GB2583357A

GB2583357A - Exhaust gas measurement

Info

Publication number: GB2583357A
Application number: GB1905753.8A
Authority: GB
Inventors: Di Chiaro Giacomo; Caslos Roman Casado Juan; Knapton Jonathan; Tentorio Luca; Zambon Filippo; Irimia Rasu
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-04-25
Filing date: 2019-04-25
Publication date: 2020-10-28
Also published as: GB201905753D0

Abstract

Apparatus for measuring a parameter, eg non-volatile PM, of exhaust gas from an engine, eg a gas turbine engine 10, or a combustor comprises an emission line 51 to transport exhaust gas from the engine to a detector or analyser 54, and at least one vibrator or ultrasonic device 1, 2 to apply vibrations or ultrasonic sound waves to the emission line 51 to promote shedding of carbonaceous material from an internal wall of the emission line. A discharge valve 53 may be operated to allow the exhaust gas to flow to the detector 54, eg via an emission rake 57, a splitter 58 and a nitrogen dilution box 59, or to an alternative discharge 55. The vibrator or ultrasonic devices 1, 2 may be electromechanical devices clamped to the emission line and may include a grounded vibrating rake 60 and small shakers 2. A blower may pump high pressure air through the emission line 51 while the vibrators 1,2 are active.

Description

EXHAUST GAS MEASUREMENT

The disclosure relates to an apparatus for measuring a parameter of exhaust gas, and a method for measuring a parameter of exhaust gas.

Background

An engine, such as a gas turbine engine, produces exhaust gas. Properties of the exhaust gas can be measured as a way of evaluating the engine.

For example, the exhaust gas from an engine comprises non-volatile particle matter (nvPM), which is a solid particulate. The concentration of nvPM in the exhaust gas can be measured so as to evaluate the engine. For example, the nvPM concentration may be required to be within certain bounds in order for the engine to be commercialised.

The nvPM concentration can be measured by transferring the exhaust gas from the engine to a measurement apparatus. However, measurements of the nvPM concentration and smoke of exhaust gas from the same engine lack repeatability.

It is an aim of the present disclosure to provide an apparatus and method for measuring a parameter of exhaust gas with increased repeatability.

Brief Summary

According to a first aspect of the disclosure there is provided an apparatus for measuring a parameter of exhaust gas from an engine or combustor, the apparatus comprising: an emission line configured to transport exhaust gas from the engine or combustor; and at least one vibrator or ultrasonic device configured to apply vibrations or ultrasonic sound waves to the emission line so as to promote shedding of carbonaceous material from an internal wall of the emission line.

In an arrangement, the at least one vibrator or ultrasonic device is attached to the exterior of the emission line. In an arrangement, the at least one vibrator is also attached to the ground and connected to the emission line.

In an arrangement, the apparatus comprises a discharge valve configured to selectively control whether the exhaust gas in the emission line flows to a detector for detecting the parameter or to an alternative discharge. In an arrangement, the apparatus comprises a controller configured to control the discharge valve based on at least an operational condition of the engine or combustor. In an arrangement, the controller is configured to control the at least one vibrator or ultrasonic device to apply vibrations or sound waves to the emission line when the exhaust gas in the emission line flows to the alternative discharge.

In an arrangement, the apparatus comprises a blower configured to selectively blow high pressure air through the emission line.

According to a second aspect of the disclosure there is provided a method for measuring a parameter of exhaust gas from an engine or combustor, the method comprising: transporting exhaust gas from the engine along an emission line; and apply vibrations or ultrasonic sound waves to the emission line so as to promote shedding of carbonaceous material from an internal wall of the emission line.

