Gas turbine engine

Abstract

GAS TURBINE ENGINE The present invention relates to a gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Illustrative embodiments described include a gas turbine engine (10) for an aircraft comprising: a central core (11) including a turbine (19), a compressor (14) and a central shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the central core (11), the fan comprising a plurality of fan blades; a nacelle (21) surrounding the central core (11) and defining a bypass duct (22) and a bypass exhaust nozzle (18);and a gearbox (30) which receives an input from the central shaft (26) and delivers an output drive to the fan (23) so as to drive the fan (23) at a lower rotational speed than the central shaft (26), in which the gas turbine engine (10) is configured such that a jet velocity ratio of a first jet velocity exiting the bypass exhaust nozzle (18) to a second jet velocity exiting an exhaust nozzle (20) of the central core (11) at idle conditions is greater by a factor of about 2 or more than the jet velocity ratio at maximum takeoff thrust conditions. [Figure 1];

Description

Title of the invention: Gas turbine engine

[0001] This description relates to a gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft.

[0002] Turbofan gas turbine engines for aircraft propulsion have many design factors that affect overall efficiency and power output or thrust. To allow higher thrust while maintaining efficiency, a larger diameter fan can be used. However, when the fan diameter is increased, the lower required fan speed tends to conflict with the requirements of the turbine component to which the central shaft is connected, typically a low-pressure turbine. A more optimal combination can be achieved by including a gearbox between the fan and the central shaft. This allows the fan to operate at a reduced rotational speed, thus permitting a larger fan, while maintaining a high rotational speed for the low-pressure turbine, thereby reducing the overall turbine diameter.

[0003] High propulsion efficiency for a geared gas turbine engine is achieved through a high mass flow rate through the engine. This can be achieved in part by increasing the engine's bypass ratio, which is the ratio of the mass flow rate of the bypass flow to the mass flow rate entering the core. To achieve a high bypass ratio with a larger fan while maintaining an optimum gear ratio and fan speed, the size of the core, particularly the low-pressure turbine, may need to be increased, which would make integrating a larger turbofan engine under an aircraft wing more difficult. A general problem to be solved, therefore, is how to achieve high propulsion efficiency for a larger geared gas turbine engine while still allowing the engine to be integrated into an aircraft.

[0004] According to a first aspect, a gas turbine engine for an aircraft is supplied, comprising:

[0005] a central core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor;

[0006] a blower located upstream of the central core, the blower comprising a plurality of blower blades;

[0007] a nacelle surrounding the central core and defining a bypass duct, a bypass exhaust nozzle; and

[0008] a gearbox which receives an input from the central shaft and delivers an output drive to the blower so as to drive the blower at a lower rotational speed than the central shaft,

[0009] in which the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting the bypass exhaust nozzle to a second jet velocity exiting a central core exhaust nozzle under idle conditions is greater by a factor of about 2 or more than the jet velocity ratio under maximum takeoff thrust conditions.

[0010] A large variation in the jet velocity ratio between idle and maximum thrust allows the core and bypass flows to be managed to maintain the engine operating with high propulsion efficiency and a high bypass ratio. This large variation can be achieved by optimizing the aerodynamic design of the fan and compressor components and / or by using other devices on the engine such as variable guide vanes and guide grids.

[0011] The jet velocity ratio, Rj, can be defined as follows:

[0012] [Math.l] £ _ VbCb J VcCcVlptUf

[0013] where VB is the first fully expanding jet velocity, CB is a thrust coefficient of the bypass nozzle, Vc is the second fully expanding jet velocity, Cc is a thrust coefficient of the central exhaust nozzle, qLPT is an isentropic efficiency of a central core minimum pressure turbine, and qF is an isentropic efficiency of air compression in the bypass duct by the blower. The fully expanding jet velocity can be defined as the axial jet velocity at the point where the exhaust jet is expanding to ambient pressure.

[0014] The gearbox may have a gear ratio between about 2.5 and about 5, or may have a gear ratio within a range as defined in more detail below.

[0015] In some examples the gas turbine engine can be configured so that the jet speed ratio is in a range of about 0.75 to about 1.3 under cruising conditions.

[0016] The gas turbine engine can be configured so that the jet speed ratio at idle conditions is between approximately 2 and 3.

