Abstract: A fluidic device for providing analogue output control includes a main channel, a first control channel, a second control channel, a comparator which receives respective input fluid flows from the main, the first and the second control channels. The first control channel is configured such that the input fluid flow therefrom carries an oscillating pressure wave signal, the second control channel includes a flow regulator controllable to vary the mass flow rate of the input fluid flow from the second control channel, and the main channel is configured such that the input fluid flow therefrom is at a reference mass flow rate. The comparator is configured such that the input fluid flows from the first control and the second control channels act in combination on the input fluid flow from the main channel to produce an output fluid flow from the comparator having a PWM mass flow rate characteristic.

Abstract: A system and method are provided for controlling turbine blade tip-to-static structure clearance in a gas turbine engine installed on an aircraft. Mode control data are processed to determine that a fuel-saving mode is enabled, and aircraft data are processed to determine that the aircraft gas turbine engine is generating a substantially constant thrust. The turbine blade tip-to-static structure clearance in the aircraft gas turbine engine is minimized upon determining that both the aircraft gas turbine engine is generating a substantially constant thrust and the fuel-saving mode is enabled. The turbine blade tip-to-static structure clearance in the aircraft gas turbine engine is then selectively increased to a predetermined clearance, and a change in aircraft gas turbine engine thrust is prevented until the predetermined clearance is achieved.

Abstract: A gas turbine engine comprising first and second axially spaced turbine rotor stages (46, 48) and a turbine casing (56) radially outside the rotor stages. A first seal segment arrangement (58) forms a cavity (64) radially between the first turbine rotor stage (46) and the turbine casing (56). A first air source (82) is coupled to the first seal segment arrangement (58). A second seal segment arrangement (70) forms a cavity (74) radially between the second turbine rotor stage (48) and the turbine casing (56). A heating chamber (84) is provided radially between the second seal segment arrangement (70) and the turbine casing (56). A duct (86) is coupled between the first air source (82) and the heating chamber (84).

Abstract: A control system for use in controlling a cabin blower system. The cabin blower system includes a gas turbine engine spool, a cabin blower compressor powered by the spool and arranged in use to compress fluid used in a cabin of an aircraft, and one or more control mechanisms via which the control system controls the power extracted by the cabin blower compressor from the spool. The control system is arranged in use to control the power extracted from the spool by the cabin blower compressor in accordance with one or more primary control parameters. The control system is arranged in use to alter the spool power extracted by the cabin blower compressor by comparison with the power that would have been extracted in accordance with the primary control parameters alone, in response to modifications in a secondary control parameter indicative of the commencement or occurrence of an engine transient.

Abstract: A fuel flow system comprising a fuel manifold and a fuel nozzle. A fluidic valve having an inlet, a first outlet and a control flow inlet. The inlet is coupled to the fuel manifold. The first outlet is coupled to the fuel nozzle. The control flow inlet is coupled to a source of control fuel. The fluidic valve is configured to selectively couple the inlet to the first outlet dependent on the control flow into the control flow inlet.

Abstract: A gas turbine engine is disclosed having a tip clearance control system. The tip clearance control system has a cabin blower system, a casing arranged in use about a rotor of a gas turbine engine and a fluid delivery passage. The cabin blower system having a cabin blower compressor arranged in use to compress fluid used in a cabin of an aircraft and to compress fluid conducted via the fluid delivery passage into heat exchange with the casing.

Abstract: A thrust demand signal is provided to a processor of a gas turbine engine and is modified, according to growth time constants of a rotor and/or a casing of the engine, in order to control the rotational speed or the rate of change of rotational speed of the engine so as to prevent contact between the rotor and the casing.

Abstract: A clearance control device including a segment having a passage to deliver fluid towards a component rotating past the segment. Also a fluid flow device having a first fluid path coupled to the passage and a second fluid path that is decoupled from the passage. A first plasma generator is located in the fluid flow device that directs fluid towards the first fluid path; a second plasma generator is located in the fluid flow device that directs fluid towards the second fluid path; and a control arrangement is configured to alternately energize the first and second plasma generators at an energizing frequency to deliver fluid to the passage at a frequency coincident with the passing frequency of the component.

Abstract: A fluidic actuator comprising: a fluid nozzle for delivering fluid and a tube having an open end and a closed end, the open end spaced from the fluid nozzle. Also a pair of electrodes mounted in the tube and spaced apart to create a spark gap therebetween. A voltage source is arranged to supply a voltage across the pair of electrodes wherein the voltage causes plasma formation in the spark gap thereby shortening the effective length of the tube.

