Gas turbine engine with relative timing of bifurcations

Abstract

A gas turbine engine for an aircraft comprises: an air intake comprising a lip, an upstream-most portion of which defines a plane of attack; an engine core comprising a compressor, a combustion chamber, and a turbine coupled to the compressor through a shaft; a fan located upstream of the engine core and adapted to rotate about a main engine axis, the fan comprising a plurality of fan blades having a respective leading edge, a trailing edge, and a tip, a forward-most portion of the leading edge of each fan blade defining a fan inlet plane; an air intake arranged upstream of, and configured to direct air toward, the fan; a plurality of fan outlet guide vanes (FOGVs) arranged downstream of the fan in a bypass duct of the gas turbine engine;and upper and lower bifurcations arranged in the bypass duct and extending along respective radial directions. Figure for abstract: Fig. 5A;

Description

Title of the invention: Gas turbine engine with relative timing of bifurcations

[0001] CROSS REFERENCE TO RELATED APPLICATION

[0002] This represents the first request relating to this subject.

[0003] SCOPE OF APPLICATION

[0004] The present invention relates to gas turbine engines, more particularly to gas turbine engines with improved relative positioning of bifurcations to reduce flow distortions and improve fan efficiency.

[0005] CONTEXT

[0006] Gas turbine engines are used to power aircraft and generally comprise, in series axial flow from front to rear, an air inlet, a fan, one or more compressors, a combustion chamber, one or more turbines, and an exhaust nozzle. Air entering the air inlet is accelerated by the fan to produce two air streams: a first air stream (engine block stream) in the compressor and a second air stream (bypass stream) which passes through a bypass duct to provide propulsive thrust. The air entering the compressor is compressed, mixed with fuel and then introduced into the combustion chamber, where combustion of the air / fuel mixture occurs.The high-temperature, high-energy exhaust fluids are then introduced into the turbine, where the fluid energy is converted into mechanical energy to drive the rotating compressor by a suitable interconnecting shaft.

[0007] The fan includes a plurality of fan blades configured to rotate about a central axis of the gas turbine engine. The fan receives and pressurizes air entering the air intake to produce thrust.

[0008] It is desirable to increase fan efficiency to increase thrust while reducing specific fuel consumption (SFC). It is known that fan efficiency depends, at least in part, on the circumferential flow profile at, or near, the trailing edge of the fan blades, also known as the intra-stage flow profile. Flow distortions, upstream and downstream of the fan, generally influence the intra-stage static pressure profile. For example, flow distortions are generally introduced into the fan system upstream through the air intake and downstream through the upper and lower branches in the bypass duct. Due to their geometric shape and position in the bypass duct, the upper and lower branches generate large potential fields, which move upstream toward the FOGVs and the fan, thus generating significant flow distortions. In particular, the flow distortions generated by the bifurcations interact with the aerodynamics of fan outlet guide vanes (FOGVs) and produce blocking effects, which reduce the flow diffusion efficiency of the FOGVs. The increase in fan and FOGV flow losses reduces the overall aerodynamic performance of the fan; which results in an increase in the engine SFC.

[0009] Furthermore, the upstream and downstream distortions generate a variable unstable pressure distribution on the fan blades which in turn causes the variation in modal force magnitude and phase, which can lead to increased vibration stresses, increased HCF (high cycle fatigue) endurance and mechanical integrity problems which will reduce the life of the fan blades.

[0010] To mitigate the blocking effect of the upper and lower bifurcations it is known to design the FOGVs to divert the flow in the bypass duct around the bifurcations, thereby mitigating the blockage, reducing the flow distortions produced by the bifurcations, and thereby increasing the static pressure profile uniformity at the trailing edge of the fan blades.

[0011] The present invention aims to reduce flow distortions and increase blower efficiency, by providing an alternative and more efficient solution for dealing with flow distortions, in particular by dealing with flow distortions produced by bifurcations downstream of the blower.

[0012] SUMMARY

[0013] Thus, there is provided a gas turbine engine for an aircraft according to claim 1.

[0014] According to one aspect there is provided a gas turbine engine for an aircraft comprising:an air intake comprising a lip, an upstream-most portion of which defines a plane of attack; an engine core comprising a compressor, a combustion chamber, and a turbine coupled to the compressor through a shaft; a fan located upstream of the engine core and adapted to rotate about a main engine axis, the fan comprising a plurality of fan blades having a respective leading edge, a trailing edge, and a tip, a forward-most portion of the leading edge of each fan blade defining a fan inlet plane; an air intake arranged upstream and configured to direct air toward the fan; a plurality of fan outlet guide vanes (FOGVs) arranged downstream of the fan in a bypass duct of the gas turbine engine;upper and lower bifurcations arranged in the bypass duct and extending along respective radial directions, wherein the lower bifurcation is arranged at a circumferential position corresponding to a; highest fan blade load area caused by flow distortions introduced by the air intake under flight conditions.

