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. Exemplary embodiments described include a gas turbine engine (10) for an aircraft 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 (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 outputs a drive to the fan (23) so as to drive the fan (23) at a lower rotational speed than the central shaft (26), wherein 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] The present disclosure relates to a gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft.

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

[0003] High propulsive 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 bypass ratio of the engine, which is the ratio of the mass flow rate of the bypass stream to the mass flow rate entering the central core. To achieve a high bypass ratio with a larger fan while maintaining an optimum gear ratio and fan speed, the size of the central core, particularly the low-pressure turbine, may need to be increased, which would make integration of a larger turbofan under an aircraft wing more difficult. A general problem that needs to be solved, therefore, is how to achieve high propulsive efficiency for a larger geared gas turbine engine while still allowing integration of the engine into an aircraft.

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

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

[0006] a fan located upstream of the central core, the fan comprising a plurality of fan 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 outputs a drive to the blower so as to drive the blower at a lower rotational speed than the central shaft,

[0009] wherein 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 center core exhaust nozzle under idle conditions is greater by a factor of about 2 or more than the jet velocity ratio at 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 keep the engine operating with high propulsive efficiency and a high bypass ratio. The 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 grilles.

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

[0012] [Math.l] £ _ VbCb J VcCcVlptUf

[0013] where VB is the first full expansion jet velocity, CB is a thrust coefficient of the bypass nozzle, Vc is the second full expansion jet velocity, Cc is a thrust coefficient of the center exhaust nozzle, qLPT is an isentropic efficiency of a center core minimum pressure turbine, and qF is an isentropic efficiency of air compression in the bypass duct by the fan. The full expansion 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 of 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 may be configured such that the jet velocity ratio is in a range of from about 0.75 to about 1.3 at cruise conditions.

[0016] The gas turbine engine may be configured such that the jet velocity ratio at idle conditions is between about 2 and 3.

[0017] The gas turbine engine may be configured such that the jet velocity ratio at maximum takeoff thrust conditions is between about 0.75 and 1.3, or optionally between about 0.8 and 1.0.

[0018] The blower may have an outer diameter of between about 220 cm and about 290 cm, and optionally between about 230 cm and about 260 cm, or may have an outer diameter within a range as defined below.

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

[0020] [Math.2] Rj idling on the ground Rj taking off •>

[0021] i.e. the jet speed ratio at ground idle conditions divided by the jet speed ratio at maximum takeoff thrust conditions.

[0022] The factor may be in a range between about 2 and about 3.5, or possibly between about 2.1 and about 3.16. Being above the lower limit of about 2 or 2.1 allows for reduced fuel burn and may be achieved by features such as a more linear fan root to maintain fan operability at lower specific thrust achieved by a larger fan diameter in combination with a smaller gearbox and center core. Above the upper limit of about 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 may be defined as operation of the engine at International Standard Atmosphere (ISA) sea level pressure and temperature + 15°C at maximum takeoff thrust at the end of the runway, which is typically defined at an aircraft speed of about 0.25 Mn, or between about 0.24 and 0.27 Mn. Maximum takeoff thrust conditions for the engine may therefore be defined as operation of the engine at maximum takeoff thrust at ISA sea level pressure and temperature + 15°C with a fan inlet speed of 0.25 Mn.

[0024] Idle conditions may be defined as engine operation at about 4% of maximum takeoff thrust under ISA pressure and temperature conditions at sea level +15°C. Alternatively, idle conditions may be defined as engine operation at minimum steady-state thrust under ISA pressure and temperature conditions at sea level +15°C.

[0025] According to a second aspect, a method of 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 fan located upstream of the central core, the fan comprising a plurality of fan 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] wherein the method comprises 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 may also be applied to the second aspect.

[0031] As indicated elsewhere herein, the present disclosure 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 fan (having fan blades) located upstream of the central core.

[0032] Arrangements of the present disclosure may be particularly, although not exclusively, advantageous for fans that are driven via a gearbox. Thus, the gas turbine engine may include a gearbox that receives an input from the central shaft and outputs a drive to the fan so as to drive the fan 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 spur shaft and / or gear. The central shaft may interlock the turbine and compressor, such that the turbine and compressor rotate at the same speed (with the fan 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, e.g., 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 include 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 shaft may be arranged to rotate at a higher rotational speed than the first central shaft.

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

[0035] The gearbox may be arranged to be driven by the central shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first central shaft in the example above). For example, the gearbox may be arranged to be driven only by the central shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. only the first central shaft, and not the second central shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by one or more shafts, e.g. the first and / or second shafts in the example above.