Optionally, the method comprises: flowing the exhaust gas in the emission line to a detector for detecting the parameter; and diverting the exhaust gas in the emission line such that instead of flowing to the detector, it flows to an alternative discharge, wherein the vibrations or ultrasonic sound waves are applied to the emission line when the exhaust gas is diverted to the alternative discharge. Optionally, the diverting step is performed at a timing based on an operational condition of the engine or combustor. After the vibrations or ultrasonic sound waves have been applied for a set time, the flow of the exhaust gas is reverted to the detector for detecting the parameter.

Optionally, the exhaust gas that enters the emission line is from a bypass duct of the engine. Optionally, the parameter is indicative of non-volatile particle matter in the exhaust gas.

Optionally, the engine is a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

Optionally, the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Brief Description of the Drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 schematically depicts an apparatus for measuring a parameter of exhaust gas; and Figure 5 is a flow chart showing steps in a method for measuring a parameter of exhaust gas.

Detailed Description

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example; the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second Zo turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

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 0% span position, to a tip at a 100% 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: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm Zo (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390 cm (around 155 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 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 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 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 320 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1600 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/Utip2, where dH is the enthalpy rise (for example the 1-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: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg-1K-1/(ms-1)2). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-'s, 105 Nkg-ls, 100 Nkg-'s, 95 Nkg-'s, 90 Nkg-ls, 85 Nkg-ls or Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea Zo level plus 15 deg C (ambient pressure 101.3kPa, temperature 30 deg C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. 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: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

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

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

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

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 16, 18, 20, or 22 fan blades.

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

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of -55 deg C. As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two Zo airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust.

The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3.

Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

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 Figure 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. 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 22 and the flow through the core 11 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 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

As explained above, the hot combustion products are finally exhausted through the core exhaust nozzle 20. The properties of the exhaust gas depend on the properties of the gas turbine engine 10. Properties of the exhaust gas can be measured so as to evaluate the engine 10.

One parameter that can be measured is the nvPM concentration. The invention is described primarily in relation to the measurement of nvPM concentration. However, additional or different parameters of the exhaust can be measured in accordance with the invention like soot emissions.

Figure 4 schematically depicts an apparatus 50 for measuring a parameter of exhaust gas from the engine 10. As shown in Figure 4, the apparatus comprises an emission line 51. The emission line 51 is configured to transport the exhaust gas from the engine 10. The emission line 51 defines a channel through which the exhaust gas flows away from the engine 10. The emission line 51 is for conveying the exhaust gas to an apparatus that can measure the parameter of the exhaust gas (e.g. the nvPM concentration). Hence, the emission line 51 extends at least part of the way between the engine 10 and a detector 54 that detects the parameter.

Optionally, the emission line 51 is at least 10 m, and optionally at least 20 m in length. This allows the emission line 51 to transport the exhaust gas well away from the engine 10 for measurement of the parameter. For example, the emission line 51 may be about 25 m in length. Optionally, the emission line 51 is at least partly flexible. As indicated in Figure 4, the emission line 51 may be attached at one or more fixed points, but otherwise hangs under gravity. The flexibility of the emission line 51 allows the path for the exhaust gas to flow through to be varied.

As shown in Figure 4, the apparatus 50 comprises at least one vibrator 1, 2 or ultrasonic device. Two different types of vibrator 1, 2 are shown in Figure 4. Figure 4 schematically shows a ground-fixed vibrator 1 which is attached to the ground 52. The ground-fixed vibrator 1 is also attached (at its other end) to the emission line 51. Meanwhile, the ungrounded vibrators 2 shown in Figure 4 are not fixed to the ground 52. Instead, they are only attached to the exterior of the emission line 51.

The size of the ungrounded vibrator 2 is not particularly limited. As an example, the ungrounded vibrator 2 may have a weight of about 3 kg and dimensions of the order of 50 to 200 mm.

The at least one vibrator 1, 2 or ultrasonic device is configured to apply vibrations or ultrasonic sound waves to the emission line 51 so as to promote shedding of carbonaceous material from an internal wall of the emission line 51.

