Patent Description:
Performance of a gas turbine engine, whether measured in terms of efficiency or specific output, may be improved by increasing a turbine gas temperature (TGT), that is, a temperature of combustion gases at a turbine of the gas turbine engine. It may therefore be desirable to operate the turbine at a highest possible temperature.

However, gas turbine engines may experience TGT cross-check issues during operation. TGT cross-check issues refer to an imbalance between average TGTs across two circumferentially extending channels. If a temperature difference between the average TGTs across the two channels exceeds a cross-check threshold of an electronic engine controller (EEC), the EEC may automatically reduce thrust produced by the gas turbine engine. This may pose a safety concern during operation of the gas turbine engine. Further, due to such TGT cross-check issues, the gas turbine engine may require an expensive engine overhaul during a maintenance.

Therefore, there may be a need to reduce or prevent TGT cross-check issues in gas turbine engines in order to improve their on-wing performance and life expectancy.

United States patent <CIT> discloses a system and a method for detecting gas turbine engine hot section condition using temperature measurements during engine operation. The system comprises a sensing unit for sensing a temperature distribution across a hot combustion gas stream generated by a gas turbine engine combustor. A signal processor receives temperature signals from the sensing unit and generates a combustor malfunction signal when the difference between a maximal temperature and a minimal temperature of the sensed temperature distribution is greater than a predetermined acceptable delta value.

European patent <CIT> discloses a method for detecting one failure in a burner of a combustor of a turbine system. The method involves providing a plurality of temperature sensors arranged annularly at the outlet of the turbine, detecting a plurality of temperatures through the plurality of temperature sensors, calculating a temperature spread indicator as a function of the plurality of temperatures, and carrying out a comparison between the temperature spread indicator and a threshold. A positive result of this comparison indicates a burner failure.

European patent application <CIT> discloses a gas turbine that has a plurality of combustion chambers, at least one fuel nozzle for each of the combustion chambers, at least one fuel line for each fuel nozzle in each of the combustion chambers, at least one heat exchanger for each fuel line configured to adjust a temperature of a fuel flow to each fuel nozzle, and a controller that is configured to control each of the heat exchangers to reduce temperature variations amongst the combustion chambers.

In a first aspect, there is provided a method of claim <NUM>.

The method of the present invention may optimise the performance of the combustion equipment by repositioning the at least some of the plurality of fuel injectors. The method may reduce a circumferential variation of the plurality of temperatures of the combustion gases downstream of the combustion equipment. Therefore, the method may reduce a circumferential variation of a turbine gas temperature (TGT).

Consequently, the method may reduce or prevent TGT cross-check issues that typically arise due an imbalance between average TGTs across two circumferentially extending channels. The method may enable an electronic engine controller (EEC) to operate the gas turbine engine at optimal power parameters without undesirably reducing thrust output of the gas turbine engine, which may otherwise occur due to a presence of the TGT cross-check issues.

The method may improve an on-wing performance of the gas turbine engine and a life expectancy of the gas turbine engine. Furthermore, the method may prevent a requirement of an expensive engine overhaul during a maintenance of the gas turbine engine.

In some embodiments, the method further includes ranking the plurality of fuel injectors from a lowest temperature fuel injector to a highest temperature fuel injector based on predetermined flow test data. Repositioning the at least some of the plurality of fuel injectors is further based on the ranking of the plurality of fuel injectors.

In some embodiments, repositioning the at least some of the plurality of fuel injectors further includes disposing the lowest temperature fuel injector at the injector position corresponding to the hottest circumferential position. Repositioning the at least some of the plurality of fuel injectors further includes disposing the highest temperature fuel injector at the injector position corresponding to the coldest circumferential position.

In some embodiments, repositioning the at least some of the plurality of fuel injectors further includes disposing intermediate fuel injectors ranked between the lowest temperature fuel injector and the highest temperature fuel injector at respective injector positions corresponding to the circumferential positions ranked between the hottest circumferential position and the coldest circumferential position.