[0015] In the present description, upstream and downstream are defined relative to the axial flow direction through the fan; and forward and aft are defined relative to the gas turbine engine, i.e., the fan being at the front and the turbine at the rear of the engine.

[0016] In this description, the upper and lower bifurcations, or splitters, in the fan bypass duct are defined as the aerodynamically shaped fairings covering various structural connections, air, oil and fuel service lines and electrical harnesses, etc. although they do not support structural loads and contribute to flow turning and lift generation. The upper and lower bifurcations have a maximum thickness that is at least two and a half fan outlet guide vane (FOGV) pitches at 50% span of the leading edge of the FOGVs, with 0% span corresponding to a radially innermost portion of the FOGVs and 100% corresponding to a radially outermost portion of the FOGVs. The pitch is defined as the tangential distance between the leading edges of two adjacent FOGVs at 50% span.When the tangential distances between adjacent FOGVs are not identical, an average distance is considered. In other words, the upper bifurcation at its maximum thickness span, viewed from the rear of the engine, is perceived as “covering” at least two and a half FOGVs; similarly, the lower bifurcation at its maximum thickness span, viewed from the rear of the engine, is perceived as “covering” at least two and a half FOGVs. The upper bifurcation and the lower bifurcation do not have to have the same maximum thickness and therefore may “cover” a different number of FOGVs. For example, the upper bifurcation may be thicker than the lower bifurcation and may cover a greater number of FOGVs, e.g., at least three, four, or five FOGVs.In other embodiments, the lower bifurcation may be thicker than the upper bifurcation and may span a greater number of FOGVs, e.g., at least three, four, or five FOGVs.

[0017] The present disclosure is based at least in part on the observation that flow distortions generated by the lower bifurcation and moving upstream toward the fan can be used to compensate for downstream moving distortions caused by the intake.

[0018] The present inventor has found that, for example in an underwing engine configuration, the wing induces an upward deflecting flow entering the intake, which in turn produces circumferential 1EO (first engine control) flow distortions upstream of the fan.

[0019] Downstream of the FOGVs, the large upper and lower bifurcation potential fields generate circumferential 2EO (second engine control) flow distortions that are largely in phase with the residual intake distortion.

[0020] In underwing mounted engines, wing upwash occurs under all nominal flight operating conditions, i.e., during all high forward speed flight conditions between maximum ground takeoff thrust and ground landing. Under all high forward speed conditions, the airflow tube that contains the air entering the air intake is well aligned with the air intake and, therefore, the flow distortions upstream of the fan caused by wing upwash can be considered constant. Furthermore, the present inventor has found that under flight conditions, the flow distortion upstream of the fan caused by wing upwash is not affected by the crosswind, which plays a secondary role in generating flow distortions in the airflow tube upstream of the fan.Therefore, for all practical design purposes, the crosswind and its effect on the flow distortion just upstream of the fan under flight conditions can be ignored, and the inlet upwash distortion profile and its relative circumferential position to the bifurcations can be considered fixed under flight conditions and determined by the wing upwash only.

[0021] Due to the wing updraft under all flight conditions, the fan experiences non-uniform loading, in particular the lowest load (and thus the highest efficiency) at top dead center (TDC), 0°, or 12 o'clock position, and the highest load (and thus the lowest efficiency) at bottom dead center (BDC), 180°, or 6 o'clock position. In fact, the present inventor has found that the lowest load position is not exactly at TDC due to fan rotation: for a clockwise rotating fan when viewed from the front, the lowest load area is located before (clockwise) TDC, at about 11 o'clock, or -30°; similarly, the highest load position is located before BDC, at about 5 o'clock, or 150°.Conversely, for a counterclockwise rotating fan when viewed from the front, the lowest load position is located after (still clockwise) TDC, at approximately 1 o'clock, or +30°; similarly, the highest load position is located after BDC, at approximately 7 o'clock, or 210°.

[0022] In detail, upstream flow distortions affect the fan aerodynamics of an axial fan operating under asymmetric flow conditions. without any mitigation measures in place. A fan blade rotating through flow distortions experiences a variation in effective blade velocity between the left and right halves of the annulus, which causes the pressure ratio across the fan blades to vary as well as the magnitude and direction of the relative inlet velocity at the fan blade leading edge. The variation in magnitude and direction of the relative velocity causes a variation in the angle of attack, which in turn leads to a variation in blade lift and thus the variation in the work done by the fan blades in different parts of the annulus.