In use (e.g. during an engine test), part of the exhaust flow of the engine 10 is collected and driven through the emission line 51. The exhaust gas is driven to the detector 54 (i.e. the gas analyser equipment) located downstream of the emission line 51. The nvPM and soot particles passing through the emission line 51 can deposit along the inner wall of the emission line 51 (also known as a sampling line). The vibrations applied by the vibrator 1, 2 or ultrasonic soundwaves applied by the ultrasonic device cause deposits of the carbonaceous material to leave the inner wall of the emission line 51. This reduces the amount of nvPM or soot deposited on the inner wall of the emission line 51. In turn, this reduces the effect of the carbonaceous material on the measurements on the nvPM and soot concentration. As a result, it is easier to predict the remaining effect of the carbonaceous material on the measurements. Additionally, the repeatability of the measurements is increased due to the increased predictability of the amount of deposition of the carbon agglomerate in the emission line 51.

The vibrator 1, 2 or ultrasonic device applies an external oscillating excitation to the emission line 51. Optionally, the external oscillating excitation is applied while air flow is flowing through the emission line 51. This causes the carbon particles attached to the inner wall to shed consistently. This reduces the scatter between different measurements for the same engine 10.

Even after the vibrations or ultrasonic soundwaves have been applied to the emission line 51, there may remain some carbonaceous material on the inner wall of the emission line 51. However, it is not necessary to completely clean the emission line 51. The inventors have found that the external oscillating excitations increase the consistency of the amount of deposited carbonaceous material for different measurements. This improves repeatability of the measurements.

The type of vibrator 1, 2 or ultrasonic device is not particularly limited. One possible device is an electromechanical vibration device 2 clamped to the emission line 51. As shown in Figure 4, the electromechanical vibration devices may each comprise a clamp 63 configured to clamp the electromechanical vibration device 2 to the exterior of the emission line 51. The vibration devices 2 are otherwise unconnected to the ground 52, such that they may be called ungrounded vibrators 2.

Additionally or alternatively, a ground mechanical vibration device 1 may be provided. The ground mechanical vibration device 1 may comprise a vibrating rake 60. The vibrating rake 60 is configured to apply vibrations to the emission line 51. The ground mechanical vibration device 1 is configured to grab or hold the emission line 51. The ground mechanical vibration device 1 may comprise a base 62. The base 62 is connected to the ground 52. The base 62 may house components of the ground mechanical vibration device 1 such as its power supply.

Although not shown in Figure 4, additionally or alternatively, an ultrasonic device can be provided. The type of ultrasonic device is not particularly limited. The ultrasonic device applies ultrasonic soundwaves to the emission line 51. Soundwaves are transmitted from the ultrasonic device to the emission line 51.

The amount of carbonaceous material that sticks to the inner wall of the emission line 51 depends on factors such as the diameter of the emission line 51, the number of bends formed in the emission line 51, the length of the emission line 51, the temperature of the exhaust gas, the flow rates of the exhaust gas and the pressure in the emission line 51. When exhaust gas is flowing through the emission line 51, some of the deposited carbonaceous material can shed from the inner wall of the Zo emission line 51. However, the timing and amount of shedding is unpredictable.

When deposits are randomly shed from the inner wall of the emission line 51 during measurement of the nvPM or soot concentration, peaks can be measured which are not representative of the true nvPM or soot concentration in the exhaust gas.

As shown in Figure 4, optionally the apparatus 50 comprises a discharge valve 53. The discharge valve 53 is configured to selectively control whether the exhaust gas in the emission line 51 flows to the detector 54 or to an alternative discharge 55. This allows the flow to switch between the detector 54 (when measurements are to be made) and the alternative discharge 55 (for when carbonaceous material is to be shaken away from the inner wall).

In use, exhaust gas is driven along the emission line 51 from the engine 10 to the detector 54. As shown in Figure 5, in step S1 the parameter (e.g. nvPM concentration) of the exhaust gas is measured when the engine 10 is in a known operational condition. The operational condition of the engine 10 can be controlled to reach a target operational condition, for example. This may relate to the power output of the engine 10.

