Abstract: A gas turbine engine control system comprising a variable stator vane schedule for normal operation of the gas turbine engine. The system is configured to generate an arm signal indicating potential shaft break. Then the system is configured to alter the variable stator vane schedule to slew each variable stator vane to decrease the available surge margin in response to the arm signal. Or the system is configured to limit a response rate of a variable stator vane actuator in response to the arm signal. Or the system is configured to alter the variable stator vane schedule and limit the response rate of the actuator.

Abstract: The present invention relates to control of engine variable of a gas turbine engine to regulate the surge margins of at least two compressors. A controller (20) receives data measured from engine sensors (22, 23) and uses said data to determine an indication of surge margin for each of at least two compressors (6, 7) of the gas turbine engine. The controller (20) uses the indications of surge margin for each of the compressors to determine a control strategy that balances the requirements of each compressor. In one embodiment a surge margin operating map divided into different control domains (40, 43, 43) is used. The indication of surge margin determined for each compressor is plotted to determine which control domains the current operating point on the surge margin operating map falls within.

Abstract: It is known that overshoot in fuel flow rate as a result of compensating for engine heat soak effects can create transient over values in desired thrust parameters. These over values may be displayed and/or utilised in other engine control systems, and although transient may create problems within the engine and apprehension with a user. By adjusting the normally calculated fuel flow rate to a fuel regulator by the introduction of a fuel flow adjustment dependent upon a variable which is related to heat soak effects it is possible to reduce the extent of overshoot as well as the transient time period of that overshoot to ongoing operational state. The variable used may be direct or indirect and will generally utilise a temperature sensor within the engine and/or other engine control procedures which may be extrapolated from such sensor values to provide a variable for heat soak compensation.

Abstract: A compressor variable stator vane arrangement comprises at least one stage of variable stator vanes. A speed sensor measures the rotational speed of the compressor rotor. A pressure sensor measures the outlet pressure of the compressor. A second pressure sensor, a temperature sensor and a third pressure sensor measure the total inlet pressure, the temperature and the ambient pressure at the inlet of the gas turbine engine. A processor determines a target operating line as a function of ambient pressure and total inlet pressure. The target operating line is defined to ensure the gas turbine engine operates simultaneously at both the minimum required compressor speed and the minimum required compressor outlet pressure when commanded to idle to minimise idle thrust and fuel burn. The processor determines if the operating point of the compressor, defined in terms of corrected outlet pressure and corrected rotational speed of the compressor rotor is above or below the target operating line of the compressor.

Abstract: It is known that power is normally extracted within a gas turbine engine from a shaft in order to drive auxiliary devices such as electrical power generators. With high extraction rates from such shafts there is mismatching in rotational speeds between the shafts which must be adjusted through leakage taken from valves 8, 9. By measuring parameters and in particular pressure at a number of positions or parts within an engine arrangement 1 and comparing those parameters either through ratios or directly with reference values a controller can be configured in order to choose the correct sequence and scheduling of the valves 8, 9 opening to balance shaft speed. In such circumstances more efficient operation of the engine arrangement can be achieved over a wider range of engine speeds including idling.

Abstract: A method of detecting an ice shedding event in a gas turbine engine, the engine having a rotor comprising a compressor drivingly connected via a shaft to a turbine; the method for detecting an ice shedding event comprising the steps of measuring the temperature at regular intervals and where a temperature drop of at least 20 degrees per second is recorded, producing a signal indicative of an ice shedding event and sending the signal to an indicator device. Advantageously, the engine may then be inspected for damage associated to an ice impact rather than other foreign object ingestion event or fuel flow irregularity. Alternatively, instead of temperature, pressure may be measured and a pressure drop of at least 20 kPa per second is indicative of an ice shedding event.

Abstract: An apparatus and method for inducing waves within a container arranged to permit a through-flow of fluid flowing generally from an inlet of the container to an exit of the container, wherein the container has wave inducing means located at the exit of the container, the wave inducing means being operable to vary the area of the exit thereby inducing waves within the container. The apparatus may form part of a combustor test rig.

Abstract: A gas turbine engine system comprises a first compression stage; a second compression stage; a combustor; a controller; a first sensor for sensing the speed of the first compression stage and providing a first indication of the sensed speed to the controller; and a second sensor for sensing the speed of the second compression stage and providing a second indication of the sensed speed to the controller, wherein the controller is operable to control the supply of fuel to the combustor in dependence upon the first indication received from the first sensor and the second indication received from the second sensor. This arrangement is particularly useful in controlling the acceleration of an aero-engine from minimum idle.

Abstract: A stall detection and recovery system (10) for a gas turbine engine (1) comprises surge detection means (50 for detecting an engine surge and providing a surge detection signal (154) for indicating said engine surge. The system also includes stall detection means to provide a stall detection signal (48) relating to a condition of gas flowing through the engine (1). The system (10) also includes stall recovery means (58) to receive the surge detection signal (154) and the stall detection signal. The stall recovery means (58) is arranged to control combustor operating means, whereby when the stall detection signal (48) indicates an engine stall, the stall recovery means controls the combustor operating means to effect recover from the stall.

Abstract: An apparatus and method for controlling a gas turbine engine involves calculating a simple corrected speed value for the engine using values of compressor speed and inlet stagnation temperature of the compressor. This simple corrected speed value is then adjusted to take into account the water vapour concentration at the inlet to the compressor. This results in a more accurate value of corrected speed for the compressor, which may be used to modulate the fuel supply to a combustor of the engine or to control the geometry of one or more variable compressors in the engine.

