The present disclosure relates to a gas turbine engine for an aircraft, and more specifically to a gas turbine engine with specified relative component dimensions.

The skilled person would appreciate that simply scaling up components of a known engine type may not provide a corresponding scaling of power/thrust and/or efficiency, and may introduce problems such as increased drag or difficulty of installation. Reconsideration of engine parameters may therefore be appropriate.

For example, the skilled person would appreciate that, if the overall size of a gas turbine engine is increased, one problem that may need to be addressed is how to reduce the overall drag produced by a correspondingly larger nacelle of the larger engine when it is in use. If the components of the engine are scaled proportionally—simply scaling up a known engine type—the increased drag may negatively affect the performance of the aircraft on which the engine is mounted. Additionally or alternatively, the engine may not fit to be mounted beneath the wing of the aircraft unless dimensions are adjusted.

As used herein, a range “from value X to value Y” or “between value X and value Y”, or the likes, denotes an inclusive range; including the bounding values of X and Y. As used herein, the term “axial plane” denotes a plane extending along the length of an engine, parallel to and containing an axial centreline of the engine, and the term “radial plane” denotes a plane extending perpendicular to the axial centreline of the engine, so including all radial lines at the axial position of the radial plane. Axial planes may also be referred to as longitudinal planes, as they extend along the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively.

According to one aspect there is provided a gas turbine engine for an aircraft, the engine comprising an engine core having a core length and comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. An engine area ratio of:

the⁢⁢fan⁢⁢face⁢⁢areathe⁢⁢turbine⁢⁢diameter⁢(at⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢rotor⁢⁢stage)×the⁢⁢core⁢⁢length

is in the range from 1.7 to 3.

The present aspect relates to a gas turbine engine with specified relative fan face and engine sizes. The skilled person will appreciate that a larger fan may provide improved propulsive efficiency. The skilled person will appreciate that a relatively small turbine diameter, compared to the fan size, may improve ease of installation. The skilled person will appreciate that a relatively short core length, and/or a relatively narrow core diameter, compared to the fan size, may benefit close coupled installation.

The skilled person will appreciate that turbine diameter×core length may provide an effective engine area in an axial plane, and that reducing this area may benefit close-coupled installation. In particular, the fan may be mounted closer to the wing when the engine area is smaller (mounted further back and further up than otherwise), so reducing the moment applied to the wing by the mass of the engine.

The skilled person will appreciate that equivalent units should be selected for the fan face area, turbine diameter and core length—e.g. if the area is given in m2, the lengths should both be provided in metres.

The engine area ratio may be higher than that of known aircraft gas turbine engines.

The engine area ratio may be in the range from 1.7 to 3.0, and optionally from 1.70 to 3.00. The engine area ratio may be in the range from 1.8 to 3, or 1.9 to 3 (or optionally to 3.0).

The engine area ratio may be in the range from 2 to 3. The engine area ratio may be in the range from 2.1 to 2.7.

The fan tip radius may be measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge—this may equivalently be described as the fan tip radius being defined as the radial distance between a centreline of the engine and an outermost tip of each fan blade at its leading edge. The fan face area may be equal to π multiplied by the square of the fan tip radius.

The fan tip radius, measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge, may be in the range from 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, or alternatively in the range from 155 cm to 200 cm. The fan tip radius may be greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip radius may be around 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip radius may be greater than 160 cm.

The fan tip radius may be in the range from 95 cm to 150 cm, optionally in the range from 110 cm to 150 cm, optionally in the range of from 110 cm to 145 cm, and further optionally in the range from 120 cm to 140 cm.

The fan tip radius may be in the range from 155 cm to 200 cm, optionally in the range from 160 cm to 200 cm, and further optionally in the range from 165 cm to 190 cm.

Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the engine area ratio may be in the range from 1.7 to 3, optionally 1.7 to 2.7, optionally from 2.1 to 2.7, and further optionally from 2.2 to 3.

Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the engine area ratio may be in the range from 2 to 3, optionally 2.2 to 3, optionally 2.3 to 2.6, and optionally from 2.5 to 2.6.

The turbine diameter at the lowest pressure rotor stage may be measured at the axial location of blade tip trailing edges of rotor blades of the lowest pressure rotor stage. In embodiments in which the rotor (of the lowest pressure rotor stage) is shrouded, the turbine diameter of the turbine at the lowest pressure rotor stage may be measured to the underside of the shroud. In embodiments in which the rotor (of the lowest pressure rotor stage) is unshrouded, the turbine diameter of the turbine at the lowest pressure rotor stage may be measured to the blade tips of the rotor. The lowest pressure rotor stage may be the most axially rearward (or most downstream) rotor stage.

The turbine diameter at the lowest pressure rotor stage may be in the range from 70 cm to 170 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 140 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 70 cm to 120 cm, for example 80 cm to 115 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 120 cm to 170 cm, for example 130 cm to 160 cm.

The ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage

(fan⁢⁢tip⁢⁢radiusturbine⁢⁢diameter)
may be in the range of 0.8 to 2.1. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm, the ratio of the fan tip radius to the turbine diameter may be in the range 0.8 to 2.1. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the ratio of the fan tip radius to the turbine diameter may be in the range of 0.9 to 1.7.

The core length may be defined as the axial distance between a forward region of the compressor and a rearward region of the turbine. The core length may be measured along a centreline of the engine from a mean radius point of the first stage of the compressor blade leading edge to a mean radius point of the lowest pressure turbine rotor stage blade trailing edge of the turbine. The core length may be in the range from 150 cm to 350 cm, and optionally 160 cm to 320 cm. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm, the core length may be in the range from 160 cm to 260 cm, for example 200 cm to 250 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the core length may be in the range from 240 cm to 320 cm, for example 260 cm to 300 cm.

The ratio of the fan tip radius to the core length

(fan⁢⁢tip⁢⁢radiuscore⁢⁢length)
may be in the range of 0.3 to 1, and optionally of 0.4 to 0.9, optionally 0.5 to 0.8. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm, the ratio of the fan tip radius to the core length may be in the range from 0.4 to 0.9, optionally 0.5 to 0.8, or 0.55 to 0.75. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the ratio of the fan tip radius to the core length may be in the range from 0.5 to 0.8, optionally 0.60 to 0.80.

The gas turbine engine may further comprise a gearbox. The gearbox may be connected between the core shaft and the fan. The gearbox may be arranged to receive an input from the core shaft, and to provide an output to drive the fan at a lower rotational speed than the core shaft. The gearbox may help to facilitate the ratios (for example the engine area ratio) described and/or claimed herein.

The fan may comprise a plurality of fan blades extending radially from a hub, each fan blade having a leading edge and a trailing edge. The lowest pressure turbine stage may comprise a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge. The gas turbine engine may have a fan tip axis that joins a radially outer tip of the leading edge of one of the plurality of fan blades and the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure stage of the turbine. The fan tip axis may lie in longitudinal plane which contains a centreline of the gas turbine engine. A fan tip axis angle may be defined as the angle between the fan tip axis and the centreline, and the fan axis angle is in a range between 10 to 20 degrees, optionally 11 to 18 degrees, 12 to 17 degrees, or 12 to 16 degrees. The fan tip axis angle may be as described below.

The engine core may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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.

According to another aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit. A bypass to core ratio of:

bypass⁢⁢exhaust⁢⁢nozzle⁢⁢pressure⁢⁢ratiocore⁢⁢exhaust⁢⁢nozzle⁢⁢pressure⁢⁢ratio
is configured to be in the range from 1.1 to 2 under aircraft cruise conditions.

The present aspect relates to a gas turbine engine with specified relative core and bypass exhaust nozzle pressure ratios. The skilled person will appreciate that a nozzle pressure ratio (NPR) is defined as:

total⁢⁢pressure⁢⁢at⁢⁢nozzle⁢⁢exitambient⁢⁢pressure⁢⁢of⁢⁢surroundings

The skilled person will appreciate that, as is standard in the field, “total pressure” at a nozzle exit is defined as the sum of the static and dynamic pressures at the nozzle exit. Given that the ambient pressure of the surroundings is equal for the core exhaust nozzle and the bypass exhaust nozzle, the bypass to core ratio therefore may be simplified as follows:

NPRbypass⁢⁢exhaust⁢⁢nozzleNPRcore⁢⁢exhaust⁢⁢nozzle=total⁢⁢pressure⁢⁢at⁢⁢bypass⁢⁢nozzle⁢⁢exit/ambient⁢⁢pressuretotal⁢⁢pressure⁢⁢at⁢⁢core⁢⁢nozzle⁢⁢exit/ambient⁢⁢pressure=total⁢⁢pressure⁢⁢at⁢⁢bypass⁢⁢nozzle⁢⁢exittotal⁢⁢pressure⁢⁢at⁢⁢core⁢⁢nozzle⁢⁢exit

The ratio of

total⁢⁢pressure⁢⁢at⁢⁢bypass⁢⁢nozzle⁢⁢exittotal⁢⁢pressure⁢⁢at⁢⁢core⁢⁢nozzle⁢⁢exit
may also be referred to as the extraction ratio. The ambient pressure (or pressure of the surroundings) may also be referred to as exit static pressure.