[0017] The gas turbine engine can be configured so that the ratio of jet speed to maximum takeoff thrust conditions is between about 0.75 and 1.3, or possibly between about 0.8 and 1.0.

[0018] The blower may have an external diameter between approximately 220 cm and approximately 290 cm, and possibly between approximately 230 cm and approximately 260 cm, or may have an external diameter within a range as defined below.

[0019] The factor relating to the difference in jet velocity ratio between idle and maximum takeoff thrust conditions can be defined as

[0020] [Math.2] RJ in slow motion on the ground, RJ taking off •>

[0021] that is to say the ratio of jet speed at idle conditions on the ground divided by the ratio of jet speed at maximum takeoff thrust conditions.

[0022] The factor can be in a range between approximately 2 and approximately 3.5, or possibly between approximately 2.1 and approximately 3.16. Being above the lower limit of approximately 2 or 2.1 allows for reduced fuel combustion and can be achieved by features such as a more linear fan foot to maintain fan operability at a lower specific thrust, obtained by a larger fan diameter in combination with a smaller gearbox and center core. Above the upper limit of approximately 3.16 or 3.5, the required fan diameter becomes increasingly unacceptable for underwing installation and would require other features to reduce drag.

[0023] Maximum Takeoff Thrust (MTO) conditions can be defined as engine operation under International Standard Atmosphere (ISA) sea level pressure and temperature conditions + 15 °C at maximum takeoff thrust at the end of the runway, which is typically defined at an aircraft speed of approximately 0.25 NM, or between approximately 0.24 and 0.27 NM. Maximum takeoff thrust conditions for the engine can therefore be defined as engine operation at maximum takeoff thrust at ISA sea level pressure and temperature + 15 °C with a fan inlet speed of 0.25 NM.

[0024] Idle conditions can be defined as engine operation at approximately 4% of maximum thrust at takeoff under ISA sea level +15°C pressure and temperature conditions. Alternatively, idle conditions can be defined as engine operation at minimum thrust in continuous operation under ISA sea level +15°C pressure and temperature conditions.

[0025] According to a second aspect, a method for operating a gas turbine engine on an aircraft is provided, the gas turbine engine comprising:

[0026] a central core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor;

[0027] a blower located upstream of the central core, the blower comprising a plurality of blower blades; and

[0028] a gearbox which receives an input from the central shaft to drive the blower at a lower rotational speed than the central shaft,

[0029] in which the method includes operating the gas turbine engine to provide propulsion such that a jet velocity ratio between a first jet velocity exiting a bypass duct of the engine and a second jet velocity exiting an exhaust nozzle of the central core varies by more than a factor of two between idle and maximum thrust at ISA conditions at sea level.

[0030] The various optional and advantageous features associated with the invention according to the first aspect can also be applied to the second aspect.

[0031] As indicated elsewhere in this document, this description may relate to a gas turbine engine. Such a gas turbine engine may include a central core comprising a turbine, a combustion chamber, a compressor, and a central shaft connecting the turbine to the compressor. Such a gas turbine engine may include a blower (having blower blades) located upstream of the central core.

[0032] Arrangements of the present description may be particularly, though not exclusively, advantageous for blowers that are driven via a gearbox. Thus, the gas turbine engine may include a gearbox that receives an input from the central shaft and delivers an output drive to the blower so as to drive the blower at a lower rotational speed than the central shaft. The input to the gearbox may be directly from the central shaft, or indirectly from the central shaft, for example, via a shaft and / or spur gear. The central shaft may connect the turbine and the compressor, such that the turbine and the compressor rotate at the same speed (with the blower rotating at a lower speed).

[0033] 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 connecting the turbines and compressors, for example, one, two, or three shafts. By way of example only, the turbine connected to the central shaft may be a first turbine, the compressor connected to the central shaft may be a first compressor, and the central shaft may be a first central shaft. The central core may further comprise a second turbine, a second compressor, and a second central shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second shaft central shafts can be arranged to rotate at a higher rotational speed than the first central shaft.

[0034] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example, directly, for example via a generally annular conduit) a flow from the first compressor.

[0035] The gearbox can be arranged to be driven by the central shaft that is configured to rotate (for example, in operation) at the lowest rotational speed (for example, the first central shaft in the example above). Alternatively, the gearbox can be arranged to be driven only by the central shaft that is configured to rotate (for example, in operation) at the lowest rotational speed (for example, only the first central shaft, and not the second central shaft, in the example above). According to another possibility, the gearbox can be arranged to be driven by one or more shafts, for example, the first and / or second shafts in the example above.

