Flow measurement for gas turbine engines

Abstract

A turbomachine having an airflow measurement system. The system includes a plurality of acoustic sensors located upstream of a compressor of the turbomachine. The acoustic sensors include at least one acoustic emitter configured to emit an acoustic waveform through an intake airflow to an acoustic receiver. The acoustic sensors are located in a single plane and mounted at angular intervals around a central axis of the intake.

Description

Technical Field

[0001] This disclosure relates to acoustic measurement systems, and more particularly to acoustic measurement systems for turbines such as axial-flow engines.

[0002] Description of related technologies

[0003] In turbines (such as gas turbine engines), the characteristics of the engine can be determined by measuring the properties of the air flowing into or through the engine. For example, the average velocity or mass flow rate of the airflow can be used to calculate the engine's performance and / or efficiency.

[0004] Measurements of total pressure and static pressure can be performed using conventional equipment such as a pitot tube. A pitot tube includes an orifice configured to face the incoming airflow and must protrude / extend into the airflow, for example, to avoid the formation of a boundary airflow layer on the inner surface of the housing / duct surrounding the airflow. Other equipment, such as static ports, must be installed within the housing.

[0005] The extent to which a Pitot tube extends into the airflow is limited in order to minimize disruption to the airflow; therefore, a Pitot tube can only sample the airflow characteristics in a limited area of ​​the airflow.

[0006] It is assumed that sampling measurements taken by the pitot tube within a limited region of the airflow are correlated with the overall (global) airflow flowing into / through the engine. However, this correlation must be modeled or estimated to calculate the overall airflow based on the sampled measurements. Furthermore, the sample in the limited region may not represent the overall airflow, for example, it may be affected by boundary airflow layers along the casing. Therefore, the correlation model may not be robust to varying operating conditions, and thus the calculation of the overall airflow may be inaccurate.

[0007] Another problem encountered when attempting to measure flow rates in axial-flow machines such as gas turbine engines is that the compressibility effect of the flow becomes relevant to high volumetric flow rates. This complicates the models that may be needed to correlate sensor readings with overall / gross flow rate values.

[0008] Conventional models rely on velocity distributions obtained during analysis or experimental calibration and do not account for unassessed conditions. Furthermore, mean pressure and temperature are used to determine mass flow rate from volumetric flow rate. However, the flow conditions inside an engine are highly dynamic.

[0009] This disclosure aims to overcome or improve one or more of the above-mentioned problems. Summary of the Invention

[0010] According to a first aspect, a turbine with an airflow measurement system is provided, the airflow measurement system comprising: a plurality of acoustic sensors, the plurality of acoustic sensors including at least one acoustic transmitter configured to transmit acoustic waveforms through the intake airflow of the flow machine to an acoustic receiver; and wherein the acoustic sensors are located in a single plane and mounted at angular intervals around the central axis of the intake airflow.

[0011] The engine may include a housing surrounding a central axis and defining an air intake, wherein acoustic sensors may be mounted on the housing.

[0012] Acoustic sensors can be configured as a circumferential array around the housing.

[0013] Acoustic sensors may not protrude into the airflow and / or be mounted behind a gas-washed surface at the air inlet.

[0014] Acoustic sensors can be located downstream of the intake throat. Acoustic sensors can be located upstream of the turbine rotor (such as a compressor / fan rotor). Acoustic sensors can be located upstream of any / all rotors of the turbine.

[0015] Acoustic sensors can be configured to generate tomographic data of airflow passing through the air inlet.

[0016] The turbine may include 2 to 40 acoustic sensors.

[0017] The turbine may include an internal component that extends axially within the airflow and intersects a single plane defined by a plurality of sensors, the internal component including additional acoustic sensors configured to communicate with the plurality of sensors.

[0018] The internal components are rotatable about an axial axis, and additional sensors are configured to rotate with the internal components.

[0019] Internal components may include a rotor hub and / or a nose cone.

[0020] The turbine may include an acoustic liner surrounding the rotor and configured to absorb vibrations generated by the rotation of the rotor, with acoustic sensors located upstream of the acoustic liner.

[0021] Acoustic sensors can be installed in the turbine nacelle and / or rotor housing.

[0022] The turbine may include internal components that extend axially within the airflow and intersect with the flow area. The internal components may include additional acoustic sensors configured to communicate with multiple sensors. The internal components may include / be a rotor, for example, arranged to rotate about a longitudinal / central axis of the inlet or flow area. The additional sensors may be configured to rotate with the internal components. The internal components may include a rotor hub and / or a nose cone.

[0023] The turbine may include a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; and a gearbox receiving input from the spindle and outputting drive to the fan to drive the fan at a lower rotational speed than the spindle.

[0024] Optionally, the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle; the engine core also includes a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

[0025] According to a second aspect, a system configured to determine an airflow through a turbine is provided, the system comprising: a plurality of acoustic sensors disposed upstream of the compressor of the turbine; the acoustic sensors including at least one acoustic transmitter configured to transmit an acoustic waveform through an intake airflow to an acoustic receiver; wherein the acoustic sensors are located in a single plane and mounted at angular intervals around a central axis of the intake; and a processing system configured to receive signals from the acoustic sensors and determine the flow rate of the intake airflow.

[0026] The processing system can be configured to determine the time of flight of an acoustic waveform between at least two acoustic sensors.

[0027] The processing system can be configured to determine the average velocity of the intake airflow.

[0028] The processing system can be configured to determine one or more of the following: mass flow rate; static temperature; or static pressure of the intake airflow.

[0029] Acoustic sensors can be configured to determine the flow profile across an airflow cross section.

[0030] Acoustic sensors can be configured to generate tomographic data of airflow passing through the air inlet.

[0031] According to a third aspect, a method for determining the airflow characteristics of an intake airflow of a turbine is provided, the method comprising the steps of: providing signal communication with a plurality of acoustic sensors located in a single plane, the single plane being disposed upstream of the compressor of the turbine; determining the time of flight of acoustic waveforms between the plurality of acoustic sensors; and using the time of flight between the plurality of sensors to determine the average velocity of the intake airflow.

[0032] This method may include using the time of flight between multiple sensors to determine the average mass flow rate of the intake airflow.

[0033] The time of flight between these multiple sensors can be used to determine the average flow velocity in the flow region. The time of flight between these multiple sensors can also be used to determine the average mass flow rate in the flow region. A flow profile above the flow region (e.g., across the flow cross-section) can be determined.

