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.

The invention relates to an aircraft propulsion system, comprising a main transmission unit (12) and at least two turbojet engines connected to the main transmission unit (12), respectively a first turbojet engine (14a) and a second turbojet engine (14b), each turbojet engine comprising a free turbine (24a, 24b), characterized in that the first turbojet engine (14a) comprises a heat exchanger (30) configured to recover some of the thermal energy from the exhaust gas at the outlet of the free turbine, and in that the propulsion system comprises at least one computer (28a, 28b) configured to control the two turbojet engines and to limit the acceleration and the deceleration of the first turbojet engine (14a) when neither of the turbojet engines is broken down, in order to limit the reactor power transients at the heat exchanger (30).

Systems and methods for controlling a gas turbine engine are provided. The system comprises an interface to a fuel flow metering valve for controlling a fuel flow to the engine in response to a fuel flow command and a controller connected to the interface and configured for outputting the fuel flow command to the fuel flow metering valve in accordance with a required fuel flow. The controller comprises a feedforward controller configured for receiving a requested engine speed and acceleration, obtaining a steady-state fuel flow for the requested engine speed and a relationship between fuel flow and gas generator speed, and determining the required fuel flow to obtain the requested engine acceleration as a function of the requested engine speed, the steady-state fuel flow, and the relationship between fuel flow and gas generator speed.

Systems and method for filling a fuel manifold comprising at least a primary and a second manifold of a gas turbine engine are described. The method comprises providing fuel flow to the secondary manifold of the gas turbine engine, the secondary manifold being partly or completely empty; monitoring at least one engine operational parameter of the gas turbine engine as fuel fills the secondary manifold; and accelerating the engine when a transition threshold is reached, the transition threshold being associated with the engine operational parameter and indicative that fuel has reached the combustor.

A system includes a gas turbine engine having a low speed spool, a high speed spool, and a combustor. The system also includes a low spool motor configured to augment rotational power of the low speed spool. The system further includes a controller configured to cause fuel flow. The controller is operable to control the low spool motor to drive rotation of the low speed spool responsive to a thrust command while the controller does not command fuel flow to the combustor.

A device for driving an integrated generator from an accessories relay box of a turbomachine. The device includes first and second electric motors arranged to transfer electric power from one to the other, one or more controllers configured for controlling said electric motors, and an epicyclic reduction gear train. The gear train includes a first element intended to be coupled to the accessories relay box, a second element intended to be coupled to the generator, and a third element driven to rotate by said first electric motor. The control means are configured to modify the speed of rotation of the third element in such a way that the second element is driven to rotate at a constant speed.

A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; and an engine controller configured to identify a condition to the effect that the engine is in an approach idle condition, and operate the first electric machine in a motor mode and operate the second electric machine in a generator mode to transfer power electrically from the LP spool to the HP spool to thereby reduce the LP spool rotational speed and increase the HP spool rotational speed.

A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; a combustion system comprising a fuel metering unit; and an engine controller configured to, in response to a change of a throttle lever angle setting indicative of an acceleration event, increase fuel flow to the combustion system by the fuel metering unit, and to operate the first electric machine in a motor mode to increase the HP spool rotational speed and engine core mass flow.

A method of operating a gas turbine engine includes commanding an acceleration of the gas turbine engine and moving a variable pitch high pressure compressor vane toward an open position thereby reducing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal. An active clearance control system of a gas turbine engine includes an engine control system configured to command an acceleration of the gas turbine engine and move a variable pitch high pressure compressor vane toward an open position thereby slowing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal located radially outboard of the high pressure turbine rotor.

A propulsion system is disclosed comprising a gas turbine engine and an acceleration schedule which determines the rate of acceleration of the gas turbine engine from an idle condition in response to a demand for increased thrust off-idle. The acceleration schedule determines the rate of acceleration in dependence upon the value of an engine parameter of the engine the value of which is substantially unaltered by variation in the magnitude of an electrical load drawn from the engine while it is operating in the idle condition.