[0036] The gearbox can be a reduction gearbox (in that the output to the blower has a lower rotational speed than the input from the central shaft). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "star" gearbox, as described in more detail elsewhere in this description. 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 in the order of 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. The gear ratio can for example be between any two of the values ​​in the preceding sentence.Strictly as an example, the reducer may be a "star" reducer with a ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

[0037] In any gas turbine engine as described and / or claimed herein, a combustion chamber may be provided axially downstream of the fan and the compressor(s). For example, the combustion chamber may be directly downstream (e.g., at the outlet) of the second compressor, where a second compressor is supplied. As a further example, the flow outlet of the combustion chamber may be supplied to the inlet of the second turbine, where a second turbine is supplied. The combustion chamber may be provided upstream of the turbine(s).

[0038] The compressor(s) (for example, the first compressor and the second compressor as described above) may comprise any number of stages, for example, several stages. Each stage may have a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other.

[0039] The turbine(s) (for example, the first turbine and the second turbine as described above) may comprise any number of stages, for example, several stages. Each stage may comprise 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.

[0040] Each fan blade can be defined as having a radial extent extending from a foot (or hub) at a radially internal gas-washed location, or 0% extent position, to a tip at a 100% extent position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip can be less than (or on the order of) any of the following values: 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 can be included in a range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range of 0.28 to 0.32. These ratios can commonly be referred to as the hub-to-tip ratio.The radius at the hub and the radius at the tip can also be measured at the leading edge (or the axially forward 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.

[0041] The radius of the blower can be measured between the centerline of the engine and the tip of a blower blade at its leading edge. The blower diameter (which can simply be double the blower radius) can be greater than (or in the order of): 220 cm, 230 cm, 240 cm, 250 cm (approximately 100 inches), 260 cm, 270 cm (approximately 105 inches), 280 cm (approximately 110 inches), 290 cm (approximately 115 inches), 300 cm (approximately 120 inches), 310 cm, 320 cm (approximately 125 inches), 330 cm (approximately 130 inches), 340 cm (approximately 135 inches), 350 cm, 360 cm (approximately 140 inches), 370 cm (approximately 145 inches), 380 cm (approximately 150 inches), 390 cm (approximately 155 inches), 400 cm, 410 cm (approximately 160 inches) or 420 cm (approximately 165 inches).The diameter of the blower can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, in the range of 240 cm to 280 cm or 330 cm to 380 cm.

[0042] The fan speed may vary during operation. Generally, the fan speed is lower for fans with a large diameter. By way of non-limiting example only, the fan speed under cruising conditions may be less than 2,500 rpm, for example, less than 2,300 rpm. By way of non-limiting example only, the fan speed under cruising conditions for an engine having a fan diameter in the range of 220 cm to 300 cm (e.g. 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of 1,700 rpm to 2,500 rpm, e.g. in the range of 1,800 rpm to 2,300 rpm, e.g. in the range of 1,900 rpm to 2,100 rpm.Strictly as a further non-limiting example, the fan speed under cruising conditions for an engine with a fan diameter in the range of 320 cm to 380 cm can be in the range of 1200 rpm to 2000 rpm, for example in the range of 1300 rpm to 1800 rpm, for example in the range of 1400 rpm to 1800 rpm.

[0043] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the fan blade tip to move at a velocity Utip. The work done by the fan blades 13 on the flow leads to an increase in the enthalpy dH of the flow. A fan tip load can be defined as dH / Utip², where dH is the increase in enthalpy (e.g., the increase in mean enthalpy 1-D) across the fan and Utip is the (translational) velocity of the fan tip, for example, at the leading edge of the tip (which can be defined as a fan tip radius at the leading edge multiplied by the angular velocity).The peak fan load under cruise conditions may be greater than (or on the order of) any of the following values: 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (with dH assigned to the unit JKg 1 and Utip assigned to the unit ms1). The peak fan load may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0044] Gas turbine engines according to this description may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flux through the bypass duct to the mass flow rate of the flux through the core under cruising conditions. In some arrangements, the bypass ratio may be greater than (or on the order of) any of the following values: 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 an inclusive range delimited by any two of the values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range of 13 to 16, or 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially external to the reactor. The radially external surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0045] The overall pressure ratio of a gas turbine engine as described and / or claimed herein can be defined as the ratio of the total pressure upstream of the fan to the total pressure at the outlet of the high-pressure compressor (before entering the combustion chamber). By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein during cruise operation may be greater than (or on the order of) any of the following values: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 50 to 70.