[0023] It is important to note that the annular, or circumferential, positions of the lowest fan blade loading (corresponding to a static pressure peak and a minimum effective fan blade speed) and the highest fan blade loading (corresponding to a static pressure minimum and a peak effective fan blade speed) depend on the direction of rotation of the fan, but not on the engine operating mode, or the aircraft on which the engine is mounted: for example, the circumferential positions of the highest and lowest fan blade loadings are independent of the fan rotation speed, aircraft speed and altitude, ambient temperature and pressure, and / or flight conditions.The annular positions of the peak and minimum static pressure may depend on the air intake geometry, for example, the lip gap angle and the lip thickness. In general, the peak and minimum static pressure are located at opposite circumferential positions. Conversely, the static pressure distortion force may vary depending on operational parameters of the aircraft on which the engine is mounted: for example, at cruise conditions, the updraft angle is limited to about 4° and the static pressure distortion force is therefore lower than at the end-of-runway (EOR) condition, in which the updraft angle can reach 15° to 18°.

[0024] While the annular position of the upper bifurcation is essentially determined by the pylon on which the engine is mounted, the present inventor has found that the lower bifurcation can be arranged in a circumferential position corresponding to the position of highest fan blade loading (or highest effective blade velocity) caused by flow distortions introduced by the air intake, such that upstream moving distortions caused by the lower bifurcation can pass through the FOGVs and counteract the intake distortions, thereby increasing fan blade loading uniformity. In other words, the lower bifurcation produces a circumferential static pressure variation of 1EO at the fan blade trailing edge having a static pressure phase essentially opposite to that caused by the intake distortions.This intra-stage flow field, . produced by means of the upstream movement of the distortions induced by the lower bifurcation, counteracts the intake distortions, generates a favorable pressure ratio for all fan blades, thereby increasing the fan blade efficiency at all annular positions. This increases the overall fan rotor and fan stage efficiency and reduces the forced 1EO fan blade response caused by asymmetric flow.

[0025] To maximize the counterbalancing of intake distortions at the highest fan blade load annular position by means of upstream moving lower bifurcation induced distortions, FOGVs may be designed to pass lower bifurcation induced distortions of controlled magnitude. For this purpose, any suitable cyclic variation of FOGV inlet and outlet angles is used by optimizing FOGV offset and camber angles, as well as the number of FOGV types.

[0026] The present invention is particularly advantageous for gas turbine engines with a short intake, since the shorter the intake, the greater the distortions caused by the intake. Air intakes are generally characterized by the ratio L / D, in which L is the intake length defined as an axial distance along the main engine axis between a plane of attack and a fan inlet plane.

[0027] The intake comprises a lip, an upstream portion of which defines the plane of attack.

[0028] The fan blades have a respective leading edge, a trailing edge, and a tip, a forward-most portion of the leading edge of the tip of each fan blade defining the fan inlet plane. In embodiments, the air intake may have an L / D ratio of less than 0.70, preferably less than 0.50, more preferably less than 0.45, even more preferably less than 0.40. The L / D ratio may be greater than 0.25, preferably greater than 0.30. For example, the L / D ratio may be in a range of 0.25 to 0.70, or in a range of 0.25 to 0.60, or in a range of 0.25 to 0.50, or in a range of 0.25 to 0.45.

[0029] In embodiments, the gas turbine engine may be an underwing mounted, clockwise rotating gas turbine engine and the lower bifurcation may be arranged at a rotation angle relative to the upper bifurcation in a range of 120° to 170°, preferably in the range of 130° to 165°, preferably in the range of 135° to 165°, preferably in the range of 140° to 160°.

[0030] In embodiments, the gas turbine engine may be an underwing mounted, counterclockwise rotating gas turbine engine and the lower bifurcation may be arranged at a rotation angle relative at the upper bifurcation in a range of 190° to 240°, preferably in the range of 190° to 230°, preferably in the range of 195° to 225°, preferably in the range of 200° to 220°.

[0031] In embodiments, the gas turbine engine may be a fuselage-mounted gas turbine engine, or a wing-mounted engine.

[0032] In embodiments, the gas turbine engine may have more than two bypass branches having different designs, shapes, and sizes.

[0033] In embodiments, the gas turbine engine may have more than two branches into core compressor ducts having different designs, shapes, and sizes.

[0034] The upper bifurcation may be arranged in a circumferential position corresponding to top dead center (TDC).

[0035] The blower may have a blower diameter greater than 210 cm, preferably greater than 215 cm, preferably greater than 220 cm.

