Patent Description:
There is a general need to improve on known techniques for determining the operating condition of a gas turbine engine for an aircraft.

<CIT> describes a system for monitoring an acoustic signal in a gas turbine engine using frequency filtering to detect possible issues in the gearbox.

<CIT> describes a method and apparatus comprising a pressure or sound sensor for determining and controlling the speed of a fan of a gas turbine engine.

<CIT> describes a method and apparatus for detecting and locating fluid leaks using a pair of geophones and a method of pulse waveform comparison.

<CIT> describes a fault detection and diagnosis system for a gas turbine engine utilising MFCC and CELP algorithms to allow matching of acoustic features from vibration or acoustic sensor readings with feature profiles stored in a memory.

According to a first aspect there is provided a method of automatically determining an operating condition of at least part of a gas turbine engine for an aircraft according to claim <NUM>.

The method may comprise measuring one or more gas pressure waves by one or more further gas pressure detectors, wherein the one or more further gas pressure detectors are located in the gas turbine engine; and said automatic determination of an operating condition is also dependent on each output signal from the one or more further the gas pressure detectors.

The method may comprise the computing system performing a Fourier Transform on the output signal of each gas pressure detector to generate a frequency domain version of the output signal.

The method may comprise determining an operating condition by comparing, by the computing system, the output signal of each gas pressure detector with one or more predetermined signals.

The method may comprise said comparison of the output signal of each gas pressure detector with one or more predetermined signals being a comparison of the frequency domain version of each output signal with one or more predetermined signals.

The method may comprise filtering, by the computing system, each output signal of a gas pressure detector.

The method may comprise monitoring, by the computing system, the operating condition of at least part of the gas turbine engine; and detecting, by the computing system, a change in the operating condition.

The method may comprise determining, by the computing system, a type of change of operating condition of at least part of the gas turbine engine in dependence on the comparison.

The method may comprise the determined type of change including a pipe cracking, bursting and/or leaking; and/or the determined type of change comprising the type of pipe that a change has occurred in, such as a cabin air pipe, anti-ice air pipe or handling bleed pipe.

The method may comprise determining the location of the change.

The method may comprise the location of the change being determined in dependence on the difference in time of arrival of gas pressure waves received at two or more of the gas pressure detectors.

The method may comprise the acoustic elements being arranged to form a substantially spherical shape such that the sensitivity of the gas pressure detector is substantially the same in all directions around the gas pressure detector.

The method may comprise each acoustic element comprising an acoustic sensor and a housing; and the housing being a horn wave guide.

The method may comprise each gas pressure detector being located in a fire zone of the gas turbine engine.

The method may comprise: impacting the gas turbine engine; wherein the determination of an operating condition of at least part of the gas turbine engine is dependent on one or more gas pressure waves caused by the impact.

According to a second aspect, there is provided a system as defined in claim <NUM>.

In the second aspect, the gas turbine engine may further comprise: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

In the second aspect, the turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft; the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

As noted elsewhere herein, the present disclosure relates to a gas turbine engine.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or <NUM>% span position, to a tip at a <NUM>% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: <NUM>, <NUM>, <NUM><NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches) cm or <NUM> (around <NUM> inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than <NUM> rpm, for example less than <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> (for example <NUM> to <NUM>) may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades <NUM> on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the <NUM>-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s or <NUM> Nkg-<NUM>s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> deg C (ambient pressure <NUM>. 3kPa, temperature <NUM> deg C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, or <NUM> fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example on the order of Mach <NUM>, on the order of Mach <NUM> or in the range of from <NUM> to <NUM>. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach <NUM> or above Mach <NUM>.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM> (around <NUM> ft), for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> (around <NUM> ft) to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example on the order of <NUM>. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to: a forward Mach number of <NUM>; a pressure of <NUM> Pa; and a temperature of -<NUM> deg C.

As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example <NUM> or <NUM>) gas turbine engine may be mounted in order to provide propulsive thrust.

The planet carrier <NUM> constrains the planet gears <NUM> to process around the sun gear <NUM> in synchronicity whilst enabling each planet gear <NUM> to rotate about its own axis.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle that is separate to and radially outside the core engine nozzle <NUM>. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct <NUM> and the flow through the core <NUM> are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine <NUM> may not comprise a gearbox <NUM>.