The skilled person will appreciate that the specified relationship between NPRs may improve engine efficiency as compared to known turbine engines, for example by improving fuel-burn. The skilled person will appreciate that the specified relationship between NPRs may allow dimension limitations for the engine core and/or the fan to be inferred—the specified relationship therefore is not limited in its application to when the aircraft is at cruise conditions; rather, whether or not an engine falls within the scope of the claims may be inferred from those dimensions when the aircraft/engine is not in use.

The bypass to core ratio may be higher than that of known aircraft gas turbine engines.

The bypass to core ratio may be in the range from 1.1 to 2.0 under aircraft cruise conditions. The bypass to core ratio may be in the range from 1.10 to 2.00 under aircraft cruise conditions. The bypass to core ratio may be above 1.15 under aircraft cruise conditions. The bypass to core ratio may be in the range from 1.2 to 1.5 under aircraft cruise conditions. The bypass to core ratio may be in the range from 1.1 to 1.6 under aircraft cruise conditions. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the bypass to core ratio may be in the range from 1.0 to 1.4; for example from 1.1 to 1.4 or from 1.0 to 1.3. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the bypass to core ratio may be in the range from 1.3 to 1.6.

Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, NPRbypass exhaust nozzlemay be in the range from 2.0 to 2.3. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, NPRbypass exhaust nozzlemay be in the range from 2.1 to 2.3.

Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, NPRcore exhaust nozzlemay be in the range from 1.7 to 1.9. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, NPRcore exhaust nozzlemay be in the range from 1.4 to 1.6.

The bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. The bypass ratio may be greater than (or on the order of) any of the following: 8, 8.5, 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass ratio may be in the range from 11 to 20, and optionally in the range from 13 to 20 or 14 to 20. The bypass ratio may be in the range of 8 to 9.5, for example for some direct drive engines (engines without a gearbox). The bypass ratio may be in the range of 9 to 16, for example for some geared engines (engines with a gearbox). Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm (which may be geared), the bypass ratio may be in the range from 9 to 15, and optionally from 13 to 15. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm (which may be geared), the bypass to core ratio may be in the range from 13 to 18, optionally 13 to 16. Cruise conditions may be as defined elsewhere herein.

The total pressure at the bypass nozzle exit may be determined at an exit plane of the bypass exhaust nozzle. The exit plane may extend from a rearmost point of the nacelle towards a centreline of the engine. The exit plane may be a radial plane.

The engine core may comprise a casing (also referred to as an inner fixed structure). The total pressure at the core nozzle exit may be determined at an exit plane of the core exhaust nozzle. The exit plane may extend from a rearmost point of the engine core casing towards a centreline of the engine. The exit plane may be a radial plane. The outer diameter of the bypass exhaust nozzle at the bypass exhaust nozzle exit may be in the range of 200 cm to 400 cm, and optionally 200 cm to 380 cm. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm (for example 110 cm to 150 cm), the outer diameter of the bypass exhaust nozzle may be in the range from 200 cm to 290 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the outer diameter of the bypass exhaust nozzle may be in the range from 290 cm to 380 cm.

The inner diameter of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The inner diameter of the bypass exhaust nozzle may be the radial distance between the outer surfaces of the engine core at the axial position of the rearmost tip of the nacelle. The inner diameter of the bypass exhaust nozzle may be in the range from 100 cm to 250 cm, and optionally from 130 cm to 220 cm. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm (for example 110 cm to 150 cm), the inner diameter of the bypass exhaust nozzle may be in the range from 130 cm to 180 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner diameter of the bypass exhaust nozzle may be in the range from 160 cm to 220 cm.

The flow area of the core exhaust nozzle at the core exhaust nozzle exit may be from 0.4 m2(600 square inches) to 1.3 m2(2000 square inches). Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm (for example 110 cm to 150 cm), the flow area of the core exhaust nozzle at the core exhaust nozzle exit may be in the range from 0.4 m2(600 square inches) to 0.6 m2(900 square inches). Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the flow area of the core exhaust nozzle at the core exhaust nozzle exit may be in the range from 0.6 m2(900 square inches) to 1.3 m2(2000 square inches).

The flow area of the bypass exhaust nozzle at the bypass duct exhaust nozzle exit may be from 1.9 m2(3000 square inches) to 5.8 m2(9000 square inches). Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm (for example 110 cm to 150 cm), the flow area of the bypass duct exhaust nozzle at the bypass duct nozzle exit may be in the range from 1.9 m2(3000 square inches) to 4.5 m2(7000 square inches). Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the flow area of the bypass duct exhaust nozzle at the bypass duct exhaust nozzle exit may be in the range from 4.5 m2(7000 square inches) to 5.8 m2(9000 square inches).

A ratio of bypass exhaust nozzle flow area to the core exhaust nozzle flow area

(bypass⁢⁢exhaust⁢⁢nozzle⁢⁢flow⁢⁢areacore⁢⁢exhaust⁢⁢nozzle⁢⁢flow⁢⁢area)
may be in the range from 4 to 6, and optionally in the range from 5 to 6. Optionally, a geared engine (with a gearbox) may have a ratio of bypass exhaust nozzle flow area to the core exhaust nozzle flow area in the range from 5 to 6.

The bypass exhaust nozzle and/or the core exhaust nozzle may be a convergent nozzle.

The gas turbine engine may further comprise a gearbox connected between the core and the fan. The gearbox may be arranged to receive an input from the core shaft and to provide an output to drive the fan at a lower rotational speed than the core shaft.

The fan tip radius, measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge, may be in the range from 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, or alternatively in the range from 155 cm to 200 cm. The fan tip radius may be greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip radius may be around 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip radius may be greater than 160 cm. The fan tip radius may be in the range from 95 cm to 150 cm, optionally in the range from 110 cm to 150 cm, optionally in the range of from 110 cm to 145 cm, and further optionally in the range from 120 cm to 140 cm. The fan tip radius may be in the range from 155 cm to 200 cm, optionally in the range from 160 cm to 200 cm, and further optionally in the range from 165 cm to 190 cm.

The engine core may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may comprise 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.

The gas turbine engine may be arranged to be mounted beneath a wing of an aircraft. A downstream blockage ratio may be defined as:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢turbine'⁢slowest⁢⁢pressure⁢⁢rotor⁢⁢stageground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
A quasi-non-dimensional mass flow rate, Q, may be defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2.
A Q ratio of:
the downstream blockage ratio×quasi non dimensional mass flow rate Q
may be in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2.

The Q ratio may be in a range from 0.005 Kgs−1N−1K1/2to 0.010 Kgs−1N−1K1/2, and optionally from 0.0050 Kgs−1N−1K1/2to 0.0110 or to 0.0100 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.006 Kgs−1N−1K1/2to 0.009 kgs−1N−1K1/2.

The downstream blockage ratio may be in a range from 0.2 to 0.3. The downstream blockage ratio may be in a range from 0.20 to 0.29. The downstream blockage ratio may be in a range from 0.22 to 0.28.

A specific thrust may be defined as net engine thrust divided by mass flow rate through the engine; and at engine cruise conditions, it may be that:0.029 Kgs−1N−1K1/2≤Q≤0.036 Kgs−N−1K1/2;70 Nkg−1s≤specific thrust≤110 Nkg−1s.

At cruise conditions, it may be that: 0.032 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2. At cruise conditions, it may be that: 0.033 Kgs−1N−1K1/2≤Q≤0.035 Kgs−1N−1K1/2.

According to another aspect, there is provided a method of operating an aircraft comprising a gas turbine engine comprising:

an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and

a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit.

The method comprises controlling the aircraft such that a bypass to core ratio of:

bypass⁢⁢exhaust⁢⁢nozzle⁢⁢pressure⁢⁢ratiocore⁢⁢exhaust⁢⁢nozzle⁢⁢pressure⁢⁢ratio
is in the range from 1.1 to 2 under aircraft cruise conditions.