[0036] The gearbox may 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 may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere in this specification. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example in the range 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 may be for example between any two of the values ​​in the preceding sentence.Strictly by way of example, the reducer may be a "star" reducer having 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 combustor may be provided axially downstream of the fan and the compressor(s). For example, the combustor may be directly downstream (e.g., at the outlet) of the second compressor, where a second compressor is provided. As a further example, the flow out of the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0038] The compressor(s) (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise 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) (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple 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 may be defined as having a radial extent extending from a root (or hub) at a radially inner 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 may 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 fan blade radius at the hub to the fan blade radius at the tip may be included in a range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.32. These ratios may be commonly 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 axially forward) portion 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 outboard of any platform.

[0041] The fan radius can be measured between the engine centerline and the tip of a fan blade at its leading edge. The blower diameter (which may simply be twice the blower radius) may be greater than (or of the order of): 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches) or 420 cm (approximately 165 inches).The diameter of the blower 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 limits), for example in the range 240 cm to 280 cm or 330 cm to 380 cm.

[0042] The rotational speed of the blower may vary in use. Generally, the rotational speed is lower for blowers with a large diameter. By way of non-limiting example only, the rotational speed of the blower 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 rotational 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 1700 rpm to 2500 rpm, e.g. in the range of 1800 rpm to 2300 rpm, e.g. in the range of 1900 rpm to 2100 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruising conditions for an engine having a fan diameter in the range of 320 cm to 380 cm may 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] When operating 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 speed Utip. The work done by the fan blades 13 on the flow results in an increase in the enthalpy dH of the flow. A fan tip load can be defined as dH / UpOint2, where dH is the enthalpy increase (e.g., the 1-D average enthalpy increase) across the fan and Utip is the (translational) velocity of the fan tip, e.g., 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 at cruise conditions may be greater than (or of 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 (all units in this paragraph being Jkg 'K ' / (ms *)2). 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 limits), for example in the range 0.28 to 0.31 or 0.29 to 0.3.

[0044] The gas turbine engines according to the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements, the bypass ratio may be greater than (or on the order of) any of the following 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 in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 13 to 16, or 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine. The radially outer 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 may 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 at cruise may be greater than (or of the order of) any one of the following values: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio 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 limits), for example, in the range of 50 to 70.

[0046] The specific thrust of an engine may 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 in the order of) any of the following values: 110 Nkg 's, 105 Nkg's, 100 Nkg's, 95 Nkg's, 90 Nkg's, 85 Nkg's or 80 Nkg's. 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 limits), for example, in the range of 80 Nkg 's to 100 Nkg 's, or 85 Nkg 's to 95 Nkg 's. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

[0047] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any 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 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 limits). By way of 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 350 kN to 400 kN. The thrust referred to above may be the thrust maximum net power under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

[0048] In use, the temperature of the flow at the high-pressure turbine inlet may be particularly high. This temperature, called TET, may be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself may be called the nozzle guide vane. Under cruising conditions, the TET may be at least (or of 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 may be within an inclusive range delimited by any two of the values ​​of the preceding sentence (i.e. the values ​​may form upper or lower limits). The maximum TET in engine use may be, for example, at least (or of 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 may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 1800 K to 1950 K. The maximum TET may occur, for example, in a high thrust condition, for example, in a maximum takeoff thrust (MTO) condition.

[0049] Maximum takeoff thrust for the engine may be defined as operating the engine within plus or minus 15 degrees C of the International Standard Atmosphere sea level pressure and temperature conditions at maximum takeoff thrust at the end of the runway, which is typically defined at an aircraft speed of about 0.25 Mn, or between about 0.24 and 0.27 Mn. Maximum takeoff thrust conditions for the engine may therefore be defined as operating the engine at maximum takeoff thrust at ISA sea level pressure and temperature with a fan inlet speed of 0.25 Mn.

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

[0051] A fan as described and / or claimed herein may include a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be secured to the central portion in any desired manner. For example, each fan blade may include a fastener that is engageable 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 is slidable into and / or engageable with a corresponding slot in the hub / disc to secure the fan blade to the hub / disc. By way of further example, the fan blades may be integrally formed with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may 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 provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the outlet of the bypass duct to be varied during use. The general principles of the present disclosure may apply to engines with or without a VAN.