[0036] The blower may have a blower diameter of less than 400 cm, preferably less than 380 cm, preferably less than 360 cm.

[0037] The blower may have a blower diameter greater than 120 cm, preferably greater than 130 cm, preferably greater than 140 cm.

[0038] The blower may have a blower diameter of less than 160 cm, preferably less than 150 cm, preferably less than 145 cm.

[0039] The gas turbine engine may include a reduction gear configured to receive an input from the shaft and output a drive to the fan, so as to drive the fan at a speed lower than that of the shaft.

[0040] The reducer may have a reducer ratio greater than 3.0, preferably greater than 3.1, preferably greater than 3.2.

[0041] The turbine may be a first turbine, the compressor may be a first compressor, and the shaft may be a first shaft. The engine core may further comprise a second turbine, a second compressor, and a second shaft connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second shaft may be configured to rotate at a higher rotational speed than the first core shaft.

[0042] The lower bifurcation may be arranged at a circumferential position corresponding to a fan blade highest loading position caused by flow distortions introduced by the air intake at cruise conditions.

[0043] Arrangements of the present disclosure may be particularly, although not exclusively, advantageous for fans that are driven via a reduction gear. Accordingly, the gas turbine engine may include a reduction gear that receives an input from the core shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the reducer may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear. The core shaft may connect the turbine and compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0044] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, e.g., one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

[0045] 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.

[0046] The reducer may be arranged to be driven by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above). For example, the reducer may be arranged to be driven only by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the reducer may be arranged to be driven by any shaft or shafts, e.g. the first and / or second shafts in the example above.

[0047] The reducer may be a reduction gearbox (in that the output to the blower has a lower rotational speed than the input from the core shaft). Any type of reducer may be used. For example, the reducer may be a "planetary" or "star" reducer, as described in more detail elsewhere in this document. The reducer 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 3 to 4.2 or 3.2 to 3.8, for example of 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 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 arrangements, the gear ratio may be outside these ranges, for example the reducer may be a "compound star" having a ratio greater than 4.2.

[0048] 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 of (e.g., at the outlet of) the second compressor, when a second compressor is provided. As a further example, the outlet flow to the combustor may be provided at the inlet of the second turbine, when a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0049] The or each compressor (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 vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0050] The or each turbine (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 vanes. The respective rows of rotor blades and stator vanes may be axially offset from each other. In each stage, the row of rotor blades may be arranged downstream of the respective row of stator vanes. Each fan blade may be defined as having a radial throw extending from a root (or hub) at a radially inner gas-washed location, or a 0% throw position, to a tip at a 100% throw position.The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or in the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the 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 of 0.28 to 0.32. These ratios can be commonly referred to as the hub to tip ratio. The radius at the . Both the hub and tip radius can be measured at the leading edge (or axially forward-most) 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.

[0051] The fan radius can be measured between the centerline of the engine 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) any of: 200 cm, 210 cm, 215 cm, 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 (approximately 160 inches) or 420 cm (approximately 165 inches). The blower diameter can be less than 420 cm (approximately 165 inches), 410 cm (approximately 160 inches), 400 cm, 390 cm (approximately 155 inches) or 380 cm (approximately 150 inches).The diameter of the blower may be within an inclusive range bounded by any two of the values ​​in the preceding two sentences (i.e., the values ​​may form upper or lower limits), for example, within the range 200 cm to 210 cm, or 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm.

[0052] The rotational speed of the blower may vary during use. Generally, the rotational speed is lower for blowers with a larger diameter. Strictly by way of non-limiting example, the rotational speed of the blower at cruising conditions may be less than 2,600 rpm, for example less than 2,500 rpm, or less than 2,300 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 200 cm to 300 cm (e.g. 210 cm to 290 cm, or 240 cm to 280 cm, or 250 cm to 270 cm) may be in the range of 1700 rpm to 2600 rpm, e.g. in the range of 1800 rpm to 2300 rpm, e.g. in the range of 1900 rpm to 2100 rpm.By way of non-limiting example only, the rotational speed of the fan under cruising conditions for an engine having a fan diameter in the range of 330 cm to 380 cm may be in the range of 1200 rpm to 2000 rpm, e.g. in the range of 1300 rpm to 1800 rpm, e.g. in the range of 1400 rpm to 1800 rpm.

[0053] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results into a displacement of the fan blade tip with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy increase dH of the flow. A fan tip load can be defined as dH / Utip2, 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 the fan tip radius at the leading edge times the angular velocity). The peak fan load at cruise conditions may be greater than (or of the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values ​​being dimensionless).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, within the range 0.28 to 0.31 or 0.29 to 0.3.

