, and the flexible, woven fabric includes a tensile strength ranging between 2700 megapascals (mPa) and 3700 mPa.

Abstract: The present application provides an inlet bleed heat control system for a compressor of a gas turbine engine. The inlet bleed heat control system provides an inlet bleed heat manifold and an ejector in communication with the inlet bleed heat manifold such that the ejector is in communication with a flow of compressor discharge air and a flow of ambient air.

Abstract: A compressor of a gas turbine engine includes a plurality of rotors rotatable about a central longitudinal axis, a plurality of stators, and a circumferential inner case wall located outwardly of the plurality of rotors and the plurality of stators. The inner case wall includes a plurality of first slots located in a first section of the circumferential inner case wall and a plurality of second slots located in at least one second section of the circumferential inner case wall. A first area of each of the first slots is greater than a second area of each of the second slots. A circumferential manifold is located outwardly of the inner case wall including a port. At least one of the plurality of second slots is located proximate to the port.

Abstract: A cooling system for providing a buffer cooled cooling air to a turbine section of a gas turbine engine is disclosed. The cooling system may comprise a first conduit configured to transmit a cooling air toward the turbine section, a heat exchanger configured to cool a bleed airflow diverted from the first conduit to provide a buffer air, and a bypass conduit configured to direct at least a portion of the buffer air through at least one passageway that bypasses a bearing compartment of the gas turbine engine. The cooling system may further comprise a manifold configured to allow the cooling air exiting the first conduit and the buffer air exiting the bypass conduit to mix and provide the buffer cooled cooling air, and a nozzle assembly configured to deliver the buffer cooled cooling air to the turbine section.

Abstract: A gas turbine engine comprises a compressor rotor including blades and a disk, with a bore defined radially inwardly of the disk, and a combustor. A tap directs the products of combustion first to a valve and then into the bore of the disk. At least two temperature sensors sense a temperature of the products of combustion downstream of the valve. A control compares sensed temperatures from the at least two temperature sensors to ensure the at least two temperature sensors are functioning properly. The sensed temperatures are utilized to control the valve. A method of operating a gas turbine engine is also disclosed.

Abstract: A system for providing combustion air and fuel gas to a premix burner includes a premix engine, a premix burner in fluid communication with an outlet of the premix engine, an exhaust flue, a flue gas recirculation line in fluid communication with the flue and an inlet of the premix engine, and a fresh air line in fluid communication with a source of fresh air and the inlet of the premix engine. A flue gas flow restrictor is installed in the flue gas recirculation line, and a fresh air flow restrictor is installed in the fresh air line. The flow restrictors are sized so that the premix engine, in operation, draws recycled flue gas and fresh air from the recycled flue gas line and fresh air line, respectively, in a predetermined proportion.

Abstract: An aircraft propulsion unit comprising an engine and a nacelle including an axisymmetric casing (16) delimiting an air flow path, wherein this casing has at least two openings closed by removable and interchangeable panels (18), at least one of these panels carrying one component (24) of the propulsion unit.

Abstract: A spoke for a mid-turbine frame on a gas turbine engine includes a cylindrical portion. A flange is attached to the cylindrical portion. At least one gusset extends between the cylindrical portion and the flange. At least one gusset extends to a perimeter of the flange.

Abstract: A case for a gas turbine engine includes a core body. The core body defines a longitudinally extending core flow path, a laterally extending bleed air duct coupling the core flow path in fluid communication with the external environment, and a structure-supporting member spanning the bleed air duct. A heating element is connected to the core body and is in thermal communication with the structure-supporting member.

Abstract: In an embodiment, a method includes flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path; evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; and modulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation.

Abstract: A method of modulating the cooling of a gas turbine component is disclosed. The method includes determining a target component temperature at which the gas turbine component can be maintained without the gas turbine component experiencing a failure over the course of an indicated life of the gas turbine component; scheduling a cooling air value to the target component temperature; and determining one or more of a demanded cooling air temperature and a demanded cooling air mass flow rate based on the scheduled cooling air value.

Abstract: A gas turbine engine is provided defining a radial direction. The gas turbine engine generally includes a compressor section and a turbine section, the compressor section and turbine section together defining a core air flowpath. The gas turbine engine also includes a sump positioned inward of the core air flowpath along the radial direction. An air pump is positioned inward of the core air flowpath along the radial direction for providing a flow of air from the sump to lower an internal pressure of the sump and reduce a likelihood of lubrication leaking from the sump.

Abstract: The present application provides a gas turbine engine for low turndown operations. The gas turbine engine may include a compressor with a compressor bleed air flow, a turbine, and a compressor bleed air flow manifold. The compressor bleed air manifold directs a variable portion of the compressor bleed air flow to the turbine.

Abstract: A passive cooling system for an auxiliary power unit (APU) installation on an aircraft is provided. The system is for an auxiliary power unit having at least a compressor portion of a gas turbine engine and an oil cooler contained separately within a nacelle. The system includes the auxiliary power unit housed within the nacelle of the aircraft, an engine exhaust opening defined in the aft portion of the nacelle and communicating with the gas turbine engine, at least a first air inlet duct communicating with a second opening defined in said nacelle and with said compressor portion and the oil cooler is located within a second duct communicating with an opening other than the engine exhaust opening of said nacelle and with the engine exhaust opening. Exterior cooling air and engine exhaust ejected through said engine exhaust opening entrain cooling air through said second duct to said oil cooler, and thus provide engine oil cooling. An exhaust eductor is also provided.

Abstract: A thermal actuator is provided and includes an expansion material disposed and configured to move a movable element from a first movable element position toward a second movable element position in accordance with an expansion condition of the expansion material. The expansion material includes an inorganic salt mixture or a metal oxide mixture.

Abstract: A gas turbine engine includes a heat exchanger, a diffuser case, a passageway and a nozzle assembly. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The diffuser case includes a plenum that receives the conditioned airflow. The passageway is fluidly connected between the heat exchanger and the diffuser case, and the conditioned airflow is communicated through the passageway and into the plenum. The nozzle assembly is in fluid communication with the plenum of the diffuser case to receive the conditioned airflow from the plenum.

Abstract: The device comprises first means for determining a pressure value relative to the pressure at the outlet of a precooler of an air bleed system and second means for calculating the air bleed level using this pressure value, said first means comprising means for calculating a first pressure value at the outlet of the precooler, using the pressure measured by a sensor, means for receiving a pressure measured by a second sensor, which represents a second pressure value at the outlet of the precooler and means for selecting one of said first and second pressure values, which is then transmitted to said second means for calculating said air bleed level.