A gas turbine engine <NUM> may comprise a large number of fluid carrying pipes/ducts. The fluids in the pipes can be at high temperatures and/or high pressures.

It is necessary to monitor the operating condition of a gas turbine engine <NUM> in order to determine if a change in operating condition, in particular a fault condition, has occurred. A fault condition may be, for example, a pipe failure such as a leak occurring or a pipe bursting. If a fault condition occurs then the fault condition should be quickly detected so that appropriate action can be taken.

It is known to detect pipe failures by monitoring the temperature within a gas turbine engine <NUM>. A large burst in a pipe can alter the temperature of the bulk of the gas turbine engine <NUM>. A fault condition can therefore be detected by detecting the temperature change of the bulk. For example, thermometers arranged in the fire zone of the gas turbine engine <NUM> can detect a fault condition whenever there is an unexpected change of the measured temperatures. However, this known technique is not able to detect a small burst in a pipe that may only cause local effects and does not significantly change the bulk temperature. In addition, a fault cannot be detected until the fault causes a temperature change and this can be a slow process.

Embodiments improve on known techniques by using the acoustics of a gas turbine engine <NUM> to determine its operating condition. The acoustics, i.e. sound, that the gas turbine engine <NUM> makes is measured and monitored. Any change of the sound can be used to determine that there has been a change in the operating condition of the gas turbine engine <NUM>.

Embodiments may also include techniques for recognising a type of fault condition that has occurred in dependence on the sound of the gas turbine engine <NUM>.

Embodiments may also include applying techniques for determining the location of a specific sound source. This may be used to determine the location of a fault, such as a burst pipe.

Embodiments are described below with reference to a microphone used to detect sound. However, the techniques of embodiments are not restricted to the sound being a human audible sound and the sound can more generally can be any type of gas pressure wave. The gas may be air and the detected waves air pressure waves. The gas pressure wave may be audible to a human or it may be, for example, ultrasonic, super-sonic or sub-sonic.

The microphone is also not restricted to being capable of only measuring/recording human audible sounds and may more generally be a gas pressure detector/transducer for detecting the gas pressure waves.

When a gas turbine engine <NUM> is operating, a failure of a pipe within the gas turbine engine <NUM> will have an acoustic effect. The acoustic effect may be the direct sound of the pipe wall rupturing and/or fluid passing out of, or into, the pipe through the hole in the pipe wall caused by the rupture.

The acoustic effects can also be used to determine the operating condition of other parts of the system and changes other than burst pipes. For example, the monitored acoustics may include any components of the power plant noise signature such as fan noise, combustor rumble and/or compressor acoustics. Any changes in the measured sound and/or differences to expected values of the measured sound can be used to detect a failure.

The techniques for measuring and monitoring sounds in a gas turbine engine <NUM> according to embodiments are not restricted to being performed when the gas turbine engine <NUM> is operating. For example, a sound may be induced in the gas turbine engine <NUM> by, for example, impacting a part of the gas turbine engine <NUM>. The condition of the gas turbine engine <NUM> may be determined in dependence on the sound generated in response to the impact.

<FIG> is a cross-section along the longitudinal axis of a gas turbine engine <NUM> according to an embodiment.

As shown in <FIG>, the gas turbine engine <NUM> comprises a bulk <NUM> that is surrounded by a bypass duct <NUM>. Within the bulk <NUM> is a fire zone <NUM> that surrounds a core engine. The core engine is also referred to herein as a core <NUM> or engine core <NUM>. There are one or more inlets <NUM> to the fire zone <NUM> and a plurality of outlets <NUM> from the fire zone <NUM>. Within the bulk <NUM> are pipes <NUM> and microphones <NUM>.

<FIG> is a cross-section of a gas turbine engine <NUM> according to an embodiment. The cross-section is orthogonal to the longitudinal axis of the gas turbine engine <NUM>.

As shown in <FIG>, there are splitters/pylons <NUM> between the bulk <NUM> and the outer surface of the gas turbine engine <NUM>. Two microphones <NUM> are positioned in the fire zone <NUM>. The microphones <NUM> are located on opposite sides of the core engine to each other.

