Patent Description:
Aircraft mounted gas turbine engines must be capable of continued operation during flight encounters with ice or liquid water. Thus, ice crystal ingestion or ice accretion, i.e. the build-up of ice on surfaces, is a problem that may arise when aircraft are required to operate in conditions where the atmosphere has a high concentration of ice. The ingestion of rain, hail or heavy snow into the engine compression system may cause a number of adverse effects within the engine, such as combustor flameout, caused by the presence of liquid water within the engine's combustor. This may be especially true at low idle power setting where the low combustion temperatures and pressures can more easily result in the combustion process being extinguished. High pressure compressor surge may also be caused by an increase in the working line of the compressor from the evaporation of water within the compressor, and the additional work done by the compressor on the water. Additionally, thermal contraction of the compressor casings may lead to mechanical interaction between the compressor rotor blades and the casings, resulting in damage to or deterioration of compressor components, and a subsequent loss of compressor efficiency or stability.

The ingestion of ice crystals into the engine core may also cause a number of adverse effects within the engine. For example, compressor surge may be caused by the build-up of ice within or at the inlet to the compression system. The build-up of ice may introduce a pressure distortion which can act to reduce the nominal surge line of the compressor. Secondly, engine rollback may be caused by the build-up of ice within or at the inlet to the compression system. The build-up of ice may act as an aerodynamic blockage, restricting the mass flow rate of air through the engine and reducing power output. Thirdly, compressor damage may be caused by the impact of ice which has built up and then shed into the downstream stages of the compressor. This is normally associated with higher power operation where the rotational speed of the compressor(s) and thus impact energies are greater. Fourthly, compressor surge may be caused by the shedding of ice into the downstream stages of the compressor. Additionally, combustor flameout may be caused by sudden shedding of ice through the compressor and into the combustion chamber, particularly at low power conditions.

The atmospheric conditions that give rise to ice crystal ingestion or ice accretion are typically found at relatively low altitudes. However, certain atmospheric phenomena, such as thunderstorm activity, can result in ice crystals being present in the air at higher altitudes causing a risk of so-called 'ice crystal icing'. Thus, the ingestion of ice crystals into the engine could adversely affect the operation, or operating efficiency of the engine.

Common approaches to engine management during such weather conditions may be to proactively detect that high flow rates of rain or hail are being ingested into the engine compression system, and subsequently change either the engine operating condition or configuration when it is determined that ice crystal ingestion or ice accretion is possible. However, detecting these conditions with any certainty can be difficult. In particular, high altitude ice crystals cannot be detected with currently available ice detection systems, so current aircraft do not have means to detect ice crystal ice in the atmosphere. The crystals can be very small, and are often present in low concentrations above powerful storm systems but cannot currently be detected directly using weather radar. Due to their small size and low radar reflectivity, clouds of ice crystals are not currently detectable using conventional aircraft weather radar. They are not easily discernible by eye within the cockpit and not possible to identify by visual means during night time flying. Additionally, ice crystal clouds can be many hundreds of miles long and occupy altitudes that are consistent with the cruising altitude of commercial aircraft. Consequently it is very difficult for commercial aircraft to avoid ice crystal conditions altogether.

When determining that ice crystal ingestion or ice accretion is possible, changes to the engine operating condition may include increasing the engine power setting to increase shaft speeds and reduce the flow of rain or hail into the engine core, vaporising water before it reaches the combustor, as well as increasing compressor and combustor stability margins. Further changes to the engine operating configuration may also include the opening of compressor handling bleeds in order to extract ice or liquid water from the compression system. Altering the engine's operating point by, for example, raising idle thrust levels to increase temperatures in the compressors, increasing engine power setting in order to increase engine shaft rotation speeds or core temperatures, or by opening handling bleed valves to eject ice crystals or ice sheds from the core of the engine, may compromise the engine's operating efficiency, increasing fuel burn and operating costs. Thus, changes to the engine operating condition are preferably initiated after detecting environmental conditions likely to contain ice, but may also be altered pre-emptively to avoid the risk of ice crystal icing, even in atmospheric conditions that are not normally associated with ice crystal ingestion or ice accretion.

Crystal icing is a relatively rare phenomenon, and therefore these precautionary measures could largely be eliminated if a reliable means of detection were available. For example, as disclosed in <CIT>, a system is disclosed for detecting the presence of ice crystals in a cloud comprising two thin walled semicylinder-shaped sensors in a leading edge of an airfoil. Additionally, <CIT> discloses an apparatus for monitoring the build-up of ice on engine and plant parts encompassing a monitoring surface area and a reference surface area located on a part of such engine. There is provided a first means which heats the reference surface area to a temperature above the freezing point. A second means illuminates the reference surface area as well as the monitoring surface area. The light reflected from both these areas is led to a light receiver which emits signals which are applied to a fourth means. This fourth means determines the difference between the output signals of the light receiver and generates accordingly a control signal.

Furthermore, it is common to make compressor rotor blades more tolerant to ice impact by increasing their thickness. Such changes, however, represent a compromising of the engine's operating efficiency and consequently degradation of fuel burn and increase in operating costs. Since flight through high concentrations of ice crystals for long durations are relatively rare events, it would be preferable if changes to engine operating condition or configuration were invoked only when ice crystal ingestion is reliably detected. Thus, it is an aim to reduce or eliminate at least some of the aforementioned problems.