Abstract: A method of controlling a rotor tip clearance arrangement of a gas turbine engine and a control system configured to control rotor tip clearance. Steps include measuring at least one engine parameter; determining engine power demand from the at least one engine parameter; and calculating rotor tip clearance given the determined engine power demand. The rotor tip clearance arrangement is controlled to increase or decrease the rotor tip clearance based on the difference between the calculated clearance and a predefined target clearance.

Abstract: A method of detecting shaft break in a gas turbine engine having a shaft system. The shaft system including a shaft that couples a compressor and a turbine. First construct a frequency model of the shaft system. Then determine a notch frequency and a first torsional frequency for the shaft system from the model. Then in real time measure a rotational speed of the shaft; detect the presence or absence of a feature at least one of the notch frequency and the first torsional frequency in the measured speed; and generate a shaft break signal in response to the absence of at least one of the features.

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 method to determine inertia of components of a rotating shaft system. The shaft system includes a shaft coupling a turbine to drive the rotation and a load to be driven by the rotation. The method includes steps to: apply a feedback to a forcing input to the shaft system; measure resonant frequency of the shaft; iterate steps 1.a) and 1.b) for different feedbacks; plot resonant frequency squared against gain; and determine inverse of gradient from the plot to give inertia of the turbine. Also a method to determine shaft stiffness using the inertia of the turbine.

Abstract: The present invention provides a method of detecting shaft break in a shaft system comprising a shaft coupled between two masses. The method comprises a number of steps. Firstly, to define a time-dependent rotational speed equation for the shaft in terms of system inertia for an engine transient event. Then to discretize the rotational speed equation in terms of a discrete time constant in the discrete domain. Then to recursively define the discretized equation to give a recursive equation and to solve the recursive equation to determine the discrete time constant. Then to define a threshold as a function of engine power and then to set a shaft break signal to TRUE if the discrete time constant is greater than the threshold. A shaft break detection system is also provided by the present invention.

Abstract: A method to determine inertia of components of a rotating shaft system. The shaft system includes a shaft coupling a turbine to drive the rotation and a load to be driven by the rotation. The method includes steps to: apply a feedback to a forcing input to the shaft system; measure resonant frequency of the shaft; iterate steps 1.a) and 1.b) for different feedbacks; plot resonant frequency squared against gain; and determine inverse of gradient from the plot to give inertia of the turbine. Also a method to determine shaft stiffness using the inertia of the turbine.

Abstract: A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The gas turbine engine further has a system (i) for cooling the turbine section and (ii) for providing tip clearance control between turbine blades of the turbine section and a plurality of circumferentially distributed segments which form an annular shroud surrounding the outer tips of the turbine blades. The system includes a turbine section cooling sub-system which diverts a first cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool components thereof. The first cooling air flow by-passes the combustor and is cooled in the heat exchanger. The turbine section cooling subsystem has a first valve arrangement which regulates the first cooling air flow. The system further includes a tip clearance control sub-system which supplies a second cooling air flow to an engine case to which the segments are mounted.

Abstract: A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The engine further has a cooling system which diverts a cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool a component thereof. The cooling air flow by-passes the combustor and is cooled in the heat exchanger. The cooling system has a valve arrangement which regulates the cooling air flow. The engine further has a closed-loop controller which estimates and/or measures one or more temperatures of the cooled component, compares values derived from the estimated and/or measured temperatures with one or more corresponding targets, and issues a demand signal to the valve arrangement based on the comparison and a value of the demand signal at a previous time step.

Abstract: A gas turbine engine comprising first and second axially spaced turbine rotor stages (46, 48) and a turbine casing (56) radially outside the rotor stages. A first seal segment arrangement (58) forms a cavity (64) radially between the first turbine rotor stage (46) and the turbine casing (56). A first air source (82) is coupled to the first seal segment arrangement (58). A second seal segment arrangement (70) forms a cavity (74) radially between the second turbine rotor stage (48) and the turbine casing (56). A heating chamber (84) is provided radially between the second seal segment arrangement (70) and the turbine casing (56). A duct (86) is coupled between the first air source (82) and the heating chamber (84).

Abstract: A method of controlling a rotor tip clearance arrangement of a gas turbine engine and a control system configured to control rotor tip clearance. Steps include measuring at least one engine parameter; determining engine power demand from the at least one engine parameter; and calculating rotor tip clearance given the determined engine power demand. The rotor tip clearance arrangement is controlled to increase or decrease the rotor tip clearance based on the difference between the calculated clearance and a predefined target clearance.

Abstract: A thrust demand signal is provided to a processor of a gas turbine engine and is modified, according to growth time constants of a rotor and/or a casing of the engine, in order to control the rotational speed or the rate of change of rotational speed of the engine so as to prevent contact between the rotor and the casing.