Patent Description:
Gas turbine engines with lower specific thrust (and therefore higher flows for a given thrust) would tend to have larger diameter fan outlet guide vane (FOGV) rings and nacelle in order to avoid high peak Mach numbers on or near FOGVs surfaces. Larger diameter FOGVs rings and nacelle may cause a number of disadvantages: higher nacelle drag and weight, lower FOGVs stiffness, and higher weight in fan system to recover such stiffness.

<CIT> relates to a propulsor for a gas turbine engine including a case including a duct disposed along an axis to define a flow path. A rotor includes a row of propulsor blades extending in a generally radial direction outwardly from a hub, the hub rotatable about the axis such that the propulsor blades deliver airflow into the flow path. A row of guide vanes are situated in the flow path. At least two of the guide vanes extend in the generally radial direction between inner and outer surfaces of the duct, extends in a chordwise direction between a first leading edge and a first trailing edge to define a vane chord dimension (VCD) at a first span position of the corresponding guide vane, and defines a vane circumferential pitch (VCP) at the first span position of the corresponding guide vane and an adjacent one of the guide vanes. The row of guide vanes has a vane solidity (VR) defined as VCD/VCP, the vane solidity (VR) being equal to or less than <NUM>.

<CIT> relates to an airfoil of a turbine engine including pressure and suction sides and extending in a radial direction from a <NUM> percent span position to a <NUM> percent span position. The airfoil has a relationship between a camber angle and span position that defines a curve with the camber angle having a positive slope from <NUM> percent span to <NUM> percent span.

According to a first aspect claimed in claim <NUM> there is provided a gas turbine engine for an aircraft comprising:.

It has been found that the lower space-chord ratio at <NUM>% of the span length results in a reduced risk of shock induced buffet, by reducing the local flow speed across the outlet guide vane.

In the present disclosure, the span length is the length of the span of the respective outlet guide vane between the bypass duct inner wall and the bypass duct outer wall.

A point at <NUM>% of the span length corresponds to a radially innermost part of the outlet guide vane at the bypass duct inner wall and a point at <NUM>% of the span length corresponds to a radially outermost part of the outlet guide vane at the bypass duct outer wall.

In the present disclosure, a true chord length for a given cross-section is the distance along a camber line (which would typically be a curved line) between a leading edge and a trailing edge of an aerofoil in that cross-section. Accordingly, the true chord length would typically be the length of a curved line. Note that this is different to what might conventionally be referred to as the chord length, which would be the length of a straight line drawn between the leading edge and the trailing edge of the aerofoil in that cross-section.

Reference to a cross-section through the outlet guide vane at a given percentage along the aerofoil span (or a given percentage span position) in the present disclosure may mean a section through the aerofoil in a plane defined by:.

It may be that at <NUM>% of the span length from the bypass duct inner wall, the space-chord ratio of the at least one outlet guide vane is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that at <NUM>% of the span length from the bypass duct inner wall, the space-chord ratio of the at least one outlet guide vane is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that at <NUM>% of the span length from the bypass duct inner wall, the space-chord ratio of the at least one outlet guide vane is more than <NUM>, or more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. It may be that at <NUM>% of the span length from the bypass duct inner wall, the space-chord ratio of the at least one outlet guide vane is more than <NUM>, or more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>.

The space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than or equal to <NUM> + <NUM> * (span height - <NUM>) where the span height is the proportion of the span length from the bypass duct inner wall at said point. For example, at a point <NUM>% of the span length from the bypass duct inner wall, the span height would be equal to <NUM> and the space-chord ratio would be equal to or less than <NUM>. It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than or equal to x + <NUM> * (span height - <NUM>), where x is <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> and the span height is the proportion of the span length from the bypass duct inner wall at said point. It may be that the space-chord ratio of at least one outlet guide vane at <NUM>% of the span length from the bypass duct inner wall, is at least <NUM> times and/or less than <NUM> times greater than the space-chord ratio of the respective outlet guide vane at <NUM>% of the span length from the bypass duct inner wall.