The aircraft may be controlled such that a Q ratio of:
the downstream blockage ratio×quasi non dimensional mass flow rate Q,
as defined for the preceding aspect, may be in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2.

The Q ratio may be in a range from 0.005 Kgs−1N−1K1/2K to 0.010 Kgs−1N−1K1/2, and optionally from 0.0050 Kgs−1N−1K1/2to 0.0100 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.006 Kgs−1N−1K1/2to 0.009 Kgs−1N−1K1/2.

The gas turbine engine may be as described for the preceding aspect.

According to another aspect, there is provided 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 comprising a plurality of fan blades extending from a hub; and 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: the gas turbine engine has an engine length and a centre of gravity position measured relative to the fan, and wherein a centre of gravity position ratio of:
the centre of gravity position/the engine length
is in a range from 0.43 to 0.6.

Defining the centre of gravity position ratio in this range may allow the centre of gravity to be located closer the front mounting position of the gas turbine engine. This may help to reduce or minimise mounting loads compared to centre of gravity position ratios found in known gas turbine engines or which would be achieved with a proportional scaling of engine architecture. Other effects such as reducing bending of the engine core and deflection of the shaft may also be provided by defining the centre of gravity position ratio as defined above. The centre of gravity position ratio may be higher than that of known aircraft gas turbine engines.

The centre of gravity position ratio may be in a range from 0.43 to 0.60, and optionally from 0.45 to 0.6 or 0.46 to 0.6 (or to 0.60). The centre of gravity position ratio may be in a range from 0.47 to 0.49, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm. The centre of gravity position ratio may be in range from 0.45 to 0.48, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The engine length may be in the range from 200 cm to 500 cm, and optionally from 230 cm to 470 cm, optionally 300 cm to 450 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm (for example 120 cm to 140 cm), the engine length may be in the range from 230 cm to 370 cm, optionally 300 to 360 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm (for example 165 cm to 190 cm), the engine length may be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The centre of gravity position may be in a range between 100 cm to 230 cm, optionally 140 cm to 220 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the centre of gravity position may be in a range between 100 cm to 180 cm, optionally 140 cm to 180 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the centre of gravity position may be in a range between 160 cm to 230 cm, optionally 180 cm to 220 cm.

The engine length may be measured as the axial distance between a forward region of the fan and a rearward region of the turbine.

The turbine may comprise a lowest pressure turbine stage having a row of rotor blades, and the engine length may be measured as the axial distance between: the intersection of the leading edge of one of the plurality of fan blades and the hub; and a mean radius point of the trailing edge of one of the rotor blades of the lowest pressure turbine stage of the turbine. The mean radius point may be the midpoint between a 0% span position and a 100% span position of the rotor blade.

The turbine may be a lowest pressure turbine of a plurality of turbines provided in the core.

The position of centre of gravity may be measured as the axial distance between the intersection of the leading edge of one of the plurality of fan blades and the hub; and the centre of gravity of the gas turbine engine.

According to an aspect, there is provided a fan speed to centre of gravity ratio of:
the centre of gravity position ratio×maximum take off rotational fan speed
may be in a range from 600 rpm to 1350 rpm. This ratio may be lower than that of known aircraft gas turbine engines. By defining the fan speed to centre of gravity ratio in this range the centre of gravity may be moved forwards relative to that of a direct drive engine whilst also providing a relatively low fan rotational speed.

The fan speed to centre of gravity ratio may be in a range from 650 rpm to 1276 rpm. The fan speed to centre of gravity ratio may be in a range from 600 rpm to 1290 rpm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the fan speed to centre of gravity ratio may be 925 rpm to 1325 rpm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the fan speed to centre of gravity ratio may be 650 rpm to 910 rpm.

The maximum take-off rotational fan speed may be in a range between 1450 rpm to 3020 rpm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the maximum take-off rotational fan speed may be in a range between 1970 rpm to 3020 rpm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the maximum take-off rotational fan speed may be in a range between 1450 rpm to 1910 rpm.

According to another aspect, there is provided a method of operating an aircraft comprising a gas turbine engine 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 comprising a plurality of fan blades extending from a hub; and 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: the gas turbine engine has an engine length and a centre of gravity position measured relative to the fan, and wherein the method comprises controlling the aircraft such that a centre of gravity position ratio of:
the centre of gravity position/the engine length
is in a range from 0.43 to 0.6, and a fan speed to centre of gravity ratio of:
the centre of gravity position ratio×maximum take off rotational fan speed
is in a range from 600 rpm to 1350 rpm.

The fan speed to centre of gravity ratio may be in a range from 650 rpm to 1276 rpm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the fan speed to centre of gravity ratio may be 925 rpm to 1325 rpm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the fan speed to centre of gravity ratio may be 650 rpm to 910 rpm.

The maximum take-off rotational fan speed may be in a range between 1450 rpm to 3020 rpm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the maximum take-off rotational fan speed may be in a range between 1970 rpm to 3020 rpm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the maximum take-off rotational fan speed may be in a range between 1450 rpm to 1910 rpm.

According to another aspect, there is provided 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 comprising a plurality of fan blades extending from a hub; and 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: the gas turbine engine has an engine length and a gearbox location relative to a forward region of the fan, wherein a gearbox location ratio of:
gearbox location/engine length
is in a range from 0.19 to 0.45.

Defining gearbox location ratio in this range may allow the gearbox to be located at or near a front mounting position of the gas turbine engine. As the gearbox is generally amongst the heaviest components within the engine its location may have a significant influence on the position of the centre of gravity. Moving the centre of gravity closer to the front mounting may help to minimise rear mounting loads. Other effects such as reducing bending of the engine core and deflection of the core connecting drive shaft may also be provided by controlling the engine centre of gravity by suitable positioning of the gearbox. The gearbox location ratio may be higher than that of known gas turbine engines.

The gearbox location ratio may be in a range from 0.19 to 0.3. The gearbox location ratio may be in a range from 0.19 to 0.23. The gearbox location ratio may be in a range from 0.19 to 0.23, for example 0.19 to 0.21, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm. The gearbox location ratio may be in a range from 0.20 to 0.25, for example equal to or around 0.23—for example, being in the range from 0.225 to 0.235—for example, for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The engine length may be in the range from 200 cm to 500 cm, and optionally from 230 cm to 470 cm, optionally 300 cm to 450 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm (for example 120 cm to 140 cm), the engine length may be in the range from 230 cm to 370 cm, optionally 300 to 360 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm (for example 165 cm to 190 cm), the engine length may be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The gearbox location may be in a range between 50 cm to 110 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the gearbox location may be in a range between 50 cm to 80 cm, optionally 55 cm to 75 cm. Optionally, for an engine with a fan tip radius in the range from 155 cm to 200 cm, the gearbox location may be in a range between 80 cm to 110 cm, optionally 85 cm to 105 cm.

The engine length may be measured as the axial distance between a forward region of the fan and a rearward region of the turbine.

The turbine may comprise a lowest pressure turbine stage having a row of rotor blades, and the engine length is measured as the axial distance between: the intersection of the leading edge of one of the plurality of fan blades and the hub; and a mean radius point of the trailing edge of one of the rotor blades of the lowest pressure turbine stage of the turbine. The mean radius point may be the midpoint between a 0% span position and a 100% span position of the rotor blade.

The turbine may be a lowest pressure turbine of a plurality of turbines provided in the core.

The gearbox location may be measured between: the intersection of a leading edge of one of the fan blades and the hub; and a radial centre plane of the gearbox, the radial centre plane being at the midpoint between the front face of a most forward gear mesh of the gearbox and the rear face of a most rearward gear mesh of the gearbox.

The gearbox may be an epicyclic gearbox comprising a ring gear; in such embodiments, the gearbox location may be measured as the axial distance between: the intersection of a leading edge of one of the fan blades and the hub; and a radial plane intersecting the axial centre point of the ring gear.

The fan blades may be formed at least partly from a composite material, and the gearbox location for such embodiments may be in a range between 50 cm and 110 cm and optionally in a range between 80 cm and 110 cm.

The fan blades may be formed at least partly from a metal or metal alloy, such as an aluminium-lithium alloy, and wherein the gearbox location may be in a range between 50 cm and 110 cm and optionally in a range between 50 cm and 80 cm.

The fan blades may be formed at least partly from a composite material and the gearbox location ratio may be in the range of from 0.02 to 0.25, for example equal to or around 0.23; for example, being in the range from 0.225 to 0.235. This may be, for example, for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The fan blades may be formed at least partly from a metal or metal alloy, such as an aluminium-lithium alloy, and the gearbox location ratio may be in a range from 0.19 to 0.25, for example 0.19 to 0.23. This may be, for example, for an engine with any fan tip radius.

