Pressure ratio distributions for a gas turbine engine

Abstract

A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P.sub.104), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Claims

1. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine, core; the fan comprises a hub and a plurality of fan blades, each fan blade having a radial blade span extending from a root to a tip, and the plurality of fan blades being formed integrally with the hub such that the plurality of fan blades are fixed relative to the hub; a gearbox that receives an input from the core shat and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and an annular splitter at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct, wherein: the ratio of the radius of each fan blade at the root (r.sub.root) to the radius of each fan blade at the tip (r.sub.tip) is less than 0.33: a fan root to tip pressure ratio of the gas turbine engine, defined as the ratio of the mean total pressure of flow at a fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of flow at the fan exit that subsequently flows through the bypass duct (P.sub.104), is in the range of from 0.83 to 0.9 at cruise conditions; and a fan root pressure ratio of the gas turbine engine, defined as the ratio of the mean total pressure of flow at the fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of flow at a fan inlet (P.sub.100), is in the range of from 1.15 to 1.25 at the cruise conditions; a fan tip loading is defined as dH/U.sub.tip.sup.2 where dH is the enthalpy rise across the fan and U.sub.tip is the velocity of the fan at the tip of each fan blade, and the fan tip loading at the cruise conditions is greater than 0.3 JKg.sup.−1/(ms.sup.−1).sup.2.

2. A gas turbine engine according to claim 1, wherein the fan root to tip pressure ratio is less than 0.88.

3. A gas turbine engine according to claim 1, wherein the fan tip loading at the cruise conditions is in the range of from 0.3 to 0.4 JKg.sup.−1/(ms.sup.−1).sup.2.

4. A gas turbine engine according to claim 1, wherein a bypass ratio is defined as the ratio of the mass flow rate of the bypass flow (B) to the mass flow rate the core flow (A) at the cruise conditions, and the bypass ratio is greater than 10.

5. A gas turbine engine according to claim 1, wherein the specific thrust at the cruise conditions is less than 100 Nkg.sup.−1 s.

6. A gas turbine engine according to claim 1, 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 are arranged to rotate at a higher rotational speed than the first core shaft.

7. A gas turbine engine according to claim 1, wherein: a combustor is provided downstream of the fan and the compressor and upstream of the turbine; an overall pressure ratio is defined as the ratio of the mean total pressure of flow at an inlet to the combustor (P.sub.106) to the mean total pressure of flow at the fan inlet (P.sub.100); and the ratio of the fan root pressure ratio to the overall pressure ratio is less than 0.04 at the cruise conditions.

8. A gas turbine engine according to claim 7, wherein the ratio of the fan root pressure ratio to the overall pressure ratio is less than 0.03 at the cruise conditions.

9. A gas turbine engine according to claim 1, wherein: the radial blade span of each fan blade extends from the root at a 0% span position to the tip at a 100% span position; and the average camber (α.sub.2−α.sub.1).sub.ave of each fan blade taken over the radially innermost 10% of the blade span is no greater than 75% of the average camber of the 10% blade span portion that has the maximum average camber (θ.sub.2−θ.sub.1).sub.max.

10. A gas turbine engine according to claim 1, wherein: the radial blade span of each fan blade extends from the root at a 0% span position to the tip at a 100% span position; and the average camber (β.sub.2−β.sub.1).sub.ave of each fan blade taken over a portion of the fan blade between the 30% span position and the 40% span position is at least 1.2 times the average camber of the radially innermost 10% of the blade span (α.sub.2−α.sub.2).sub.ave.

11. A gas turbine engine according to claim 1, wherein the fan diameter is greater than 250 cm.

12. A gas turbine engine according to claim 1, wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85.

13. A gas turbine engine according to claim 1, wherein the forward speed of the gas turbine engine at the cruise conditions is Mn 0.8.

14. A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m.

15. A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions at an altitude of 11000 m.

16. A gas turbine engine according to claim 1, wherein the cruise conditions correspond to: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of −55 deg C.

17. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; the fan comprises a plurality of fan blades and a hub, each fan blade having a radial blade span extending from a root to a tip, and the plurality of fan blades being fixed pitch fan blades that are unable to rotate relative to the hub; 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; and an annular splitter at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct, wherein: the ratio of the radius of each fan blade at the root (r.sub.root) to the radius of each fan blade at the tip (r.sub.tip) is less than 0.33; a fan root to tip pressure ratio of the gas turbine engine, defined as the ratio of the mean total pressure of flow at a fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of flow at the fan exit that subsequently flows through the bypass duct (P.sub.104), is in the range of from 0.83 to 0.9 at cruise conditions; and a fan root pressure ratio of the gas turbine engine, defined as the ratio of the mean total pressure of flow at the fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of flow at a fan inlet (P.sub.100), is in the range of from 1.15 to 1.25 at the cruise conditions; a fan tip loading is defined as dH/Ut.sub.tip.sup.2, where dH is the enthalpy rise across the fan and Ut.sub.tip is the velocity of the fan at the tip of each fan blade, and the fan tip loading at the cruise conditions is greater than 0.3 JKg.sup.-1/(ms.sup.-1).sup.2.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine in accordance with the present disclosure;

(3) FIG. 2 is a close-up side view of an inlet region of a gas turbine engine in accordance with the present disclosure;

(4) FIG. 3 is another side view of a fan blade for use with examples of the present disclosure;

(5) FIG. 4 is a schematic of chord of the fan blade for use with examples of the present disclosure in the radially inner 10% of blade span; and

(6) FIG. 5 is a schematic of chord of the fan blade for use with examples of the present disclosure between the 30% span position and the 40% span position.

DETAILED DESCRIPTION

(7) With reference to FIG. 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, a gearbox 14, an intermediate pressure compressor 15, a high-pressure compressor 16, combustion equipment 17, a high-pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 (see FIG. 2) and defines the intake 12. The nacelle 21 may be, or may comprise, a fan containment case 23. The nacelle 21 and the fan case 23 may be separate components.

(8) The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated and compressed by the fan 13 to produce two air flows: a first air flow A into the engine core (indicated generally by reference numeral 24) and a second air flow B which passes through a bypass duct 22 to provide propulsive thrust. The first and second airflows A, B split at a generally annular splitter 40, for example at the leading edge of the generally annular splitter 40 at a generally circular stagnation line. The engine core 24 includes an intermediate pressure compressor 15 (which may be referred to herein as a first compressor 15) which compresses the air flow directed into it before delivering that air to the high pressure compressor 16 (which may be referred to herein as a second compressor 16) where further compression takes place.

(9) The compressed air exhausted from the high-pressure compressor 16 is directed into the combustion equipment 17 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high pressure turbine 18 (which may be referred to as a second turbine 18) and the low pressure turbine 19 (which may be referred to as a first turbine 19) before being exhausted through the nozzle 20 to provide additional propulsive thrust. The intermediate pressure compressor 15 is driven by the low pressure turbine 19 by a first (or low pressure) shaft 32. The high pressure compressor 16 is driven by the low pressure turbine 18 by a second (or high pressure) shaft 34. The first shaft 32 also drives the fan 13 via the gearbox 14. The gearbox 14 is a reduction gearbox in that it gears down the rate of rotation of the fan 13 by comparison with the intermediate pressure compressor 15 and low pressure turbine 19. The gearbox 14, may be any suitable type of gearbox, such as an epicyclic planetary gearbox (having a static ring gear, rotating and orbiting planet gears supported by a planet carrier and a rotating sun gear) or a star gearbox. Additionally or alternatively the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

(10) The first and second compressors 15, 16, first and second turbines 19, 18, first and second shafts 32, 34, and the combustor 17 may all be said to be part of the engine core 24.

(11) 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 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle 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 24 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.

(12) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction 300 (which is aligned with the rotational axis 11), a radial direction 400, and a circumferential direction 500 (shown perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions 300, 400, 500 are mutually perpendicular.

(13) FIG. 2 shows a close-up side (axial-radial) view of the intake region of the gas turbine engine 10. At cruise conditions (as defined elsewhere herein), the total mass flow into the engine is (A+B). The mass flow into the engine core 24 is A, and the mass flow into the bypass duct is B. A stagnation streamline 110 stagnates on the leading edge of the splitter 40. The stagnation streamlines 110 around the circumference of the engine 10 form a streamsurface 110. All of the flow A radially inside this streamsurface 110 ultimately flows through the engine core 24. The streamsurface 110 forms a radially outer boundary of a streamtube that contains all of the flow that ultimately flows through the engine core 24, which may be referred to as the core flow A. All of the flow B radially outside the streamsurface 110 ultimately flows through the bypass duct 22. The streamsurface 110 forms a radially inner boundary of a streamtube that contains all of the flow B that ultimately flows through the bypass duct 22, which may be referred to as the bypass flow B.

(14) The ratio of the mass flow rate of the bypass flow B to the core flow A may be as described and/or claimed herein, for example at least 10, 11, 12, or 13.