In some embodiments, the method further includes grouping the plurality of circumferential positions into a first channel and a second channel. Each of the first channel and the second channel circumferentially extends by <NUM> degrees with respect to the principal rotational axis. The method further includes determining a first average temperature of the temperatures determined by the temperature measurement devices corresponding to the first channel. The method further includes determining a second average temperature of the temperatures determined by the temperature measurement devices corresponding to the second channel. The method further includes determining a temperature difference between the first average temperature and the second average temperature. Repositioning the at least some of the plurality of fuel injectors is further based on the temperature difference.

Repositioning the at least some of the plurality of fuel injectors based on the temperature difference may reduce the temperature difference to less than or equal to a cross-check threshold of the EEC. As a result, the TGT cross-check issues may be obviated, and the EEC may operate the gas turbine engine at the optimal power parameters.

In some embodiments, after repositioning of the at least some of the plurality of fuel injectors, the temperature difference between the first average temperature and the second average temperature is less than or equal to <NUM> Kelvin. In some examples, the cross-check threshold may be <NUM> Kelvin.

In some embodiments, each injector position and the corresponding circumferential position are angularly offset with respect to each other by a predetermined angle. Such a relation between the plurality of injector positions and the plurality of circumferential positions may be referred to as a flow clocking relationship. The flow clocking relationship may be present due to a swirl in the combustion gases as they flow downstream of the combustion equipment.

In some embodiments, the predetermined angle is based on turning of a flow of the combustion gases by one or more vane assemblies disposed upstream of the plurality of measurement devices. The predetermined angle may depend upon a design of the gas turbine engine and may vary between different gas turbine engines.

In some embodiments, the plurality of temperature measurement devices is circumferentially disposed on a nozzle guide vane (NGV) assembly of a turbine of the gas turbine engine. In some examples, the NGV assembly may be of a low pressure turbine of the gas turbine engine.

In some embodiments, the method further includes communicably coupling the plurality of temperature measurement devices with a measurement harness. The measurement harness may enable recording or measuring individual readings from the plurality of temperature measurement devices.

In some embodiments, repositioning the at least some of the fuel injectors further includes detaching the at least some of the fuel injectors from the combustion equipment and attaching the at least some of the fuel injectors to the combustion equipment at corresponding injector positions of the second circumferential arrangement. The repositioning of the at least some of the fuel injectors may be performed manually by a maintenance engineer. For example, the maintenance engineer may manually detach the at least some of the fuel injectors and attach the at least some of the fuel injectors at the corresponding injector positions of the second circumferential arrangement.

In some embodiments, the fuel flow is provided to the combustion equipment during a maintenance of the gas turbine engine. Therefore, the method may be conveniently performed by the maintenance engineer during the maintenance of the gas turbine engine.

The gas turbine engine as described herein may have any suitable general architecture.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than <NUM>, for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, for example on the order of or at least <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from <NUM> or <NUM> to <NUM>. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described herein, a combustor may be provided axially downstream of the fan and compressor(s).

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), for example in the range of from <NUM> to <NUM>. 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>, <NUM>, <NUM>, <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, <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches) 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), for example in the range of from <NUM> to <NUM> or <NUM> to <NUM>.

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> or <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 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>, <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), for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>.

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>, <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), for example in the range of from <NUM> to <NUM>, <NUM> to <NUM>, or <NUM> to <NUM>. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. 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 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 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), for example in the range of from <NUM> to <NUM>.

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 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), for example in the range of from <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s, or <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described 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). Purely by way of example, a gas turbine as described herein may be capable of producing a maximum thrust in the range of from 330kN to <NUM> kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> degrees C (ambient pressure <NUM>. 3kPa, temperature <NUM> degrees 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), for example in the range of from <NUM> to <NUM>. 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 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 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 bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. 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 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 herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which <NUM>% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint - in terms of time and/or distance - between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e., maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO <NUM> at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

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 part of 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 (according to the International Standard Atmosphere, ISA) 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 an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (<NUM>). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of <NUM> ft (<NUM>).

In use, a gas turbine engine described 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.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The engine core <NUM> comprises, in axial flow series, a low pressure compressor <NUM>, a high pressure compressor <NUM>, combustion equipment <NUM>, a high pressure turbine <NUM>, a low pressure turbine <NUM>, and a core exhaust nozzle <NUM>.

The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines <NUM>, <NUM> before being exhausted through the core exhaust nozzle <NUM> to provide some propulsive thrust.