[0034] Acoustic sensors can be installed at or near the air intake of the flow machine.

[0035] Acoustic sensors can be mounted in the nacelle and / or rotor housing of a turbine. An acoustic liner can surround the rotor and is configured to absorb vibrations generated by the rotation of the rotor (e.g., a compressor / fan). The acoustic sensor can be located upstream of the acoustic liner.

[0036] As described elsewhere herein, this disclosure relates to gas turbine engines. Such gas turbine engines may include an engine core comprising a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such gas turbine engines may include a fan (with fan blades) located upstream of the engine core.

[0037] The arrangement disclosed herein can be particularly, but not exclusively, advantageous to a gearbox-driven fan. Thus, the gas turbine engine may include a gearbox that receives input from the spindle and outputs drive to the fan to drive the fan at a lower rotational speed than the spindle. The input to the gearbox may be directly from the spindle or indirectly from the spindle, for example via a spur shaft and / or gears. The spindle rigidly connects the turbine and compressor such that the turbine and compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0038] The gas turbine engine described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and compressor, such as one shaft, two shafts, or three shafts. By way of example only, the turbine connected to the mandrel may be a first turbine, the compressor connected to the mandrel may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0039] In such an arrangement, the second compressor may be axially positioned downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., directly, or via a generally annular duct).

[0040] The gearbox can be arranged to be driven by a spindle configured (e.g., in use) to rotate at a minimum rotational speed (e.g., the first spindle in the example above). For example, the gearbox can be arranged to be driven only by a spindle configured (e.g., in use) to rotate at a minimum rotational speed (e.g., in the example above, only the first spindle, not the second spindle). Alternatively, the gearbox can be arranged to be driven by any one or more shafts, such as the first shaft and / or the second shaft in the example above.

[0041] The gearbox can be a reduction gearbox (because the output to the fan rotates at a lower rate than the input from the spindle). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere in this document. The gearbox can 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 2.5, for example, in the range of 3 to 4.2, or 3.2 to 3.8, for example, approximately or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1, or 4.2. For example, the gear ratio can be between any two values ​​in the preceding sentence. By way of example only, the gearbox can be a "stellar" gearbox with a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0042] In any gas turbine engine as described and / or claimed herein, the burner may be axially positioned downstream of the fan and one or more compressors. For example, if a second compressor is provided, the burner may be located directly downstream of the second compressor (e.g., at its outlet). In another example, if a second turbine is provided, the flow at the burner outlet may be directed to the inlet of the second turbine. This burner may be positioned upstream of one or more turbines.

[0043] The compressor, or each compressor (e.g., the first and second compressors as described above), may include any number of stages, such as multiple stages. Each stage may include a row of rotor blades and a row of stator blades, which may be variable stator blades (because the angle of incidence of the row of stator blades may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other.

[0044] The turbine, or each turbine (e.g., the first and second turbines as described above), may include any number of stages, such as multiple stages. Each stage may include a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades may be axially offset from each other.

[0045] Each fan blade may be defined as having a radial span extending from the root (or hub) at a radially inner gas scrubbing position or a 0% span position to the 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 approximately) any of the following: 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 within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.32. These ratios may generally be referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip may be measured at the leading edge (or axially foremost) portion of the blade. Of course, the hub-to-tip ratio refers to the gas-washing portion of the fan blades, that is, the portion radially outside any platform.

[0046] The radius of the fan can be measured between the engine centerline and the tip of the leading edge of the fan blades. The fan diameter (which may be only twice the fan radius) can be greater than (or approximately) any of the following: 220cm, 230cm, 240cm, 250cm (approximately 100 inches), 260cm, 270cm (approximately 105 inches), 280cm (approximately 110 inches), 290cm (approximately 115 inches), 300cm (approximately 120 inches), 310cm, 320cm (approximately 125 inches), 330cm (approximately 130 inches), 340cm (approximately 135 inches), 350cm, 360cm (approximately 140 inches), 370cm (approximately 145 inches), 380cm (approximately 150 inches), 390cm (approximately 155 inches), 400cm, 410cm (approximately 160 inches), or 420cm (approximately 165 inches). The fan diameter can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 240cm to 280cm or 330cm to 380cm.

[0047] The fan speed can vary during use. Generally, for fans with larger diameters, the speed is lower. By way of non-limiting example only, the fan speed under cruising conditions may be less than 2500 rpm, for example, less than 2300 rpm. By way of another non-limiting example only, for engines with fan diameters in the range of 220cm to 300cm (e.g., 240cm to 280cm or 250cm to 270cm), the fan speed under cruising conditions may be in the range of 1700rpm to 2500rpm, for example, in the range of 1800rpm to 2300rpm, or for example, in the range of 1900rpm to 2100rpm. By way of another non-limiting example only, for engines with fan diameters in the range of 330cm to 380cm, the fan speed under cruising conditions may be in the range of 1200rpm to 2000rpm, for example, in the range of 1300rpm to 1800rpm, or for example, in the range of 1400rpm to 1800rpm.

[0048] When using a gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to move at a speed U. 尖端 Movement. The work done by the fan blades 13 in convection results in an enthalpy increase of dH in the flow. The fan tip load can be defined as dH / U 尖端 2 Where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端 This is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip load under cruise conditions can be greater than (or approximately) any of the following: 0.28, 0.29, 0.30, 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 are in J / kg). - 1 K -1 / (ms -1 ) 2 The fan tip load can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0049] The gas turbine engine according to this disclosure may have any desired bypass ratio, wherein 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 under cruise conditions. In some arrangements, the bypass ratio may be greater than (or approximately) 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, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 12 to 16, or 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be located radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or fan casing.

[0050] The total pressure ratio of the gas turbine engine described and / or claimed herein can be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the outlet of the highest-pressure compressor (before entering the combustor). By way of non-limiting example, the total pressure ratio of the gas turbine engine described and / or claimed herein may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 50 to 70.

[0051] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. Under cruise conditions, the specific thrust of the engine described and / or claimed herein may be less than (or approximately) any of the following: 110 Nkg -1 s, 105Nkg -1 s, 100Nkg -1 s, 95Nkg -1 s, 90Nkg -1 s, 85Nkg -1 s or 80Nkg -1 s. This specific thrust can be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, at 80 Nkg. -1 s to 100Nkg -1 s, or 85Nkg -1 s to 95Nkg -1 Within the range of s. Compared to traditional gas turbine engines, this type of engine may be particularly efficient.