[0046] 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 cruising conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any one of the following values: 110 Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s, or 80 Nkg⁻¹s. The specific thrust may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 80 Nkg⁻¹s to 100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. These engines may be particularly efficient compared with conventional gas turbine engines.

[0047] A gas turbine engine as described and / or claimed herein can have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein can be capable of producing a maximum thrust of at least (or on the order of) any one of the following values: 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 can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds). As an example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, for example, from 350 kN to 400 kN. The thrust mentioned above may be the thrust maximum net under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure of 101.3 kPa, temperature of 30 degrees C), with the engine stationary.

[0048] During operation, the flow temperature at the inlet of a high-pressure turbine can be particularly high. This temperature, known as TET, can be measured at the combustion chamber outlet, for example, immediately upstream of the first turbine blade, which itself can be called the nozzle guide blade. Under cruising conditions, the TET can be at least (or on the order of) any one of the following values: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K, or 1650 K. The TET under cruising conditions can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits). The maximum TET in motor use can be, for example, at least (or on the order of) any of the following values: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K.The maximum TET can be within an inclusive range bounded by any two of the values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range of 1800 K to 1950 K. The maximum TET can occur, for example, under a high-thrust condition, for example, under a maximum takeoff thrust (MTO) condition.

[0049] Maximum takeoff thrust for the engine can be defined as engine operation at plus or minus 15 degrees Celsius of sea-level pressure and temperature conditions of the International Standard Atmosphere (ISA) at maximum takeoff thrust at the end of the runway, which is typically defined at an aircraft speed of approximately 0.25 NM, or between approximately 0.24 and 0.27 NM. Maximum takeoff thrust conditions for the engine can therefore be defined as engine operation at maximum takeoff thrust at ISA sea-level pressure and temperature with a fan inlet speed of 0.25 NM.

[0050] A portion of a fan blade and / or airfoil described and / or claimed herein may be made from any suitable material or any suitable combination of materials. For example, at least a portion of the fan blade and / or airfoil may be made at least partially from a composite material, for example, a metal matrix composite and / or an organic matrix composite, such as carbon fiber. By way of further example, at least a portion of the fan blade and / or airfoil may be made at least partially from 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 blades may comprise at least two regions made using materials different. For example, a fan blade may have a protective leading edge, which can be made of a material more resistant to impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge could, for example, be made of titanium or a titanium-based alloy. Thus, strictly as an example, a fan blade could have a body made of carbon fiber or aluminum (such as a lithium-aluminum alloy) with a titanium leading edge.

[0051] A blower as described and / or claimed herein may include a central portion from which the blower blades may extend, for example, in a radial direction. The blower blades may be attached to the central portion in any desired manner. For example, each blower blade may include a fastener that can engage with a corresponding slot in the hub (or disc). Strictly by way of example, such a fastener may be in the form of a dovetail that can slide into and / or engage with a corresponding slot in the hub / disk to attach the blower blade to the hub / disk. As a further example, the blower blades may be formed as a single unit with a central portion. Such an arrangement may be referred to as a bladed disc or bladed ring.Any suitable process can be used to manufacture such a bladed disc or bladed ring. For example, at least part of the blower blades can be machined from a block and / or at least part of the blower blades can be attached to the hub / disc by welding, such as linear friction welding.

[0052] The gas turbine engines described and / or claimed herein may or may not be equipped with a variable area nozzle (VAN). Such a variable area nozzle can allow the outlet size of the bypass duct to be varied during operation. The general principles of this description can be applied to engines with or without a VAN.

[0053] The blower of a gas turbine as described and / or claimed herein may have any number of blower blades, for example 14, 16, 18, 20, 22, 24 or 26 blower blades.