[0054] Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), 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. In some arrangements the bypass ratio at cruise conditions may be greater than (or of the order of) any of the following: 10, 10.5, 11, 11.1, 11.2, 11.3, 11.4, 11.5, 11.6, 11.7, 11.8, 11.9, 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 at cruise conditions 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, within the range from 10 to 16, or 12 to 16, or 13 to 15, or 10 to 12, or 12 to 14, or 13 to 14.The bypass duct may be substantially annular. The bypass duct may be radially outside the engine block. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0055] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the 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 conditions may be greater than (or in the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., the values ​​can form upper or lower limits), for example in the range 50 to 70.

[0056] 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. At cruise 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: 110 Nkg 's, 105 Nkg 's, 100 Nkg 's, 99 Nkg 's, 98 Nkg 's, 97 Nkg 's, 96 Nkg 's, 95 Nkg 's, 94 Nkg 's, 93 Nkg 's, 92 Nkg 's, 91 Nkg 's, 90 Nkg 's, 85 Nkg 's or 80 Nkg 's. The specific thrust 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 of 80 Nkg's to 100 Nkg's, or 85 Nkg's to 100 Nkg's, or 92 Nkg's to 98 Nkg's, or 85 Nkg's to 95 Nkg's. Such engines may be particularly efficient compared to conventional gas turbine engines.

[0057] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any of the following: 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 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 130 kN to 145 kN, or 155 kN to 170 kN, or 330 kN to 420 kN, or 350 kN to 400 kN.The above mentioned thrust may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

[0058] In use, the flow temperature at the inlet of the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself may be referred to as the nozzle guide vane. At 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, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K or 1650 K. The TET at cruising conditions 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, 1530 K to 1600 K. The maximum TET in engine use may be, for example, at least (or in the order of) any one 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 1700 k to 1800 k, or 1800 K to 1950 K, or 1900 K to 2000 K. The maximum TET may occur, for example, in a condition high thrust, for example in a maximum takeoff thrust (MPT) condition.

[0059] A fan blade and / or airfoil portion of a fan blade 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, 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 an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions fabricated using different materials.For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to withstand impact (e.g., by birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade may have a carbon fiber or aluminum-based (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0060] A blower as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be connected to the central portion in any desired manner. For example, each fan blade may comprise a fastening element which may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail which may notch into and / or engage a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be integrally formed a central portion. Such an arrangement may be referred to as a bladed disc or 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 joined to the hub / disc by welding, such as a linear friction stir weld.

[0061] 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 area of ​​the bypass duct to be varied during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0062] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

[0063] As used herein, cruise conditions have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, those skilled in the art would 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 to which the gas turbine engine is designed to be attached. In this sense, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the beginning of descent has been burned (which may be approximated by 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 that provides thrust that would ensure steady-state operation (i.e., maintaining a constant altitude and a constant Mach number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided on that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half the total thrust that would be required for steady-state operation of that aircraft at mid-cruise.

[0064] In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady-state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the standard atmosphere international standard 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 operating point of the engine at cruise conditions is clearly defined.

[0065] Strictly by way of example, the forward speed at the cruise condition 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 condition. For a certain aircraft, the cruise conditions may be outside these ranges, e.g., below Mach 0.7 or above Mach 0.9.

[0066] Strictly by way of example, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range from 10,000 m to 15,000 m, for example in the range from 10,000 m to 12,000 m, for example in the range from 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range from 10,500 m to 11,500 m, for example in the range from 10,600 m to 11,400 m, for example in the range from 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range from 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

[0067] Strictly by way of example, cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of 20 kN to 40 kN.

[0068] Strictly by way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is of course dependent on the engine and its intended application and may be, for example, a value ranging from 35 kN to 65 kN.

[0069] In use, a gas turbine engine described and / or claimed herein may operate at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0070] 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 to which the gas turbine engine has been designed to be attached. Thus, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

[0071] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be at cruise conditions as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).

[0072] 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) mid-cruise operation of the aircraft, as defined elsewhere herein.

