Abstract: A gas turbine engine system for detecting control sensor override includes a plurality of temperature sensors coupled to the gas turbine engine system. The temperature sensors are configured to generate a plurality of signals representative of exhaust gas temperatures of the gas turbine engine. The system includes an on-site monitoring system coupled in communication to the plurality of temperature sensors. The on-site monitoring system has a processor programmed to continuously receive the plurality of signals from the temperature sensors. In addition, the processor is programed to analyze the plurality of signals to verify the accuracy of the exhaust gas temperatures associated with the plurality of signals, and to detect a jumpered temperature sensor of the plurality of temperature sensors.

Abstract: A control system (50) for a turbopropeller engine (2) of an aircraft (1) having a gas turbine (11) and a propeller assembly (3) coupled to the gas turbine (11), the gas turbine (11) having a compressor (12) coupled to an air intake (13) and a temperature sensor (22) being arranged in the air intake (13) to measure the temperature of engine intake air and provide a sensed temperature (T1sens); the control system envisages: a compensation system (40) to receive the sensed temperature (T1sens) from the temperature sensor (22) and to add to the sensed temperature (T1sens) a compensation quantity (comp) to compensate for a delay introduced by the time constant (?) of the temperature sensor (22) and generate a compensated temperature (T1comp); and a control unit (20) to perform engine control operations based on the compensated temperature (T1comp).

Abstract: A center plug for attenuating noise in a gas turbine engine includes an inner skin, the inner skin having a substantially cylindrical shape and extending along an axial centerline; a forward bulkhead disposed proximate a forward end of the inner skin, the forward bulkhead connected to and extending radially outward from the inner skin; an aft bulkhead disposed proximate an aft end of the inner skin, the aft bulkhead connected to and extending radially outward from the inner skin; and an intermediate bulkhead, the intermediate bulkhead positioned intermediate the forward bulkhead and the aft bulkhead, the intermediate bulkhead spaced a distance from the inner skin and extending radially outward from the inner skin, the distance defining a radially inner portion of the intermediate bulkhead and an annular opening.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

Abstract: A method of distributing power within a gas turbine engine is disclosed. In various embodiments, the method includes driving a high pressure turbine having a first stage and a second stage with an exhaust stream from a combustor, the first stage connected to a high pressure turbine first stage spool and the second stage connected to a high pressure turbine second stage spool; driving a high pressure compressor connected to a high pressure compressor spool via a differential system, the differential system having a first stage input gear connected to the high pressure turbine first stage spool, a second stage input gear connected to the high pressure turbine second stage spool and an output gear assembly connected to the high pressure compressor spool; and selectively applying an auxiliary input power into at least one of the high pressure compressor spool and the high pressure turbine.

Abstract: An aircraft gas turbine engine comprises at least a first main engine shaft an electric machine coupled to the main engine shaft, and a lubrication pump configured to provide oil to at least a part of the electric machine. The lubrication pump is controlled by a controller. The controller is configured to operate the lubrication pump in a main gas turbine engine running mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a first flow rate relative to the rotational rate of the main engine shaft, and a fault mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a second, higher flow rate relative to the rotational rate of the main engine shaft.

Abstract: An electronic controller for an engine and a propeller, a control system and related methods are described herein. The control system comprises the controller having a first channel and a second channel independent from and redundant to the first channel. Each channel comprises a control processor configured to receive first engine and propeller parameters and to output, based on the first engine and propeller parameters, at least one engine control command and at least one propeller control command. Each channel also comprises a protection processor configured to receive second engine and propeller parameters and to output, based on the second engine and propeller parameters, at least one engine protection command and at least one propeller protection command. The control system comprises sensors for measuring the parameters of the engine and/or the propeller and effectors configured to control the engine and the propeller.

Abstract: Electrical harness for a turbomachine, comprising a first termination connected to an apparatus (3) capable of emitting an electrical signal, and a second termination connected to a signal processing apparatus (4), characterised in that it comprises an electronic circuit for filtering the signal, said circuit comprising at least one electronic component associated with an operating temperature threshold, and one thermally triggered electrical interruption means connected in series or in parallel with said electronic component, said electrical interruption means being capable of changing from an open state in which a current cannot flow through it, to a closed state in which a current can flow through it, or reciprocally from a closed state to an open state, from a tripping temperature lower than or equal to the operating temperature threshold of said electronic component.