It may be that the at least one outlet guide vane comprises each and every outlet guide vane of the plurality of outlet guide vanes of the outlet guide vane assembly in the bypass duct.

According to a second aspect there is provided a gas turbine engine for an aircraft comprising:.

wherein a space-chord ratio of at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, wherein the space-chord ratio is defined by an average spacing of outlet guide vanes at a respective span height divided by a true chord length of the respective outlet guide vane at the respective span height, wherein the average spacing of outlet guide vanes at the respective span height is defined as the circumference of the bypass duct at the respective span height (2πr, where r is the radius at the respective span height from a principal rotational axis of the gas turbine engine) divided by the number of outlet guide vanes (NV), (2πr / NV).

It has been found that the lower space-chord ratio at a point from <NUM>% to <NUM>% of the span length results in a reduced risk of shock induced buffet, by reducing the local flow speed across the outlet guide vane.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that every point from <NUM>% to <NUM>% of the span length of the at least one outlet guide vane from the bypass duct inner wall has a space-chord ratio of less than <NUM>. It may be that every point from <NUM>% to <NUM>% of the span length of the at least one outlet guide vane from the bypass duct inner wall has a space-chord ratio of less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

For example, the space-chord ratio of at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, may be in the range of from <NUM> to <NUM>, or from <NUM> to <NUM>. The space-chord ratio of at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, may be in the range of from <NUM> to <NUM>, or from <NUM> to <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. It may be that the space-chord ratio of the at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio of the at least one outlet guide vane, at every point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is more than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than or equal to <NUM> + <NUM> * (span height - <NUM>) where the span height is defined as a proportion of the span length from the bypass duct inner wall at said point. For example, at a point <NUM>% of the span length from the bypass duct inner wall, the span height would be <NUM> and the space-chord ratio would be equal or less than <NUM>. It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>-<NUM>% of the span length from the bypass duct inner wall, is less than or equal to x + <NUM> * (span height - <NUM>), where x is <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> and the span height is the proportion of the span length from the bypass duct inner wall at said point.

According to a third aspect, there is provided a gas turbine engine for an aircraft comprising:.

wherein a space-chord ratio of at least one outlet guide vane at <NUM>% of the span length from the bypass duct inner wall, is at least <NUM> times, and less than <NUM> times, greater than the space-chord ratio of the respective outlet guide vane at <NUM>% of the span length from the bypass duct inner wall, wherein the space-chord ratio is defined by an average spacing of outlet guide vanes at a respective span height divided by a true chord length of the respective outlet guide vane at the respective span height, wherein the average spacing of outlet guide vanes at the respective span height is defined as the circumference of the bypass duct at the respective span height (2πr, where r is the radius at the respective span height from a principal rotational axis of the gas turbine engine) divided by the number of outlet guide vanes (NV), (2πr / NV).

It has also found that having a space-chord ratio at <NUM>% of the span length from the bypass duct inner wall which is at least <NUM> times, and less than <NUM> times, greater than the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall reduces the peak Mach number of flow, in use.

It may be that the space-chord ratio of at least one outlet guide vane at <NUM>% of the span length from the bypass duct inner wall, is at least <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> times greater than the space-chord ratio of the respective outlet guide vane at <NUM>% of the span length from the bypass duct inner wall,
It may be that at <NUM>% of the span length from the bypass duct inner wall, the space-chord ratio of the at least one outlet guide vane is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> times greater than the space-chord ratio of the respective outlet guide vane at <NUM>% of the span length from the bypass duct inner wall.

It may be that the space-chord ratio of the at least one outlet guide vane at <NUM>% of the span length from the bypass duct inner wall is less than <NUM>, preferably less than <NUM>, more preferably less than <NUM>.

It has been found that the lower space-chord ratio at a point between <NUM>-<NUM>%, and/or at <NUM>% and/or <NUM>% of the span length results in a reduced risk of shock induced buffet, by reducing the local flow speed across the outlet guide vane.