(15) In use, the fan blades 130 of the fan 13 do work on the flow, thereby raising the total pressure of the flow. A fan root pressure ratio is defined as the mean total pressure of the flow at the fan exit that subsequently flows (as flow A) through the engine core 24 to the mean total pressure at the inlet to the fan 13. With reference to FIG. 2, the mean total pressure of the flow at the fan exit that subsequently flows through the engine core is the mean total pressure over the surface 102 that is just downstream of the fan 13 and radially inside the streamsurface 110, and may be designated as P.sub.102. Also in FIG. 2, the mean total pressure at the inlet to the fan 13 is the mean total pressure over the surface 100 that extends across the engine and is immediately upstream of the fan 13, and may be designated as P.sub.100.

(16) In some arrangements, the value of the fan root pressure ratio (P.sub.102/P.sub.100) may be as described and/or claimed herein, for example less than 1.25, and/or less than 1.22.

(17) A fan root to tip pressure ratio is defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows (as flow A) through the engine core 24 to the mean total pressure of the flow at the fan exit that subsequently flows (as flow B) through the bypass duct. With reference to FIG. 2, the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct 22 is the mean total pressure over the surface 104 that is just downstream of the fan 13 and radially outside the streamsurface 110, and may be designated as P.sub.104.

(18) The fan root to tip pressure ratio (P.sub.102/P.sub.104) is as described and/or claimed herein, for example less than 0.9, and/or less than 0.89, 0.88, 0.87, 0.86 or 0.85. This ratio may alternatively be expressed simply as the ratio between the mean total pressure (P.sub.102) of the flow at the fan exit that subsequently flows (as flow A) through the engine core 24 to the mean total pressure (P.sub.104) of the flow at the fan exit that subsequently flows (as flow B) through the bypass duct 22.

(19) For completeness, it will be appreciated that a fan tip pressure ratio may be defined as the mean total pressure of the flow at the fan exit that subsequently flows (as flow B) through the bypass duct 22 to the mean total pressure at the inlet to the fan 13.

(20) Referring back to FIG. 1, the exit from the high pressure (or second) compressor 16 is indicated by reference numeral 106. This may additionally or alternatively referred to as the location 106 at the inlet to (or immediately upstream of) the combustor 17. Accordingly, the core flow at the location 106 may have been compressed by all compression stages 15, 16 in the engine 10.

(21) An overall pressure ratio (P.sub.106/P.sub.100) may be defined as the mean total pressure immediately upstream of the combustor 17 (P.sub.106) divided by the mean total pressure at the fan inlet (P.sub.100). The ratio of the fan root pressure ratio to the overall pressure ratio (which may be referred to as the ratio of the mean total pressure (P.sub.102) of the flow at the fan exit that subsequently flows through the engine core 24 divided by the mean total pressure immediately upstream of the combustor 17 (P.sub.106)) may be in the ranges described and/or claimed elsewhere herein, for example less than 0.04.

(22) FIG. 3 shows the fan blade 130 in greater detail. The fan blade 130 extends from a root 132 to a tip 134 in a substantially radial spanwise direction 400. The root 132 may be defined by the radially innermost gas-washed points of the blade 130 and/or may be defined as an intersection between the fan blade 130 and a surface (for example a conical and/or cylindrical surface) from which the fan blades 13 extend. The fan blade 130 has a leading edge 136 and a trailing edge 138. The leading edge 136 may be defined as the line defined by the axially forwardmost points of the fan blade 130 from its root 132 to its tip 134. The fan blade 130 may (or may not) have a fixture portion (not shown) radially inboard of the root, which may be used to fix the fan blade 130 to the rest of the engine.

(23) The radius of the leading edge 136 of the fan blade 130 at its root 132 is designated in FIG. 3 as r.sub.root. The radius of the leading edge 136 of the fan blade 130 at its tip 134 is designated in FIG. 3 as r.sub.tip. The ratio of the radius of the leading edge 136 of the fan blade 130 at its root 132 to the radius of the leading edge 136 of the fan blade 130 at its tip 134 (r.sub.root/r.sub.tip) may be as described and/or claimed herein, for example less than 0.35 and/or less than 0.33 and/or less than 0.28.