According to another aspect, there is provided 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 comprising a plurality of fan blades extending from a hub, the fan blades being formed from a metal or metal alloy (optionally an aluminium-lithium alloy); and 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: the gas turbine engine has an engine length and a gearbox location relative to a forward region of the fan, wherein a gearbox location ratio of:
gearbox location/engine length
wherein the gearbox location ratio may be in a range from 0.19 to 0.3, optionally 0.19 to 0.25.

This may be, for example, for an engine with a fan tip radius in the range from 110 cm to 150 cm.

According to another aspect, there is provided 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 comprising a plurality of fan blades extending from a hub, the fan blades being formed from at least partly from a composite material; and 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: the gas turbine engine has an engine length and a gearbox location relative to a forward region of the fan, wherein a gearbox location ratio of:
gearbox location/engine length
wherein the gearbox location ratio may be in the range of from 0.20 to 0.25, for example equal to or around 0.23; for example, being in the range from 0.225 to 0.235.

This may be, for example, for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The gas turbine engine of these aspects may further have a centre of gravity position ratio as defined in any of the statements of the previous two aspects.

According to another aspect there is provided 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 comprising a plurality of fan blades, wherein a fan tip radius of the fan is measured/defined between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and a nacelle surrounding the fan and the engine core and defining a bypass exhaust nozzle, the bypass exhaust nozzle having an outer radius. An outer bypass to fan ratio of:

the⁢⁢outer⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius

is in the range from 0.6 to 1.05.

The present aspect relates to a gas turbine engine comprising a fan and nacelle with specified relative shapes and/or sizes. The skilled person would appreciate that having a relatively narrow bypass exhaust nozzle, as compared to fan size, may reduce drag produced by the engine in use. Further, the skilled person would appreciate that the relatively narrow bypass exhaust nozzle may create a more compact exhaust system, which may allow or facilitate under-wing installation of a larger engine on an aircraft. The outer bypass to fan ratio may be lower than that of known aircraft gas turbine engines.

The outer bypass to fan ratio may be in the range from 0.60 to 1.05. The outer bypass to fan ratio may be in the range from 0.65 to 1.00. The outer bypass to fan ratio may be lower than 1.05, optionally lower than 1.02, and further optionally lower than 1.00. The outer bypass to fan ratio may be in the range from 0.80 to 1.00. The outer bypass to fan ratio may be in the range from 0.9 to 1.0, and optionally in the range from 0.90 to 1.00.

The fan tip radius may be in the range from 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, or alternatively in the range from 155 cm to 200 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the outer bypass to fan ratio may be in the range from 0.95 to 1.00, for example equal to or around 0.97, for example being in the range from 0.96 to 0.98. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, outer bypass to fan ratio may be in the range from 0.91 to 0.98, for example equal to or around 0.95, for example being in the range from 0.94 to 0.96.

The bypass exhaust nozzle may have an exit plane, which may be a radial plane. The outer radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle.

The outer radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The outer radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an inner surface of the nacelle at the axial position of the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be in the range of 100 cm to 200 cm, and optionally 100 cm to 190 cm. The outer radius of the bypass exhaust nozzle may be defined as half the diameter of the bypass exhaust nozzle as described above. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the outer radius of the bypass exhaust nozzle may be in the range from 100 cm to 145 cm, for example 110 cm to 140 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the outer radius of the bypass exhaust nozzle may be in the range from 145 cm to 190 cm.

The bypass exhaust nozzle may have an inner radius. An inner bypass to fan ratio of:

the⁢⁢inner⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius

may be in the range from 0.4 to 0.65.

The inner bypass to fan ratio may be lower than that for known aircraft gas turbine engines. The inner bypass to fan ratio may be in the range from 0.5 to 0.6, and optionally in the range from 0.50 to 0.60. The inner bypass to fan ratio may be in the range from 0.40 to 0.65. The inner bypass to fan ratio may be lower than 0.65, and optionally lower than 0.64, and optionally lower than 0.62. The inner bypass to fan ratio may be in the range from 0.54 to 0.64. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner bypass to fan ratio may be in the range of from 0.57 to 0.63, for example 0.57 to 0.62, for example around 0.59, for example being in the range from 0.58 to 0.60. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner bypass to fan ratio may be in the range from 0.5 to 0.6, and optionally from 0.52 to 0.58.

The bypass exhaust nozzle may have an exit plane, which may be a radial plane. The inner radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle. The inner and outer radii may therefore be measured in the same radial plane. The inner and outer radii of the bypass exhaust nozzle may therefore be measured at the exit of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The inner radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an outer surface of the engine core at the axial position of the rearmost tip of the nacelle. The inner radius of the bypass exhaust nozzle is half the inner diameter of the bypass exhaust nozzle. The inner radius of the bypass exhaust nozzle may be in the range from 50 cm to 125 cm, and optionally from 65 cm to 110 cm, optionally from 75 cm to 110 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner radius of the bypass exhaust nozzle may be in the range from 70 cm to 90 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner radius of the bypass exhaust nozzle may be in the range from 80 cm to 120 cm, for example 90 cm to 120 cm.

The nacelle may define a bypass duct located radially outside of the engine core. The bypass exhaust nozzle may be an exit of the bypass duct, for example forming a rearmost portion of the bypass exhaust duct. The nacelle may comprise an inner surface at least partly defining the bypass duct.

The engine may further comprise a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle. The bypass duct outlet guide vane may extend between a radially inner tip and a radially outer tip and may have a leading edge and a trailing edge relative to the direction of gas flow through the bypass duct. An outer wall axis may be defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle. The outer wall axis may lie in a longitudinal plane containing the centreline of the gas turbine engine. An outer bypass duct wall angle may be defined as the angle between the outer wall axis and the centreline.

The outer bypass duct wall angle may be in a range between −15 degrees and 1 degrees. The outer bypass duct wall angle may be in a range between −10 degrees and 0 degrees, for example −5 degrees and 0 degrees. The outer bypass duct wall angle may be in a range between −4 degrees and −1 degrees. The outer bypass duct wall angle may be in the range of −0.5 degrees to −4 degrees; this may be for an engine with a fan tip radius in the range from 110 cm to 150 cm. The outer bypass duct wall angle may be in a range between −1 degrees and −5 degrees, for example −2.5 degrees to −4 degrees; this may be for an engine with a fan tip radius in the range from 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond to the outer wall axis sloping towards the centreline of the gas turbine engine.

The gas turbine engine may comprise 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. The gearbox may have a gear ratio in the ranges defined elsewhere herein, for example 3.2 to 5 or 3.2 to 3.8.

The gas turbine engine may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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.

According to another aspect there is provided 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 comprising a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and a nacelle surrounding the fan and the engine core and defining a bypass exhaust nozzle, the bypass exhaust nozzle having an inner radius. An inner bypass to fan ratio of:

the⁢⁢inner⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius

is in the range from 0.4 to 0.65.

The present aspect relates to a gas turbine engine comprising a fan and engine core with specified relative shapes and/or sizes. The skilled person would appreciate that having a relatively narrow bypass exhaust nozzle, as compared to fan size, may reduce drag produced by the engine in use. The skilled person would appreciate that the engine core is situated radially within the bypass exhaust nozzle, and that the inner radius of the bypass exhaust nozzle may therefore equivalently be thought of as an outer radius of the engine core. The skilled person would appreciate that having a relatively narrow engine core, as compared to fan size, may reduce drag produced by the engine in use. The inner bypass to fan ratio may be lower than that for known aircraft gas turbine engines.

The inner bypass to fan ratio may be in the range from 0.40 to 0.65. The inner bypass to fan ratio may be lower than 0.65 and optionally lower than 0.64. The inner bypass to fan ratio may be in the range from 0.54 to 0.64. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner bypass to fan ratio may be in the range of from 0.57 to 0.63, for example 0.57 to 0.62, for example around 0.59, for example being in the range from 0.58 to 0.60. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner bypass to fan ratio may be in the range from 0.5 to 0.6, and optionally from 0.52 to 0.58, optionally 0.53 to 0.55.

The bypass exhaust nozzle may have an exit plane, which may be a radial plane. The inner radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The inner radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an outer surface of the engine core at the axial position of the rearmost tip of the nacelle.

The inner radius of the bypass exhaust nozzle may be in the range from 50 cm to 125 cm, and optionally from 65 cm to 110 cm, optionally from 75 cm to 110 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner radius of the bypass exhaust nozzle may be in the range from 70 cm to 90 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner radius of the bypass exhaust nozzle may be in the range from 80 cm to 120 cm, for example 90 cm to 120 cm.