[0054] As used here, the term cruise conditions has a classical meaning and will be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, those skilled in the art will immediately recognize that cruise conditions mean the engine's operating point at mid-cruise of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft on which the gas turbine engine is intended to be installed. In this respect, the term mid-cruise means the point in an aircraft flight cycle at which 50% of the total fuel burned between the end of the climb and the beginning of the descent has been burned (this which can roughly correspond to the midpoint, in terms of time and / or distance, between the end of the climb and the beginning of the descent. Cruise conditions thus define an operating point for the gas turbine engine that provides sufficient thrust to ensure continuous operation (i.e., maintaining a constant altitude and Mach number) at mid-cruise of the aircraft on which it is intended to be installed, taking into account the number of engines supplied to that aircraft. For example, when an engine is designed to be installed on an aircraft with two engines of the same type, under cruise conditions the engine provides half the total thrust that would be required for continuous operation of that aircraft at mid-cruise.

[0055] In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the engine operating point that provides a specified thrust (required to provide, in combination with any other engine on the aircraft, continuous operation of the aircraft on which it is intended to be installed at a given Mach number at mid-cruise) under mid-cruise atmospheric conditions (defined by the international standard atmosphere according to ISO 2533 at mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions, and Mach number are known, and therefore the engine operating point under cruise conditions is clearly defined.

[0056] By way of example only, forward speed under cruise conditions can be any point in the range of Mach 0.7 to 0.9, for example from 0.75 to 0.85, for example from 0.76 to 0.84, for example from 0.77 to 0.83, for example from 0.78 to 0.82, for example from 0.79 to 0.81, for example in the order of Mach 0.8, in the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

[0057] By way of example only, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude in the range of 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 ft), 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 ft) 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, on the order of 11,000 m. Cruising conditions can correspond to standard atmospheric conditions at any given altitude within these ranges.

[0058] By way of example only, cruise conditions may correspond to an engine operating point that provides a known required thrust level (for example, a value in the range of 30 kN to 35 kN) at a Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 ft (11,582 m). By way of further example only, cruise conditions may correspond to an engine operating point that provides a known required thrust level (for example, a value in the range of 50 kN to 65 kN) at a Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 ft (10,668 m).

[0059] In use, a gas turbine engine described and / or claimed herein can operate under the cruise conditions defined herein. Such cruise conditions can be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one (e.g., two or four) gas turbine engine can be mounted to provide propulsive thrust.

[0060] According to one aspect, an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. The aircraft according to this aspect is the aircraft on which the gas turbine engine is intended to be installed. Thus, the cruise conditions according to this aspect correspond to the mid-cruise conditions of the aircraft, as defined elsewhere in this document, and / or the maximum takeoff thrust conditions refer to the maximum takeoff thrust conditions of the aircraft.

[0061] According to one aspect, a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. Operation according to this aspect may include (or may be) operation at mid-cruise and / or maximum takeoff thrust of the aircraft, as defined elsewhere in this document.

[0062] Those skilled in the art will understand that, except for mutual exclusivity, a feature or parameter described in relation to any of the above aspects may be applied to any other aspect. Furthermore, except for mutual exclusivity, any feature or parameter described herein may be applied to any aspect and / or associated with any other feature or parameter described herein.

[0063] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0064] [Fig. 1] is a side cross-sectional view of a gas turbine engine;

[0065] [Fig.2] is a close-up side cross-sectional view of an upstream part of an engine gas turbine;

[0066] [Fig.3] is a partially cutaway view of a gearbox for a gas turbine engine;

[0067] [Fig.4] is an example of a fuel combustion change plot in function of the jet velocity ratio;

[0068] [Fig. 5] is a schematic drawing of an aircraft having a mounted gas turbine engine on this one; and

[0069] [Fig.6] is a schematic drawing illustrating the concept of a jet velocity in total expansion.

[0070] The [Fig. Figure 1 illustrates a gas turbine engine 10 having a rotational axis 9. The engine 10 includes an air inlet 12 and a propulsion fan 23 that generates two airflows: a main airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the airflow A. The central core 11 includes, 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 central 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.