[0052] The gas turbine engine described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, the gas turbine described and / or claimed herein may produce a maximum thrust of at least (or approximately) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit). By way of example only, the gas turbine described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, for example, 350 kN to 400 kN. The thrust mentioned above can be the maximum net thrust under standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), when the engine is stationary.

[0053] During operation, the temperature of the flow at the inlet of the high-pressure turbine may be particularly high. This temperature, referred to as TET, can be measured at the combustor outlet, for example, just upstream of the first turbine blade, which may be referred to as the nozzle guide vane. At cruising speed, this TET can be at least (or approximately) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K, or 1650K. The cruising TET can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit). The maximum TET of the engine during operation can be, for example, at least (or approximately) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 1800K to 1950K. Maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0054] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein may be made of any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoil may be made at least partially of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as carbon fiber. As another example, at least a portion of the fan blades and / or airfoil may be made at least partially of a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions made of different materials. For example, the fan blade may have a protective leading edge made of a material that is better resistant to impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge may be made, for example, of titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may have carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0055] The fan described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a retainer that engages with a corresponding slot in a hub (or disc). By way of example only, such a retainer may be in the form of a dovetail, which may be inserted into and / or engage with a corresponding slot in the hub / disc to secure the fan blade to the hub / disc. In another example, the fan blade may be integrally formed with the central portion. Such an arrangement may be referred to as a blade disc or blade ring. Such blade discs or blade rings may be manufactured using any suitable method. For example, at least a portion of the fan blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0056] The gas turbine engines described and / or claimed herein may or may not be equipped with variable area nozzles (VANs). Such variable area nozzles allow the outlet area of ​​the bypass duct to vary during use. The general principles of this disclosure can be applied to engines with or without VANs.

[0057] The fan of the gas turbine described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24 or 26 fan blades.

[0058] As used herein, cruise conditions may refer to the cruise conditions of an aircraft to which a gas turbine engine is attached. Such cruise conditions can generally be defined as conditions during the mid-cruise phase, such as the conditions experienced by the aircraft and / or engine at the midpoint between the peak of climb and the start of descent (in terms of time and / or distance).

[0059] By way of example only, the forward speed under cruise conditions can be any point in the range from Mach 0.7 to Mach 0.9, such as 0.75 to 0.85, 0.76 to 0.84, 0.77 to 0.83, 0.78 to 0.82, 0.79 to 0.81, approximately Mach 0.8, approximately Mach 0.85, or within the range of 0.8 to 0.85. Any single speed within these ranges can be considered cruise conditions. For some aircraft, cruise conditions may exceed these ranges, such as below Mach 0.7 or above Mach 0.9.

[0060] By way of example only, cruise conditions may correspond to standard atmospheric conditions at altitudes within the following ranges: 10,000 m to 15,000 m, for example, in the range of 10,000 m to 12,000 m, for example, in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example, in the range of 10,500 m to 11,500 m, for example, in the range of 10,600 m to 11,400 m, for example, in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example, in the range of 10,800 m to 11,200 m, for example, in the range of 10,900 m to 11,100 m, for example, approximately 11,000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0061] By way of example only, cruise conditions could correspond to: a forward Mach number of 0.8; a pressure of 23,000 Pa; and a temperature of -55°C. Also by way of example only, cruise conditions could correspond to: a forward Mach number of 0.85; a pressure of 24,000 Pa; and a temperature of -54°C (which could be standard atmospheric conditions at 35,000 feet).

[0062] As used anywhere in this document, “cruising” or “cruising conditions” may refer to an aerodynamic design point (ADP). Such an aerodynamic design point (or ADP) may correspond to the conditions under which a fan is designed to operate (including, for example, one or more of Mach number, environmental conditions, and thrust requirements). For example, this might refer to the conditions under which a fan (or gas turbine engine) is designed to have optimal efficiency.

[0063] In use, the gas turbine engines described and / or claimed herein may operate under cruise conditions defined elsewhere herein. Such cruise conditions can be determined by the cruise conditions of the aircraft (e.g., intermediate cruise conditions), on which at least one (e.g., two or four) gas turbine engines may be mounted to provide propulsive thrust.

[0064] Those skilled in the art will understand that, unless mutually exclusive, the features or parameters described with respect to any one of the foregoing aspects may be applied to any other aspect. Furthermore, unless 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. Attached Figure Description

[0065] The implementation scheme will now be described by way of example only, with reference to the accompanying drawings, in which:

[0066] Figure 1 This is a cross-sectional side view of a gas turbine engine;

[0067] Figure 2 This is a close-up cross-sectional side view of the upstream section of a gas turbine engine;

[0068] Figure 3 This is a partial cross-sectional view of a gearbox used in a gas turbine engine;

[0069] Figure 4 This is a schematic front view of the air intake of a gas turbine engine with an acoustic flow sensor.

[0070] Figure 5 This is a schematic cross-sectional view of the air intake of a gas turbine engine with an acoustic flow sensor.

[0071] Figure 6 An exemplary acoustic sensor array and the paths between them are shown;

[0072] Figure 7 Another exemplary acoustic sensor array for different flow paths is shown;

[0073] Figure 8 Another exemplary acoustic sensor array is shown;

[0074] Figure 9 Another exemplary acoustic sensor array adapted to relative rotation is shown;

[0075] Figure 10 This is a schematic cross-sectional view of another example of the air intake of a gas turbine engine with an acoustic flow sensor;

[0076] Figure 11 A schematic cross-sectional view through the acoustic sensor array is shown, indicating the parameters used in the example of flow measurement during use.

[0077] Figure 12 A flowchart for estimating one or more engine parameters is shown.

[0078] Figure 13 This is a cross-sectional side view of the bypass of a gas turbine engine. Detailed Implementation

[0079] The implementation scheme will now be described by way of example only, with reference to the accompanying drawings.

[0080] Figure 1 A gas turbine engine 10 with a main axis of rotation 9 is shown. The engine 10 includes an air intake 12 and a propulsion fan 23, which receives an intake airflow 48 and generates two airflows: a core airflow A and a bypass airflow B. The intake airflow 48 includes the sum of air flowing into the operating upstream end of the engine 10, wherein the sum of the core airflow A and the bypass airflow B is substantially equal to the intake airflow 48.