The bypass exhaust nozzle may have an outer radius, and an outer bypass to fan ratio of:

the⁢⁢outer⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius

may be in the range from 0.6 to 1.05. The outer bypass to fan ratio may be lower than that for known aircraft gas turbine engines.

The outer bypass to fan ratio may be in the range from 0.9 to 1.0, and optionally from 0.90 to 1.00. The outer bypass to fan ratio may be lower than 1.05, optionally lower than 1.02, and further optionally lower than 1.00. The outer bypass to fan ratio may be in the range from 0.8 to 1.00, and optionally in the range from 0.90 to 1.00. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the outer bypass to fan ratio may be in the range from 0.95 to 1.00, for example equal to or around 0.97, for example being in the range from 0.96 to 0.98. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, outer bypass to fan ratio may be in the range from 0.91 to 0.98, for example equal to or around 0.95, for example being in the range from 0.94 to 0.96.

The bypass exhaust nozzle may have an exit plane, which may be radial. The outer radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle. The outer radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The outer radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an inner surface of the nacelle at the axial position of the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be in the range of 100 cm to 200 cm, and optionally 100 cm to 190 cm. The outer radius of the bypass exhaust nozzle may be defined as half the diameter of the bypass exhaust nozzle as described above. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm, the outer radius of the bypass exhaust nozzle may be in the range from 100 cm to 145 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the outer radius of the bypass exhaust nozzle may be in the range from 145 cm to 190 cm.

The nacelle may define a bypass duct located radially outside of the engine core. The nacelle may comprise an inner surface at least partly defining the bypass duct.

The bypass exhaust nozzle may be an exit of the bypass duct, for example forming a rearmost portion of the bypass exhaust duct.

The engine may further comprise a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle. The bypass duct outlet guide vane may extend between a radially inner tip and a radially outer tip and may have a leading edge and a trailing edge relative to the direction of gas flow through the bypass duct. An outer wall axis may be defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle. The outer wall axis may lie in a longitudinal plane containing the centreline of the gas turbine engine. An outer bypass duct wall angle may be defined as the angle between the outer wall axis and the centreline.

The outer bypass duct wall angle may be in a range between −15 to 1 degrees. The outer bypass duct wall angle may be in a range between −10 degrees and 0 degrees, for example −5 degrees and 0 degrees. The outer bypass duct wall angle may be in a range between −4 degrees and −1 degrees. The outer bypass duct wall angle may be in the range of −0.5 degrees to −4 degrees; this may be for an engine with a fan tip radius in the range from 110 cm to 150 cm. The outer bypass duct wall angle may be in a range between −1 degrees and −5 degrees, for example −2.5 degrees to −4.0 degrees; this may be for an engine with a fan tip radius in the range from 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond to the outer wall axis sloping towards the centreline of the gas turbine engine.

The gas turbine engine may further comprise 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. The gearbox may have a gear ratio in the ranges defined elsewhere herein, for example 3.2 to 5 or 3.2 to 3.8.

The gas turbine engine may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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 are arranged to rotate at a higher rotational speed than the first core shaft.

According to another aspect, there is provided 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 comprising a plurality of fan blades; a nacelle surrounding the gas turbine engine, the nacelle comprising an inner surface at least partly defining a bypass duct located radially outside of the engine core; and a bypass duct outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle, wherein the bypass duct outlet guide vane extends between a radially inner tip and a radially outer tip and has a leading edge and a trailing edge relative to the direction of gas flow through the bypass duct, an outer wall axis is defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle, wherein the outer wall axis lies in a longitudinal plane containing the centreline of the gas turbine engine, an outer bypass duct wall angle is defined as the angle between the outer wall axis and the centreline, and the outer bypass duct wall angle is in a range between −15 to 1 degrees.

By providing the bypass duct wall angle in this range a more compact exhaust system may be provided. By using the angle range above the length of the nacelle between the bypass OGV and the rearmost inner tip of the nacelle inner wall may be reduced. This may provide a shorter length of nacelle which may provide a reduction in drag compared to known gas turbine engines, or which would be achieved if the nacelle dimensions were scaled proportionately when increasing the size of the gas turbine engine. The bypass duct wall angle may be lower (i.e. more negative) than that of known gas turbine engines.

The outer bypass duct wall angle may be in a range between −10 degrees and 0 degrees, for example −5 degrees and 0 degrees. The outer bypass duct wall angle may be in a range between −4 degrees and −1 degrees. The outer bypass duct wall angle may be in the range of −0.5 degrees to −4 degrees; this may be for an engine with a fan tip radius in the range from 110 cm to 150 cm. The outer bypass duct wall angle may be in a range between −1 degrees and −5 degrees, for example −2.5 degrees to −4 degrees; this may be for an engine with a fan tip radius in the range from 155 cm to 200 cm.

A negative value of the outer bypass duct wall angle may correspond to the outer wall axis sloping towards the centreline of the gas turbine engine.

A bypass duct outlet guide vane radius, measured radially between the engine centreline and the radially outer tip of the trailing edge of the bypass OGV, may be in a range from 90 cm to 210 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the bypass duct outlet guide vane radius may be in the range from 90 cm to 150 cm, optionally 110 cm to 135 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the bypass duct outlet guide vane radius may be in the range from 160 cm to 210 cm, optionally 170 cm to 200 cm.

The gas turbine engine may further comprise 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. The gearbox may have a gear ratio in the ranges defined elsewhere herein, for example 3.2 to 5 or 3.2 to 3.8.

The rearmost inner tip of the nacelle inner wall may be movable in use of the gas turbine engine to provide a variable area bypass duct exhaust nozzle (also referred to as a fan nozzle). The outer wall axis may be defined based on the position of the rearmost tip of the inner surface of the nacelle during cruise conditions. Cruise conditions may be as defined in connection with any other aspect.

The gas turbine engine of this aspect may further have an outer bypass to fan ratio and/or inner bypass to fan ratio as defined in the previous relevant aspects of any of the above statements.

A fan tip radius of the fan may be measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and the nacelle may surround the fan and the engine core and define a bypass exhaust nozzle, the bypass exhaust nozzle having an outer radius.

An outer bypass to fan ratio of:

the⁢⁢outer⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius

may be in the range from 0.6 to 1.05. The outer bypass to fan ratio may be in the range from 0.60 to 1.05. The outer bypass to fan ratio may be in the range from 0.65 to 1.00. The outer bypass to fan ratio may be lower than 1.05, optionally lower than 1.02, and further optionally lower than 1.00. The outer bypass to fan ratio may be in the range from 0.80 to 1.00. The outer bypass to fan ratio may be in the range from 0.9 to 1.0, and optionally in the range from 0.90 to 1.00. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the outer bypass to fan ratio may be in the range 0.95 to 1.00, optionally equal to or around 0.97, for example being in the range from 0.96 to 0.98. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, outer bypass to fan ratio may be in the range 0.91 to 0.98, optionally equal to or around 0.95, for example being in the range from 0.94 to 0.96. The outer bypass to fan ratio may be lower than that for known aircraft gas turbine engines.

The bypass exhaust nozzle may have an exit plane, which may be a radial plane. The outer radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle. The outer radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle.

The outer radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an inner surface of the nacelle at the axial position of the rearmost tip of the nacelle. The outer radius of the bypass exhaust nozzle may be in the range of 100 cm to 200 cm, and optionally 100 cm to 190 cm. The outer radius of the bypass exhaust nozzle may be defined as half the diameter of the bypass exhaust nozzle as described above. Optionally, for example for an engine with a fan tip radius in the range from 95 cm to 150 cm, the outer radius of the bypass exhaust nozzle may be in the range from 100 cm to 145 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the outer radius of the bypass exhaust nozzle may be in the range from 145 cm to 190 cm.

The bypass exhaust nozzle may have an inner radius. An inner bypass to fan ratio of:

the⁢⁢inner⁢⁢radius⁢⁢of⁢⁢the⁢⁢bypass⁢⁢exhaust⁢⁢nozzlethe⁢⁢fan⁢⁢tip⁢⁢radius
may be in the range from 0.4 to 0.65. The inner bypass to fan ratio may be lower than that for known aircraft gas turbine engines. The inner bypass to fan ratio may be in the range from 0.5 to 0.6, and optionally in the range from 0.50 to 0.60. The inner bypass to fan ratio may be in the range from 0.40 to 0.65. The inner bypass to fan ratio may be lower than 0.65, and optionally lower than 0.64, and optionally lower than 0.62. The inner bypass to fan ratio may be in the range from 0.54 to 0.64. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner bypass to fan ratio may be in the range of from 0.57 to 0.63, for example 0.57 to 0.62, for example around 0.59, for example being in the range from 0.58 to 0.60. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner bypass to fan ratio may be in the range from 0.5 to 0.6, and optionally from 0.52 to 0.58, optionally 0.53 to 0.55.