One of the pipes <NUM> shown in <FIG> has burst and this has generated pressure waves <NUM> that propagate through the gas in the fire zone <NUM>.

Each microphone <NUM> measures the sound of its environment and outputs an electric signal that is generated in dependence on the measured sound. Each microphone <NUM> is in communication with a computing system. The output signal from each microphone <NUM> is transmitted to the computing system.

The computing system may record each signal received from a microphone <NUM>. The computing system can detect changes to the operating condition of a part of the gas turbine engine <NUM> and/or the entire gas turbine engine <NUM> by, for each microphone <NUM>, comparing the most recently received signal to previously received signals.

The computing system may analyse each signal that it receives from a microphone <NUM> by performing a Fourier transform on the signal. The Fourier transform may be, for example, a fast Fourier transform and/or a discrete Fourier transform. The Fourier transform generates a frequency domain representation of the signal. This can be used to determine if the sound comprises components within specific frequency ranges that may be an indication of an incorrect operating condition. For example, a burst pipe may cause super-sonic screech noise to be generated. The burst pipe can then be detected by the computing system whenever frequency components corresponding to super-sonic screech noise are present in a received signal from a microphone <NUM>.

The computing system may determine the operating condition of a part of the gas turbine engine <NUM> and/or the entire gas turbine engine <NUM> by comparing the received signal from each microphone <NUM> to predetermined values/waveforms of signals.

The computing system may store, or have access over a network to, a library of predetermined sound profiles with each sound profile corresponding to one of a plurality of types of fault condition. The fault conditions may include cracked, burst and/or leaking pipes as well as other events that may occur. The computing system may therefore be able to determine, from a comparison of a signal received from a microphone <NUM> and the sound profiles, the type of fault condition that has occurred. If there are predetermined sound profiles for different types of pipe, the computing system may be able to determine the type of pipe that has failed. For example, cabin air, anti-ice air and handling bleed pipes may all have different sound profiles when they burst and the type of pipe that has burst can therefore be automatically determined by the computing system. The sound profiles may be generated, for example, empirically or through modelling.

The comparison of a signal received by the computing system from a microphone <NUM> and sound profile may be performed in either the time domain or the frequency domain. If it is performed in the frequency domain then this will allow events that are characterised by the components of their frequency spectrum to be easily compared. Each signal received by the computing system from a microphone <NUM> may also have other process performed on it, such as filtering operations to prevent aliasing.

The computing system may be able to determine the location and/or direction of a sound source caused by an event occurring, such as a hole occurring in a pipe. If there is only one microphone <NUM>, the location of the sound source may be determined if the microphone <NUM> is directional and/or if the sound profile is dependent on the distance between the microphone <NUM> and the sound source.

When more than one microphone <NUM> is used, as shown in <FIG>, in addition to the above techniques for a single microphone <NUM> being used, the difference in phase, i.e. time of arrival, of the sound signal received by each microphone <NUM> can be used to determine the location of the sound source This can also assist the determination of the type of event that has occurred.

In the example shown in <FIG>, the burst pipe is a sound source. A first microphone <NUM> above the core engine is closer to the burst pipe than a second microphone <NUM> below the core engine. The first microphone <NUM> will therefore detect the sound from the sound source before the second microphone <NUM>. The time difference between when the first and second microphones <NUM> detect the same sound from the sound source can be used to determine the location of the sound source.

Embodiments are not restricted to the microphones <NUM> being provided in the locations, and with the relative orientations, shown in <FIG> and <FIG>. Embodiments include microphones <NUM> being provided in other locations and with different relative orientations. For example, in <FIG> there may alternatively be three or more microphones <NUM> in the fire zone <NUM> around the circumference of the core engine. The number of microphones <NUM> is also not restricted. For example, there may be nine or more microphones <NUM>. The microphones <NUM> may also be provided in other parts of the gas turbine engine <NUM> than the fire zone <NUM>.