[0081] The gas turbine engine 10 includes a core 11 that receives core airflow A. The engine core 11 includes, in axial-flow series, a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 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. Bypass airflow B flows through the bypass duct 22. A fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and a rotary gearbox 30.

[0082] In operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion device 16, where it is mixed with fuel and the mixture is burned. The resulting thermal combustion products then expand through the high-pressure turbine and low-pressure turbines 17, 19 before being discharged through nozzle 20, thereby driving the high-pressure turbine and low-pressure turbine to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 typically provides most of the propulsive thrust. The rotary gearbox 30 is a reduction gearbox.

[0083] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown. Low-pressure turbine 19 (see [reference]). Figure 1 A drive shaft 26 is connected to the sun gear or sun gear 28 of the planetary gear arrangement 30. A plurality of planet gears 32 mesh with the sun gear 28 radially outward, and are connected together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously around the sun gear 28, while simultaneously causing each planet gear 32 to rotate about its own axis. The planet carrier 34 is connected to a fan 23 via a connecting rod 36 to drive the fan to rotate about the engine axis 9. A ring gear or ring gear 38 meshes with the planet gears 32 radially outward, and is connected to a fixed support structure 24 via a connecting rod 40.

[0084] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" as used herein may refer to the lowest-pressure turbine stage and the lowest-pressure compressor stage (i.e., excluding fan 23), and / or the turbine stage and compressor stage connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving fan 23). In some literature, the terms "low-pressure turbine" and "low-pressure compressor" mentioned herein may alternatively be referred to as "intermediate-pressure turbine" and "intermediate-pressure compressor." In the case of such alternative nomenclature, fan 23 may be referred to as the first or lowest-pressure compression stage.

[0085] exist Figure 3 The rotary gearbox 30 is shown in more detail by way of example. Each of the sun gear 28, planetary gear 32, and ring gear 38 includes teeth surrounding its periphery for meshing with other gears. However, for clarity, Figure 3 Only exemplary portions of the teeth are shown. Four planetary gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planetary gears 32 can be provided within the scope of the claimed invention. Practical applications of the planetary gearbox 30 typically include at least three planetary gears 32.

[0086] exist Figure 2 and Figure 3 The planetary gearbox 30 shown by way of example is a planetary type, in which the planet carrier 34 is connected to the output shaft via a connecting rod 36, and the ring gear 38 is fixed. However, any other suitable type of planetary gearbox 30 can be used. As another example, the planetary gearbox 30 can be a stellar arrangement, in which the planet carrier 34 remains fixed, allowing the ring gear (or gear ring) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As yet another alternative example, the gearbox 30 can be a differential gearbox, in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0087] It should be understood that Figure 2 and Figure 3 The arrangement shown is merely exemplary, and various alternatives are within the scope of this disclosure. By way of example only, any suitable arrangement can be used to position the gearbox 30 within the engine 10 and / or for connecting the gearbox 30 to the engine 10. As another example, the connection between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, output shaft, and mounting structure 24) (such as...) Figure 2The connecting rods 36 and 40 in the example can have any desired level of stiffness or flexibility. In another example manner, any suitable arrangement of bearings between the rotating and stationary parts of the engine (e.g., between the input and output shafts from the gearbox and a stationary structure such as the gearbox housing) can be used, and this disclosure is not limited to... Figure 2 The exemplary arrangement is as follows. For example, in the case where the gearbox 30 has a stellar arrangement (as described above), those skilled in the art will readily understand that the arrangement of the output link, support link, and bearing positions is generally different from that of the gearbox 30. Figure 2 The arrangement structure is shown as an example.

[0088] Therefore, this disclosure extends to gas turbine engines having any arrangement of gearbox type (e.g., star or planetary gear), support structure, input and output shaft arrangement, and bearing location.

[0089] Optionally, the gearbox may drive additional and / or alternative components (e.g., a medium-pressure compressor and / or a booster compressor).

[0090] Other gas turbine engines to which this disclosure is applicable 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. In another example, Figure 1 The gas turbine engine shown has split nozzles 18 and 20, meaning that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of this disclosure can also be applied 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 mixing nozzle. One or both nozzles (whether mixing or splitting) may have a fixed or variable area. While the described example relates to a turbofan engine, this disclosure is applicable to, for example, any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0091] The geometry of the gas turbine engine 10 and its components are defined by a conventional shaft system, including the axial direction (aligned with the axis of rotation 9) and the radial direction (in... Figure 1 The direction from bottom to top) and the circumferential direction (perpendicular to) Figure 1 (Page in the view). The axial, radial, and circumferential directions are perpendicular to each other.

[0092] This disclosure will now continue with respect to gas turbine engines; however, it should be understood that this disclosure can be applied to other types of axial flow engines / machines.

[0093] like Figure 4 and Figure 5 As shown, the gas turbine engine 10 includes a plurality of acoustic sensors 42. Preferably, the acoustic sensors 42 include ultrasonic sensors. The ultrasonic sensors 42 are disposed on a housing 44 that surrounds and defines the air intake 12 and the air intake flow 48 of the engine 10. The housing 44 may include the inner surface of the nacelle 21. The air intake should be considered any area upstream of the engine compressor (e.g., upstream of the fan 23).

[0094] In one example, two to forty sensors are provided. Typically, more than four sensors will be used for suitable coverage above the flow area. The sensors may be spaced apart about an axis of rotation, for example, as a circumferential array, and may be equidistant or unequally spaced. However, it should be understood that increasing the number of sensors can increase the fidelity and accuracy of the measurement. Therefore, the invention is not limited to such examples, and any number of sensors may be used as needed, depending on the application.

[0095] The ultrasonic sensor 42 includes an ultrasonic transmitter and / or an ultrasonic receiver. Each individual sensor 42 may include a transmitter, a receiver, or a transmitter / receiver pair.

[0096] The ultrasonic transmitter includes an ultrasonic transducer configured to emit ultrasonic waves 46 into an intake airflow 48.

[0097] An ultrasound receiver is configured to receive and detect ultrasound waves 46 emitted by an ultrasound transmitter. In one example, the receiver is located in substantially the same location as the transmitter and / or is integrally formed with the transmitter (i.e., they form part of the same component). In other examples, the receiver is located in a different location from the transmitter and / or is formed as a separate component from the transmitter.