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>. In addition, the present invention is equally applicable to aero gas turbine engines, marine gas turbine engines, and land-based gas turbine engines.

The geometry of the gas turbine engine <NUM>, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the principal rotational axis <NUM>), a radial direction (in the bottom-to-top direction in <FIG>), and a circumferential direction (perpendicular to the page in the <FIG> view). The axial, radial, and circumferential directions are mutually perpendicular.

<FIG> illustrates a portion of the gas turbine engine <NUM> of <FIG>. As shown in <FIG>, the gas turbine engine <NUM> may include a pre-diffuser <NUM> disposed upstream of the combustion equipment <NUM>. The core airflow A (also shown in <FIG>) may pass into the pre-diffuser <NUM> before entering the combustion equipment <NUM>. The pre-diffuser <NUM> may slow down the core airflow A in order to promote efficient combustion and avoid large total pressure losses.

A fuel flow is provided into the combustion equipment <NUM> via a plurality of fuel injectors <NUM> (shown in <FIG>). The fuel is mixed with air (e.g., the core airflow A) and the mixture is combusted to produce combustion gases <NUM>.

The gas turbine engine <NUM> may further include an intermediate pressure turbine <NUM> disposed between the high pressure turbine <NUM> and the low pressure turbine <NUM> along the principal rotational axis <NUM>. The gas turbine engine <NUM> may further include a high pressure turbine (HPT) nozzle guide vane assembly <NUM> disposed upstream of the high pressure turbine <NUM> and downstream of the combustion equipment <NUM>, an intermediate pressure turbine (IPT) nozzle guide vane assembly <NUM> disposed upstream of the intermediate pressure turbine <NUM> and downstream of the high pressure turbine <NUM>, and a low pressure turbine (LPT) nozzle guide vane assembly <NUM> disposed upstream of the low pressure turbine <NUM> and downstream of the intermediate pressure turbine <NUM>.

The combustion gases <NUM> may flow through, and drive, the high pressure turbine <NUM>, the intermediate pressure turbine <NUM>, and the low pressure turbine <NUM>. A temperature of the combustion gases <NUM> measured at any one of the turbine stages may be referred to as a turbine gas temperature (TGT). In some cases, the TGT may be refer to the temperature of the combustion gases <NUM> at the low pressure turbine <NUM>. In such cases, the TGT may be measured using temperature measurement devices circumferentially disposed on the LPT nozzle guide vane assembly <NUM>. It may be noted that the TGT may vary circumferentially. The circumferential variation of the TGT may depend on specific engine design, manufacturing tolerances, and the like.

<FIG> illustrates a schematic front sectional view of the combustion equipment <NUM>.

The combustion equipment <NUM> includes the plurality of fuel injectors <NUM> circumferentially disposed about the principal rotational axis <NUM> at a plurality of injector positions <NUM> in a first circumferential arrangement <NUM>. The plurality of fuel injectors <NUM> and the plurality of injector positions <NUM> are depicted by respective circles in <FIG>. Specifically, each of the plurality of fuel injectors <NUM> is schematically depicted by a bigger circle. Further, each of the plurality of injector positions <NUM> is schematically depicted by a smaller circle. In <FIG>, the fuel injectors 50A-<NUM> are circumferentially disposed about the principal rotational axis <NUM> at respective injector positions 51A-<NUM>. This arrangement of the fuel injectors 50A-<NUM> may be referred to as the first circumferential arrangement <NUM>.

Each fuel injector <NUM> from the plurality of fuel injectors <NUM> may be fluidly coupled to a fuel line and may receive a fuel from a fuel tank (not shown) via the fuel line. The plurality of fuel injectors <NUM> may collectively discharge the fuel into the combustion equipment <NUM>. In the combustion equipment <NUM>, the fuel is mixed with air, and the mixture is combusted to produce the combustion gases <NUM> (shown in <FIG>).

As discussed above, the TGT may vary circumferentially. In other words, the temperature of the combustion gases <NUM> (shown in <FIG>) may vary circumferentially. Each fuel injector <NUM> may affect the TGT at a circumferential position of the turbine stage that corresponds to the respective injector position <NUM>. The effect on the TGT may be dependent on an air/fuel ratio of the respective fuel injector <NUM>. Due to manufacturing variability, design tolerances, and the like, the effect on the TGT may vary between different fuel injectors <NUM>.

