Abstract: Methods and systems for slowing down an aircraft having a propeller. The method comprises operating the propeller at a reference speed with the propeller blades in a first position, applying a load to the propeller to slow down a rotational speed of the propeller when the propeller is in a windmilling state or just before the propeller enters the windmilling state, adjusting the propeller blades to a second position in response to the load being applied, to increase the rotational speed of the propeller back towards the reference speed, and operating the propeller at the reference speed with the propeller blades in the second position.

Abstract: Herein provided are methods and systems for controlling an engine having a variable geometry mechanism. A pressure ratio between a first pressure at an inlet of the engine and a predetermined reference pressure is determined. A fuel flow rate to the engine is determined. The fuel flow rate is adjusted based at least in part on the pressure ratio to obtain a corrected fuel flow rate. A position control signal for the variable geometry mechanism of the engine is generated based on the corrected fuel flow rate. The position control signal is output to a controller of the engine to control the variable geometry mechanism.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a compressor, and a turbine section where the turbine shaft is configured to compound power with the engine shaft. The turbine section may include a first stage turbine and a second stage turbine. The turbine shaft is rotationally supported by a plurality of bearings all located on a same side of the compressor rotor(s) and all located on a same side of the turbine rotor(s), for example all located between the compressor rotor(s) and the turbine rotor(s), such that the compressor rotor(s) and the turbine rotor(s) are cantilevered. A method of driving a rotatable load of an aircraft is also discussed.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a turbine section, and a compressor having an outlet in fluid communication with an inlet of the engine core. A casing is connected to the turbine section, the compressor and the engine core. A mount cage is connected to mounts attached to the casing between the compressor and a hot zone including the turbine section and the engine core. The mount cage includes a plurality of struts all extending from the mounts away from the turbine section and engine core. The casing may be a gearbox module casing through which the turbine shaft in engaged with the engine shaft. The mount cage may be completely contained within an axial space with the turbine section and engine core being located outside of the axial space.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a turbine section, and a compressor having an outlet in fluid communication with an inlet of the engine core. A casing is connected to the turbine section, compressor and engine core. A mount cage is connected to mounts attached to the casing between the compressor and a hot zone including the turbine section and exhaust pipe(s). The struts are separated from the hot zone by at least one firewall. The mount cage may include a plurality of struts all extending from the mounts away from the turbine section and engine core. The casing may be a gearbox module casing through which the turbine shaft in engaged with the engine shaft. The mount cage may be completely contained within an axial space with the turbine section and exhaust pipe(s) being located outside of the axial space.

Abstract: Systems and methods for controlling an gas turbine engine are provided. The method comprises receiving a requested engine speed and obtaining a shaft inertia of the engine, a steady state fuel flow for the requested engine speed, and a relationship between fuel flow and acceleration power generated by the fuel flow. A required fuel flow to obtain an engine acceleration is determined as a function of the requested engine speed, the shaft inertia of the engine, the steady state fuel flow for the requested engine speed, and the relationship between fuel flow and acceleration power generated by the fuel flow. A command to a fuel flow metering valve is output in accordance with the required fuel flow.

Abstract: A compound engine assembly with an inlet duct having an inlet surrounded by an inlet lip including at least one conduit extending therethrough, a compressor, an engine core including at least one internal combustion engine, a turbine section having a turbine shaft in driving engagement with the engine shaft, and an exhaust conduit providing a fluid communication between the outlet of the turbine section and the conduit(s) of the inlet lip. An exhaust duct and ant exhaust conduit providing a fluid communication between the outlet of the turbine section and the exhaust duct may also be provided. The internal combustion engine(s) may be rotary engine(s). A method of driving a rotatable load of an aircraft is also discussed.

Abstract: Herein provided are methods and systems for setting an acceleration schedule for engine start of a gas turbine engine. A rotational acceleration measurement of the engine is obtained after the engine is energized in response to a start request. The rotational acceleration measurement of the engine is compared to a threshold value within an acceleration band having a maximum threshold and a minimum threshold. An acceleration schedule is determined based on a position of the rotational acceleration measurement of the engine in the acceleration band relative to the threshold value.

Abstract: An engine assembly including an engine core with at least one internal combustion engine, a first casing, a turbine module including a second casing located outside of the first casing, and a compressor module including a third casing located outside of the first and second casings. The turbine shaft extends into the first casing, is rotationally supported by a bearings all contained within the first casing, and is free of rotational support within the second casing. The first casing may be a gearbox module casing through which the turbine shaft is in driving engagement with the engine shaft. A method of driving a rotatable load of an aircraft, and an engine assembly with a rotary engine core, a gearbox module with a first casing, and a second module including a second casing located outside of the first casing and detachably connected to the first casing are also discussed.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a turbine section including a turbine shaft in driving engagement with the engine shaft, and a compressor, and a firewall. The compressor is located on one side of the firewall, and the turbine section and the engine core are located on the other side. The assembly may include a gearbox module with the turbine section and the engine core located on a same side of the gearbox module casing and the compressor located on the opposite side of the gearbox module casing, and with the firewall extending from the gearbox module casing. One or more rotatable accessory may be located on a same side of the firewall as the compressor. A method of reducing fire hazard in a compound engine assembly is also discussed.