The bypass exhaust nozzle may have an exit plane, which may be a radial plane. The inner radius of the bypass exhaust nozzle may be measured at the axial position of the exit plane of the bypass exhaust nozzle. The inner and outer radii may therefore be measured in the same radial plane. The inner and outer radii of the bypass exhaust nozzle may therefore be measured at the exit of the bypass exhaust nozzle.

The inner radius of the bypass exhaust nozzle may be measured at the axial position of the rearmost tip of the nacelle. The inner radius of the bypass exhaust nozzle may be the radial distance between the centreline of the engine and an outer surface of the engine core at the axial position of the rearmost tip of the nacelle. The inner radius of the bypass exhaust nozzle may be in the range from 50 cm to 125 cm, and optionally from 65 cm to 110 cm, optionally from 75 cm to 110 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the inner radius of the bypass exhaust nozzle may be in the range from 70 cm to 90 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the inner radius of the bypass exhaust nozzle may be in the range from 80 cm to 120 cm, for example 90 cm to 120 cm.

According to another aspect, there is provided 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 comprising a plurality of fan blades extending radially from a hub, each fan blade having a leading edge and a trailing edge; and wherein: the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge; the gas turbine engine has a fan tip axis that joins: a radially outer tip of the leading edge of one of the plurality of fan blades; and the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure stage of the turbine, the fan tip axis lying in a longitudinal plane which contains a centreline of the gas turbine engine, and a fan tip axis angle is defined as the angle between the fan tip axis and the centreline, and the fan tip axis angle is in a range between 10 to 20 degrees.

By providing a fan tip axis angle (also referred to as a fan axis angle) in this range the gas turbine engine may have a large fan diameter to provide improved propulsive efficiency, whilst also having a relatively small diameter core. The may help to aid installation of the engine beneath the wing of an aircraft. A relatively smaller diameter core may allow the engine to be mounted relatively further aft on the aircraft, and so may allow the centre of gravity of the engine to be moved closer to the wing structure. The range of the fan tip axis angle defined above may allow improved engine mounting compared to known gas turbine engines, or where engine components are scaled proportionally with increasing fan diameter. The fan axis angle may be higher than that of known gas turbine engines.

The fan axis angle may be in a range of 12 degrees and 17 degrees, for example 13 degrees to 16 degrees.

The fan tip radius, measured as the radial distance from the engine centreline to the radially outer tip of the leading edge of one of the plurality of fan blades, may be in a range from 95 cm to 200 cm, and optionally may be from 110 cm to 150 cm or from 155 cm to 200 cm.

The fan axis angle may be in a range between 13 and 15 degrees—this may be for an engine with a fan tip radius in the range from 110 cm to 150 cm. The fan axis angle may be in a range between 13.5 degrees and 15.5 degrees—this may be for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The turbine radius at the lowest pressure rotor stage, measured/defined as the radial distance from the engine centreline to the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure stage of the turbine, may be in the range from 45 cm to 85 cm. Optionally, for example for an engine with a fan tip radius in a range from 110 cm to 150 cm, the turbine radius at the lowest pressure rotor stage may be in the range from 50 cm to 60 cm. Optionally, for example for an engine with a fan tip radius in a range from 155 cm to 200 cm, the turbine radius at the lowest pressure rotor stage may be in the range from range from 60 cm to 85 cm.

The gas turbine engine may further comprise 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.

The gas turbine engine may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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 are arranged to rotate at a higher rotational speed than the first core shaft.

According to another aspect, there is provided 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; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending radially from a hub, each fan blade having a leading edge and a trailing edge; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan (23) at a lower rotational speed than the core shaft (26); and wherein: the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge; a fan-turbine radius difference is measured as the radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades, and a fan speed to fan-turbine radius ratio defined as:

the⁢⁢maximum⁢⁢take⁢-⁢off⁢⁢rotational⁢⁢speedof⁢⁢the⁢⁢fan⁢⁢(in⁢⁢rpm)fan⁢-⁢turbine⁢⁢radius⁢⁢difference⁢⁢(in⁢⁢mm)
is in a range of 0.8 rpm/mm to 5 rpm/mm.

By configuring the geometry of the gas turbine engine so that the fan-turbine radius difference is within the range above the loading on the pylon which connects the gas turbine engine to the wing of an aircraft may be reduced. By defining the fan speed to fan-turbine radius difference ratio in this way a reduction in fan rotational speed may reduce the restraining torque required for the pylon to restrain the gas turbine engine for relative rotation. Moreover, defining the ratio in this range may also provide a smaller relative core diameter, allowing the gas turbine engine to be mounted closer to the wing. This may also reduce the restraining torque on the pylon. The fan speed to fan-turbine radius may be lower than that of known gas turbine engines.

The fan speed to fan-turbine radius ratio may be in a range between 1.5 rpm/mm to 4.0 rpm/mm. The fan speed to fan-turbine radius ratio may be in a range between 1.5 rpm/mm to 3.6 rpm/mm. The fan speed to fan-turbine radius ratio may be in a range between 2.9 rpm/mm and 3.8 rpm/mm, for example 3.4 rpm/mm and 3.6 rpm/mm—this may be for an engine with a fan tip radius in the range from 110 cm to 150 cm. The fan speed to fan-turbine radius ratio may be in a range between 1.2 rpm/mm and 2 rpm/mm, for example 1.5 rpm/mm and 1.7 rpm/mm—this may be for an engine with a fan tip radius in the range from 155 cm to 200 cm.

The fan-turbine radius difference may be in a range between 55 cm (i.e. 550 mm) to 120 cm (i.e. 1200 mm). Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the fan-turbine radius difference may be in a range between 55 cm (i.e. 550 mm) to 90 cm (i.e. 900 mm), for example 65 cm to 85 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the fan-turbine radius difference may be in a range between 90 cm (i.e. 900 mm) to 120 cm (i.e. 1200 mm), for example 95 cm to 115 cm.

The maximum take-off rotational fan speed may be in a range between 1450 rpm to 3020 rpm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the maximum take-off rotational fan speed may be in a range between 2100 rpm to 3000 rpm, optionally 2300 rpm to 2900 rpm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the maximum take-off rotational fan speed may be in a range between 1450 rpm to 1910 rpm, optionally 1500 rpm to 1800 rpm.

In embodiments in which the rotor (of the lowest pressure rotor stage) is shrouded, the radially outer tip of the trailing edge of each of the rotor blades may be defined at the underside of the shroud. In embodiments in which the rotor (of the lowest pressure rotor stage) is unshrouded, the radially outer tip of the trailing edge of each of the rotor blades may be the blade tip of the rotor blade.

The turbine may be the lowest pressure turbine of a plurality of turbines provided in the core. The turbine stage may be the axially most rearward turbine stage and/or the most downstream turbine stage.

The gearbox may have a gear ratio in the range of from 3.2 to 5, for example 3.2 to 4.2, for example 3.2 to 3.8.

The gas turbine engine of this aspect may further have a fan axis angle as defined in the previous relevant aspect or any of the above statements.

In another aspect, the present application provides a method of operating an aircraft comprising a 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 extending radially from a hub, each fan blade having a leading edge and a trailing edge; and wherein: the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge; a fan-turbine radius difference is measured as the radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades, wherein the method comprises controlling the aircraft such that a fan speed to fan-turbine radius ratio is defined as:

the⁢⁢maximum⁢⁢take⁢-⁢off⁢⁢rotational⁢⁢speed⁢⁢of⁢⁢the⁢⁢fanfan⁢-⁢turbine⁢⁢radius⁢⁢difference
is in a range between 0.8 rpm/mm to 5 rpm/mm.

The gas turbine engine of this aspect may further have a fan axis angle as defined in the previous relevant aspect or any of the above statements

According to another aspect there is provided a gas turbine engine for an aircraft, and arranged to be mounted beneath a wing of the aircraft, the engine comprising:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub; and

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. A downstream blockage ratio of:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotor⁢⁢stageground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance

is in the range from 0.2 to 0.3.

