GAS TURBINE ENGINE CORE ARRANGEMENT

Abstract

A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan including a plurality of fan blades, wherein the gas turbine engine is configured such that a flow velocity ratio between a first flow velocity at an exit of the engine core and a second flow velocity at an inlet of the engine core is within a range from around 0.82 to around 1.1 at cruise conditions.

Claims

1-10. (canceled)

11. A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein the method comprises operating the gas turbine engine to provide propulsion under cruise conditions such that a flow velocity ratio between a first flow velocity at an exit of the engine core and a second flow velocity at an inlet of the engine core is within a range from 0.82 to 1.1.

12. The method of claim 11, wherein cruise conditions correspond to a forward Mach number of between 0.7 and 0.9 at an altitude of between 10000 m and 15000 m, optionally either: a forward Mach number of 0.85 and international standard atmospheric conditions at an altitude of 35000 ft (10668 m); or a forward Mach number of 0.8 and international standard atmospheric conditions at an altitude of 38000 ft (11582 m).

13. The method of claim 11 wherein the second flow velocity is no greater than Mach 0.7 under cruise conditions.

14. The method of claim 11 wherein a bypass ratio of the engine is in a range from 10 to 20.

15. The method of claim 11 wherein the gas turbine engine comprises a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein, optionally, the gear ratio is in a range from 3.2 to 4.2.

16. The method of claim 11 wherein a pressure ratio between a first pressure at an inlet of the compressor and a second pressure at an outlet of the compressor is in a range from 40 to 70, under maximum take-off conditions.

17. The method of claim 11, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft rotate at a higher rotational speed than the first core shaft.

18. The method of claim 17 wherein a pressure ratio between a first pressure at an inlet of the fan and a second pressure at an outlet of the second compressor is in a range from 40 to 75, under maximum take-off conditions.

Description

[0052] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0053] FIG. 1 is a sectional side view of a gas turbine engine;

[0054] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0055] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine; and

[0056] FIG. 4 is a schematic drawing of an aircraft having a gas turbine engine mounted thereon.

[0057] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0058] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0059] 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 shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0060] Note that the terms low pressure turbine and low pressure compressor as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the low pressure turbine and low pressure compressor referred to herein may alternatively be known as the intermediate pressure turbine and intermediate pressure compressor. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

[0061] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in 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 a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0062] The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0063] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

[0064] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

[0065] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0066] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0067] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0068] Referring to FIGS. 1 and 2, a first flow velocity V.sub.C is defined at an exit or exhaust nozzle 20 of the engine core 11 and a second flow velocity V.sub.A is defined at an inlet of the engine core 11. The first flow velocity V.sub.C is the average axial flow velocity through the core 11 immediately downstream of the final set of rotor blades of the low pressure turbine 19. The second flow velocity V.sub.A is the average axial flow velocity into the core 11 at the inlet of the core 11, i.e. at the front face of the core 11 upstream of the compressor 14, immediately before the ESS 24 (i.e. immediately upstream of the first stator vane in the engine core 11). In a general aspect therefore, the first flow velocity at the exit of the engine core may be defined as an average axial flow velocity through the engine core immediately downstream of a final set of rotor blades of the engine core turbine. The second flow velocity at the inlet of the engine core may be defined as an average axial flow velocity into the engine core at a front face of the engine core upstream of the compressor.

[0069] A preferred range for a flow velocity ratio between the first and second flow velocities is from around 0.82 or 0.85 to around 1.1. This enables a high speed turbine with a high exit flow velocity, which results in an improved efficiency turbine. In combination with this, a low flow velocity into the core allows for a larger lower speed fan (possibly enabled by a gearbox linking the engine core shaft to the fan), which increases propulsive efficiency and reduces fuel burn. Speeding up the flow through the turbine 19 allows the turbine to be made smaller, which improves integration with an aircraft, especially for a larger engine with an increased size of fan. Given that a stress limit on the components making up the turbine is proportional to the area of the outlet 20 and to the square of the rotational speed of the low pressure turbine, the optimum way of improving integration and reducing fuel burn is to increase the exit flow velocity and reduce the area to allow for higher speed flow.