It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall, is less than or equal to <NUM> + <NUM> * (span height - <NUM>) where the span height is defined as the proportion of the span length from the bypass duct inner wall at said point. For example, at a point <NUM>% of the span length from the bypass duct inner wall, the span height would be <NUM> and the space-chord ratio would be equal or less than <NUM>. It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>-<NUM>% of the span length from the bypass duct inner wall, is less than or equal to x + <NUM> * (span height - <NUM>), where x is <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> and the span height is the proportion of the span length from the bypass duct inner wall at said point.

In any of the aspects described above, one or more of the following features may apply. It may be that the at least one outlet guide vane comprises a majority of outlet guide vanes of the plurality of outlet guide vanes of the outlet guide vane assembly in the bypass duct.

It may be that the gas turbine engine is in accordance with any of the first, second or third aspect, and it may be that the gas turbine engine further comprises a gearbox configured to receive an input from the core shaft and output 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 a range of from <NUM> to <NUM>, or in the range of from <NUM> to <NUM>, or in the range of from <NUM> to <NUM>.

It may be that the gas turbine engine is in accordance with any of the first, second or third aspect, and it may be that the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;.

In any of the first, second, and third aspect the fan may have a diameter in a range of from <NUM> to <NUM>, or in a range of from <NUM> to <NUM>, or in a range of from <NUM> to <NUM>, or in a range of from <NUM> to <NUM>, or in a range of from <NUM> to <NUM>. In any of the first, second, and third aspect the outlet guide vane assembly may be located axially between the fan and a bifurcation located in the bypass duct. The bifurcation may have a leading edge and a trailing edge and the outlet guide vane assembly may be located axially between the fan and bifurcation trailing edge, or between the fan and the bifurcation leading edge.

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 by the first core shaft, and not the second core shaft, in the example above).

The gearbox may be 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 <NUM>, for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, for example on the order of or at least <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from <NUM> or <NUM> to <NUM>. In some arrangements, the gear ratio may be outside these ranges.

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 <NUM>% span position, to a tip at a <NUM>% 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. 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), for example in the range of from <NUM> to <NUM>. 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches) or <NUM> (around <NUM> 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), for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, or <NUM> to <NUM>.

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 <NUM> rpm, for example less than <NUM> rpm, or less than <NUM> rpm, or less than <NUM> rpm, or less than <NUM> 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 <NUM> to <NUM> (for example <NUM> to <NUM> or <NUM> to <NUM>) may be in the range of from <NUM> rpm to <NUM>, for example in the range of from <NUM> rpm to <NUM> rpm, or in the range of from <NUM> rpm to <NUM> rpm, or in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> 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 <NUM> to <NUM> may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> 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 <NUM> on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the <NUM>-D average enthalpy rise) across the fan and Utip is 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all values being dimensionless). 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), for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>.

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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. 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), for example in the range of form <NUM> to <NUM>, <NUM> to <NUM>, or <NUM> to <NUM>. 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. 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), for example in the range of from <NUM> to <NUM>.

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: <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s or <NUM> Nkg-<NUM>s. 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), for example in the range of from <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s, or <NUM> Nkg-<NUM>s to <NUM> Nkg-<NUM>s. 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: 130kN, 135kN, 140kN, 145kN, 150kN, 155kN, 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. 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). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 155kN to 170kN, 330kN to <NUM> kN, or 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> degrees C (ambient pressure <NUM>. 3kPa, temperature <NUM> degrees 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. 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), for example <NUM> to <NUM>. The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. 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), for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>. 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 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 bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a 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 <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> 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 <NUM>% 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 <NUM> 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 <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example on the order of Mach <NUM>, on the order of Mach <NUM> or in the range of from <NUM> to <NUM>. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach <NUM> or above Mach <NUM>.

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 <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM> (around <NUM> ft), for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> (around <NUM> ft) to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example on the order of <NUM>. 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 a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (<NUM>). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of from 20kN to 40kN.

Purely by way of further example, the cruise conditions may correspond to a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (<NUM>). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of from 35kN to 65kN.