Abstract: A gas turbine engine system comprises a first compression stage; a second compression stage; a combustor; a controller; a first sensor for sensing the speed of the first compression stage and providing a first indication of the sensed speed to the controller; and a second sensor for sensing the speed of the second compression stage and providing a second indication of the sensed speed to the controller, wherein the controller is operable to control the supply of fuel to the combustor in dependence upon the first indication received from the first sensor and the second indication received from the second sensor. This arrangement is particularly useful in controlling the acceleration of an aero-engine from minimum idle.

Abstract: A gas turbine engine (10) comprises a fan rotor (26) having a plurality of fan blades (24) and an apparatus (34) for detecting damage to the fan blades (24). The apparatus (34) comprises a transducer (36) arranged to detect the pressure in the gas flow around the fan blades (34) and to produce a pressure signal. A speed sensor (48) is arranged to measure the speed of rotation of the fan rotor (26) and to produce a speed signal. A processor unit (40) analyses the pressure signal and the speed signal to detect changes in the amplitude of the pressure signal, which occur at multiples of the rotational frequency of the fan rotor (24). The changes indicate differences in pressure between the gas flow around a damaged fan blade (24) and the gas flow around the remainder of the fan blades (24). The processor unit (40) sends a signal indicative of damage to a fan blade (24), if the difference in pressure is above the predetermined level, to an indicator device (44, 46).

Abstract: An apparatus and method for controlling a gas turbine engine involves calculating a simple corrected speed value for the engine using values of compressor speed and inlet stagnation temperature of the compressor. This simple corrected speed value is then adjusted to take into account the water vapour concentration at the inlet to the compressor. This results in a more accurate value of corrected speed for the compressor, which may be used to modulate the fuel supply to a combustor of the engine or to control the geometry of one or more variable compressors in the engine.

Abstract: A gas turbine engine (10) comprises a fan rotor (26) having a plurality of fan blades (24) and an apparatus (34) for detecting damage to the fan blades (24). The apparatus (34) comprises a transducer (36) arranged to detect the pressure in the gas flow around the fan blades (34) and to produce a pressure signal. A speed sensor (48) is arranged to measure the speed of rotation of the fan rotor (26) and to produce a speed signal. A processor unit (40) analyses the pressure signal and the speed signal to detect changes in the amplitude of the pressure signal, which occur at multiples of the rotational frequency of the fan rotor (24). The changes indicate differences in pressure between the gas flow around a damaged fan blade (24) and the gas flow around the remainder of the fan blades (24). The processor unit (40) sends a signal indicative of damage to a fan blade (24), if the difference in pressure is above the predetermined level, to an indicator device (44, 46).

Abstract: A control system for a ducted fan gas turbine engine (10) comprises a control unit (29) that receives input signals from a pressure transducer (27) located within the engine's fan duct (11) downstream of the fan (13) and a rotational speed transducer (32) mounted adjacent the shaft (24) carrying the fan (13). In the event of the control unit (29) detecting a fall in the air pressure downstream of the fan (13) and an increase in fan speed, it commands a temporary reduction in the fuel flow to the engine (10). Such a decrease in the air pressure downstream of the fan (13) and increase in fan speed is indicative of the fan (13) entering a stall condition. The temporary reduction in fuel flow to the engine (10) enables the fan (13) to recover from that stall condition.

Abstract: If the fan of a ducted fan gas turbine engine is damaged, the power output of the engine may not be sufficient for a given throttle setting. The present invention provides a control system which detects fan damage and ensures that there is sufficient fuel flow to the engine to ensure that the overall engine power output is consistent with that commanded by the throttle setting.

Abstract: A fuel control system for a premixed lean burn combustor in a combustion turbine engine, the combustor having at least two stages (P,S,T), the control system including; an air flow evaluation module having means utilising an iterative algorithm to derive an accurate value of total air flow to the combustor, the evaluation module further having means dividing the derived total air flow value into the proportions (Mp, Ms, Mt) required by the combustor stages (p,s, t) and outputting signals in parallel representing the air flow to respective combustor stages (p,s,t), and a control module associated with each stage (p,s,t) of the combustor and responsible for setting the fuel demand signal for the associated combustor stage (p,s,t), each control module being adapted to set the fuel demand signal with reference to an appropriate one of said output signals.

Abstract: A multispool gas turbine engine for an aircraft application includes a transmission system which is operative to transfer power between relatively rotatable engine spools. In a first mode the transmission system is operative to transfer power from the engine's low pressure spool to its high pressure spool and thereby improve the engine's in-flight relight characteristics; in a second mode it is operative to transfer power from the engine's intermediate pressure spool to its high pressure spool to improve engine part speed performance; and in a third mode to transfer power from the engine's high pressure spool to its intermediate pressure spool to assist in ground starting. With each shaft is an associated flow displacement machine (64,66,62) operable as a pump or motor. In other embodiments, a differential is employed for the power transfer between spools.

Abstract: A ducted fan gas turbine engine is provided with means for measuring the actual temperature of air exhausted from its high pressure compressor and means for computing a theoretical value of that temperature.

Abstract: A ducted fan gas turbine engine is provided with a plurality of sensors which detect the presence of liquid water in air exhausted from the high pressure compressor of the engine into its combustion system. In the event that liquid water is detected, the fuel flow to the combustion system is increased to cause the high pressure compressor to rotate faster and thereby vaporize the liquid water. Such liquid water can be present in the air exhausted from the high pressure compressor when the engine is operating in very heavy rain and can adversely affect normal engine operation.