[0073] Those skilled in the art would 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 in association with any other feature or parameter described herein. Brief description of the drawings

[0074] Embodiments will now be described by way of example only, with reference to the accompanying drawings, in which:

[0075] [Fig.l] is a sectional side view of a gas turbine engine;

[0076] [Fig.2] is a close-up sectional side view of an upstream portion of a motor gas turbine;

[0077] [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine;

[0078] [Fig.4] is a sectional side view of a gas turbine engine in a typical under / over wing mounting configuration;

[0079] Figures 5a and 5b are a schematic front view of the gas turbine engine of [Fig.4] in an underwing configuration, showing the circumferential positions upper and lower bifurcation with the blower rotating clockwise and counterclockwise, respectively;

[0080] [Fig.6] is a schematic graph showing the circumferential variations of normalized static pressure at different axial positions and the fan blade loading of the gas turbine engine of [Fig.4]. DETAILED DESCRIPTION

[0081] [Fig.l] illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 comprises an air intake 12 and a propulsion fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 which receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the duct 22 and the bypass exhaust nozzle 18. bypass 22. The blower 23 is attached to and driven by the low pressure turbine 19 by means of a first shaft 26 and an epicyclic reduction gear 30.

[0082] 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 takes place. 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 some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a second shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic reduction gear 30 is a reduction gearbox.

[0083] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in [Fig. 2]. The low-pressure turbine 19 (see [Fig. 1]) drives the first shaft 26, which is coupled to a sun wheel, 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 forces the planet gears 32 to change orientation about the sun gear 28 in synchronism while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via of links 36 to the fan 23 in order to cause its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and meshing therewith, there is a ring or crown gear 38 which is coupled, via links 40, to a stationary support structure 24.

[0084] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the first shaft 26 with the lowest rotational speed in the engine (i.e., not including the reduction gear output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as an "intermediate pressure turbine" and "intermediate pressure compressor".When such alternative nomenclature is used, the blower 23 may be designated as the first compression stage or the lower pressure compression stage.

[0085] The epicyclic reduction gear 30 is shown by way of example in more detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic planetary reduction gear 30 generally include at least three planet gears 32.

[0086] The epicyclic reduction gear 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, with the ring gear 38 fixed. However, any other suitable type of epicyclic reduction gear 30 may be used. As a further example, the epicyclic reduction gear 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38.As an alternative example, the epicyclic reducer 30 may be a two-stage star arrangement, in which the first-stage sun gear is coupled to the first shaft 26, the second-stage sun gear is coupled to the first-stage ring gear, and the fan is coupled to the second-stage ring gear via suitable linkages. As another alternative example, the reducer 30 may be a reducer . differential in which the crown gear 38 and the planet carrier s 34 are both allowed to rotate.

[0087] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the reducer 30 to the motor 10. As a further example, the connections (such as the links 36, 40 in the example of [Fig. 2]) between the reducer 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 degree of rigidity or flexibility.As a further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g., between input and output shafts from the reducer and fixed structures, such as the reducer housing) may be used, and the description is not limited to the exemplary arrangement of [Fig. 2]. For example, where the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of the output and support links and bearing locations would typically be different from that shown as an example in [Fig. 2].

[0088] Thus, the present disclosure extends to a gas turbine engine having any arrangement of reduction gear styles (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing locations.

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

[0090] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have another number of compressors and / or turbines and / or another number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.l] has a split-flow nozzle 18, 20 which means that the flow through the bypass duct 22 has its own nozzle 18 which is independent of, and radially outward of, the engine block nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed-flow nozzle.The one or both nozzles (whether mixed or split flow) may have a fixed or variable area. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[0091] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the main rotation axis 9), a radial direction (in the bottom-up direction in [Fig.l]), and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial, and circumferential directions are mutually perpendicular.

[0092] [Fig.4] illustrates the exemplary gas turbine engine 10 of [Fig.l] with an inclined air inlet 12 in an under / over wing mounting configuration under flight conditions. The fan 23 includes a plurality of fan blades 43, each fan blade 43 having a leading edge 44, a trailing edge 45, and a tip 46. A forwardmost portion of the leading edge of the fan blades 43 defines a fan inlet plane at an axial position identified by the letter E. The rearmost portion of the leading edge of the fan blades 43 defines a fan outlet plane at an axial position identified by the letter F. The wing (not shown) produces an upwash flow entering the intake 12, which in turn produces circumferential first engine control flow distortions (1EO) at the fan inlet plane.As explained previously, the crosswind effect on the flow distortions at the fan inlet plane under flight conditions is negligible and can be ignored.

[0093] In other words, each of the blades 43 of the fan 23 experiences a non-uniform load as it passes through the circumferential flow distortions of 1EO produced by the inlet 12, in particular a lower load in the upper half of the ring and a higher load in the lower half of the ring in [Fig.4].