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 <NUM> or <NUM>) 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 engine <NUM> comprises an air intake <NUM> and a propulsive fan <NUM>, comprising a plurality of blades, that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine <NUM> comprises a core <NUM> that receives the core airflow A downstream of the fan. A nacelle <NUM> surrounds the gas turbine engine <NUM> and defines a bypass exhaust nozzle <NUM>. The engine <NUM> comprises a bypass duct <NUM>, located radially outwardly from the core and downstream of the fan, and delimited by a bypass duct inner wall <NUM> and a bypass duct outer wall <NUM>. The fan <NUM> is attached to and driven by the low pressure turbine <NUM> via a core shaft <NUM> and an epicyclic gearbox <NUM>. The fan <NUM> has a fan diameter in a range of from <NUM> to <NUM>; in examples the fan <NUM> has a fan diameter in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, or <NUM> to <NUM>.

An outlet guide vane assembly <NUM> is located in the bypass duct <NUM>, and comprises a plurality of outlet guide vanes 46a, which are distributed circumferentially within the bypass duct <NUM>. Each outlet guide vane 46a extends radially between the bypass duct inner wall <NUM> and the bypass duct outer wall <NUM>, along a span, and at least partially axially from a leading edge <NUM> to a trailing edge <NUM>. The length of the outlet guide vane 46a along the span (i.e., between the bypass duct inner wall <NUM> and the bypass duct outer wall <NUM>) is referred to herein as the span length.

A bifurcation <NUM> (e.g., a flow splitter) is disposed in the bypass duct <NUM> downstream of the outlet guide vane assembly <NUM>, such that the outlet guide vane assembly <NUM> is disposed axially between the fan <NUM> and the bifurcation <NUM>. The bifurcation <NUM> has a leading edge and a trailing edge and the outlet guide vane assembly <NUM> is disposed axially between the fan <NUM> and the bifurcation leading edge. In other embodiments, the outlet guide vane assembly <NUM> is disposed axially between the fan <NUM> and the bifurcation trailing edge. The bifurcation <NUM> may be an upper bifurcation; the engine <NUM> may further comprise a lower bifurcation arranged in the bypass duct <NUM> downstream of the outlet guide vane assembly <NUM> and radially opposite to the upper bifurcation <NUM>. <FIG> shows a cross-sectional view of an outlet guide vane 46a at a particular point along its span. A true chord length of the outlet guide vane 46a, for a given cross-section at a point along the span length, is the distance along a camber line <NUM> (which would typically be a curved line) between the leading edge <NUM> and the trailing edge <NUM> in that cross-section. Accordingly, the true chord length would typically be the length of a curved line. Note that this is different to what might conventionally be referred to as the chord length, which would be the length of a straight line drawn between the leading edge and the trailing edge in that cross-section.

Peak surface Mach numbers in use across the outlet guide vanes 46a is typically reduced, to avoid aerodynamic buffet, by increasing the span length, which increases weight and drag of the nacelle. In this invention, the cross-sectional geometry of each outlet guide vane 46a along the span length is designed to reduce peak surface Mach numbers, without the need to increase the span length. The cross-sectional geometry of each outlet guide vane is defined in relation to its space-chord ratio (s/c), which varies along its span. A profile of space-chord ratio (s/c) against percentage of span length from the bypass duct inner wall <NUM> may be substantially identical for each outlet guide vane 46a in the outlet guide vane assembly <NUM>. In other examples, a plurality of outlet guide vanes in the outlet guide vane assembly <NUM> may comprise an identical profile, and a few outlet guide vanes may comprise alternative profiles for providing additional support, for accommodating other support structures in the bypass duct <NUM>, or for different camber standards (including directing the flow around the bifurcation(s) <NUM> whilst limiting incidence onto the vanes).