[0071] In operation, the central airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where further compression occurs. The compressed air discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then expand, driving the high-pressure and low-pressure turbines 17, 19 before being discharged through the nozzle 20 to provide propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by means of a suitable interconnecting shaft 27. The blower 23 typically provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0072] An exemplary arrangement for a gas turbine engine 10 is illustrated in [Fig. 2]. The low-pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outward from the sun wheel 28 and meshing with it are a plurality of satellites 32 which are coupled together by a satellite carrier 34. The satellite carrier 34 constrains the satellites 32 to rotate about the sun wheel 28 in synchronicity while allowing each satellite 32 to rotate about its own axis. The satellite carrier s 34 is coupled via links 36 to the fan 23 to drive its rotation around the motor axis 9. Radially outwards from the satellites 32 and meshing with them is a ring or a planetary ring 38 which is coupled, via links 40, to a fixed support structure 24.

[0073] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" can be used to designate the minimum-pressure turbine stages and the minimum-pressure compressor stages (i.e., not including the blower 23) respectively, and / or the turbine and compressor stages connected by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the blower 23). In some documents, the "low-pressure turbine" and "low-pressure compressor" referred to in this description may alternatively be called "intermediate-pressure turbine" and "intermediate-pressure compressor." Where such alternative nomenclature is used, the blower 23 may be called the first compression stage, or the minimum-pressure compression stage.

[0074] The epicyclic gear set 30 is shown in more detail by way of example in [Fig. 3]. Each of the sun gear 28, the planet gears 32, and the ring gear 38 has teeth around its periphery for meshing with the other gears. However, for clarity, only representative portions of the teeth are shown in [Fig. 3]. Four planet gears 32 are shown, but the reader familiar with the technique will understand that the number of planet gears 32 may vary depending on the application of the claimed invention. Practical applications of a planetary epicyclic gear set 30 generally include at least three planet gears 32.

[0075] 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 links 36, the ring gear 38 being fixed. However, any other suitable type of epicyclic gearbox 30 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 gear (or ring) 38 able to rotate. In such an arrangement, the blower 23 is driven by the ring gear 38. By way of further example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 may rotate.

[0076] It is understood that the arrangement illustrated in Figures 2 and 3 is provided by way of example only, and various alternatives are within the scope of this description. By way of example only, any suitable arrangement may be used to place the gearbox 30 in the motor 10 and / or to connect the gearbox 30 to the motor 10. By way of further example, the connections (such as the connections 36, 40 in the example in [Fig. 2]) between the gearbox 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) can have any desired stiffness or flexibility. By way of further example, any suitable arrangement of bearings between rotating and fixed parts of the motor (for example, between the input and output shafts of the gearbox and fixed structures, such as the gearbox housing) can be used, and the invention is not limited to the example arrangement in [Fig. 2]. For example, where the gearbox 30 has a star arrangement (described above), those skilled in the art will readily understand that the arrangement of the output and support connections and the bearing locations are typically different from those shown as an example in [Fig. 2].

[0077] Accordingly, the present invention extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangements, and bearing locations.

[0078] Optionally, the gearbox can drive additional and / or alternative components (for example, the intermediate pressure compressor and / or a pre-compressor).

[0079] Other gas turbine engines to which the present description may apply 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 illustrated in [Fig. 1] has a split-flow nozzle 18, 20, i.e., the flow through the bypass duct 22 has its own nozzle 18 which is separated radially to and outside the reactor nozzle 20. However, this is not limiting, and any aspect of the invention may also apply to engines in which the flow in the bypass duct 22 and the flow in the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be called a mixed-flow nozzle. One or both nozzles (mixed or split flow) may have a fixed or variable surface area.

[0080] The geometry of the gas turbine engine 10, and of its components, is defined by a conventional axial system, comprising an axial direction (which is aligned with the axis of rotation 9), a radial direction (from bottom to top in [Fig. 1]), and a circumferential direction (perpendicular to the page in the view of [Fig. 1]). The axial, radial, and circumferential directions are perpendicular to each other.

[0081] Figure 4 illustrates an example of a fuel combustion change plot, AFB, as a function of the jet velocity ratio, Rj, with other factors being constant. The change in contribution to fuel combustion from the efficiency The propulsion value of the 401 is determined relative to an optimal jet velocity ratio of approximately 1.0, with an increase in fuel combustion both above and below this value. Factors that can affect the jet velocity ratio include the relative rotational speeds of the fan and turbine and the exhaust nozzle areas for the bypass and center exhausts. A lower gearbox gear ratio, i.e., a gear ratio of approximately 3.4 or lower, will tend to result in jet velocity ratios of 1.0 or higher. To keep fuel combustion loss to approximately 0.5% or less, it can be seen from [Fig. 4] that the jet velocity ratio should be between approximately 1.0 and approximately 1.3. As the jet velocity ratio increases, the increase in fuel combustion becomes more significant.Another preferred upper limit for the jet velocity ratio is about 1.2, which keeps the fuel combustion increase to about 0.25 to 0.3%.