[0094] Furthermore, the fan blade 43 rotating through the flow distortions caused by the wing updraft experiences a variation in the effective blade velocity between different parts of the ring, particularly between the left and right halves of the ring (when viewed from the front of the engine), which in turn causes a variation in the total pressure ratio and the magnitude and direction of relative inlet velocity, and ultimately the variation in work performed by the fan blade 43 in different parts of the ring. Accordingly, for the gas turbine engine 10 with the fan 23 rotating clockwise when viewed from the front, the highest load annular position is arranged before (clockwise) the bottom dead center (or 180°), in a circumferential position corresponding to a range of 120° to 170°.

[0095] For gas turbine engines with the fan rotating counterclockwise when viewed from the front, the load positions lowest and highest are arranged after (always clockwise) top dead center (or 0°) and bottom dead center, respectively. In these embodiments, the position of the highest load may be arranged in a circumferential position corresponding to a range of 190° to 240°, or 190° to 230°, or 195° to 225°, or 200° to 220°, for example at about 210°.

[0096] The gas turbine engine 10 comprises a row of fan outlet guide vanes (FOGVs) 47 arranged downstream of the fan 23 in the bypass duct 22. The FOGVs 47 each comprise a leading edge 48 and a trailing edge 49.

[0097] Downstream of the FOGVs 47, the gas turbine engine 10 includes an upper bifurcation 50 and a lower bifurcation 51. In the embodiment of Figure 5a, the gas turbine engine 10 is illustrated in an underwing mounting configuration and the upper bifurcation 50 is arranged at top dead center and proximate to a wing 52. In alternative embodiments not illustrated, the upper bifurcation may be arranged differently, depending on the structural mounting arrangement of the engine on the wing or on the fuselage.

[0098] The lower bifurcation 51 is arranged distal to the wing 52.

[0099] The lower bifurcation 51 is arranged in rotation relative to the bottom dead center, in a circumferential position corresponding to the circumferential position of the highest fan blade loading caused by flow distortions introduced by the air intake under flight conditions, for example at cruise conditions.

[0100] In other words, the lower bifurcation 51 is arranged radially along a direction 53 at a rotation angle α relative to a direction 54 defined by the upper bifurcation 50. In the illustrated embodiment, the direction 54 is vertical and passes through the top dead center and the bottom dead center. In the embodiment illustrated in Figure 5a, the fan rotates in a clockwise direction 60 when viewed from the front and therefore the lower bifurcation 51 is arranged in a circumferential position corresponding to a rotation angle α in the range of 120° to 170°.In embodiments, the lower bifurcation 51 is arranged in a circumferential position corresponding to a rotation angle α in the range of 130° to 165°, or in a range of 135° to 165°, or in a range of 140° to 160°, for example at about 150°, to correspond to the annular position of the highest fan blade load.

[0101] Figure 5b illustrates an embodiment which differs from the embodiment of Figure 5a by the direction of rotation 60' of the fan 23 and the position of the lower bifurcation 51'. In detail, the fan rotates counterclockwise when viewed from the front and the lower bifurcation 51' is arranged in a circumferential position corresponding to a rotation angle a in the range of 190° to 240°, or 190° to 230°, or 195° to 225°, or 200° to 220°, for example at about 210°, corresponding to the annular position of the highest load position of fan blades 43.

[0102] In other embodiments not shown where the gas turbine engine is mounted on the fuselage, the upper bifurcation (i.e., the bifurcation near the fuselage) does not extend along a vertical direction; however, the rotation angle α at which the lower bifurcation (i.e., the bifurcation at a position distal to the fuselage) is arranged relative to the upper bifurcation is within the same ranges as described above to correspond to the annular position of the highest fan blade loading.

[0103] The upper and lower bifurcations 50, 51, 51' produce distortions by moving upstream toward the fan 23. In embodiments, the FOGVs may be designed to promote the passage of pressure distortions produced by the lower bifurcation 51, 51'. To this end, the FOGVs 47 may have appropriate offset and camber angles, as well as a number of vane types to achieve a 0-pk amplitude of lower bifurcation distortions with an opposite phase of 0-pk amplitude of intake distortions at the fan vane trailing edge 45.

[0104] In embodiments, the FOGVs 47 may be designed to mask, e.g., entirely mask, the pressure distortions produced by the upper bifurcation 50, using any type of FOGV, and any suitable offset and camber angles.

[0105] This produces a circumferential static pressure variation of 1EO at the fan outlet plane having a pressure distortion phase opposite, or approximately opposite, that of the intake distortion. This intentionally produced pressure field at the fan outlet effectively counteracts the intake distortions; thereby recovering a significant portion of fan stage efficiency and reducing the forced 1EO fan blade response caused by asymmetric flows.