The present aspect relates to a gas turbine arranged to provide a specified downstream blockage when mounted on the wing of an aircraft, the downstream blockage providing a measure of how much of the vertical space beneath a wing of an aircraft is taken up by the gas turbine engine. The skilled person would appreciate that having a relatively low turbine diameter towards the rearward end of the engine, as compared to distance from the aircraft wing on which the engine is to be mounted to the ground plane, may allow more space for a pylon structure arranged to mount the engine to the wing. A lower downstream blockage may therefore be preferable. In general, aircraft engines are mounted under aircraft wings, by means of a pylon (also referred to as an airframe strut) connected between a rearward portion of the engine and an underside of the wing. The turbine diameter at the axial position of the lowest pressure rotor may be used as a measure of engine size towards the rear of the engine.

The skilled person would appreciate that the ground plane is the plane on which the aircraft would land—i.e. the plane of the ground when the aircraft has landed/before take-off/whilst taxiing. The distance between the ground plane and the wing may therefore be measured between the ground (e.g. a runway) and the wing whilst the aircraft is parked. The skilled person would appreciate that the ground plane is generally the plane in which the lowest part of each landing gear is located whilst the aircraft is on the ground. The skilled person would appreciate that the distance between the ground plane and the wing may be measured to the leading edge centre line of the aircraft wing.

The downstream blockage ratio may be lower than that of known aircraft gas turbine engines. The downstream blockage ratio may be in the range from 0.20 to 0.30. The downstream blockage ratio may be in the range from 0.20 to 0.29. The downstream blockage ratio may be in the range from 0.20 to 0.28, and optionally from 0.22 to 0.28. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the downstream blockage ratio may be in the range from 0.23 to 0.27. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, downstream blockage ratio may be in the range from 0.24 to 0.28.

The distance between the ground plane and the wing may be measured to centre point of a leading edge of the wing. The distance between the ground plane and the wing may be measured along a line perpendicular to the ground plane and passing through and perpendicular to an axial centre line of the engine.

The turbine diameter at the axial location of the lowest pressure rotor stage may be measured adjacent blade tips of rotor blades of the lowest pressure rotor stage. The turbine diameter at the axial location of the lowest pressure rotor stage may be in the range from 70 cm to 170 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 100 cm to 130 cm, optionally 110 cm to 120 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 120 cm to 170 cm.

The hub and fan blades of the fan together define a fan face having a fan diameter, and the engine has an engine length. An engine ratio may be defined as:

(2×the⁢⁢fan⁢⁢radius/the⁢⁢engine⁢⁢length)the⁢⁢downstream⁢⁢blockage⁢⁢ratio

The engine ratio may be in the range from 2.5 to 4. The engine ratio may be in the range from 2.5 to 4.0. The engine ratio may be greater than 2.5. The engine ratio may be greater than 3.0. The engine ratio may be in the range of from 2.7 to 3.7, for example 2.8 to 3.5.

The engine length may be measured as the axial distance between a forward region of the fan and a rearward region of the turbine. The engine length may be measured along a centreline of the engine from an axial position of an intersection of the leading edge of each fan blade and the hub of the fan to an axial position of a trailing edge mean radius point of one of the rotor blades provided in the lowest pressure stage of the turbine.

The engine length may be in the range from 200 cm to 500 cm, and optionally from 230 cm to 470 cm, optionally 300 cm to 450 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm (for example 120 cm to 140 cm), the engine length may be in the range from 230 cm to 370 cm, optionally 300 to 360 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm (for example 165 cm to 190 cm), the engine length may be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The fan radius (also referred to as a fan tip radius) may be measured between an engine centreline and a tip of a fan blade at its leading edge and the fan diameter may be twice the radius of the fan. The diameter of the fan, equal to twice the radius of the fan may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm, 260 cm, 270 cm, 280 cm, 290 cm, 300 cm, 310 cm, 320 cm, 330 cm, 340 cm, 350 cm, 360 cm, 370 cm, 380 cm or 390 cm. In some embodiments, the fan diameter may be in the range from 220 cm to 300 cm. In some embodiments, the fan diameter may be in the range from 310 cm to 400 cm.

The turbine diameter at the lowest pressure rotor stage may be measured at the axial location of blade tip trailing edges of rotor blades of the lowest pressure rotor stage. In embodiments in which the rotor (of the lowest pressure rotor stage) is shrouded, the turbine diameter of the turbine at the lowest pressure rotor stage may be measured to the underside of the shroud (as this defines the edge of the gas flow path). In embodiments in which the rotor (of the lowest pressure rotor stage) is unshrouded, the turbine diameter of the turbine at the lowest pressure rotor stage may be measured to the blade tips of the rotor.

The gas turbine engine may comprise more than one turbine. The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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.

A quasi-non-dimensional mass flow rate, Q, may be defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2.

A Q ratio of:
the downstream blockage ratio×quasi non dimensional mass flow rate Q
may be in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2.

The Q ratio may be in a range from 0.0050 Kgs−1N−1K1/2to 0.0110 Kgs−1N−1K1/2, or optionally from 0.005 Kgs−1N−1K1/2to 0.010 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.006 Kgs−1N−1K1/2to 0.009 Kgs−1N−1K1/2.

A specific thrust may be defined as net engine thrust divided by mass flow rate through the engine; and at engine cruise conditions, it may be that:0.029 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2; and70 Nkg−1s≤specific thrust≤110 Nkg−1s.

At cruise conditions, it may be that: 0.032 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2. At cruise conditions, it may be that: 0.033 Kgs−1N−1K1/2≤Q≤0.035 Kgs−1N−1K1/2.

According to another aspect, there is provided an aircraft comprising a wing and a gas turbine engine mounted beneath the wing of the aircraft. The engine comprises:

an engine comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a plurality of rotor stages including a lowest pressure rotor located furthest downstream, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and

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. A downstream blockage ratio of:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotorground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance

is in the range from 0.2 to 0.3.

The engine may be as described for the preceding aspect.

According to another aspect, there is provided a gas turbine engine for an aircraft and arranged to be mounted beneath a wing of the aircraft. The engine has an engine length and comprises:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades of the fan together defining a fan face having a fan tip radius; and

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. A downstream blockage ratio is defined as:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotorground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
An engine blockage ratio of:

(2×the⁢⁢fan⁢⁢tip⁢⁢radius/the⁢⁢engine⁢⁢length)the⁢⁢downstream⁢⁢blockage⁢⁢ratio

is in the range from 2.5 to 4.

The engine blockage ratio may equivalently be represented as:

2×the⁢⁢fan⁢⁢tip⁢⁢radius×ground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distanceengine⁢⁢length×turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢locationof⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢rotor

The skilled person would appreciate that a larger fan diameter (and therefore a larger fan tip radius) may improve propulsive efficiency, but that simply mounting a larger fan on a known gas turbine engine may cause other difficulties or inefficiencies. Having a shorter engine length may allow an engine with a larger fan to be installed closer to a wing of the aircraft, i.e. further rearwards; this may reduce the moment exerted on the wing by the (potentially larger and heavier) engine. Similarly, reducing the turbine diameter may allow an engine with a larger fan to be installed higher up (maintaining intake ground clearance) whilst maintaining adequate vertical depth for a pylon structure arranged to connect the engine to the wing. The engine blockage ratio may be higher than that of known aircraft gas turbine engines.

The engine blockage ratio may be in the range from 2.5 to 4.0, optionally 2.7 to 3.7. The engine blockage ratio may be greater than 2.5. The engine blockage ratio may be greater than 3.0.

The downstream blockage ratio may be as described in the preceding aspect. The downstream blockage ratio may be in the range from 0.2 to 0.3. The downstream blockage ratio may be in the range from 0.20 to 0.30. The downstream blockage ratio may be in the range from 0.20 to 0.29. The downstream blockage ratio may be in the range from 0.22 to 0.28.

The distance between the ground plane and the wing may be measured to centre point of a leading edge of the wing. The distance between the ground plane and the wing may be measured along a line perpendicular to the ground plane and passing through and perpendicular to an axial centre line of the engine.

The turbine diameter at the axial location of the lowest pressure rotor stage may be measured adjacent blade tips of rotor blades of the lowest pressure rotor stage. The turbine diameter at the axial location of the lowest pressure rotor stage may be in the range from 70 cm to 170 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 100 cm to 130 cm, optionally 110 cm to 120 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 120 cm to 170 cm.

The engine length may be measured as the axial distance between a forward region of the fan and a rearward region of the turbine. The engine length may be measured along a centreline of the engine from an axial position of an intersection of the leading edge of each fan blade and the hub of the fan to an axial position of a trailing edge mean radius point of one of the rotor blades provided in the lowest pressure stage of the turbine.