[0098] Sensors 42 are removed from the intake airflow 48 (i.e., they do not protrude into the airflow 48, obstruct the airflow 48, or otherwise interfere with the airflow 48). Sensors 42 may be mounted on the outer surface of housing 44 or behind it (i.e., within the nacelle 21 housing) and / or may be mounted flush with the inner surface of housing 44 (e.g., such that the edges / sides of the sensor are flush with the gas-washed surface).

[0099] In other examples, sensor 42 may protrude into the airflow. This can intentionally generate turbulence, for example, to study airflow characteristics within an engine.

[0100] The ultrasonic sensors 42 are preferably spaced circumferentially around the housing 44 in a uniformly distributed manner. The ultrasonic sensors 42 may be spaced around the housing 44 such that each sensor 42 is radially opposite to another sensor 42. Alternatively, the sensors 42 may be non-uniformly distributed, for example, clustered. These sensors may be spaced at an angle of 90° around an axis.

[0101] The ultrasonic sensors 42 are located and / or oriented in a single plane (e.g., each individual sensor is located in an imaginary plane defined by the other sensors). This plane may be substantially flat. In other examples, the plane is arcuate, curved, etc. The exact form of the plane is irrelevant to this disclosure; however, it should be understood that providing sensors in a single plane means that these sensors are circumferentially spaced around the air inlet. Therefore, no two sensors within a given system are placed in the same circumferential position (i.e., no two sensors are spaced apart only axially without circumferential spacing).

[0102] In this example, the plane is substantially orthogonal to the net direction of the local airflow 48. In other embodiments, the entire plane and / or portions of the plane are not orthogonal to the net direction of the local airflow 48.

[0103] In one example, the plane is essentially orthogonal to the main engine axis 9. However, if desired, the plane can be deviated from the direction orthogonal to the airflow / axis, for example, by tilting.

[0104] In the example of a flat plane, one or more lines of sight between multiple corresponding sensors 42 can be oriented to lie within a single plane. All transmitters / receivers can be arranged to transmit / receive signals within this single plane.

[0105] The ultrasonic sensor 42 is located in the upstream portion of the engine 10, that is, upstream of the compressor stage of the engine. The sensor 42 is also located upstream of the fan 23. The sensor 42 may be located at, adjacent to, or immediately behind the air intake 12. In this example, the sensor 42 is located between the air intake throat 50 (i.e., the narrowest point of the air intake 12) and the fan housing 52 (i.e., the portion of the housing 44 surrounding the fan 23).

[0106] In one example, sensor 42 is offset in a downstream direction relative to the intake throat 50. In other examples, sensor 42 is located upstream of the intake throat. In any example, the upstream and / or downstream directions can be assumed to be along or parallel to axis 9.

[0107] Engine 10 may include an acoustic liner 54 surrounding housing 44 and configured to reduce acoustic vibrations therein. The acoustic liner is positioned adjacent to and / or upstream of the upstream side of fan housing 52. Sensor 42 may be located upstream of acoustic liner 54.

[0108] Sensor 42 may be located upstream or downstream of other sensing devices located on housing 44. These other sensing devices may include one or more of the following: pitot tube 56; or hydrostatic tube 58; or temperature probe.

[0109] The engine may include internal components located within housing 44, i.e., solid regions within the flow field. These internal components may extend axially through or within housing 44. For example, the internal components may include a portion of a rotor hub (for attaching fan blades 23 to the shaft); a rotator / nose cone 60; or a static portion / housing of the engine core 11.

[0110] In one example, sensor 42 is located upstream of the rotor hub and the upstream end / tip of the cone, such that the nose cone 60 does not intersect the plane of sensor 42. Figure 6 As shown, the positioning of sensor 42 prevents the line of sight 62 between multiple sensors 42 from being interrupted, such that each sensor 42 has a line of sight 62 to each of the other sensors 42, thereby allowing ultrasonic communication between sensors 42 along a substantially straight path.

[0111] Each line of sight 62 between the sensors 42 can be sampled to determine the average characteristics of the airflow 48 along the line of sight 62 (i.e., the ultrasonic sensors measure the average of the airflow characteristics between these sensors, rather than just at a single point near the sensor 42 itself). The characteristics of the airflow 48 can be sampled along multiple lines of sight 62 between each of the multiple sensors 42 to provide multiple samples across the airflow 48. The airflow characteristics can be sampled at multiple spatially separated points within the airflow, for example, including the boundary flow layer adjacent to the surface of the housing 44.

[0112] like Figure 6 As shown, multiple samples provide a sample "grid" across airflow 48, which spans many regions of the flow field / area.

[0113] In different examples, sensor 42 is located downstream of the tip of cone 60, such that cone 60 or rotor hub exists (i.e., through) the plane of sensor 42. For example... Figure 7 As shown, the line of sight 62 between the relative sensors 42 is interrupted in the area defined by the solid region of the rotating cone 60, thereby preventing ultrasonic communication between those relative sensors 42.

[0114] Due to the interruption of communication between the relative sensors 42, a dead zone 64 is formed (e.g., around the solid body of the cone 60), in which measurements of the characteristics of the airflow 48 cannot be performed. It should be understood that such a problem exists when the internal components include other parts of the engine 10 (e.g., the engine core 11). This dead zone 64 may be acceptable or unacceptable in different specific implementations. For example, if the number of sensors 42 mounted around the housing is increased, the dead zone area can be sufficiently reduced.

[0115] Figure 8 An alternative exemplary arrangement is shown when the solid internal component or body is located in the plane of the sensor. At least one additional acoustic sensor 66 is disposed on the internal component 60. This additional sensor may include an ultrasonic sensor 66 configured to operatively communicate with an ultrasonic sensor 42 disposed on the housing 44. In one example, one to six additional sensors 66 are provided. However, it should be understood that any number of additional sensors 66 may be provided depending on the application.

[0116] The additional sensors 66 can be evenly distributed around the circumference of the internal components. The additional sensors 66 can be removed from the airflow 48 in a similar manner to the sensors 42 on the housing 44.

[0117] like Figure 8 As shown, the line of sight 62 is maintained between sensor 42 on housing 44 and sensor 66 on internal component, thereby eliminating and / or reducing the size of dead zone 64. The line of sight between sensors 42 and 66 passes through the boundary layer flow at the surface of internal component 60, and thus can accommodate the contribution of the boundary layer to the overall flow profile.