As shown in Figure 5, in step S2 it may be determined that the engine 10 is again entering into the same operational condition. When the engine 10 enters for the second time in the operational condition (which may have been proven to be poorly repeatable in previous tests that were not in accordance with the invention), the exhaust gas coming from the emission line 51 is diverted from the detector 54 by means of the discharge valve 53. The conditions at which this occurs can be variable and not always the same. Hence, in step S3 the exhaust gases are diverted away from the measuring hardware.

In step S4 vibrational excitations are applied to the emission line 51. This may be done by controlling the vibrators 1, 2. For example, the small electromechanical vibration devices (i.e. shakers) 2 may be switched on so that they vibrate the emission line 51. The vibrating rake 60 of the grounded vibrator 1 may be switched on. During this time, air flow continues to flow through the emission line 51. Optionally, the mass flow through the emission line 51 remains substantially constant.

The vibrators 1, 2 are clamped to the line (or activated if they are already attached to the emission line 51). The frequency and/or amplitude of the vibrational excitations can be controlled in order to shed the carbonaceous deposits in the emission line 51.

The air flow through the emission line 51 transports the carbonaceous deposits through the emission line 51 to the alternative discharge 55. Having a mass flow passing through the line is beneficial as it enhances the shedding process. . In step S5, the vibrations are deactivated after a given amount of time. This may be done by detaching the vibrators 1, 2 or deactivating them. The discharge valve 53 is controlled so as to redirect the flow back to the emission hardware, i.e. the detector 54.

In step S6, measurements of the parameter of the exhaust gas are continued.

Optionally, the apparatus 50 comprises a controller configured to control the discharge valve 53 based on at least an operational condition of the engine 10 or combustor. Sensors are used to detect the operational condition of the engine 10. The sensors are connected to the controller so that the controller can determine the operational condition of the engine 10. The controller is connected to the discharge valve 53 so as to control it depending on the operational condition of the engine 10. Alternatively, the discharge valve 53 can be operated manually.

Optionally, the controller is configured to control the at least one vibrator 1, 2 and/or ultrasonic device to apply vibrations and/or soundwaves to the emission line 51 when the exhaust gas in the emission line 51 flows to the alternative discharge 55. This allows the carbonaceous material to be purged between measurements. This increases the repeatability of the measurements. Alternatively, the vibrations or soundwaves can be applied even when the parameter of the exhaust gas is being measured.

Optionally, the apparatus 50 comprises a blower. The blower is configured to selectively blow high pressure air through the emission line 51. For example, the blower can be used to blow air through the emission line 51 during purging of the carbonaceous material in the emission line 51.

Optionally, the blower is configured to pump high pressure air flow through the emission line 51 while the vibrators 1, 2 are in action. This means that the engine 10 can be kept in its lowest power condition while the purging is in process. This helps to save fuel.

As shown in Figure 4, optionally an emission rake 57 is provided at the output of the engine 10. The emission rake 57 is, for example, a fixed multi-arm and multi-point probe with rakes. The emission rake 57 is configured to collect a sample of exhaust gas averaged radially and circumferentially.

As shown in Figure 4, optionally the apparatus 50 comprises a splitter 58. The splitter 58 is configured to split the exhaust gas into multiple flows (e.g. smoke, gas) for measuring.

As shown in Figure 4, optionally the apparatus 50 comprises a dilution box 59. The dilution box is configured to dilute the exhaust gas, for example by particle-free dry nitrogen. The dilution is for suppressing the potential for water condensation, particle coagulation, gas-to-gas particle conversion and volatile particle formation within the exhaust gas.