Each of the plurality of fuel injectors <NUM> may undergo a flow test (e.g., by its manufacturer) to determine the effect on the TGT due to the respective fuel injector <NUM>. Therefore, the plurality of fuel injectors <NUM> may be ranked from a lowest temperature fuel injector <NUM> to a highest temperature fuel injector <NUM> based on predetermined flow test data. The lowest temperature fuel injector <NUM> may refer to one of the plurality of fuel injectors <NUM> that causes a lowest TGT rise. The highest temperature fuel injector <NUM> may refer to one of the plurality of fuel injectors <NUM> that causes a highest TGT rise.

<FIG> illustrates a schematic front sectional view of the LPT nozzle guide vane assembly <NUM> (also shown in <FIG>).

The gas turbine engine <NUM> further includes a plurality of temperature measurement devices <NUM> circumferentially disposed about the principal rotational axis <NUM> at a plurality of circumferential positions <NUM>. Each of the plurality of temperature measurement devices <NUM> is schematically illustrated by two adjacent dashed lines in <FIG>.

Further, each of the plurality of circumferential positions <NUM> is depicted by a square between the two adjacent dashed lines of the corresponding temperature measurement device <NUM>. Moreover, in <FIG>, the temperature measurement devices 60A-<NUM> are circumferentially disposed about the principal rotational axis <NUM> at the respective circumferential positions 61A-<NUM>. The temperature measurement devices <NUM> may include, for example, thermocouples.

The plurality of temperature measurement devices <NUM> may be circumferentially disposed on a nozzle guide vane (NGV) assembly of a turbine of the gas turbine engine <NUM>. Specifically, as shown in <FIG>, the plurality of temperature measurement devices <NUM> may be circumferentially disposed on the LPT nozzle guide vane assembly <NUM>. Alternatively, in some examples, the plurality of temperature measurement devices <NUM> may be circumferentially disposed on the HPT nozzle guide vane assembly <NUM> or the IPT nozzle guide vane assembly <NUM> (shown in <FIG>).

The plurality of circumferential positions <NUM> may be grouped into two channels. Specifically, the plurality of circumferential positions <NUM> may be grouped into a first channel <NUM> and a second channel <NUM>. For example, the circumferential positions <NUM> and 61A-61E may be grouped into the first channel <NUM>, and the circumferential positions 61F-<NUM> may be grouped into the second channel <NUM>. Each of the first channel <NUM> and the second channel <NUM> may circumferentially extend by <NUM> degrees with respect to the principal rotational axis <NUM>.

During operation of the gas turbine engine <NUM> (shown in <FIG>), an electronic engine controller (EEC) <NUM> (shown schematically by a block in <FIG>) may determine a first average temperature of temperatures determined by the plurality of temperature measurement devices <NUM> (i.e., 60A-60E, <NUM>) corresponding to the first channel <NUM>, and a second average temperature of temperatures determined by the plurality of temperature measurement devices <NUM> (i.e., 60F-<NUM>) corresponding to the second channel <NUM>. The EEC <NUM> may further determine a temperature difference between the first average temperature and the second average temperature. In some cases, the EEC <NUM> may reduce engine power parameters if the temperature difference exceeds a cross-check threshold of the EEC <NUM>, thereby reducing thrust produced by the gas turbine engine <NUM>. Therefore, it may be important that the temperature difference remains below the cross-check threshold. In some embodiments, the cross-check threshold may be <NUM> Kelvin. The EEC <NUM> may be communicably coupled with the plurality of temperature measurement devices <NUM> to determine the first average, the second average temperature, and the temperature difference between the first average temperature and the second average temperature. In <FIG>, the EEC <NUM> is shown to be communicably coupled to the temperature measurement devices 60B, 60C for clarity purposes only. It may be noted that the EEC <NUM> may be communicably coupled to each of the plurality of temperature measurement devices <NUM>. The EEC <NUM> may include one or more processors communicably coupled to a memory for performing various computations and control operations.