[0118] In one example, an additional sensor 66 is attached to the cone 60 of the rotor hub, such that the additional sensor 66 rotates with the rotation of the hub while maintaining operative communication with the sensor 42 disposed on the housing 44.

[0119] like Figure 9 As shown, at time "t", the additional sensor 66 is operatively communicating with multiple sensors 42a, 42b, and 42c on the housing 44. The line of sight 62 between the sensors is shown as a thick dashed line. At time "t", the characteristics of the airflow 48 along the line of sight 62 can be sampled using sensors 42a, 42b, and 42c.

[0120] As the additional sensor 66 rotates with the hub, it moves to a new angular position at time "t+dt" (dt being an arbitrary time step). This additional sensor 66 communicates with the same plurality of different sensors 42a, 42b, 42c disposed on the housing 44; however, the line of sight 63 (shown as a thin dashed line) has moved to a new position. The same sensors can be used to sample the characteristics of the airflow 48 along the line of sight 63, but the portion of airflow 48 sampled at "t+dt" differs from the portion sampled at time "t".

[0121] As the additional sensor 66 continues to rotate, it communicates with a series of subsequent sensors 42 arranged sequentially around the housing (i.e., 42b, 42c, 42d, then 42c, 42d, 42e, etc.). As the additional sensor 66 rotates, a line of sight 63 sweeps across the airflow 48 within the housing 44. The characteristics of the airflow 48 are sampled throughout the rotation of the additional sensor 66, thus providing a “sweep scan” of the airflow 48 around the hub.

[0122] The cone 60 or hub may include multiple additional sensors 66, in a manner similar to the sensors 42 on the housing 44. Figure 8 The additional sensor 66 is shown. Multiple additional sensors 66 may be evenly distributed around the circumference of the hub 60. This sensor arrangement allows for simultaneous scanning of the entire flow area during each sweep / rotation.

[0123] In some examples, the static line of sight between the static sensors 42 can also be used for readings (e.g., as...). Figure 7 and Figure 8 (As shown). Therefore, the total readings may include, in part, readings obtained for a static reference frame and, in part, readings from a rotating reference frame.

[0124] like Figure 10 As shown, sensor 42 is operatively connected to power supply 74 and signal processing system 72. Static sensor 42 may have a direct electrical (e.g., wired) connection to signal processing unit 72.

[0125] Engine 10 includes an additional sensor system 76, such as a telemetry system, operatively connected to power supply 78 and signal processing system 72. This additional sensor system may be located upstream of sensor 42, for example, in the nacelle and / or engine air intake. The additional sensor system 76 is configured to wirelessly receive data measured by rotation sensor 66. This additional sensor system then forwards the data to signal processing box 72.

[0126] A rotary electrical coupler 68 may be provided for the additional sensor 66. The rotary electrical coupler 68 connects the additional sensor 66 to the power supply 70 and / or the signal processing system 72, and allows the additional sensor 66 to rotate relative to the power supply 70 and / or the signal processing system 72 while maintaining the connection between them. The rotary electrical coupler 68 can provide a physical connection (e.g., a wire). In other examples, the rotary electrical coupler 68 includes wireless transmission (e.g., wireless power or signal transmission).

[0127] In some implementations, an additional sensor 66 may be directly and operably connected to the signal processing system 72, thereby reducing the need for an additional sensor system 76.

[0128] In one example, engine 10 includes a second set of sensors. The second set of sensors may be located on engine 10 at a different axial position than the first set of sensors.

[0129] In some examples, the second plurality of sensors may be located in the downstream portion of engine 10, preferably at the outlet nozzle of engine 10. The second plurality of sensors may be axially spaced from each other and may not be circumferentially spaced from each other. The second plurality of sensors can be implemented in substantially the same manner as described in EP 3255438 A1, which is incorporated herein by reference.

[0130] The processing system 72 is configured to receive signals from one or more of the following: a first plurality of sensors 42; an additional sensor 66; or a second plurality of sensors. The processing system 72 includes one or more computer processors configured to process the signals to calculate the airflow velocity profile, volumetric flow rate, and / or mass flow rate of the inlet airflow 48.

[0131] The processing system 72 can be configured to provide signals to the ultrasonic sensor to start / stop ultrasonic transmission and / or reception.

[0132] The processing system 72 is operatively in communication with an additional sensor system 76. The additional system 76 can provide values ​​of one or more operating parameters (i.e., values ​​of one or more variable operating parameters) required to calculate the volumetric flow rate and / or mass flow rate of the inlet gas flow 48.

[0133] The processing system 72 can be configured to record airflow velocity, volumetric flow rate, and / or mass flow rate data over a given time period. The processing system 72 can analyze the data based on engine 10 or specific parameters of engine use 10 (e.g., specific power or thrust output or throttle valve settings of engine 10) to provide trends or patterns therein (e.g., using regression analysis).

[0134] The processing system 72 may have an output interface configured to transmit data related to any processing inputs or outputs described herein to another system, such as a monitoring and / or control system for the engine or its sub-components. This other system may be onboard to the engine or aircraft, for example, connected via a data bus or local wired or wireless network or a remote monitoring facility. The output of the processing system 72 may be used for: providing feedback to a user, such as a user interface in the aircraft cockpit; serving as input to an operational control system; and / or as input to an equipment health monitoring system.

[0135] Alternatively or in addition, the processing system 72 includes onboard non-volatile memory for data storage.

[0136] In some examples, additional conventional measuring equipment may be provided to determine airflow characteristics in the engine. This conventional equipment can be used in conjunction with the system to detect / measure any differences between the two measuring techniques.

[0137] Calculation of intake air mass flow rate

[0138] The following mathematical formula estimates the velocity and / or volumetric flow rate of the intake airflow 48 of engine 10 with a known stagnation temperature. The mass flow rate can be estimated by further knowledge of the stagnation pressure.

[0139] name:

[0140] U: The speed of the acoustic signal along the line of sight between the transmitter and receiver

[0141] V: Flow velocity

[0142] m: mass-flow rate

[0143] M: Mach number of the flow rate

[0144] T: Flow temperature

[0145] h: enthalpy

[0146] C: Correction factor

[0147] α: speed of sound

[0148] β: Angle

[0149] s: distance

[0150] γ: Adiabatic index

[0151] A = Area

[0152] ρ = density

[0153] p = pressure

[0154] R: Molar gas constant per mole of air mass

[0155] () t Overall or hysteresis characteristics, such as pressure and / or temperature.