The present invention provides an apparatus, a vane, a strut, and a gas turbine engine, as set out in the appended claims.

According to a first aspect there is provided an apparatus for detecting water or ice crystal ingestion, or ice accretion on an aerofoil as defined in appended claim <NUM>.

Thus, relative to previously known arrangements, the invention may be increasingly sensitive to ice or liquid water ingestion within a gas turbine engine. Furthermore, when located within a compressor section, the arrangement is capable of directly detecting the presence of ice or liquid water within the compressor, without inference of compressor ice or liquid water flow rates from intake mounted probes and knowledge of other factors such as engine fan speed. Additionally, if embedded within an aerofoil or other static structure, the arrangement may provide the potential to reduce compressor aerodynamic loss and engine specific fuel consumption, whilst increasing the protection of sensors from matter ingested into the engine.

The arrangement also provides the freedom to install the detection system over a range of axial positions within the compression system, offering the ability to install the sensor system in a region of the compressor which is reliably free from super cooled liquid water ice accretion. Such a region may be downstream of, for example, the third stage of the compression system. The arrangement also provides the ability to install the sensor system in a region of the compressor which is reliably exposed to super cooled liquid water ice accretion, and in particular, in a region where the maximum temperature depressions at the outer wall are typically seen. Such regions may be established by the engine manufacturer based on analytical or experimental knowledge of compressor behaviour.

The apparatus has a first heater for applying heat to the first region of the aerofoil and a second heater for applying heat to the second region of the aerofoil.

Thus, the arrangement may provide two or more sensors positioned within a main gas path. Each of the two or more sensors may comprise a heater for heating the area adjacent to or around the each respective sensor. Each sensor may be either or both of at least partially embedded within, or integrated with, the turbomachinery casing or statics in order to protect the probe from erosion and impact related deterioration. The arrangement may therefore provide a means of detecting whether ice or liquid water is being ingested into a gas turbine engine compressor. Such detection may measure, either directly or indirectly, the change in radial (spanwise) temperature profile, within a given stage, or stages, of the compressor. The temperature profile may be caused by the sublimation, melting or evaporation of liquid water or ice together with the centrifugal action of the rotors on the ice or water.

At least one sensor may be positioned towards the outer annulus line. By towards the outer annulus line, it meant that at least one sensor is relatively closer to the outer annulus line than, or radially offset from, a further sensor. As such, at least one sensor may be situated in a radially outwardly region that is reliably subjected to an intercooling effect of one or more of rain, hail or ice crystals. Such intercooling occurs as ice or water moves through the compression system, as it adopts a bias towards the outer annulus line due to the centrifugal action of the compressor rotors on the ice crystals, water droplets or water film. Thus, thermal energy is given up by the air and turbomachinery surfaces as ice is melted, or water is evaporated within the compressor. This leads to a temperature depression within the outer radial section of the compressor. Such a region may be referred to as a 'wet' region. The sensor within the wet region may be referred to as the wet sensor.

At least one sensor may be positioned towards the inner annulus line. By towards the inner annulus line, it meant that at least one sensor is relatively closer to the inner annulus line than, or radially offset from, a further sensor. As such, at least one sensor may be situated in a radially inward region that is reliably free from an intercooling effect of one or more of rain, hail or ice crystals. Such a region may be referred to as a dry region. The sensor within the dry region may be referred to as the dry sensor. As the wet sensor is cooled by the ingested water or ice, a difference in power consumption between the wet sensor and the dry sensor may be compared and may be used for detection of ice formation, and the subsequent activation of engine protection systems via the EEC.

In the event of ice build-up on any part/region of a component, the heat energy applied may be largely absorbed in melting the ice rather than in raising the temperature of the component. The measured temperature will thus tend to be driven close to zero degrees Celsius. Comparing the measured temperature with a reference value, for example a temperature limit in the range of zero to five degrees Celsius such as one or two degrees Celsius, can provide an indication of ice accretion on a region of the component.

The comparison may be repeated, such that a rate of heating can be compared to a set of reference temperature values forming a heating profile. If the determined temperature is seen to be moving or increasing very slowly, or more slowly than expected, then there is an indication that it is being lagged or buffered by the presence ice.

The reliable detection of ice helps to avoid the need for unnecessary changes in engine operating point, thus increasing efficiency. In addition, compressor rotor blades and other engine components need not be made so robust if there is increased confidence that ice accretion can be avoided, so material and manufacturing costs can also be reduced.

Providing two or more radially temperature sensors, each paired with an electrical heating element, allows the apparatus to monitor respective sensor outputs relative to environmental conditions and/or ambient temperatures, and hence provides the ability to reliably detect ice formation, or ice crystal ingestion. For example, where the apparatus is located in a gas compressor stream, the two temperature readings will show a stable relationship with compressor gas stream temperature during normal operation. Continued monitoring of the temperature difference between the two sensors may provide an indication of when the temperature difference departs from the expected relationship. This information can then be interpreted, potentially with confirmation from other engine sensors, as indicating either of both of ice crystal ingestion, the accretion of ice on a component, or the potential for such ingestion or accretion to occur.