Abstract: Herein provided are methods and systems for setting an acceleration schedule for engine start of a gas turbine engine. A rotational acceleration measurement of the engine after the engine is energized in response to a start request is obtained. The rotational acceleration measurement of the engine is compared to an acceleration band having a maximum threshold and a minimum threshold. An acceleration schedule is determined based on a position of the rotational acceleration measurement of the engine in the acceleration band.

Abstract: A compound engine assembly with an inlet duct, a compressor, an engine core including at least one internal combustion engine, and a turbine section including a turbine shaft configured to compound power with the engine shaft. The turbine section may include a first stage turbine and a second stage turbine. The turbine shaft and the engine shaft are parallel to each other. The turbine shaft, the engine shaft and at least part of the inlet duct are all radially offset from one another. A method of driving a rotatable load of an aircraft is also discussed.

Abstract: Systems and methods for directing fuel flow to an engine when the engine is in an electronic manual override mode are described herein. In accordance with an aspect, a commanded fuel flow to the engine is determined from a fuel schedule based on the position on an engine control lever; a limit is applied on the commanded fuel flow when the commanded fuel flow exceeds a maximum fuel flow threshold; and fuel flow is directed to the engine based on the commanded fuel flow.

Abstract: Systems and methods for limiting power of a gas turbine engine for an aircraft are described herein. A blade angle of a propeller blade of the engine and a commanded power for the engine are obtained. A thrust transition direction is determined. The commanded power is compared to a selected threshold based on the blade angle and the thrust transition direction. Power to the engine is limited when the commanded power exceeds the selected threshold.

Abstract: Herein provided are methods and systems for detecting a high temperature condition of a gas turbine engine. A fuel flow to a combustor of the engine and a compressor outlet pressure of the engine are obtained. A ratio of the fuel flow to the compressor outlet pressure is determined. The ratio is compared to a threshold and a high temperature condition of the engine is detected when the ratio exceeds the threshold.

Abstract: Herein provided are methods and systems for setting an acceleration schedule for engine start of a gas turbine engine. A rotational acceleration measurement of the engine after the engine is energized in response to a start request is obtained. The rotational acceleration measurement of the engine is compared to an acceleration band having a maximum threshold and a minimum threshold. An acceleration schedule is determined based on a position of the rotational acceleration measurement of the engine in the acceleration band.

Abstract: Herein provided are methods and systems for setting a fuel flow schedule for starting a gas turbine engine of an aircraft. Aircraft speed and engine rotational speed are obtained. A compressor inlet recovered pressure is estimated by combining a first component influenced by the aircraft speed and a second component influenced by the engine rotational speed, and a fuel flow schedule is selected for engine start in accordance with the estimated compressor inlet recovered pressure.

Abstract: Systems and methods for detecting and accommodating for loss of a torque signal of a gas turbine engine are described herein. An engine deterioration offset may be determined while the torque signal of the engine is available. Then, in the event that the torque signal is lost, a predicted operating offset may be determined. A synthesized torque signal may be generated when the torque signal is lost at least in part from the engine deterioration offset and the predicted operating offset.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a compressor, and a turbine section where the turbine shaft is configured to compound power with the engine shaft. The turbine section may include a first stage turbine and a second stage turbine. The turbine shaft is rotationally supported by a plurality of bearings all located on a same side of the compressor rotor(s) and all located on a same side of the turbine rotor(s), for example all located between the compressor rotor(s) and the turbine rotor(s), such that the compressor rotor(s) and the turbine rotor(s) are cantilevered. A method of driving a rotatable load of an aircraft is also discussed.

Abstract: Systems and methods for controlling an gas turbine engine are provided. The method comprises receiving a requested engine speed and obtaining a shaft inertia of the engine, a steady state fuel flow for the requested engine speed, and a relationship between fuel flow and acceleration power generated by the fuel flow. A required fuel flow to obtain the requested engine speed is determined as a function of the requested engine speed, the shaft inertia of the engine, the steady state fuel flow for the requested engine speed, and the relationship between fuel flow and acceleration power generated by the fuel flow. A command to a fuel flow metering valve is output in accordance with the required fuel flow.