[0082] For a higher gear ratio, i.e., about 3.3 to 3.4 and above, for example up to about 3.8 or, in some cases, even higher, the jet velocity ratio tends to be about 1.0 or lower. As the jet velocity ratio decreases, the contribution of fuel combustion by the propulsion efficiency 401 increases, and at a higher rate than for the portion above 1.0. To keep this loss at plus or minus about 0.5%, it can be seen from [Fig. 4] that the jet velocity ratio must be kept at plus or minus about 0.8 up to about 1.0, and for a ratio of about 0.75 and below, the contribution of fuel combustion by the propulsion efficiency becomes dominant, amounting to about 0.7% and above.Another preferred lower limit for the jet velocity ratio of approximately 0.85 or 0.90 can be used to keep the fuel combustion contribution from propulsion efficiency to approximately 0.25% or less. However, further reductions in the jet velocity ratio allow for the use of a higher gear ratio and / or a lower pressure ratio on the IP turbine, enabling the use of a smaller, faster, and / or lighter IP turbine, resulting in a lower contribution of fuel combustion loss 402 from the IP turbine. A range of approximately 0.75 to approximately 0.82 for the jet velocity ratio is therefore advantageous.

[0083] For a given set of gears composing an epicyclic gearbox, a planetary drive arrangement will produce a higher gear ratio than a star drive arrangement. A star arrangement is therefore generally preferred in combination with a jet speed ratio of about 1.0 and above, and a planetary arrangement for a jet speed ratio of about 1.0 and below.

[0084] In general terms, for this reason, the gas turbine engine can be configured so that the jet speed ratio is in a range of about 0.75 to about 1.3 under cruising conditions.

[0085] Figure 5 illustrates an example of an aircraft 50 having a gas turbine engine 10 attached to each wing 51a, 51b thereof. When the aircraft is in flight under cruise conditions, as defined herein, each gas turbine engine 10 operates such that a jet velocity ratio between a first jet velocity exiting a bypass duct of the engine 10 and a second jet velocity exiting an exhaust nozzle 20 of the central core is in a range of approximately 0.75 to approximately 1.3.

[0086] Figure 6 illustrates an example of an exhaust nozzle 60 of a gas turbine engine. The pressure Pj at the outlet or at the throat 61 of the exhaust nozzle 60 is greater than the ambient pressure Pa around the engine. At a certain distance from the nozzle outlet 61, the jet pressure is equal to the ambient pressure, i.e., Pj = Pa. The fully expanding jet velocity is defined as the jet velocity 62 at this point, i.e., the jet velocity along the engine axis at a minimum distance from the exhaust nozzle where the pressure is equal to the ambient pressure.

[0087] The parameters that can be adjusted to obtain a jet velocity ratio within the desired range can include the LPT blade exit angle, LPT exit direction, and LPT rotation speed.

[0088] The following table illustrates examples of parameters for two examples of motors, Example 1 being for a relatively low-power motor, or lower, and Example 2 for a relatively high-power motor, or higher. A small motor may, for example, have a fan diameter of between approximately 200 and 280 cm and / or a maximum net thrust of between approximately 160 and 250 kN or as defined elsewhere in this document. A large motor may, for example, have a fan diameter of between approximately 310 and 380 cm and / or a maximum net thrust of between approximately 310 and 450 kN or as defined elsewhere in this document.