[0070] In preferred embodiments, the flow velocity ratio is between around 0.82 or 0.85 and around 1.1 at cruise conditions, and may be defined as:

[00001]$\mathrm{Core}.Math..Math.\mathrm{Velocity}.Math..Math.\mathrm{Ratio}=\frac{\mathrm{LPT}.Math..Math.\mathrm{Exit}.Math..Math.\mathrm{Axial}.Math..Math.\mathrm{Mn}}{\mathrm{ESS}.Math..Math.\mathrm{Inlet}.Math..Math.\mathrm{Axial}.Math..Math.\mathrm{Mn}}=\frac{\mathrm{LPT}.Math..Math.\mathrm{Exit}.Math..Math.\mathrm{Flow}.Math..Math.\mathrm{Velocity}}{\mathrm{ESS}.Math..Math.\mathrm{Inlet}.Math..Math.\mathrm{Flow}.Math..Math.\mathrm{Velocity}}$

where the LPT Exit Axial Mn is the Mach number of the axial flow at the low pressure turbine exit, i.e. the lowest pressure turbine at the downstream end of the core, the ESS Inlet Axial Mn is the Mach number of the flow into the engine section stator inlet. The ratio between the LPT exit flow velocity and the ESS axial inlet flow velocity will be the same as that for the Mach number ratio. The ESS is the stationary engine supporting structure 24, as for example shown in FIG. 2. The ESS inlet is therefore the inlet to the core immediately upstream of the ESS 24.

[0071] To ensure that the LPT exit axial flow velocity is high, the ESS axial inlet flow velocity is preferably at Mach 0.7 or lower.

[0072] Parameters that may be adjusted to achieve a core velocity ratio within the above range may include the fan blade exit angle, LPT blade exit angle, ESS inlet area, LPT exit area, a ratio of the ESS inlet area to LPT exit area, the fan rotation speed and the LPT rotation speed.

[0073] The following table illustrates example parameters for two engine examples, example 1 being for a relatively small, or lower power, engine and example 2 for a relatively large, or higher power, engine. A small engine may for example have a fan diameter of between around 200 and 280 cm and/or a maximum net thrust of between around 160 and 250 kN or as defined elsewhere herein. A large engine may for example have a fan diameter of between around 310 and 380 cm and/or a maximum net thrust of between around 310 and 450 kN or as defined elsewhere herein.

TABLE-US-00001 Example 1 Example 2 Parameter (small engine) (large engine) Fan diameter (cm) 215 320 LPT Exit Total Pressure at maximum 130 130 flow (kPa) Maximum LPT Exit Mass Flow (kg/s) 50 100 LPT Final Rotor Area (m.sup.2) 0.38 or less, 0.75 or less, for example for example 0.25 to 0.38 0.5 to 0.75 ESS Inlet Total Pressure at maximum 140 140 flow (kPa) ESS Inlet Mass Flow (kg/s) 50 100 ESS Inlet Rotor Area (m.sup.2) 0.275 or 0.55 or greater, greater, for example for example 0.27-0.3 0.55-0.6

[0074] The above parameters relating to LPT exit total pressure at maximum flow, maximum LPT exit mass flow and LPT final rotor area together determine the exit flow velocity of the LPT, i.e. the flow velocity at an exit of the engine core. The ESS inlet total pressure at maximum flow, maximum ESS inlet mass flow and ESS inlet rotor area together determine the velocity at the inlet of the engine 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.

[0075] To reduce the inlet Mach number, the ESS inlet average radius may be increased, which may be done while retaining a given ESS inlet span. A further advantage of this is to create additional space for a gearbox. Alternatively, or additionally, the fan aerodynamic design may be adjusted to reduce the fan root pressure ratio, which has the advantage of improving fan operability. The fan root may be defined as a portion of the fan that drives incoming air into the ESS inlet.

[0076] FIG. 4 illustrates an example aircraft 40 having a gas turbine engine 10 attached to each wing 41a, 41b thereof. When the aircraft is flying under cruise conditions, as defined herein, each gas turbine engine 10 operates such that a flow velocity ratio between a first flow velocity at an exit of the engine core and a second flow velocity at an inlet of the engine core is within a range from around 0.82 to around 1.1.

[0077] It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.