Again, repeated measurements and comparisons may be performed to allow a comparison of heating profiles.

The apparatus may be used for the detecting ice crystal ingestion or ice accretion on an aerofoil of a gas turbine engine. In some examples, the first region may, for example, comprise a platform of the aerofoil. The first region may comprise a trailing edge. The first region may comprise a leading edge. The first region may comprise a suction surface. The first region may comprise a pressure surface. In some examples, the second region may, for example, comprise a platform of the aerofoil. The second region may comprise a trailing edge. The second region may comprise a leading edge. The second region may comprise a suction surface. The second region may comprise a pressure surface.

Each of the first and second heater is located on or at least partially within the aerofoil. For example, the or each of the first and second heater may be mounted on the pressure surface, suction surface, leading edge or trailing edge of a vane or strut in a region of interest. For example, the or each heater may be located on a so-called 'intercase strut', i.e. a strut associated with a compressor intermediate casing that sits between an intermediate pressure compressor and a high pressure compressor.

The or each heater comprises an electrical heating element.

The electrical property may be resistance. The electrical property may be impedance.

The arrangement may comprise two or more sets of the respective heaters and sensors as previously described. The second or further set of respective heaters and sensors may be axially offset from the first set of respective heaters and sensors. Thus, the second or further set of respective heaters and sensors may be in communication with the comparator for comparing the temperatures determined by the first set of respective heaters and sensors and the second or further set of respective heaters and sensors. For example, a bleed offtake, or further such component or engine feature, may be situated between the first set and second set of respective heaters and sensors. The power consumption requirements from the sensors may be used to determine, how much water/ice is being extracted by the offtake and, for example, how quickly the intercase plenum needs to be evacuated. Alternatively, the introduction of a second or further set of respective heaters and sensors in more than one stage of compression could be used to better resolve the location within the compressor that ice crystal accretion is occurring. Thus, the or each set of sensors, or the overall sensor arrangement, may be located within a compressor section.

A vane or a strut for use in a gas turbine engine, such as an intercase strut as described above, may comprise apparatus as previously described, and a gas turbine engine may comprise such a vane or strut.

The apparatus has a controller configured to read computer readable instructions to execute steps to determine a first temperature value at the first radial position within the first region with the first temperature sensor; determine a second temperature value at the second radial position within the second region with the second temperature sensor; and, compare the first temperature value with the second temperature value; wherein ice crystal ingestion or ice accretion is detected based on the comparison of the first temperature value with the second temperature value, during use.

The controller may be configured to read computer readable instructions to control the first and/or second heater to apply heat constantly during use of the aerofoil.

The controller may be configured to read computer readable instructions to control the first and/or second heater to apply heat to heat a region to an equilibrium temperature, and the comparing step may be performed only once said region has reached said equilibrium temperature.

The apparatus may further comprise a memory, and the controller may be configured to read computer readable instructions to execute the additional steps to monitor the rate of temperature increase of the first region as the heat is applied to provide a measured heating profiles, store the measured heating profiles in the memory, and compare the measured heating profiles, in the comparing step, to reference temperature profiles.

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 <NUM> that is separate to and radially outside the core exhaust 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>.

Unless otherwise stated, the terms "axial" or "axially" refer to the principal and rotational axis <NUM>, describing a dimension along a longitudinal axis of the gas turbine engine <NUM>. The terms "aft" or "downstream", unless otherwise stated, refers to a direction towards either or both of the rear and outlet of the gas turbine engine <NUM> relative to the principal and rotational axis <NUM>. The terms "forward" or "upstream", unless otherwise stated, refers to a direction towards either or both of the front and inlet of the gas turbine engine <NUM> relative to the principal and rotational axis <NUM>, or refer to a component being relatively closer to the inlet of the gas turbine engine <NUM> as compared to a further component.

Unless otherwise stated, the terms "radial" or "radially" refer to a dimension extending between the principal and rotational axis <NUM> and an outwardly displaced circumference therefrom. The terms "proximal" or "proximally," unless otherwise stated, refers to a direction towards the principal and rotational axis <NUM>, or a component being relatively closer to the principal and rotational axis <NUM> as compared to a further component. The use of the terms "distal" or "distally," unless otherwise stated, refers to a direction towards the outwardly displaced circumference, or a component being relatively closer to the outwardly displaced circumference as compared to a further component. Directional references (i.e., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, aft, etc.) are to aid the reader's understanding of the arrangement and are, unless otherwise stated, not intended to limit the position, orientation, or use. Connection references (i.e., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members and relative movement between elements unless otherwise stated. Connection references are, unless otherwise stated, not intended to infer that two elements are directly connected to, or in fixed relation to each other. Furthermore, the exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings may vary.

<FIG> shows a plan perspective of an aerofoil <NUM> known within the art. The aerofoil <NUM> comprises a suction surface <NUM> and a pressure surface <NUM>. The suction surface <NUM>, pressure surface <NUM>, leading edge <NUM> and the trailing edge <NUM> extend between a root and a tip. The aerofoil <NUM> also comprises a camber line <NUM>, which represents the mean camber extending between the leading edge <NUM> and the trailing edge <NUM>. The external surface of the aerofoil <NUM>, being that of one or more of the suction surface <NUM>, pressure surface <NUM>, and leading edge <NUM>, is also shown to comprise a normal axis <NUM> extending in a direction perpendicular to a tangent <NUM> of the external surface. The normal direction <NUM> is shown, for example, to extend perpendicularly relative to the pressure surface <NUM>, but may extend perpendicularly relative to any such further external surface.