The engine length may be in the range from 200 cm to 500 cm, and optionally from 230 cm to 470 cm, optionally 300 cm to 450 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm (for example 120 cm to 140 cm), the engine length may be in the range from 230 cm to 370 cm, optionally 300 to 360 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm (for example 165 cm to 190 cm), the engine length may be in the range from 370 cm to 470 cm, optionally 390 cm to 450 cm.

The fan tip radius may be measured between an engine centreline and a tip of a fan blade at its leading edge. A fan diameter equal to twice the fan tip radius may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm, 260 cm, 270 cm, 280 cm, 290 cm, 300 cm, 310 cm, 320 cm, 330 cm, 340 cm, 350 cm, 360 cm, 370 cm, 380 cm or 390 cm.

The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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.

A quasi-non-dimensional mass flow rate, Q, may be defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2.
A Q ratio of:
the downstream blockage ratio×quasi non dimensional mass flow rate Q
may be in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−K1/2.

The Q ratio may be in a range from 0.005 Kgs−1N−1K1/2to 0.010 Kgs−1N−1K1/2, from 0.0050 Kgs−1N−1K1/2to 0.0110 Kgs−1N−1K1/2, or optionally from 0.0050 Kgs−1N−1K1/2to 0.0100 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.006 Kgs−1N−1K1/2to 0.009 Kgs−1N−1K1/2.

A specific thrust may be defined as net engine thrust divided by mass flow rate through the engine; and at engine cruise conditions, it may be that:0.029 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2; and70 Nkg−1s≤specific thrust≤110 Nkg−1s.

At cruise conditions, it may be that: 0.032 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2. At cruise conditions, it may be that: 0.033 Kgs−1N−1K1/2≤Q≤0.035 Kgs−1N−1K1/2.

According to another aspect, there is provided an aircraft comprising a wing and a gas turbine engine mounted beneath the wing of the aircraft. The engine has an engine length and comprises:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades of the fan together defining a fan face having a fan tip radius; and

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. A downstream blockage ratio is defined as:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotorground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
An engine blockage ratio of:

(2×the⁢⁢fan⁢⁢tip⁢⁢radius/the⁢⁢engine⁢⁢length)the⁢⁢downstream⁢⁢blockage⁢⁢ratio
is in the range from 2.5 to 4.

The downstream blockage ratio may be in the range from 0.2 to 0.3. The engine may be as described for the preceding aspect.

In another aspect, there is provided a gas turbine engine for an aircraft and arranged to be mounted beneath a wing of the aircraft, the gas turbine engine comprising:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, an annular fan face being defined at a leading edge of the fan; and wherein:

a downstream blockage ratio is defined as the ratio of:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotor⁢⁢stageground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
and a quasi-non-dimensional mass flow rate Q is defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2
and wherein a Q ratio of:
the downstream blockage ratio×Q
is in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2. The values of Q used to calculate the Q ratio as referred to herein may be at cruise conditions.

By defining the Q ratio in this range a large mass flow may be achieved while also minimising the downstream blockage. The Q ratio may be lower than that of a known gas turbine engine. The Q ratio may be in a range from 0.0050 Kgs−1N−1K1/2to 0.0110, or to 0.0100 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.005 Kgs−1N−1K1/2to 0.010 Kgs−1N−1K1/2. The Q ratio may be in a range from 0.006 Kgs−1N−1K1/2to 0.009 Kgs−1N−1K1/2.

The downstream blockage ratio may be in a range from 0.2 to 0.3. The downstream blockage ratio may be in a range from 0.20 to 0.29. The downstream blockage ratio may be in a range from 0.22 to 0.28.

The distance between the ground plane and the wing may be measured to a centre point of a leading edge of the wing. The distance between the ground plane and the wing may be measured along a line perpendicular to the ground plane and passing through and perpendicular to an axial centreline of the engine.

The turbine diameter at the axial location of the lowest pressure rotor stage may be measured adjacent blade tips of rotor blades of the lowest pressure rotor stage. The turbine diameter at the axial location of the lowest pressure rotor stage may be in the range from 70 cm to 170 cm. Optionally, for example for an engine with a fan tip radius in the range from 110 cm to 150 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 100 cm to 120 cm. Optionally, for example for an engine with a fan tip radius in the range from 155 cm to 200 cm, the turbine diameter at the lowest pressure rotor stage may be in the range from 120 cm to 170 cm. The turbine diameter at the lowest pressure rotor stage may be measured at the axial location of blade tip trailing edges of rotor blades of the lowest pressure rotor stage. The turbine diameter of the turbine at the lowest pressure rotor stage may be measured:(i) when the rotor is shrouded, to the underside of the shroud; or(ii) when the rotor is unshrouded, to the blade tips of the rotor.

A specific thrust may be defined as net engine thrust divided by mass flow rate through the engine; and at engine cruise conditions it may be that:0.029 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2; and/or70 Nkg−1s specific thrust≤110 Nkg−1s.

It may be that at cruise conditions: 0.032 Kgs−1N−1K1/2≤Q≤0.036 Kgs−1N−1K1/2. It may be that at cruise conditions: 0.033 Kgs−1N−1K1/2≤Q≤0.035 Kgs−1N−K1/2.

The ratio of the radius of fan blade at its hub to the radius of the fan blade at its tip may be less than 0.33.

At cruise conditions, the specific thrust may be less than 100 Nkg−s. Cruise conditions may be as defined elsewhere herein.

The gas turbine engine may comprise 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.

The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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.

In another aspect, there is provided an aircraft comprising a wing and a gas turbine engine mounted beneath the wing of the aircraft, the engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, an annular fan face being defined at a leading edge of the fan wherein: a downstream blockage ratio is defined as the ratio of:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotor⁢⁢stageground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
and a quasi-non-dimensional mass flow rate Q is defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2
and wherein a Q ratio of:
the downstream blockage ratio×Q
is in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2.
The aircraft may comprise two wings, with one or more gas turbine engines mounted beneath each wing; the or each gas turbine engine may have the features of any of the statements related to the previous aspect.
According to another aspect, there is provided a method of operating an aircraft comprising a gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, an annular fan face being defined at a leading edge of the fan, wherein: a downstream blockage ratio is defined as the ratio of:

the⁢⁢turbine⁢⁢diameter⁢⁢at⁢⁢an⁢⁢axial⁢⁢location⁢⁢of⁢⁢the⁢⁢lowestpressure⁢⁢rotor⁢⁢stageground⁢⁢plane⁢⁢to⁢⁢wing⁢⁢distance
and a quasi-non-dimensional mass flow rate Q is defined as:

Q=W⁢T⁢⁢0P⁢⁢0·Aflow
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa; and
Aflowis the flow area of the fan face in m2
the method comprises controlling the aircraft such that a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 Kgs−1N−1K1/2to 0.011 Kgs−1N−1K1/2.

The gas turbine engine may have the features of any of the statements related to the previous two aspects.

The skilled person will appreciate that a feature described above in relation to any one of the aspects may be applied, mutatis mutandis, to any other aspect of the invention. For example, in various embodiments any two or more of the conditions for ratios as defined above, and optionally all specified ratio ranges, may apply to any given aspect or embodiment. All aspects may apply to an engine of some embodiments. Furthermore, any feature described below may apply to any aspect and/or may apply in combination with any one of the claims.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise 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. The input to the gearbox 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 rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

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, for example one, two or three shafts. Purely by way of example, 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 comprise 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.

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

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a “planetary” or “star” gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 5, for example on 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 previous sentence. The gear ratio may be in the range of from 3.2 to 4.2, and optionally in the range 3.2 to 3.8, 3.3 to 3.8 or 3.4 to 3.7. A higher gear ratio may be more suited to “planetary” style gearbox. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example 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.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on 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 radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390 cm (around 155 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 320 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utipis the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg−1K−1/(ms−1)2). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass ratio may be in the range of from 11 to 20, and optionally in the range from 13 to 20 or 14 to 20.

The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio 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 exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg−1s, 105 Nkg−1s, 100 Nkg−1s, 95 Nkg−1s, 90 Nkg−1s, 85 Nkg−1s or 80 Nkg−1s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kN, 1701N, 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 in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 deg C.), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds) The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein according to any aspect may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured 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 resist impact (for example from 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, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan 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 attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

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 exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

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.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise 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 regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint—in terms of time and/or distance—between top of climb and start of descent). Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i e maintaining a constant altitude and 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 to 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 of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

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 the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein 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 may be mounted in order to provide propulsive thrust.

According to an 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 for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