Abstract: A hybrid aircraft propulsion system. The system comprises a gas turbine engine comprising a compressor, a combustor, one or more turbines, a shaft coupled to one of the turbines, and a bypass fan mechanically driven by the shaft. The system further comprises an electrical generator mechanically coupled to the shaft, and an auxiliary propulsor mechanically coupled to an electric motor and electrically coupled to the electric generator. At maximum power, the gas turbine engine is configured to produce a turbine entry temperature at maximum power between 1800 Kelvin and 2000 Kelvin, the engine comprises a fan bypass ratio of between 4:1 and 13:1, and the generator is configured to absorb between 10% and 60% of the mechanical power generated by the turbine.

Abstract: An apparatus and method relating an airfoil for a turbine engine with an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The airfoil includes a first cooling passage extending in the span-wise direction within the interior and a second cooling passage defining an impingement surface and located proximate the first cooling passage, an interior wall separating the first cooling passage from the second cooling passage, and at least one cooling hole passing through the interior wall.

Abstract: Flow path assemblies having features for positioning the assemblies within a gas turbine engine are provided. For example, a flow path assembly comprises an inner wall and a unitary outer wall that includes an integral combustion portion and turbine portion, the combustor portion extending through a combustion section of the gas turbine engine and the turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The flow path assembly further comprises at least two positioning members for radially centering the flow path assembly within the gas turbine engine. The positioning members extend to the flow path assembly from one or more structures external to the flow path assembly, constrain the flow path assembly tangentially, and allow radial and axial movement of the flow path assembly. Other embodiments for positioning flow path assemblies also are provided.

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: Flow path assemblies and methods for assembling a flow path assembly of a gas turbine engine are provided. For example, a flow path assembly comprises a unitary outer wall including combustor and turbine portions that are integrally formed as a single unitary structure; a single piece, generally annular inner band; and a plurality of nozzle airfoils extending from the unitary outer wall to the inner band. Each nozzle airfoil interfaces with the inner band to position the inner band within the assembly. An exemplary assembly method comprises inserting an inner band into a flow path having a unitary outer wall as its outer boundary; inserting a plurality of nozzle airfoils into the flow path; and securing the nozzle airfoils with respect to the unitary outer wall. The inner band interfaces with an inner end of each nozzle airfoil to radially locate the inner band within the flow path.

Abstract: A propulsion system includes a main gas turbine engine adapted for generating propulsive thrust during subsonic and supersonic flight operations and a supplementary propulsion unit adapted for generating additional thrust. The supplementary propulsion unit has an air intake and an exhaust for gas accelerated by the supplementary propulsion unit to provide the additional thrust and is adapted to generate the additional thrust during a limited range of subsonic flight operations, and to be dormant during other flight operations. The propulsion system has housing for the supplementary propulsion unit, including intake and exhaust covers which are moveable between deployed and stowed configurations. During the limited range of subsonic flight operations the intake and exhaust cover are moved to the deployed configuration to open the intake and the exhaust. During other flight operations the intake and exhaust cover are moved to the stowed configuration to close the intake and the exhaust.

Abstract: A combustor includes a circumferential main combustion chamber and a plurality of air inlets, the plurality of air inlets dispersing air into the circumferential main combustion chamber. The combustor also includes at least one swirl-stabilized pilot burner, which disperses combustion gases into the circumferential main combustion chamber. The air inlets and the swirl-stabilized pilot burner are aligned at least partially in a radial direction and at least partially in a circumferential direction.

Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Abstract: A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a tip speed of from 1 to 3 times that of the outer fan.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a rate of rotation of from 3 to 8 times that of the outer fan.

Abstract: An auxiliary power unit that can reduce a flow loss of gas includes: a compressor, a combustion chamber, a turbine, a turbine outlet, and a bypass duct, wherein the turbine outlet comprises an exhaust diffuser and a guide portion, wherein the bypass duct connects the compressor with the guide portion, wherein the guide portion is a channel for an air or gas and is extended radially from an outer circumferential surface of the exhaust diffuser and communicates with an inside of the exhaust diffuser via an opening, and wherein the exhaust diffuser has a first portion adjacent to a front end of the opening and a second portion adjacent to a rear end of the opening, and a radius of the second portion is larger than a radius of the first portion so that there is formed a step difference between the first portion and the second portion.