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

[0054] As used herein, the term cruise conditions has a conventional 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 readily recognize that cruise conditions mean the mid-cruise engine operating point of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft on which the gas turbine engine is intended to be installed. In this regard, the term mid-cruise means the point in an aircraft flight cycle at which 50% of the total fuel burned between the end of climb and the beginning of descent has been burned (this which may correspond approximately to the midpoint, in terms of time and / or distance, between the end of climb and the beginning of descent. Cruise conditions thus define an operating point of the gas turbine engine which provides thrust that ensures continuous operation (i.e. maintaining a constant altitude and a constant Mach number) at mid-cruise of an aircraft on which it is intended to be installed, taking into account the number of engines provided in that aircraft. For example, when an engine is designed to be installed on an aircraft having two engines of the same type, at 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 operating point of the engine that provides a specified thrust (required to provide, in combination with any other engine in 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 the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and therefore the operating point of the engine under cruise conditions is clearly defined.

[0056] By way of example only, the forward speed under cruise conditions may be any point in the range of Mach 0.7 to 0.9, e.g., 0.75 to 0.85, e.g., 0.76 to 0.84, e.g., 0.77 to 0.83, e.g., 0.78 to 0.82, e.g., 0.79 to 0.81, e.g., of the order of Mach 0.8, of the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, e.g., 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 (about 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 (about 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 of the order of 11,000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0058] By way of example only, the cruise conditions may correspond to an engine operating point that provides a known required thrust level (e.g., 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, the cruise conditions may correspond to an engine operating point that provides a known required thrust level (e.g., 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 may operate under the cruise conditions defined herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0060] According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. 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 of the aircraft, as defined elsewhere herein, and / or the maximum takeoff thrust conditions relate to the maximum takeoff thrust conditions of the aircraft.

[0061] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. 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 herein.

[0062] Those skilled in the art will understand that, unless mutually exclusive, a feature or parameter described in connection with any of the above 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 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 sectional view of a gas turbine engine;

[0065] [Fig.2] is a close-up side sectional view of an upstream portion of a motor 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 plot of the fuel combustion change in function of jet velocity ratio;

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

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

[0070] The [Fig. 1] illustrates a gas turbine engine 10 having an axis of rotation 9. The engine 10 comprises an air inlet 12 and a propulsion fan 23 which generates two airflows: a main airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 which receives the airflow A. The central core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19 and a 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 use, the core air stream 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 combusted. The resulting hot combustion products then expand, and thereby drive 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 fan 23 generally 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 gear, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of and meshing with the sun gear 28 are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to rotate about the sun gear 28 in synchronicity while allowing each planet gear 32 to rotate about its own axis. The planet carrier s 34 is coupled via links 36 to the fan 23 to drive its rotation about the engine axis 9. Radially outwardly of the planets 32 and meshing therewith is a ring or a planetary ring gear 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" may be used to refer to the minimum pressure turbine stages and the minimum pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some documents, the "low pressure turbine" and the "low pressure compressor" referred to in the present description may alternatively be referred to as "intermediate pressure turbine" and "intermediate pressure compressor". In the case where such alternative nomenclature is used, the fan 23 may be referred to as the first compression stage, or the minimum pressure compression stage.

[0074] The epicyclic gearbox 30 is shown in exemplary detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32, and the ring gear 38 includes teeth around their periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. Four planet gears 32 are illustrated, but the skilled reader will understand that the planet gears 32 may be provided in more or less number within the scope of the claimed invention. Practical applications of an epicyclic planetary gearbox 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. As a 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 being rotatable. In such an arrangement, the fan 23 is driven by the ring gear 38. As a further example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are rotatable.

[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 disclosure. By way of example only, any suitable arrangement may be used for locating the gearbox 30 in the motor 10 and / or for connecting the gearbox 30 to the motor 10. By way of further example, the connections (such as links 36, 40 in the example of [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) may have any desired stiffness or flexibility. As a further example, any suitable arrangement of bearings between rotating and fixed parts of the motor (e.g. between the input and output shafts of the gearbox and the fixed structures, such as the gearbox housing) may be used, and the invention is not limited to the exemplary arrangement of [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 links and the locations of the bearings are typically different from those shown by way of example in [Fig. 2].

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

[0078] Optionally, the gearbox may drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a precompressor).

[0079] Other gas turbine engines to which the present disclosure 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. As a 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 toward and radially outwardly from the reactor nozzle 20. However, this is not limiting, and any aspect of the invention may also be applied 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 referred to as a mixed-flow nozzle. One or both nozzles (mixed or split flow) may have a fixed or variable area.

[0080] The geometry of the gas turbine engine 10, and 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.l]), and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial and circumferential directions are perpendicular to each other.

[0081] [Fig.4] illustrates an example plot of fuel combustion change, AFB, versus jet velocity ratio, Rj, with other factors held constant. The change in fuel combustion contribution from efficiency propulsion 401 is determined relative to an optimum value for the jet velocity ratio of about 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 areas of the exhaust nozzles for the bypass and center exhausts. A lower gear ratio of the gearbox, i.e., a gear ratio of about 3.4 or lower, will tend to give values ​​for the jet velocity ratio of 1.0 or more. To keep the fuel combustion loss to plus or minus about 0.5% or less, it can be seen from [Fig. 4] that the jet velocity ratio should be between about 1.0 and about 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 burn increase to about 0.25 to 0.3%.