The space-chord ratio (s/c) is defined by an average spacing of outlet guide vanes 46a at a respective span height divided by the true chord length of the respective outlet guide vane 46a at the respective span height. The average spacing of outlet guide vanes at the respective span height is defined as the circumference of the bypass duct at the respective span height (2πr, where r is the radius at the respective span height from the centre of the gas turbine engine <NUM> (i.e., the principal rotational axis <NUM>)) divided by the number of outlet guide vanes (Nv). Therefore, the space-chord ratio at a respective span height is defined as 2πr / Nv.

<FIG> shows a chart <NUM> of space-chord ratio (s/c) against percentage height along the span length from the bypass duct inner wall <NUM>. For example, at <NUM>% of the span length from the bypass duct inner wall <NUM>, the percentage height is <NUM>%.

Two plots are shown on the chart <NUM>. A first plot <NUM> (shown with a dotted line) shows a relationship between the space-chord ratio and the percentage height of a previously considered outlet guide vane. A second plot <NUM> (shown in a solid line) is a first example relationship between the space-chord ratio and the percentage height of an outlet guide vane 46a according to the invention. It can be seen that the space-chord ratio between approximately <NUM>-<NUM>% of the span length is lower than the space-chord ratio of the previously considered design. The applicant has found that this lower space-chord ratio between <NUM>-<NUM>% of the span length (resulting from a longer true chord length at this span height, giving a bulged profile shown in the axial cross-section of <FIG>) results in lower peak Mach numbers of flow, in use.

In this example, it can be seen that the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall is approximately <NUM>, and that the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall is approximately <NUM>. It can be seen that every point between <NUM>-<NUM>% of the span length, the space-chord ratio is less than <NUM> in this example.

The applicant has found more broadly that having a space-chord ratio of less than <NUM> at <NUM>% of the span length and/or of less than <NUM> at <NUM>% of the span length reduces the peak Mach number of flow in use. Without being bound by theory, the applicant believes that this is due to the lift force being distributed over a larger surface area. The applicant has found that there is a trade-off between (i) added weight for each outlet guide vane 46a (due to increasing the chord length locally) and increased losses over the outlet guide vanes against (ii) reduced peak Mach numbers when reducing the space-chord ratio between <NUM>-<NUM>% of the span length. The longer the true chord length between <NUM>-<NUM>% of the span length, the smaller the space-chord ratio (s/c) and the lower the peak Mach number, but the more material is required to on either side of this range for sufficient stiffness of the entire outlet guide vane, increasing the weight of each outlet guide vane further.

The space-chord ratio at <NUM>% of the span length may therefore be less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>, or even lower. In some examples, the space-chord ratio at <NUM>% of the span height may be larger than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> or higher. The space-chord ratio at <NUM>% of the span length may therefore be less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>, or even lower. In some examples, the space-chord ratio at <NUM>% of the span length may be larger than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> or higher.

The applicant has also found that, having a space-chord ratio of at least one point between <NUM>% and <NUM>% of the span length less than <NUM> also reduces the peak Mach number of flow, in use. In other examples, the space-chord ratio anywhere between <NUM>% and <NUM>% of the span length may be less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>, or even lower. It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall <NUM>, is less than <NUM>, <NUM>, <NUM>, or <NUM> or even lower. In some examples, each and every point between <NUM>-<NUM>% of the span length from the bypass duct inner wall <NUM> may have a space-chord ratio less than <NUM>, or less than <NUM>, <NUM>, <NUM>, <NUM> or <NUM> or even lower. It may be that the space-chord ratio of the at least one outlet guide vane, at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall <NUM>, is less than <NUM>, <NUM>, <NUM>, or <NUM> or even lower. In some examples, each and every point between <NUM>% and <NUM>% of the span length from the bypass duct inner wall <NUM> may have a space-chord ratio less than <NUM>, or less than <NUM>, <NUM>, <NUM>, <NUM> or <NUM> or even lower.

In other examples the space-chord ratio of at least one point between <NUM>% and <NUM>% of the span length from the bypass duct inner wall <NUM>, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

In other examples the space-chord ratio anywhere between <NUM>% and <NUM>% of the span length from the bypass duct inner wall, is less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>.