[0156] () s Static characteristics, such as pressure and / or temperature

[0157] () TOF Average flight time

[0158] () m Average mass

[0159] () eng Overall engine parameters

[0160] () cr Core engine parameters

[0161] () aux auxiliary

[0162] () thm Thermodynamic average

[0163] Figure 11 A schematic diagram shows a plurality of sensors 42 configured to measure the velocity flow rate of the inlet airflow 48 within the housing 44.

[0164] The ultrasonic transmitter 80 emits an ultrasonic wave 46 into the airflow 48. The ultrasonic wave 46 interacts with the airflow 48, and the speed at which the wave travels through the airflow 48 varies depending on various physical properties of the airflow 48, as will be described below.

[0165] The ultrasonic receiver 82 is located within the line of sight 62 of the transmitter. The ultrasonic receiver 82 receives the ultrasonic wave pattern 46, and the processing system 72 calculates the time between the emitted ultrasonic wave pattern 48 and the received waveform to provide the measured time of flight (t). TOF ).

[0166] Given the line-of-sight distance D between the ultrasonic transmitter 80 and the receiver 82, and the time of flight (t) for the measurement of the acoustic signal. TOF ), time-of-flight average velocity (V TOF It can be calculated as:

[0167]

[0168] Given Model air as a perfect gas.

[0169] refer to Figure 12In the first step 200, formula (1) is used to determine the average flight time and average velocity along the respective line of sight 62 between the transmitter 80 and the receiver 82. This step is repeated as needed along each line of sight 62 between all the respective transmitters 80 and receivers 82 in the plane.

[0170] In the second step 202 and the third step 204, once the desired selection / number of line-of-sight time-of-flight average velocities have been calculated, tomography is applied to derive the line-of-sight average velocities at one or more nodes; these nodes are defined by the intersection of two or more lines of sight 62.

[0171] The output of tomography is a spatial velocity profile across the sensor plane. For example, tomography provides a mapped view of the velocity profile across the sensor plane. The velocity at each node is in time-of-flight currency.

[0172] Given a velocity profile, a weighted correction can be applied to the nodes with derived velocities to convert to the appropriate thermodynamic currency. In step 206, the velocity V at each node is used... TOF The thermodynamic weighted rate (V) is defined by the correction factor C1. thm For example, quality-weighted:

[0173] V thm =C1·V TOF Formula (2)

[0174] In step 208, the stagnation temperature T obtained from formula (2) is used. t and flow velocity V thm The knowledge derives the static temperature T at each node. s The calculation of the stagnation temperature T at the inlet. t Based on flight conditions or measurements of the aircraft or engine, which are known. Stagnation temperature T t The advantage is that it has a uniform profile across the inlet and sensor planes in the absence of exhaust gas re-intake.

[0175]

[0176] Steps 200 to 208 can be repeated interactively for each node until convergence to the tolerance.

[0177] In step 6.210, the local flow Mach number M at each node is known by applying its defining equation:

[0178]

[0179] In step 212, the mass flow rate of the inlet is estimated by spatial integration of the dimensionless mass flow rate formula across the sensor plane in the inlet:

[0180]

[0181] In step 8.214, given that these nodes are discrete sampling points within the profile, the estimated inlet mass flow rate is corrected for sampling errors using a factor C2. The sampling correction can be calculated based on a database from during or after flight. The sampling correction C2 can be calculated using computational methods such as computational fluid dynamics (CFD) or other methods designed to resolve the flow state to the required accuracy.

[0182]

[0183] Calculation of bypass airflow quality

[0184] See Figure 13 This system can be used to determine the mass flow rate of the bypass airflow B passing through engine 10. The bypass mass flow rate can be determined by understanding the mass flow rate (m³) of the core airflow A. cr From engine mass flow rate (m eng The core mass flow rate during flight can be estimated based on sea-level calibration methods and / or other conventional methods. Therefore, the bypass flow is the difference between the total mass flow rate entering the engine and the core mass flow rate.

[0185] m 125 =m eng -m cr Formula (7)

[0186] Where m 125 This is the bypass airflow at station 125. Station 125 is located downstream of fan 23, preferably between fan 23 and outlet guide vane 84.

[0187] Site 150 is located downstream of OGV 84. The mass flow rate of bypass air B at site 150 is equal to the mass flow rate measured at site 125 by adding / subtracting the mass flow rate due to sources or sinks (e.g., leaks, flow additions or reductions from auxiliary systems, etc.). The mass flow rates of such sinks, sources, and leaks can be modeled and are considered to be known.

[0188] Therefore, the mass flow rate measured at station 150 can be determined in the following way:

[0189] m 150 =m 125 -m leak,1 -m aux Formula (8)

[0190] Calculate the bypass stagnation pressure at the filling plane.

[0191] The mass flow rate at station 150 can be used to determine the stagnation pressure p via formulas (9) and (10). t :

[0192]

[0193] The stagnation pressure p, which varies with other parameters, can then be determined. t :

[0194] p t =f(m,p s ,T t Formula (10) ,A,γ,R)

[0195] • The mass flow rate m is derived by measuring the wind speed at the air inlet using the aforementioned method.

[0196] Static pressure p s Measurements are taken at site 150 (e.g., using a Pitot 86 or one or more embedded static sensors).

[0197] • Stagnation temperature T downstream of fan 23 t It can be derived from the engine analysis of shaft power, which is based on fan characteristic assumptions or measured using conventional techniques.

[0198] γ,R represents the known gaseous properties of the gas (usually air).

[0199] • A represents the geometric (cross-sectional) area at station 150, which can be measured or known from design parameters, etc. Corrections for flight conditions can be applied to account for expansion and / or contraction due to thermal or mechanical stresses, etc.

[0200] Given that the geometric area at station 150 is considered in Equation (10), the derived stagnation pressure is the average stagnation pressure across the channel and all associated flow characteristics (i.e., including the boundary layer and secondary flows, if any)).

[0201] This completes the estimation of the stagnation pressure on the filling plane. The total thrust can be calculated using various published gas path methods.

[0202] Calculation of nozzle emission factor during flight

[0203] Alternatively, a new gas approach is shown below, which focuses on the derivation of nozzle emission coefficients during engine operation.

[0204] Station 180 is located at the outlet of the bypass airflow, for example, at the throat of the bypass nozzle. Therefore, the mass flow rate m at stations 150 and 180... 150 and m 180 The ratio can be determined in the following ways:

[0205]

[0206] in

[0207] ·C d,180 This represents the nozzle emission coefficient.