In order to individually obtain readings from the plurality of temperature measurement devices <NUM>, a measurement harness <NUM> may be used. Specifically, the measurement harness <NUM> may be communicably coupled with the plurality of temperature measurement devices <NUM> to obtain individual readings from the plurality of temperature measurement devices <NUM>. In <FIG>, the measurement harness <NUM> is schematically illustrated by a block and communicably coupled to the temperature measurement devices <NUM>, <NUM> for clarity purposes only. It may be noted that the measurement harness <NUM> may be communicably coupled to each of the plurality of temperature measurement devices <NUM>. The measurement harness <NUM> may be mounted outside of the gas turbine engine <NUM> (shown in <FIG>).

The plurality of circumferential positions <NUM> may be ranked based on the plurality of temperatures of the combustion gases <NUM> (shown in <FIG>) determined using the plurality of temperature measurement devices <NUM>. Specifically, the plurality of circumferential positions <NUM> may be ranked from a hottest circumferential position <NUM> to a coldest circumferential position <NUM>. The hottest circumferential position <NUM> may refer to a circumferential position at which the combustion gases <NUM> has a highest temperature. The coldest circumferential position <NUM> may refer to a circumferential position at which the combustion gases <NUM> has a lowest temperature.

The plurality of circumferential positions <NUM> corresponds to the plurality of injector positions <NUM> (shown in <FIG>), i.e. each injector position <NUM> may have a corresponding circumferential position <NUM>. The fuel injector <NUM> disposed at the injector position <NUM> may have a direct effect on the TGT at the corresponding circumferential position <NUM>.

<FIG> schematically illustrates a flow clocking relationship between the plurality of injector positions <NUM> and the corresponding plurality of circumferential positions <NUM>.

Each of the plurality of injector positions <NUM> is depicted by a circle, and each of the plurality of circumferential positions <NUM> is depicted by a square in <FIG>. Referring to <FIG> and <FIG>, the combustion gases <NUM> flowing from the combustion equipment <NUM> to the low pressure turbine <NUM> may swirl circumferentially as opposed to flowing straight therethrough, resulting in the flow clocking relationship.

As a result, in some embodiments, each injector position <NUM> and the corresponding circumferential position <NUM> may be angularly offset with respect to each other by a predetermined angle α. For example, as shown in <FIG>, the injector position 51A and the corresponding circumferential position 61A are angularly offset with respect to each other by the predetermined angle α. Similarly, the injector positions 51B-<NUM> and the corresponding circumferential positions 61B-<NUM> may be angularly offset with respect to each other by the predetermined angle α.

The predetermined angle α may depend on various factors, such as a design of the gas turbine engine <NUM> (shown in <FIG>). In some embodiments, the predetermined angle α is based on turning of a flow of the combustion gases <NUM> (shown in <FIG>) by one or more vane assemblies disposed upstream of the plurality of measurement devices <NUM>. Specifically, the one or more vane assemblies are disposed between the combustion equipment <NUM> and the plurality of measurement devices <NUM> with respect to the principal rotational axis <NUM>. For example, as shown in <FIG>, the one or more vane assemblies disposed upstream of the of plurality of measurement devices <NUM> may include the HPT nozzle guide vane assembly <NUM>, the IPT nozzle guide vane assembly <NUM>, and the LPT nozzle guide vane assembly <NUM>. Each of the HPT nozzle guide vane assembly <NUM>, the IPT nozzle guide vane assembly <NUM>, and the LPT nozzle guide vane assembly <NUM> may turn the combustion gases <NUM> by a respective angle.

The predetermined angle α may therefore depend on turbine design parameters, for example, a number of turbine stages of the gas turbine engine <NUM> and flow turning provided by each of the one or more vane assemblies disposed upstream of the plurality of measurement devices <NUM>. The predetermined angle α may typically range from <NUM> to <NUM> degrees. In some cases, the predetermined angle α may be from <NUM> degrees to <NUM> degrees. In some specific cases, the predetermined angle α may be <NUM> degrees.

<FIG> illustrates a method <NUM> of optimising the performance of combustion equipment of a gas turbine engine having a principal rotational axis in accordance with an embodiment of the present disclosure. For example, the method <NUM> may be used to optimise the performance of the combustion equipment <NUM> (shown in <FIG>) of the gas turbine engine <NUM>.