[0089] [Tables 1] Parameter Example 1 (small motor) Example 2 (large motor) Blower diameter (cm) 215 320 Total outlet pressure LPT at maximum flow (kPa) 130 130 Maximum outlet mass flow rate LPT (kg / s) 50 100 Final rotor area LPT (m²) 0.38 or less, e.g., 0.25 to 0.38 0.75 or less, e.g., 0.5 to 0.75 Total inlet pressure ESS at maximum flow (kPa) 140 140 Inlet mass flow rate ESS (kg / s) 50 100 Inlet rotor area ESS (m²) 0.275 or more, e.g., 0.27 to 0.3 0.55 or more, e.g., 0.55 to 0.6

[0090] The above parameters relating to the total outlet pressure (LPT) at maximum flow, the maximum outlet mass flow rate (LPT), and the final rotor area (LPT) together determine the outlet flow velocity of the LPT, i.e., the flow velocity at the outlet of the central core. The total inlet pressure (ESS) at maximum flow, the maximum inlet mass flow rate (ESS), and the inlet rotor area (ESS) together determine the velocity at the inlet of the central core. The axial velocity of the exhaust flow from the bypass exhaust nozzle can be determined, at least in part, by the area of ​​the bypass exhaust nozzle outlet.

[0091] To achieve a jet velocity ratio within the desired range, the blower may be equipped with features such as a more linear blower foot. Compressors, particularly the high-pressure compressor, may be equipped with features to manage their operability, allowing them to operate at the lower power required to meet the defined ratios. These features may include, for example, devices such as variable guide vanes. This changes the effect of the flow on the vanes and helps maintain an operability margin that prevents the compressor from surging or stalling when operating at lower speeds.

[0092] It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements may be made without departing from the concepts described herein. Except as mutually exclusive, any feature may be used separately or in combination with other features and the description extends to and includes all combinations and sub-combinations of one or more characteristics described here.

Claims

Demands

1. Method of operating a gas turbine engine (10) on an aircraft (50), the gas turbine engine (10) comprising: a central core (11) comprising a turbine (19), a compressor (14) and a central shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the central core, the fan comprising a plurality of fan blades; a nacelle (21) surrounding the central core (11) and defining a bypass duct (22) and a bypass exhaust nozzle (18); and a gearbox (30) which receives input from the central shaft (26) to drive the fan (23) at a rotational speed lower than that of the central shaft (26), wherein at least one of the LPT blade exit angle, LPT exit area, and LPT rotational speed is selected so as to achieve operation of the gas turbine engine (10) providing propulsion with a jet velocity ratio between a first jet velocity exiting the bypass exhaust nozzle (18) and a second jet velocity exiting an exhaust nozzle (20) of the central core (11) under idle conditions that is greater by a factor of about 2 or more than the jet velocity ratio under maximum takeoff thrust conditions, and in which the jet velocity ratio, Rj, is defined as follows: [Math.l] R _ VbCb where VB is the first fully expanding jet velocity, CB is a thrust coefficient of the bypass exhaust nozzle (18), Vc is the second fully expanding jet velocity, Cc is a thrust coefficient of the core exhaust nozzle (20), qLPT is the isentropic efficiency of a central minimum pressure turbine (11) and qF is the isentropic efficiency of the fan tip; and in which The operation of the gas turbine engine under cruising conditions results in a load at the end of the fan comprising between 0.28 and 0.34, the load at the end of the fan being defined as dH / U end2, where dH is the enthalpy increase across the fan and U end is the translational velocity at the leading edge of the end.

2. A method according to claim 1, wherein the operation of the gas turbine engine under cruising conditions results in a load at the end of the fan of between 0.30 and 0.

34.

3. A method according to claim 1 or claim 2, wherein the operation of the gas turbine engine under cruising conditions results in a load at the end of the fan of between 0.28 and 0.

31.

4. 4 Method according to any one of the preceding claims, wherein the operation of the gas turbine engine under cruising conditions results in a load at the end of the fan of between 0.29 and 0.

3.

5. Method according to claim 1, wherein: at least a portion of each fan blade is made from carbon fiber.

6. A method according to claim 5, wherein the operation of the gas turbine engine under cruising conditions results in a load at the end of the fan of between 0.31 and 0.

34.

7. A method according to any one of the preceding claims, wherein the gearbox has a transmission ratio between 3.1 and 3.

8.

8. A method according to any one of the preceding claims, wherein the factor is between 2 and 3.

5.

9. A method according to any one of claims 1 to 7, wherein: the factor is between 2.1 and 3.16; the transmission ratio is between 3.1 and 3.8; and the engine bypass ratio under cruising conditions is between 13 and 16.

10. A method according to any one of claims 1 to 6, wherein: the factor is between 2 and 3.5; the transmission ratio is between 3.0 and 3.9; The engine bypass ratio under cruising conditions is between 11 and 16; and The diameter is between 200 and 280 cm.

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