[0208] ·Static pressure p at site 150 s,150 Measured using conventional methods.

[0209] ·Static pressure p at station 180 s,180 This is commonly referred to as the "nozzle reference pressure." The nozzle reference pressure can be considered equal to the ambient static pressure. Alternatively, a correction can be applied to the ambient static pressure.

[0210] • Stagnation pressure p at stations 150 and 180 t,150 and p t,180 They are considered equal according to the conventions in existing gas path methods.

[0211] • Geometric area A at stations 150 and 180 g,150 and A g,180 Measurements can be taken on the ground or based on known design parameters. Corrections for in-flight conditions can be applied to account for expansion and / or contraction due to thermal or mechanical stresses, etc.

[0212] Stagnation temperature T at stations 150 and 180 t,150 and T t,180 It is conserved in the absence of heat transfer. Alternatively, any heat source or sink between stations 150 and 180 can be considered as usual.

[0213] Quality flow ratio between sites 150 and 180 The extent of leakage is determined by any mass source and / or sink between these sites. In typical civil turbofan applications, leakage may occur, for example, through thrust reversers and / or nacelle seals. The amount of leakage can be determined by nacelle leakage tests conducted on the ground.

[0214] Therefore, the nozzle emission factor C d,180 This can be determined during testing or flight. The determined value can then be compared with the calculated or modeled value. Therefore, the difference between the measured and expected values ​​can indicate the effect caused by the nozzle, for example:

[0215] • External aerodynamic effects between the wings and the engine (also known as mounting effects), such as nozzle damping effect.

[0216] • Internal aerodynamic effects. These effects can be differences between the profile observed in the engine environment and the profile tested on the drilling rig, or different levels of turbulence intensity, etc.

[0217] Calculation of thrust during flight

[0218] Then, based on the published gas path method, the bypass thrust FG can be calculated using the thermodynamic parameters of the filling plane and the downstream nozzle performance coefficient. The mass flow rate is independently known according to formula (8), the stagnation pressure is derived from formula (10), and the in-flight nozzle emission coefficient is derived from formula (11).

[0219] Given the one-to-one correlation between velocity and the ratio of dimensionless mass flow rate to nozzle pressure under any given flight conditions, any representation of the flow parameters involved in Equations (1) to (6) can be used as a power setting parameter representing thrust.

[0220] advantage

[0221] This disclosure provides an apparatus for measuring the airflow characteristics of an axial-flow engine with minimal intrusive airflow.

[0222] This disclosure allows for the use of a greater number of sensors to improve the accuracy of airflow characteristic measurements.

[0223] This disclosure provides an airflow measurement system with reduced sensitivity to the aerodynamic characteristics of airflow (i.e., the variability of radial and circumferential profiles, the amount of turbulence, etc.).

[0224] This disclosure provides a measurement system that better represents the average characteristics of airflow passing through an engine.

[0225] This invention provides a cross-sectional profile of the airflow through the system. This allows for tomographic imaging of the airflow profile using a series of measurements.

[0226] Single-row / planar sensors simplify installation requirements and need to be installed at the machine's air inlet to accurately understand flow thermodynamic characteristics, such as stagnation pressure and temperature.

[0227] Mass flow rate can be determined in a practical and effective manner using a single-row acoustic sensor.

[0228] The location of the acoustic sensor near the acoustic liner of the fan or upstream of the acoustic liner of the fan can help filter out a portion of the pressure wave emitted from the tip of the rotating fan.

[0229] Positioning the flow sensor after the throat of the air inlet can facilitate a more uniform flow profile.

[0230] Using the known stagnation temperature and pressure upstream of the compressor (i.e., near the inlet) allows for accurate determination of the mass flow rate and other airflow characteristics across the inlet.

[0231] The entry structure described in this article may be part of a pod installation or part of an installation embedded in the fuselage structure.

[0232] This disclosure provides a non-invasive device for measuring airflow.

[0233] Although the system and method are described in relation to gas turbine engines, they can be applied to the walls / inlets of any other suitable turbine, such as axial-flow machines, typically involving high subsonic flow rates and stringent requirements for aerodynamic efficiency.

[0234] It should be understood that the present invention is not limited to the embodiments described above, and various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and this disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A turbine having an airflow measurement system, the airflow measurement system comprising: Multiple acoustic sensors are located upstream of the compressor of the turbine; The acoustic sensor includes at least one acoustic transmitter configured to transmit an acoustic waveform to an acoustic receiver via an intake airflow. The acoustic sensors are located in a single plane and mounted at angular intervals around the central axis of the air inlet; and The turbine has an internal component extending axially within the airflow, wherein the internal component intersects a single plane defined by the plurality of acoustic sensors, the internal component has additional acoustic sensors configured to communicate with the plurality of acoustic sensors, the internal component rotates about an axial axis, and the additional acoustic sensors are configured to rotate with the internal component. The turbine also includes an acoustic liner surrounding the rotor and configured to absorb vibrations generated by the rotation of the rotor, with the acoustic sensor located upstream of the acoustic liner, wherein the rotor is the turbine's fan or compressor.

2. The turbine of claim 1, wherein the engine includes a housing surrounding the central axis and defining the air intake, wherein the acoustic sensor is disposed on the housing.

3. The turbine of claim 2, wherein the acoustic sensors are configured as a circumferential array around the housing.

4. The turbine of claim 1, wherein the acoustic sensor does not protrude into the airflow and / or is mounted behind the gas-washed surface of the air inlet.

5. The turbine of claim 1, wherein the acoustic sensor is located downstream of the inlet throat.

6. The turbine of claim 1, wherein the internal components include a rotor hub and / or a nose cone.

7. The turbine of claim 1, wherein the acoustic sensor is mounted in the nacelle and / or rotor housing of the turbine.

8. The turbine of claim 1, wherein the airflow measurement system measures airflow velocity, volumetric flow rate and / or mass flow rate.

9. The turbine according to any one of claims 1 to 8, wherein the turbine comprises a gas turbine engine for an aircraft, the gas turbine engine comprising: An engine core, the engine core including a turbine, a compressor and a spindle connecting the turbine to the compressor; A fan, located upstream of the engine core, comprising a plurality of fan blades; A gearbox that receives input from the spindle and outputs drive to the fan to drive the fan at a lower rotational speed than the spindle.

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