Each microphone <NUM> comprises a plurality of acoustic elements <NUM>. As shown in <FIG>, each acoustic element <NUM> may comprise an acoustic sensor <NUM> and a housing <NUM> of the acoustic sensor <NUM>. The housing <NUM> both protects the acoustic sensor <NUM> and provides the acoustic sensor <NUM> with directional sensitivity. The housing <NUM> may be a horn waveguide. Each acoustic element <NUM> has a directional sensitivity to sound that is dependent on the orientation of the horn waveguide. The shape of the cross-section of the end of the horn waveguide may be a regular polygon. A plurality of acoustic elements <NUM> may be arranged together in a tessellated manner as shown in <FIG>. The shape formed by combining the acoustic elements <NUM> may be, for example, be a dodecahedron. The microphone <NUM> therefore has a substantially spherical shape. The microphone <NUM> comprises a plurality of acoustic elements <NUM> with orientations that allow the microphone <NUM> to detect sounds with substantially the same sensitivity in all directions.

Each acoustic element <NUM> generates and outputs an electric signal that is generated in dependence on the measured sound. The signal output from the microphone <NUM> that is transmitted to the computing system may comprise a plurality of signals, with each of the plurality of signals being an output signal from one of the acoustic elements <NUM> comprised by the microphone <NUM>. Alternatively, the plurality of electric signals from the acoustic elements <NUM> may be combined with each other at the microphone <NUM> to generate a single electric signal that is transmitted to the computing system.

Embodiments improve the determination of the operating condition of a gas turbine engine <NUM> over known techniques based on thermal detection. In particular, a one microphone <NUM> can detect sound changes over a large region that would require a plurality of thermal detectors, a change of condition can be detected instantly (there is no thermal lag), the location and/or direction of a sound source can be determined, microphones <NUM> are not expensive and the microphones <NUM> can detect changes in other parts of the overall system that contribute to the sound profile of the system.

In an alternative to the above-described techniques, one or more of the microphones <NUM> may be a single horn, or other shape, so that a directional signal is measured. Embodiments include detecting any type of fault event and well as the general engine health, such as engine degradation and deterioration.

Embodiments may also be integrated with an engine vibration monitoring system and/or thermal detection system.

<FIG> is a flowchart of a process for automatically determining an operating condition of at least part of a gas turbine engine <NUM> for an aircraft according to an embodiment.

In step <NUM>, one or more gas pressure waves are measured by a gas pressure detector <NUM>, wherein the gas pressure detector <NUM> is located in the gas turbine engine <NUM>.

In step <NUM>, a computing system automatically determines an operating condition of at least part of a gas turbine engine <NUM> in dependence on an output signal of the gas pressure detector <NUM>.

Embodiments are not restricted to all of the microphones <NUM> being located within the gas turbine engine <NUM>. Embodiments also include one or more microphones <NUM> being located outside of the gas turbine engine <NUM>. There may be microphones located both inside and outside of the gas turbine engine <NUM> or all of the microphones may be located outside of the gas turbine engine <NUM>.

Embodiments may be used for detecting burst ducts in an automatic thrust pull back system. Other applications that embodiments may be used for include determining if operations are being correctly performed during a pilot shutdown operation and determining if ventilation systems, cooling systems and/or pressure relief panels are operating correctly.

When incorrect operation is detected, embodiments include automatically generating messages to maintenance teams and dispatch notes for reporting the incorrect operation.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the scope of the invention as defined by the claims.

Claim 1:
A method of automatically determining an operating condition of at least part of a gas turbine engine (<NUM>) for an aircraft, the method comprising:
measuring one or more gas pressure waves by a gas pressure detector (<NUM>), wherein the gas pressure detector (<NUM>) is located in the gas turbine engine (<NUM>); and
automatically determining, by a computing system, an operating condition of at least part of a gas turbine engine (<NUM>) in dependence on an output signal of the gas pressure detector (<NUM>);
wherein
the gas pressure detector (<NUM>) comprises a plurality of acoustic elements (<NUM>); characterised in that;
each acoustic element (<NUM>) has directional sensitivity; and
all of the acoustic elements (<NUM>) have different orientations such that the gas pressure detector (<NUM>) is able to detect gas pressure waves with substantially the same sensitivity in all of the directions that an acoustic element (<NUM>) is orientated in.