The aerofoil <NUM> comprises a leading edge <NUM>, and a trailing edge <NUM> which is aft, or downstream, of leading edge <NUM>. The leading edge is defined, in use, by a stagnation point, the stagnation point being a region of the aerofoil <NUM> when incident flow splits in order to flow over either the pressure surface or the suction surface of the aerofoil <NUM>. Thus, the leading edge <NUM> is the portion of the aerofoil which first meets the gas flow <NUM>. In some examples, the leading edge <NUM> may represent a region of the external surface which is substantially perpendicular to the gas flow direction <NUM>. In some examples, the leading edge <NUM> may represent a region of the external surface which is immediately adjacent to the area substantially perpendicular to the gas flow direction <NUM>. In some examples, the trailing edge <NUM> may represent a region of the external surface where the pressure surface <NUM> meets the suction surface <NUM>. In some examples, the trailing edge <NUM> may represent a region of the external surface, aft of the leading edge <NUM>, where the fluid flow separated by the leading edge <NUM> rejoins. The trailing edge <NUM> is aft of the leading edge <NUM>, so being spaced from the leading edge <NUM> in a chordwise direction <NUM>. The distance between the leading edge <NUM> and the trailing edge <NUM> may be expressed as a chord length which extends along a chord line <NUM> between the leading edge <NUM> and the trailing edge <NUM>. The chord length is the distance between the trailing edge <NUM> and the point on the leading edge <NUM> where the chord intersects the leading edge <NUM>. The distance between the leading edge <NUM> and the trailing edge <NUM> may be expressed as a length which extends in a chordwise direction <NUM>. The aerofoil <NUM> is shown to be at angled at a particular angle of attack <NUM>, which extends parallel to the gas flow direction <NUM>, and which may vary according to requirements. In some examples, the angle of attack may be measured relative to the chord line <NUM>. In further examples, the angle of attack may be measured relative to the chordwise direction <NUM>.

<FIG> shows shows a perspective view of a component comprising an ice crystal ingestion or ice accretion detection arrangement according to examples. The component shown is an aerofoil <NUM>. The aerofoil <NUM> forms part of a vane within a given stage of a compression system <NUM>,<NUM> shown in <FIG>. The vane may be a stator. In further examples, the aerofoil <NUM> may be a turbine or a compressor blade for use in a defined stage of the aforementioned subassemblies, or a further aerofoil for use in a gas turbine engine <NUM>. The component comprising the ice crystal ingestion or ice accretion detection arrangement may, in further examples, be formed with one or more parts of an engine casing, or any such further component comprising a gas-washed surface. In further examples, the aerofoil <NUM> comprises a wall comprising an external surface. The wall may be any one of a suction surface <NUM>, a pressure surface <NUM>, a leading edge <NUM> and a trailing edge <NUM> spaced from the leading edge <NUM>, in a chordwise direction <NUM> along a chord line 50a. The aerofoil <NUM> may comprise a cavity (not shown) defined by an internal surface. The cavity may be configured to receive a cooling fluid, in use. The suction surface <NUM>, pressure surface <NUM>, leading edge <NUM> and the trailing edge <NUM> extend between a root <NUM> and a tip <NUM>. The tip <NUM> is spaced from the root <NUM> in a spanwise direction <NUM>. The spanwise direction <NUM> extends from the root <NUM> towards the tip <NUM>. In some examples, the spanwise direction <NUM> extends perpendicularly from the root <NUM> towards the tip <NUM>.

A first sensor 60a and a second sensor 60b are shown to be formed within the body of the aerofoil <NUM>. It will be appreciated that, in further examples, the first sensor 60a and the second sensor 60b may be located on or at least partially within the platform <NUM>. In further examples, the first sensor 60a and the second sensor 60b may be located on or at least partially within an external surface of the aerofoil <NUM>. In further examples, the first sensor 60a and the second sensor 60b may be located on or at least partially within an internal surface of the aerofoil <NUM>. The location of the first sensor 60a and the second sensor 60b within the body of the aerofoil <NUM> may be preferred due to ease of manufacture and placement of the first sensor 60a and the second sensor 60b. Such location may also protect the first sensor 60a and the second sensor 60b from environmental or thermal damage, which may be expected from location upon an external surface of the aerofoil <NUM>. The two or more temperature sensors 60a,60b may include, for example, one or more of a thermocouple or a resistance temperature detectors (RTDs).