[0106] This is illustrated in [Fig. 6], wherein, for the clockwise rotating gas turbine engine 10 of Fig. 5a, the normalized static pressure resulting from the upstream moving distortions introduced by the bifurcations 50, 51 at the fan outlet plane (position F in [Fig. 4]) is illustrated by the dotted line 110, the normalized static pressure from the downstream moving inlet distortions at the fan inlet plane (position E in [Fig. 4]) is illustrated by the dotted line 120, and the effective circumferential fan blade loading variation resulting from the combination of the two within the fan 23 is illustrated by solid line 130. All normalized static pressures and fan blade loading are illustrated as a function of circumferential position.

[0107] The effective fan blade loading is more uniform along the annulus than the fan blade loading produced by the intake distortions alone, thereby increasing fan efficiency and reducing the forced 1EO fan blade response and vibrational stresses.

[0108] 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. The scope of protection is defined in the appended claims.

Claims

Claims

1. A gas turbine engine for an aircraft comprising: an air intake comprising a lip, an upstream-most portion of which defines a plane of attack, an engine core comprising a compressor, a combustion chamber, and a turbine coupled to the compressor through a shaft; a fan located upstream of the engine core and adapted to rotate about a main engine axis, the fan comprising a plurality of fan blades having a respective leading edge, a trailing edge, and a tip, a forward-most portion of the leading edge of each fan blade defining a fan inlet plane; an air intake arranged upstream of, and configured to direct air toward, the fan; a plurality of fan outlet guide vanes (FOGVs) arranged downstream of the fan in a bypass duct of the gas turbine engine;and upper and lower bifurcations arranged in the bypass duct and extending along respective radial directions; wherein the lower bifurcation is arranged at a circumferential position corresponding to a fan blade highest loading position caused by flow distortions introduced by the air intake under flight conditions.;

2. A gas turbine engine according to claim 1, wherein the fan has a fan diameter D, and the air inlet has a length L measured along the main engine axis between the leading plane and the fan inlet plane, and wherein the L / D ratio is less than 0.

70.

3. A gas turbine engine according to claim 2, wherein the L / D ratio is less than 0.

50.

4. A gas turbine engine according to claim 2, wherein the L / D ratio is less than 0.

45.

5. A gas turbine engine according to claim 2, wherein the L / D ratio is less than 0.

40.

6. A gas turbine engine according to claim 2, wherein the L / D ratio is greater than 0.25, preferably greater than 0.

30.

7. A gas turbine engine according to any preceding claim, wherein the gas turbine engine is an underwing gas turbine engine and has a clockwise rotating fan, the lower bifurcation being arranged at a rotation angle α relative to the upper bifurcation 50 in a range of 120° to 170°, preferably in the range of 130° to 170°.

8. A gas turbine engine according to claim 7, wherein the rotation angle α is in a range of 130° to 165°, preferably 135° to 165°, more preferably in a range of 140° to 160°.

9. A gas turbine engine according to any one of claims 1 to 6, wherein the gas turbine engine is an underwing gas turbine engine and has a counterclockwise rotating fan, the lower bifurcation being arranged at a rotation angle α relative to the upper bifurcation 50 in a range of 190° to 240°, preferably in the range of 190° to 230°.

10. A gas turbine engine according to claim 9, wherein the rotation angle α is in a range of 195° to 225°, or preferably in the range of 200° to 220°.

11. A gas turbine engine according to any preceding claim, wherein the upper bifurcation is arranged in a circumferential position corresponding to top dead center (TDC).

12. A gas turbine engine according to any preceding claim, wherein the fan has a fan diameter greater than 120 cm, preferably greater than 130 cm, preferably greater than 140 cm.

13. A gas turbine engine according to any preceding claim, wherein the fan has a fan diameter of less than 400 cm, preferably less than 380 cm, preferably less than 360 cm.

14. A gas turbine engine according to any preceding claim, comprising a reduction gear adapted to receive an input from the shaft and output a drive to the fan, so as to drive the fan at a speed lower than that of the shaft.

15. A gas turbine engine according to claim 11, wherein the reduction gear has a reduction ratio greater than 3.0, preferably greater than 3.1, preferably greater than 3.

2.

16. A gas turbine engine according to any preceding claim, wherein the reduction gear has a reduction ratio greater than 3.

3.

17. A gas turbine engine according to any preceding claim, wherein the turbine is a first turbine, the compressor is a first compressor, and the shaft is a first shaft; the engine core further comprises a second turbine, a second compressor, and a second shaft connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second shaft are arranged to rotate at a rotational speed greater than that of the first core shaft.

18. A gas turbine engine according to any preceding claim, wherein the lower bifurcation is arranged at a circumferential position corresponding to a fan blade highest loading position caused by flow distortions introduced by the air intake at cruise conditions.