A multi-shaft gas turbine engine has plural engine spools. A first spool of the engine and a second spool of the engine are operatively connected by an electrical machine that transfers power from one of the spools to the other.

An inlet door state control system for an auxiliary power unit (APU) includes an inlet door configured to allow air into an inlet of the APU. The system also includes an actuator configured to open the inlet door. The system also includes an APU threshold switch connected to the actuator, the APU threshold switch configured to automatically change from an open position to a closed position when a predetermined APU operating condition is met. In the system, the actuator opens the inlet door when the APU threshold switch is in the closed position.

A liquid fuel rocket engine according to one example includes a combustor, a liquid fuel repository connected to the combustor via a fuel line and a first valve, an oxidizer repository connected to the combustor via an oxidizer line and a second valve, a valve controller configured to output a valve control current to the first valve, the valve controller storing instructions for determining at least one actual minimum impulse bit of a valve based on a current profile and a voltage profile of a single operation of the first valve, and to adjust valve controls to account for the at least one actual minimum impulse bit.

A method of managing the thermal output of a gas combustion engine having a plurality of combustion zones uses a temperature control system integrated with the engine to provide digital thermal signals to a control unit. The control unit determines when to adjust an engine operating parameter based on an analysis of the thermal signals. The temperature control system includes a plurality of smart thermal sensors with at least one sensor for each combustion zone, a communication bus, and a control unit. Each smart thermal sensor converts the measured temperature from an analog thermal signal to a digital thermal signal.

Systems and methods are provided for a dual-pump fuel delivery system. The system includes a primary pump to deliver a first fluid flow to the engine via a pump outlet. A secondary pump delivers a second fluid flow and a fuel metering valve to return the second fluid flow to a pump inlet in a first position, and to channel the second fluid flow to the pump outlet in a second position

A method of stopping an engine of a rotorcraft in overspeed, the engine comprising a gas generator and a power assembly. When the engine is in operation, a relationship is established giving a limit derivative that varies as a function of the speed of rotation of the power assembly. The speed of rotation, referred to as the current speed, reached by the power assembly is measured and the time derivative of the speed of rotation is determined and referred to as the current derivative. The engine is stopped automatically when the limit derivative corresponding to the current speed as determined by the relationship is less than or equal to the current derivative.

A variable vane control system for controlling the angle of rotation of a circumferential row of variable vanes of a gas turbine engine. The control system includes a mechanical linkage operable to rotate the variable vanes, one or more actuators for operating the linkage and one or more position sensors for detecting the respective actuation positions of the one or more actuators. The control system further includes a linkage position signalling switch for signalling that the mechanical linkage is at a calibration position corresponding to a predetermined rotation angle of the vanes. The control system further includes a controller for controlling the one or more actuators and thereby controlling the angle of rotation of the vanes, based on the detected actuation positions which the controller correlates with vane rotation angle. The controller further corrects the correlation between the detected actuation positions and vane rotation angle on receipt of a signal from the linkage position signalling switch indicating that the mechanical linkage is at the calibration positionvanes.

A solid state spark device that operates as a two terminal spark gap in a CDI exciter of an aircraft ignition system. The device includes a triggering transformer, a triggering circuit, and a control circuit. The triggering circuit is electrically connected to a first coil of the transformer and includes circuit elements connected to supply current to the first coil upon charging of the triggering circuit up to a triggering voltage. This current through the first coil of the triggering transformer induces an output in a second coil of the transformer. The control circuit is electrically connected to the second coil and includes a switch controlled by the output from the second coil. The switch, when activated by the triggering circuit, discharges energy from the exciter into an igniter of the aircraft ignition system.

A turbo-machine, which can be operated in an optimized driving range is provided. To this end, a method for operating a turbo-machine having at least one turbo-machine stage, which has at least one rotary shaft is disclosed. According to the method, the following method steps are carried out: a) determining a desired efficiency characteristic value .sub.soll of the turbomachine stage; b) determining an actual efficiency characteristic value .sub.ist of the turbo-machine-stage; c) determining a comparison efficiency characteristic value of the turbo-machine stage by comparing the actual efficiency characteristic value .sub.ist and the desired efficiency characteristic value .sub.soll to one another; and d) changing at least one operating parameter of the turbo-machine stage subject to the comparison efficiency characteristic value .sub.vgl, wherein in order to determine the actual efficiency characteristic value .sub.ist, a measuring of a torque of the rotary shaft of the turbo-machine-stage is carried out.

A turbine engine including a core engine cowl including a compartment, a cooling airflow source positioned within the compartment, and a full authority digital engine control (FADEC) system coupled in communication with the cooling airflow source. The FADEC system is configured to determine a flight status of the turbine engine, and actuate the cooling airflow source when the turbine engine is not in flight, and before the turbine engine has been shut down, such that heat is exhausted from the compartment.