According to the example shown, both the first sensor60a and second sensor60b are located at a first axial location 51a, measured along the chord line 50a relative to the leading edge <NUM> of the aerofoil <NUM>. In some examples, either or both of the first sensor 60a and the second sensor 60b may be located in the suction surface <NUM>, leading edge <NUM>, or trailing edge <NUM> of the aerofoil <NUM> at a single first axial location 51a. In some examples, either or both of the first sensor 60a and the second sensor 60b may be located in the suction surface <NUM>, leading edge <NUM>, or trailing edge <NUM> of the aerofoil <NUM> at a second axial location (not shown). The second axial location may be measured along the chord line 50a relative to the leading edge <NUM>, and offset relative to the first location 51a, or vice versa. Thus, the first sensor 60a may be offset from the second sensor 60b in a chordwise direction.

The chord length is traditionally measured over a chord line 50a at a specific spanwise location, and can vary in length with radius. The chord length is defined as minimum distance from the leading edge to the trailing edge at a specified spanwise height. In <FIG>, the chord length is measured over a chord line 50a at the root <NUM> of the aerofoil. Additionally or alternatively, the chord length may be measured over a chord line 50b at the tip <NUM> of the aerofoil. In further examples, the aerofoil <NUM> may comprise a number of further chord lines at various distances in the spanwise direction <NUM> measured relative to the root <NUM>, on either or both of the suction surface <NUM> and the pressure surface <NUM>. Such chord lines may include a chord line measured at, for example, the median location of the aerofoil <NUM>. The chord length of the further chord lines may vary according to the degree of curvature and the shape of the respective aerofoil <NUM>. Thus, the chordwise location of the first and second sensors 60a,60b may be expressed as a percentage value of the chord length.

<FIG> shows a sectional side view of a compression section, including the aerofoil <NUM> previously shown in <FIG>, which in turn forms part of a compressor within the compression system <NUM>,<NUM> of the gas turbine engine <NUM>. The first sensor 60a is located at a first radial position 61a, measured along the spanwise direction <NUM> from either or both of the principal and rotational axis <NUM> and the base of the aerofoil <NUM> extending from the platform <NUM> towards the tip <NUM>. Thus, the first sensor 60a is configured to measure the temperature of a first region of the aerofoil <NUM>. The second sensor 60b is located at a second radial position 61b, measured along the spanwise direction <NUM> and extending from the root <NUM> towards the tip <NUM>. Thus, the second sensor60b is configured to measure the temperature of a second region of the aerofoil <NUM>. Thus, the first sensor 60a is offset from the second sensor 60b in a spanwise direction <NUM>. In this way, the first region is offset from the second region in a spanwise direction <NUM>. In some examples, the first region is distinct from the second region. In further examples, the first region may at least partially overlap with the second region. Thus, the first radial position 61a is radially offset from the second radial position 61b, relative to either or both of the principal and rotational axis <NUM>, and the base of the aerofoil <NUM> extending from the platform <NUM> towards the tip <NUM>.

Also shown is radial section A-A, located immediately upstream of aerofoil <NUM>. Adjacent the stator aerofoil <NUM> is shown an aerofoil 240a of an upstream rotor and an aerofoil 240b of a downstream rotor, relative to the direction of airflow <NUM>, parallel to a principal and rotational axis <NUM>. Also shown are the first sensor 60a and the second sensor 60b, which are formed within the body of the aerofoil <NUM>. Thus, in some examples, the first sensor 60a and the second sensor 60b are shown to be positioned aft of the leading edge <NUM> of the aerofoil <NUM>. <FIG> also shows a cross-sectional perspective of ingested ice or liquid water <NUM> within a gas flow <NUM>. As the ingested ice or liquid water <NUM> flows through the compression system <NUM>,<NUM>, it adopts a bias towards the radially outer region (outer span) of the aerofoil <NUM> due to the centrifugal action of the upstream and downstream rotors 240a,240b on the ingested ice or liquid water <NUM> present. In the region adjacent the aerofoil 240a of an upstream rotor, the ingested ice or liquid water <NUM> is shown to be distributed over a substantial portion of the aerofoil 240a before being centrifuged towards the outer annulus line due to passing through and being centrifugally accelerated by the rotating stage comprising the aerofoil 240a of the upstream rotor. The extent to which the ingested ice or liquid water <NUM> is centrifuged towards the outer annulus line depends of one or more of, for example, the rotational speed of the or each rotor 240a,240b, the geometry and solidity of the or each rotor 240a,240b, the type and size of water or ice droplets or particles contained within the ingested ice or liquid water <NUM>, and the incoming velocity of the ingested ice or liquid water <NUM> within the gas flow <NUM>. Thus, at the aerofoil <NUM> and the aerofoil 240b of a downstream rotor, the ingested ice or liquid water <NUM> is shown become concentrated towards the outer radial section of the compressor. Thus, the ingested ice or liquid water <NUM> may form a slurry, a water film, or may fully vaporise depending on operational conditions.