In use, the sample collected by the emission rake 57 is fed to the splitter 58, which splits the sample into multiple flows. The exhaust gas to be measured is fed to the dilution box 59 where it is diluted in particle-free dry nitrogen. The dilution box 59 is connected to the emission line 51. The diluted exhaust gas is transported through the emission line 51 to the detector 54.

As shown in Figure 4, optionally a secondary line 56 is provided for connecting the discharge valve 53 to the alternative discharge 55. During a purging process, the exhaust gas (or other air mass flow) is transported to the alternative discharge 55 through the secondary line 56.

Optionally, the exhaust gas that enters the emission line 51 is from the core exhaust nozzle 20. Alternatively, the air coming out from the bypass duct 22 of the engine 10 can be used. The gas that enters the emission line 51 may be from the bypass duct 22 (e.g. from the bypass nozzle 18) of the engine 10. This reduces the possibility of additional layers of carbonaceous material building up inside the emission line 51 during the purging process. This is because the bypass flow does not take part in the combustion process. Hence, the input to the emission line 51 could be switched between the core exhaust nozzle 20 and the bypass nozzle 18 depending on whether a measurement is being performed or a purging process is being performed.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS1. An apparatus (50) for measuring a parameter of exhaust gas from an engine (10) or a combustor, the apparatus (50) comprising: an emission line (51) configured to transport exhaust gas from the engine (10) or combustor; and at least one vibrator (1, 2) and/or ultrasonic device configured to apply vibrations and/or ultrasonic sound waves to the emission line (51) so as to promote shedding of carbonaceous material from an internal wall of the emission line (51). 2. 3. 4. 5. 6. 7.
The apparatus (50) of Claim 1, wherein the at least one vibrator (1, 2) or ultrasonic device is attached to the exterior of the emission line (51).
The apparatus (50) of Claim 2, wherein the at least one vibrator (1) is also attached to the ground (52).
The apparatus (50) of any one of the preceding claims, comprising a discharge valve (53) configured to selectively control whether the exhaust gas in the emission line (51) flows to a detector (54) for detecting the parameter or to an alternative discharge (55).
The apparatus (50) of Claim 4, comprising a controller configured to control the discharge valve based on at least an operational condition of the engine ( 10) or combustor.
The apparatus (50) of Claim 5, wherein the controller is configured to control the at least one vibrator (1, 2) and/or ultrasonic device to apply vibrations and/or sound waves to the emission line (51) when the exhaust gas in the emission line (51) flows to the alternative discharge (55).
The apparatus (50) of any one of the preceding claims, comprising a blower configured to selectively blow high pressure air through the emission line (51).
8. A method for measuring a parameter of exhaust gas from an engine (10) or a combustor, the method comprising: transporting exhaust gas from the engine (10) or combustor along an emission line (51); and apply vibrations and/or ultrasonic sound waves to the emission line (51) so as to promote shedding of carbonaceous material from an internal wall of the emission line (51).
9. The method of Claim 8, comprising: flowing the exhaust gas in the emission line (51) to a detector (54) for detecting the parameter; and diverting the exhaust gas in the emission line (51) such that instead of flowing to the detector (54), it flows to an alternative discharge (55), wherein the vibrations or ultrasonic sound waves are applied to the emission line (51) when the exhaust gas is diverted to the alternative discharge (55).
10. The method of Claim 9, wherein the diverting step is performed at a timing based on an operational condition of the engine (10) or combustor.
11. The method of Claim 9 or Claim 10, wherein after the vibrations or ultrasonic sound waves have been applied for a set time, the flow of the exhaust gas is reverted to the detector (54) for detecting the parameter.
12. The method of any one of Claims 8 to 11, wherein the exhaust gas that enters the emission line (51) is from a bypass duct (22) of the engine (10).
13. The method of any one of Claims 8 to 12, wherein the parameter is indicative of non-volatile particle matter in the exhaust gas.
14. The method of any one of Claims 8 to 13, wherein the engine is a gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
15. The method of Claim 14, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); so the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.