At step <NUM>, the method <NUM> includes providing a fuel flow into the combustion equipment via a plurality of fuel injectors circumferentially disposed about the principal rotational axis at a plurality of injector positions in a first circumferential arrangement. The fuel is mixed with air and the mixture is combusted to produce combustion gases. Referring to <FIG>, for example, the method <NUM> may include providing a fuel flow into the combustion equipment <NUM> via the plurality of fuel injectors <NUM> circumferentially disposed about the principal rotational axis <NUM> at the plurality of injector positions <NUM> in the first circumferential arrangement <NUM>.

At step <NUM>, the method <NUM> further includes determining a plurality of temperatures of the combustion gases at a plurality of circumferential positions downstream of the combustion equipment using a plurality of temperature measurement devices. The plurality of circumferential positions corresponds to the plurality of injector positions. Referring to <FIG> and <FIG>, for example, the method <NUM> may further include determining the plurality of temperatures of the combustion gases <NUM> at the plurality of circumferential positions <NUM> downstream of the combustion equipment <NUM> using the plurality of temperature measurement devices <NUM>. The plurality of circumferential positions <NUM> corresponds to the plurality of injector positions <NUM>. In some embodiments, the plurality of circumferential positions <NUM> downstream of the combustion equipment <NUM> may correspond to a plurality of circumferential positions at a nozzle guide vane assembly (e.g., the LPT nozzle guide vane assembly <NUM>).

At step <NUM>, the method <NUM> further includes ranking the plurality of circumferential positions based on the plurality of temperatures of the combustion gases determined using the plurality of temperature measurement devices. The plurality of circumferential positions is ranked from a hottest circumferential position to a coldest circumferential position. Referring to <FIG> and <FIG>, for example, the method <NUM> may further include ranking the plurality of circumferential positions <NUM> based on the plurality of temperatures of the combustion gases <NUM> determined using the plurality of temperature measurement devices <NUM>. The plurality of circumferential positions <NUM> is ranked from the hottest circumferential position to the coldest circumferential position.

At step <NUM>, the method <NUM> further includes repositioning at least some of the plurality of fuel injectors between the plurality of injector positions based at least on the ranking of the plurality of circumferential positions. After repositioning of the at least some of the plurality of fuel injectors, the plurality of fuel injectors is disposed at the plurality of injector positions in a second circumferential arrangement different from the first circumferential arrangement. Referring to <FIG>, for example, the method <NUM> may further include repositioning at least some of the plurality of fuel injectors <NUM> between the plurality of injector positions <NUM> based at least on the ranking of the plurality of circumferential positions <NUM>.

In some embodiments, the method <NUM> may further include ranking the plurality of fuel injectors from a lowest temperature fuel injector to a highest temperature fuel injector based on predetermined flow test data. Referring to <FIG>, for example, the method <NUM> may further include ranking the plurality of fuel injectors <NUM> from the lowest temperature fuel injector <NUM> to the highest temperature fuel injector <NUM> based on the predetermined flow test data. The predetermined flow test data may be available from a manufacturer of the plurality of fuel injectors <NUM> or suitable tests may be performed on the plurality of fuel injectors <NUM> to determine the flow test data.

Exemplary rankings of the plurality of plurality of fuel injectors <NUM> of <FIG> and the plurality of circumferential positions <NUM> of <FIG> are provided in Table <NUM> below for explanatory purposes.

For the purposes of explanation, it is assumed that each injector position 51A-<NUM> corresponds to the respective circumferential position 61A-<NUM>. As discussed above, each injector position 51A-<NUM> and the corresponding circumferential position 61A-<NUM> may be angularly offset with respect to each other by the predetermined angle α. Based on the rankings provided above in Table <NUM>, one example of how the plurality of fuel injectors <NUM> may be repositioned is provided below in Table <NUM>.

<FIG> illustrates a schematic front sectional view of the combustion equipment <NUM> with the plurality of fuel injectors <NUM> after repositioning thereof based on the method <NUM> of <FIG> in accordance with an embodiment of the present disclosure.

Referring to Table <NUM> and <FIG>, after repositioning of the at least some of the plurality of fuel injectors <NUM>, the plurality of fuel injectors <NUM> is disposed at the plurality of injector positions <NUM> in a second circumferential arrangement <NUM> different from the first circumferential arrangement <NUM> (shown in <FIG>).