<FIG> shows a schematic of operational radial air temperature profiles along section A-A, previously shown in <FIG>. <FIG> shows a first radial air temperature profile <NUM> along section A-A associated with compressor operation in which no ice or liquid is being ingested. During normal operation, in dry air, the temperatures detected by the temperature sensors 60a,60b will show a stable relationship with compressor gas stream temperature, so will track compressor temperatures in a predictable manner. If ice crystal ingestion or ice accretion begins to occur on the aerofoil <NUM>, it will preferentially build up in the radially outwardly wet region of the second sensor 60b. The melting of the ice build-up will act as a buffer, absorbing the heat from the heating element <NUM>,<NUM> associated with that sensor 60a,60b and driving the temperature on that surface close to zero degrees Celsius. The difference in measured temperature between the two temperature sensors 60a,60b can then be interpreted, potentially with confirmation from other engine sensors, indicating ice crystal ingestion or ice accretion within the gas turbine engine <NUM>. To this extent, <FIG> also shows a first radial air temperature profile <NUM> along section A-A, caused by the sublimation, melting or evaporation of ice or liquid water towards the outer annulus line. In the case of the first radial air temperature profile <NUM>, as ice sublimes or melts, or water is evaporated, heat is supplied to overcome the latent heat of the phase change. This heat is provided by the surrounding air or surfaces within the gas turbine engine <NUM>. This leads to an air temperature depression towards the outer radial section of the compressor and a subsequent change to the first radial temperature profile <NUM>.

Referring once again to <FIG>, the first sensor 60a is positioned at the first radial position 61a of the aerofoil <NUM>, which is free from the temperature depression of the first radial air temperature profile <NUM> generated when ice or liquid water is ingested into the gas turbine engine <NUM>. The second sensor 60b is positioned at the second radial position 61b of the aerofoil <NUM>, which is reliably subjected to the temperature depression of the first radial air temperature profile <NUM> generated when ice or liquid water is ingested into the gas turbine engine <NUM>. Thus, according to examples, at least one second sensor 60b is positioned towards the outer annulus line, relative to one or more of the or each first sensor 60a, the principal and rotational axis <NUM> and the base of the aerofoil <NUM> extending from the platform <NUM> towards the tip <NUM>. In this way, at least one second sensor 60b is situated in a second region that is reliably subjected to the temperature depression of the first radial air temperature profile <NUM> generated when ice or liquid water is ingested into the compressor. Additionally, at least one first sensor 60a is positioned towards the inner annulus line, relative to one or more of the or each second sensor 60b, the principal and rotational axis <NUM> and the base of the aerofoil <NUM> extending from the platform <NUM> towards the tip <NUM>. In this way, at least one first sensor 60a is situated in a first region that is reliably free from the temperature depression of the first radial air temperature profile <NUM> generated when ice or liquid water is ingested into the compressor. Thus, the second radial position 61b of the second sensor 60b is radially offset from the first radial position 61a of the first sensor 60a. It will be appreciated that many such configurations may exist for the specific placement of the first sensor 60a and the second sensor 60b, providing the testable conditions outlined above are met.

<FIG> shows a plan perspective of an aerofoil <NUM>, and an alternate example of the arrangement shown in <FIG>. The aerofoil <NUM> comprises a pressure surface <NUM> and a suction surface <NUM>. The first temperature sensor 60a is provided between the pressure surface <NUM> and the suction surface <NUM>. A first electrical heating element <NUM> is located adjacent the pressure surface <NUM>. A second temperature sensor 60b is provided between the pressure surface <NUM> and the suction surface <NUM>. The second electrical heating element <NUM> is located adjacent the suction surface <NUM>. Each temperature sensor 60a,60b is heated by its adjacent electrical heater element <NUM>,<NUM>. In the example shown, the spacing between the distinct regions helps to avoid the heating of one region directly influencing the temperature of the other. It will however be appreciated that in further examples, the sensors 60a, 60b and regions may be on equivalent surfaces, or in equivalent chordwise locations of the aerofoil <NUM>. In further examples, the first temperature sensor 60a and the second temperature sensor 60b may be located off-centre, i.e. off the camber line <NUM>. Thus, either or both of the first temperature sensor 60a the second temperature sensor 60b may be located towards the pressure surface <NUM>, i.e. off the camber line <NUM>. In further examples, either or both of the first temperature sensor 60a the second temperature sensor 60b may be located towards the suction surface <NUM>, i.e. off the camber line <NUM>, according to requirements.

In the example shown, the first temperature sensor 60a and the second temperature sensor 60b are offset in the chordwise direction. It will also be appreciated that the electrical heater elements <NUM>,<NUM> are offset in the chordwise direction. Thus, the second axial location 51b may be measured along the chord line 50a relative to the leading edge <NUM>, and offset relative to the first location 51a, or vice versa. Although not shown, it will be appreciated that the first temperature sensor 60a and the second temperature sensor 60b are also offset in the spanwise direction. It will also be appreciated that the electrical heater elements <NUM>,<NUM> are offset in the spanwise direction. Thus, in some examples, the first and second heated regions may be both distinct, and offset in the spanwise direction such that the first temperature sensor 60a and the second temperature sensor 60b are radially offset.

The first and second heated regions refer to the areas of the aerofoil, or aerofoil <NUM>, heated by the respective heating elements <NUM>,<NUM>. The size of the respective heated regions may be enlarged or reduced according to, for example, the size, shape, location, power requirements or input of the respective heating elements <NUM>,<NUM>. The first electrical heater element <NUM> and the second electrical heater element <NUM> are also shown to be offset in the chordwise direction <NUM>. The first sensor 60a is configured to measure the temperature of a first region of the aerofoil <NUM>. The second sensor 60b is configured to measure the temperature of a second region of the aerofoil <NUM>. In some examples, the first region is distinct from the second region. In further examples, the first region may at least partially overlap with the second region. Thus, in some examples, the first temperature sensor 60a and the second temperature sensor 60b are axially offset.