The applicant has also found that a space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>-<NUM>% of the span length from the bypass duct inner wall <NUM>, being less than or equal to <NUM> + <NUM> * (span height - <NUM>), where the span height is the proportion of the span length from the bypass duct inner wall <NUM>, reduces peak Mach numbers of flow in use. For example, at a point <NUM>% of the span length from the bypass duct inner wall, the span height would be <NUM> and the space-chord ratio would less than or equal to <NUM>. This is shown as the dash-dot line <NUM> in <FIG>. It may be that the space-chord ratio (s/c) of the at least one outlet guide vane, at a point from <NUM>-<NUM>% of the span length from the bypass duct inner wall, is less than or equal to x + <NUM> * (span height - <NUM>), where x is <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> or even lower. The applicant has also found that having a space-chord ratio at <NUM>% of the span length from the bypass duct inner wall <NUM> which is at least <NUM> times and/or less than <NUM> times greater than the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall <NUM> reduces the peak Mach number of flow, in use. In some examples, the space-chord ratio at <NUM>% may be a factor of <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or even higher, larger than the space-chord ratio at <NUM>% of the span length. In the example of <FIG> the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall <NUM> is <NUM> times greater than the space-chord ratio at <NUM>% of the span length from the bypass duct inner wall <NUM>. In some examples, the space-chord ratio at <NUM>% of the span length of the outlet guide vane may be less than <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> times greater than the space-chord ratio of the respective outlet guide vane at <NUM>% of the span length from the bypass duct inner wall <NUM>.

Referring back to <FIG>, in use, the core airflow A is accelerated and compressed by the low pressure compressor <NUM> and directed into the high pressure compressor <NUM> where further compression takes place. The high pressure turbine <NUM> drives the high pressure compressor <NUM> by a suitable second core shaft <NUM>.

The low pressure turbine <NUM> (see <FIG>) drives the core shaft <NUM>, which is coupled to a sun wheel, or sun gear, <NUM> of the epicyclic gear arrangement <NUM>.

Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan <NUM>) respectively and/or the turbine and compressor stages that are connected together by the core shaft <NUM> with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan <NUM>).

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

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

Claim 1:
A gas turbine engine (<NUM>) for an aircraft comprising:
an engine core (<NUM>) comprising a turbine (<NUM>), a compressor (<NUM>), and a core shaft (<NUM>) connecting the turbine to the compressor;
a fan (<NUM>) located upstream of the engine core (<NUM>), the fan (<NUM>) comprising a plurality of fan blades;
a bypass duct (<NUM>) delimited by a bypass duct inner wall (<NUM>) and a bypass duct outer wall (<NUM>) and located radially outwardly from the engine core (<NUM>) and downstream of the fan (<NUM>); and
an outlet guide vane assembly (<NUM>), located within the bypass duct (<NUM>) and, comprising a plurality of outlet guide vanes (46a) distributed circumferentially within the bypass duct (<NUM>), each outlet guide vane (46a) extending radially along a span between the bypass duct inner wall (<NUM>) and the bypass duct outer wall (<NUM>),
characterised in that,
a space-chord ratio (s/c) of at least one outlet guide vane (46a), at <NUM>% of the span length from the bypass duct inner wall (<NUM>), is less than <NUM>,
wherein the space-chord ratio (s/c) is defined by an average spacing of outlet guide vanes (46a) at a respective span height divided by a true chord length of the respective outlet guide vane (46a) at the respective span height, wherein the average spacing of outlet guide vanes (46a) at the respective span height is defined as the circumference of the bypass duct (<NUM>) at the respective span height (2πr, where r is the radius at the respective span height from a principal rotational axis (<NUM>) of the gas turbine engine) divided by the number of outlet guide vanes (Nv), (2πr / Nv), and
wherein a space-chord ratio (s/c) of the at least one outlet guide vane (46a), at a point from <NUM>% to <NUM>% of the span length from the bypass duct inner wall (<NUM>), is less than<MAT>
wherein the span height is defined as a proportion of the span length from the bypass duct inner wall (<NUM>) at said point.