(24) In use of the gas turbine engine 10, the fan 13 (with associated fan blades 130) rotates about the rotational axis 11. This rotation results in the tip 134 of the fan blade 130 moving with a velocity U.sub.tip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. Accordingly, a fan tip loading may be defined as dH/U.sub.tip.sup.2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U.sub.tip is the velocity of the fan tip (which may be defined as fan tip radius at leading edge multiplied by rotational speed). As noted elsewhere herein, the fan tip loading at cruise conditions may be greater than (or on the order of) 0.3, for example greater than (or on the order of) 0.31, for example greater than (or on the order of) 0.32, for example greater than (or on the order of) 0.33, for example greater than (or on the order of) 0.34, for example greater than (or on the order of) 0.35, for example greater than (or on the order of) 0.36, for example in the range of from 0.3 to 0.4 (all figures having units JKg.sup.−1K.sup.−1/(ms.sup.−1).sup.2.

(25) The specific thrust of the gas turbine engine 10 may be in the ranges described and/or claimed herein.

(26) A cross-sectional plane P through the blade 130 may be defined by an extrusion in the circumferential direction of a straight line formed between a point on the leading edge 136 that is at a given percentage X of the span s from the root 132 (i.e. at a radius of (r.sub.root+X/100*(r.sub.tip−r.sub.root))), and a point on the trailing edge that is at the same radial percentage X of a trailing edge radial extent t along the trailing edge 138 from the root 132 at the trailing edge 138. The circumferential direction of the extrusion may be taken at the leading edge position of the plane P. In other words, reference to a cross-section through the blade 130 at a given percentage along the blade span (or a given percentage span position) may mean a section through the aerofoil in a plane defined by: a line that passes through the point on the leading edge that is at that percentage of the span s along the leading edge from the leading edge root and points in the direction of the tangent to the circumferential direction at that point on the leading edge; and a point on the trailing edge that is at that same percentage along the trailing edge 138 from the trailing edge root.

(27) An example of a cross-section A-A taken through the blade 130 in such a plane P.sub.A is shown in FIG. 4. The plane P.sub.A is as defined above, with the percentage X of the span s from the root 132 being no greater than 10% (the 10% span position being indicated by the line S10 in FIG. 3). The cross-section A-A has a camber line 142. In general, a camber line 142, 144 (see FIGS. 4 and 5, described below) may be defined as the line formed by the points in a given cross-section that are equidistant from a pressure surface 139 and the suction surface 137 for that cross-section. The camber of the blade 130 at a given cross-section (which may be referred to as the camber of the blade at a given span position) may be defined as the change in angle of the camber line 142 between the leading edge 136 and the trailing edge 138, that is to say (α.sub.2−α.sub.1) in the FIG. 4 example.

(28) FIG. 5 shows, again purely by way of example, a possible cross-section B-B taken through the blade 130 in a plane P.sub.B. As indicated in FIG. 3, the P.sub.B is as defined above, with the percentage X of the span s from the root 132 being no less than 30% (the 30% span position being indicated by the line S30 in FIG. 3) and no greater than 40% (the 40% span position being indicated by the line S40 in FIG. 3). The camber line 144 of the blade 130 in this cross-section is indicated by the dashed line 144 in FIG. 5. The camber of the blade 130 at the cross-section B-B may be defined as the change in angle of the camber line 144 between the leading edge 136 and the trailing edge 138, that is to say (β.sub.2−β.sub.1).

(29) The average (which, as used herein, may be taken as the mean) camber taken over all cross-sections A-A within the radially innermost 10% of the blade span (that is, between the root 132 and the 10% span line S10) may be indicated as (α.sub.2−α.sub.1).sub.ave. The average camber taken over all cross-sections B-B between the 30% span position S30 and the 40% span position S40 may be indicated as (β.sub.2−β.sub.1).sub.ave. The camber may, of course, be averaged over any 10% portion of the blade span, for example between 13% and 23%, 51% and 61%, 76% and 86% or any other 10% span range. There will be one 10% span portion for which the average camber is higher than any other 10% span portion. This may be referred to as the 10% blade span portion that has the maximum average camber (θ.sub.2−θ.sub.1).sub.max.

(30) In some arrangements, the average camber (β.sub.2−β.sub.1).sub.ave of each fan blade 130 taken over a portion of the blade between the 30% span position and the 40% span position may be at least 1.2 times the average camber of the radially innermost 10% of the blade span (α.sub.2−β.sub.1).sub.ave.

(31) In some arrangements, the average camber (α.sub.2−α.sub.1).sub.ave of each fan blade taken over the radially innermost 10% of the blade span may be no greater than 75% of the average camber of the 10% blade span portion that has the maximum average camber (θ.sub.2−θ.sub.1).sub.max.

(32) In use, the gas turbine engine 10 may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine 10 may be mounted in order to provide propulsive thrust.

(33) It will be understood that the disclosure 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.