The two temperature sensors 60a,60b provide temperature information in the regions of interest. Thus, the second sensor 60b is located, during use, in the wet region. Conversely, the first sensor 60a is situated in a radially inward region that is reliably not subjected to an intercooling effect of one or more of rain, hail or ice crystals. Thus, the second sensor 60b is located, during use, in the dry region.

Referring again to <FIG>, a controller <NUM> and a memory <NUM> are also shown. The controller may comprise a comparator to perform the comparison step between the temperature measured by the temperature sensors 60a,60b or a reference value, which can be stored in the memory <NUM>, or directly between the readings of the first and second temperature sensors 60a,60b. The memory <NUM> may also record measured temperature values over time to create a measured temperature profile for comparison with a reference profile.

Controller <NUM> can control various aspects of the apparatus, including one or more of heater operation, temperature measurement intervals, and data recording and comparison. In use, computer readable instructions are provided to the controller <NUM>, which may form part of a standard engine controller, for example a full authority digital engine/electronics control (FADEC), or may be provided as a stand-alone unit.

The ability to detect and respond to ice crystal ingestion or ice accretion will eliminate a fuel burn penalty currently caused by the need to defend engines in all conditions in which crystal icing could occur.

As ice or liquid water is ingested into the compressor the temperature measured by the second sensor 60b is appreciably lower than that by the first sensor 60a. The difference in measured temperature between the first sensor 60a and the second sensor 60b is compared against a threshold which may vary with compressor operating condition. If the difference in measured temperature exceeds the threshold, the ingestion of ice or liquid water ingestion may be confirmed. During use, when no ice or liquid water is being ingested, a small difference in the measured temperature between the first sensor 60a and second sensor 60b may exist due to the aerodynamic design of the compressor. If necessary, such differences may be accounted or corrected for, using knowledge of the compressor behaviour or operational characteristics. Thus, in a first example, a method of detecting ice crystal ingestion or ice accretion on an aerofoil <NUM>, during use, is illustrated in <FIG>. In a first step <NUM>, the temperature of the first region is monitored during use at <NUM>, and compared at step <NUM>. The value calculated at step <NUM> is compared with a predetermined reference, temperature value, temperature profile, or threshold at step <NUM>. At step <NUM>, if the measured difference exceeds the predetermined reference, temperature value, temperature profile, or threshold of step <NUM>, at step <NUM>, ingestion of ice or liquid water may be confirmed. At step <NUM>, if the measured difference does not exceed the predetermined reference, temperature value, temperature profile, or threshold of step <NUM>, ingestion of ice or liquid water remains unconfirmed.

At low air temperatures, such as in the front stages of a compression system during high altitude flight, the saturation vapour pressure of the air is low. Thus, the temperature depression generated by the ingestion of ice or liquid water into the compressor may be small and may be difficult to reliably detect given the measurement accuracy of the first sensor 60a and the second sensor 60b, which may be either or both of thermocouples or RTD sensors.

In further examples, either or both of the first sensor 60a and the second sensor 60b may be replaced with one or more electrically heated sensors which increase the observable temperature depression local to each respective sensor. In such examples, one or more heated sensors may be positioned on the surface of, or may be at least partially embedded within, a compressor stator vane, within a given stage of the compression system. The or each sensor may be heated using an electrical power supply to maintain a constant temperature. This temperature may vary according to gas turbine engine <NUM> operating condition. Thus, as the second sensor is cooled by the ingested ice or liquid water the electrical power required to keep it at its constant temperature increases. In this way, electrical power requirements of each sensor may be compared against each other, with the difference in electrical power required between each sensor being compared against a threshold which may vary with compressor operating condition. If the measured difference exceeds the threshold then the ingestion of ice or liquid water may be confirmed. Thus, in a second example, a method of detecting ice crystal ingestion or ice accretion on a aerofoil <NUM>, during use, is illustrated in <FIG>. In a first step <NUM>, heat is applied to a first region of a aerofoil <NUM> during use. The temperature of the first region is monitored during use at <NUM>, and compared with a second, reference, temperature value or temperature profile at step <NUM>. If the comparison <NUM> shows no difference, then the heating, monitoring and comparing steps <NUM>,<NUM>,<NUM> are repeated until interrupted by a user, or stopped based on a predetermined time or temperature threshold. If the comparison <NUM> shows a difference at step <NUM>, then an indication of ice crystal ingestion or ice accretion is provided at step <NUM>. Ice crystal ingestion or ice accretion may be detected based simply on a comparison, at step <NUM>, of the monitored temperatures of the first region <NUM> with a reference temperature close to zero degrees Celsius, for example in the range of zero to five degrees Celsius such as one or two degrees Celsius.

Alternatively, heat may also be applied to second region of the aerofoil <NUM>, during use, at step <NUM>, and the temperature of the second region monitored during use at <NUM> to provide the reference temperature value or update a reference temperature profile <NUM> for use in the comparison step <NUM>. Heat may be applied directly to the first region of the aerofoil <NUM> at step <NUM> and/or to the second region of the aerofoil <NUM> at step <NUM>, for example using an electrical heating element.