[0082] For 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 fuel combustion contribution 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 lower, the fuel combustion contribution by the propulsion efficiency becomes dominant, rising to about 0.7% and above.Another preferred lower limit for the jet velocity ratio of about 0.85 or 0.90 may be used to maintain the fuel combustion contribution by propulsion efficiency at about 0.25% or lower. However, further decreasing the jet velocity ratio allows for a higher gear ratio and / or a lower pressure ratio on the IP turbine, thereby allowing a smaller, faster, and / or lighter IP turbine to be used, resulting in a lower contribution of fuel combustion loss 402 by the IP turbine. A range of about 0.75 to about 0.82 for the jet velocity ratio is therefore advantageous.

[0083] For a given set of gears comprising 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 velocity ratio of about 1.0 and above, and a planetary arrangement for a jet velocity ratio of about 1.0 and below.

[0084] In a general aspect, for this reason, the gas turbine engine may be configured such that the jet velocity ratio is in a range of from about 0.75 to about 1.3 at cruise conditions.

[0085] [Fig. 5] illustrates an exemplary 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 from about 0.75 to about 1.3.

[0086] [Fig. 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 jet velocity in total expansion 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] Parameters that may be adjusted to achieve a jet velocity ratio within the desired range may include LPT vane exit angle, LPT exit angle, and LPT rotational speed.

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

[0089] [Tables 1] Parameter Example 1 (small engine) Example 2 (large engine) Fan diameter (cm) 215 320 Total LPT outlet pressure at maximum flow (kPa) 130 130 Maximum LPT outlet mass flow rate (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 ESS inlet total pressure at maximum flow (kPa) 140 140 ESS inlet mass flow rate (kg / s) 50 100 ESS inlet rotor area (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 LPT outlet pressure at maximum flow, maximum LPT outlet mass flow rate and LPT final rotor area together determine the LPT outlet flow velocity, i.e. the flow velocity at an outlet of the central core. The total ESS inlet pressure at maximum flow, the maximum ESS inlet mass flow rate and the ESS inlet rotor area together determine the velocity at the inlet of the central core. The axial exhaust flow velocity from the bypass exhaust nozzle may be determined, at least in part, by the area of ​​the bypass exhaust nozzle outlet.

[0091] In order to achieve a jet velocity ratio within the desired range, the blower may be provided with features such as a more linear blower root. Compressors, particularly the high pressure compressor, may be provided with features to manage their operability to allow the compressors to operate at a low power requirement to meet the defined ratios, which may, for example, include features such as variable guide vanes. This changes the flow incidence on the vanes and helps maintain an operability margin preventing 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 where mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Claims

1. A 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) that receives an input from the central shaft (26) for driving the fan (23) at a lower rotational speed than the central shaft (26), wherein the method comprises adjusting at least one of the following parameters: LPT blade exit angle, LPT exit angle, and LPT rotational speed, so as to achieve operation of the gas turbine engine (10) providing propulsion with 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 that is greater by a factor of about 2 or more than the jet velocity ratio at maximum takeoff thrust conditions, and in which the jet velocity ratio, Rj, is defined as: [Math.l] R _ VbCb where VB is the first full expansion jet velocity, CB is a thrust coefficient of the bypass exhaust nozzle (18), Vc is the second full expansion jet velocity, Cc is a thrust coefficient of the center core exhaust nozzle (20), qLPT is an isentropic efficiency of a center core minimum pressure turbine (11) and qF is an isentropic efficiency of the fan tip.

2. The method of claim 1, wherein the gearbox has a gear ratio of between about 2.5 and about 5, optionally in the range of 3.2 to 3.

8.

3. A method according to any one of claims 1 to 2, wherein the jet velocity ratio is: in a range of about 0.75 to about 1.3 at cruise conditions; and / or between about 2 and 3 at idle conditions; and / or between about 0.75 and 1.3, optionally between about 0.8 and 1.0, at maximum takeoff thrust conditions.

4. A method according to any one of claims 1 to 3, wherein: the blower (23) has an outer diameter of between about 220 cm and about 290 cm, optionally 230 cm to 280 cm; or the blower (23) has an outer diameter of between about 320 cm and about 420 cm, optionally 330 cm to 380 cm.

5. A method according to any one of claims 1 to 4, wherein the factor is in a range of about 2 to about 3.5, optionally in a range of about 2.1 to about 3.16.

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