In some embodiments, repositioning the at least some of the plurality of fuel injectors <NUM> may be further based on the ranking of the plurality of fuel injectors <NUM>. For example, the repositioning of the plurality of fuel injectors <NUM> as per the arrangement shown in Table <NUM> is based on the ranking of the plurality of fuel injectors <NUM>.

In some embodiments, repositioning the at least some of the plurality of fuel injectors further includes disposing the lowest temperature fuel injector at the injector position corresponding to the hottest circumferential position, and disposing the highest temperature fuel injector at the injector position corresponding to the coldest circumferential position.

Referring to Tables <NUM>, <NUM> and <FIG>, for example, repositioning the at least some of the plurality of fuel injectors <NUM> may further include disposing the lowest temperature fuel injector <NUM> (50D according to Table <NUM>) at the injector position <NUM> (<NUM> according to Table <NUM>) corresponding to the hottest circumferential position <NUM> (<NUM> according to Table <NUM>). Moreover, repositioning the at least some of the plurality of fuel injectors <NUM> may further include disposing the highest temperature fuel injector <NUM> (<NUM> according to Table <NUM>) at the injector position <NUM> (51J according to Table <NUM>) corresponding to the coldest circumferential position (61J according to Table <NUM>).

Referring to Tables <NUM>, <NUM> and <FIG>, for example, repositioning the at least some of the plurality of fuel injectors <NUM> may further include disposing intermediate fuel injectors <NUM> ranked between the lowest temperature fuel injector <NUM> (50D according to Table <NUM>) and the highest temperature fuel injector <NUM> (<NUM> according to Table <NUM>) at respective injector positions <NUM> corresponding to the circumferential positions <NUM> ranked between the hottest circumferential position <NUM> (<NUM> according to Table <NUM>) and the coldest circumferential position <NUM> (61J according to Table <NUM>).

In the illustrated example of <FIG>, the intermediate fuel injector 50B is already at the injector position 51B, and hence does not require repositioning. However, the intermediate fuel injector <NUM> has to be repositioned from the injector position <NUM> to the injector position 51C. Therefore, in some cases, only a subset of the plurality of fuel injectors <NUM> may need to be repositioned as per the method <NUM>, and one or more fuel injectors <NUM> may already be disposed at desired injector positions <NUM>.

In some embodiments, the method <NUM> further includes grouping the plurality of circumferential positions into a first channel and a second channel. Each of the first channel and the second channel circumferentially extends by <NUM> degrees with respect to the principal rotational axis. The method <NUM> further includes determining a first average temperature of the temperatures determined by the temperature measurement devices corresponding to the first channel. The method <NUM> further includes determining a second average temperature of the temperatures determined by the temperature measurement devices corresponding to the second channel. The method <NUM> further includes determining a temperature difference between the first average temperature and the second average temperature. Repositioning the at least some of the plurality of fuel injectors is further based on the temperature difference.

Referring to <FIG>, for example, the method <NUM> may include grouping the plurality of circumferential positions <NUM> into the first channel <NUM> and the second channel <NUM>. As discussed above, each of the first channel <NUM> and the second channel <NUM> may circumferentially extend by <NUM> degrees with respect to the principal rotational axis <NUM>. The method <NUM> may further include determining a first average temperature of the temperatures determined by the temperature measurement devices <NUM> corresponding to the first channel <NUM> (for example, the circumferential positions <NUM> and 61A-61E in <FIG>). The method <NUM> may further include determining a second average temperature of the temperatures determined by the temperature measurement devices <NUM> corresponding to the second channel <NUM> (for example, the circumferential positions 61F-<NUM> in <FIG>). The method <NUM> may further include determining a temperature difference between the first average temperature and the second average temperature. Further, repositioning the at least some of the plurality of fuel injectors <NUM> (shown in <FIG>) is further based on the temperature difference.

<FIG> illustrates a graph <NUM> depicting a temperature difference 210T between the first average temperature measured at the first channel <NUM> and the second average temperature measured at the second channel <NUM> before repositioning of the plurality of fuel injectors <NUM> (shown in <FIG>).