Heat may be applied <NUM>,<NUM> to the first and/or second region constantly during use of the aerofoil <NUM>, or may be applied <NUM>,<NUM> to heat the first and/or second region to an equilibrium temperature. In this case, the method includes the step of checking whether an equilibrium temperature has been reached, at step <NUM>, and the comparing step <NUM> is performed only once the first and/or second region has reached said equilibrium temperature.

As a further alternative, the rate of temperature increase of a region may be monitored as the heat is applied, by monitoring the temperature of the first region over time at step <NUM>. This provides a measured heating profile which is compared to a reference heating/temperature profile at step <NUM>.

Heat may be applied only once, periodically, at irregular intervals or on demand, for example when the gas turbine engine <NUM> operating point changes, or based on some other trigger. The heat may be applied for a predetermined amount of time or until a set temperature or gas turbine engine <NUM> operating point is reached, or the application of heat may be entirely controlled by a user.

<FIG> shows a schematic view of a further example of an aerofoil <NUM>. In <FIG>, the first temperature sensor 60a and first electrical heating element <NUM> are provided at a first radial position 61a on the aerofoil <NUM>. Furthermore, the second temperature sensor 60b and second electrical heater element <NUM> are provided at a second radial position 61b on the aerofoil <NUM>. Thus, the aerofoil <NUM> is an electrically conductive member.

The aerofoil <NUM> may be configured so that upon application of a current by the first sensor 60a or the second sensor 60b, an electrical property of the aerofoil <NUM> may be monitored to determine a temperature. The electrical property may be resistance. In further examples, either or both of the first sensor 60a and the second sensor 60b may be configured to determine the impedance of the aerofoil <NUM>. Thus, the electrical property may be impedance.

The temperature of the aerofoil <NUM> at either or both of the first region and the second region may monitored or determined in this way in an equivalent manner to that described in relation to <FIG>. It will also be appreciated that the arrangement may be equally applied to any further structure described herein.

According to a further example, a hot wire may be used, in place of either or both of the temperature sensors 60a,60b and a heating element, with a current being passed through the wire and its resistance being measured. It will be appreciated that a hot wire is a sensor. The hot wire sensor may be made from a length of resistance wire. Furthermore, the hot wire may be, for example, circular in section. Since resistance is proportional to temperature, this would also be effective in showing the presence or absence of ice accretion. Thus, a measured temperature difference between the first and second temperature sensors 60a,60b is indicative of ice crystal ingestion or ice accretion. Again, a controller <NUM> and memory <NUM>, as previously described, are provided in the systems of <FIG>.

In further examples, a 'dry' probe, such as that described in relation to <FIG> may be replaced with a synthesized/modelled value, such as the arrangements and methods described and claimed in <CIT>. Such a synthesized/modelled value may be generated, for example, by the engine controller, from one or more of a compressor shaft speed and a gas turbine engine <NUM> temperature. In yet further examples, the rate of change of power consumption of the wet sensor may be used to infer the occurrence of ice crystal ingestion or ice accretion. In this way, in further examples, the first sensor 60a may be replaced by a calculated (synthesised) value within the engine control system. This synthesised value may be developed based on the engine manufacturer's knowledge of the compressor behaviour. Such an arrangement may further reduce the cost and complexity of incorporating the first sensor 60a and may offer an advantage in increasing system reliability.

Claim 1:
An apparatus for detecting either or both of water or ice crystal ingestion and ice accretion on an aerofoil (<NUM>), within a gas turbine engine (<NUM>) having a principal rotational axis (<NUM>), the apparatus comprising:
a first temperature sensor (60a) for determining a first temperature value at a first radial position (61a) within a first region of the aerofoil, relative to the principal rotational axis of the gas turbine engine;
a second temperature sensor (60b) for determining a second temperature value at a second radial position (61b) within a second region of the aerofoil, relative to the principal rotational axis of the gas turbine engine; and,
a controller (<NUM>);
wherein the second radial position of the second temperature sensor is radially offset from the first radial position of the first temperature sensor; and,
either or both of the first temperature sensor and the second temperature sensor are positioned aft of a leading edge (<NUM>) of the aerofoil;
characterised in that the aerofoil (<NUM>) is electrically conductive, the aerofoil being configured so that upon application of a current, the resistance or the impedance of the aerofoil is monitored by the first temperature sensor (60a) and the second temperature sensor (60b), which are paired with a first heater and a second heater respectively, each of the first heater and the second heater comprising an electrical heating element (<NUM>, <NUM>), each of the first heater and the second being located on or at least partially within the aerofoil; and
the controller (<NUM>) has a comparator for comparing the first temperature value with the second temperature value and the controller is configured to read computer readable instructions to execute steps to:
determine the first temperature value (<NUM>) at the first radial position within the first region of the aerofoil, with the first temperature sensor (60a);
determine the second temperature value (<NUM>) at the second radial position within the second region of the aerofoil, with the second temperature sensor (60b); and,
compare the first temperature value with the second temperature value (<NUM>);
wherein ice crystal ingestion or ice accretion is detected (<NUM>) based on the comparison of the first temperature value with the second temperature value, during use.