<FIG> illustrates a graph <NUM> depicting a temperature difference 260T between the first average temperature and the second average temperature after repositioning of the plurality of fuel injectors <NUM> (shown in <FIG>) based on the method <NUM> of <FIG>.

Referring to <FIG>, <FIG>, <FIG>, and <FIG>, the temperature difference 210T before repositioning of the plurality of fuel injectors <NUM> may be greater than the cross-check threshold. This may cause TGT cross-check issues. After repositioning of the plurality of fuel injectors <NUM> based on the method <NUM>, as shown in <FIG>, the temperature difference 260T may be less than the cross-check threshold. In some embodiments, after repositioning of the at least some of the plurality of fuel injectors <NUM>, the temperature difference 260T between the first average temperature and the second average temperature is less than or equal to <NUM> Kelvin. Therefore, repositioning the plurality of fuel injectors <NUM> based on the method <NUM> may reduce or prevent TGT cross-check issues.

Referring back to <FIG>, in some embodiments, the method <NUM> further includes communicably coupling the plurality of temperature measurement devices with a measurement harness. Referring to <FIG>, for example, the method <NUM> may further include communicably coupling the plurality of temperature measurement devices <NUM> with the measurement harness <NUM>.

The method <NUM> may be performed during a maintenance of the gas turbine engine. In some embodiments, the fuel flow may be provided to the combustion equipment during a maintenance of the gas turbine engine. Referring to <FIG>, for example, the fuel flow may be provided to the combustion equipment <NUM> during a maintenance of the gas turbine engine <NUM>.

In some embodiments, repositioning the at least some of the fuel injectors may further include detaching the at least some of the fuel injectors from the combustion equipment and attaching the at least some of the fuel injectors to the combustion equipment at corresponding injector positions of the second circumferential arrangement.

Referring to <FIG>, <FIG>, and <FIG>, for example, repositioning the at least some of the fuel injectors <NUM> may further include detaching the at least some of the fuel injectors <NUM> from the combustion equipment <NUM> and attaching the at least some of the fuel injectors <NUM> to the combustion equipment <NUM> at corresponding injector positions <NUM> of the second circumferential arrangement <NUM>. In some embodiments, repositioning of the at least some of the fuel injectors <NUM> may be manually carried out by a maintenance engineer.

In some cases, each of the at least some of the fuel injectors <NUM> may be conveniently detached from the combustion equipment <NUM> by removing/loosening one or more respective fasteners. Further, each of the at least some of the fuel injectors <NUM> may be conveniently attached to the corresponding injector positions <NUM> of the second circumferential arrangement <NUM> via the one or more respective fasteners.

Claim 1:
A method (<NUM>) of optimising the performance of combustion equipment (<NUM>) of a gas turbine engine (<NUM>) having a principal rotational axis (<NUM>), the method (<NUM>) comprising the steps of:
providing a fuel flow into the combustion equipment (<NUM>) via a plurality of fuel injectors (<NUM>) circumferentially disposed about the principal rotational axis (<NUM>) at a plurality of injector positions (<NUM>) in a first circumferential arrangement (<NUM>), wherein the fuel is mixed with air and the mixture is combusted to produce combustion gases (<NUM>);
determining a plurality of temperatures of the combustion gases (<NUM>) at a plurality of circumferential positions (<NUM>) downstream of the combustion equipment (<NUM>) using a plurality of temperature measurement devices (<NUM>), wherein the plurality of circumferential positions (<NUM>) corresponds to the plurality of injector positions (<NUM>); and
ranking the plurality of circumferential positions (<NUM>) based on the plurality of temperatures of the combustion gases (<NUM>) determined using the plurality of temperature measurement devices (<NUM>), wherein the plurality of circumferential positions (<NUM>) is ranked from a hottest circumferential position (<NUM>) to a coldest circumferential position (<NUM>); characterised by
repositioning at least some of the plurality of fuel injectors (<NUM>) between the plurality of injector positions (<NUM>) based at least on the ranking of the plurality of circumferential positions (<NUM>), wherein, after repositioning of the at least some of the plurality of fuel injectors (<NUM>), the plurality of fuel injectors (<NUM>) is disposed at the plurality of injector positions (<NUM>) in a second circumferential arrangement (<NUM>) different from the first circumferential arrangement (<NUM>).