Patent Description:
The design of a gas turbine engine must balance a number of competing factors. In general, it is desirable to minimize fuel burn and weight. However, gas turbine engines have been used and developed for many years, and so the underlying designs are mature. This high level of design maturity means that advances in, for example, the reduction of fuel burn and/or weight have been relatively small and incremental over recent years.

It is desirable to improve the rate of development of gas turbine engines.

European patent application <CIT> discloses a turbofan engine including a fan having a plurality of rotatable fan blades and defining a fan pressure ratio during operation of the turbofan engine. The turbofan engine also includes a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath. The turbofan also includes an outer nacelle at least partially surrounding the fan and the turbomachine, the outer nacelle defining a bypass passage with the turbomachine. A bypass ratio of an amount of airflow through the bypass passage to an amount of airflow through the core air flowpath during operation of the turbofan is less than or equal to about <NUM> and wherein the fan pressure ratio is less than or equal to about <NUM>.

US patent application <CIT> discloses a Ceramic Matrix Composite (CMC) airfoil segment for a gas turbine engine which includes a box-shape fibre geometry which defines a rectilinear pressure side bond line and a rectilinear suction side bond line.

US patent application <CIT> discloses a gas turbine engine which includes an engine case, a retention block attached to the engine case, and a blade outer air seal (BOAS). The BOAS includes a plurality of layers formed of a ceramic matrix composite (CMC) material. At least one of the plurality of layers provides a slot receiving a portion of the retention block.

US patent application <CIT> discloses a turbofan engine which includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in the ambient air and a power turbine that converts the gas stream flow into shaft power. The turbofan engine further includes a propulsor section including a fan driven by the power turbine through a geared architecture at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. An Engine Unit Thrust Parameter defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about <NUM> at a take-off condition.

US patent application <CIT> discloses a gas turbine engine for an aircraft which includes a compressor section where at least one of the airfoil members defines a vane exit vector extending tangentially from a curved surface of the airfoil member adjacent a trailing edge of the airfoil member, a projection of the vane exit vector in a longitudinal plane perpendicular to a radial direction of the engine extending at an airfoil angle from the longitudinal axis. A bleed slot defined through the casing wall and providing fluid communication between the core air passage and the bleed duct extends along a slot axis. A projection of the slot axis in the longitudinal plane extends at a slot angle with respect to the longitudinal axis. The slot angle is different from the airfoil angle.

European patent application <CIT> discloses a gas turbine engine comprising a fan section that includes a fan and a compressor section including a low pressure compressor. A turbine section includes a fan drive turbine for driving the fan and the low pressure compressor. A speed reduction device connects the fan drive turbine to the fan and the low pressure compressor. The speed reduction device includes a sun gear driven by an inner shaft. A plurality of intermediate gears surround the sun gear. A carrier supports the plurality of intermediate gears for driving the low pressure compressor. A ring gear is located radially outward from the intermediate gears and includes a forward portion for driving a fan drive shaft and an aft portion.

US patent application <CIT> discloses an industrial gas turbine engine with a high spool and a low spool in which low pressure compressed air is supplied to the high pressure compressor, and where a portion of the low pressure compressed air is bled off for use as cooling air for hot parts in the high pressure turbine of the engine. Annular bleed off channels are located in the LPC diffuser. The bleed channels bleed off around <NUM>% of the core flow and pass the bleed off air into a cooling flow channel that then flows into the cooling circuits in the turbine hot parts.

US patent application <CIT> discloses a gas turbine engine which includes a shaft defining an axis of rotation and a fan driving turbine configured to drive the shaft. The fan driving turbine comprises a plurality of stages that are spaced apart from each other along the axis. Each stage includes a turbine disk comprised of a disk material and a plurality of turbine blades comprised of a blade material. The disk material and the blade material for one of the plurality of stages is selected based on a location of the one stage relative to the other stages of the plurality of stages.

According to the invention there is provided gas turbine engine for an aircraft according to independent claim <NUM>, comprising:
an engine core comprising:.

The second turbine comprises at least one ceramic matrix composite component.

The mass of ceramic matrix composite in the second turbine is in the range of from <NUM>% to <NUM>% of the total mass of the second turbine.

The turbine entry temperature, defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is at least <NUM>; characterised in that
the second turbine (<NUM>) comprises first and second rotor blade and stator vane rows, wherein the most axially upstream row of stator vanes (<NUM>) and rotor blades (<NUM>) of the second turbine (<NUM>) are metallic, whilst the second most axially upstream row of stator vanes (<NUM>) and rotor blades (<NUM>) of the second turbine (<NUM>) comprise Ceramic Matrix Composites (CMCs).

Conventionally, components in a turbine section of a gas turbine engine have been manufactured using a metal alloy, such as a nickel alloy. However, in order to achieve greater engine efficiency, it has been found to be desirable to increase the temperature of the core gas flow entering into the turbine from the combustor. Typically, in operation, the temperature of the gas flowing past some of the components in the turbine is near to or above the melting point of those components. Thus, in order to ensure that such components have sufficient operating life, they require significant cooling. Such cooling is typically provided using air from the compressor that bypasses the combustor. The cooling flow that bypasses the combustor results in reduced engine efficiency, because that flow is simply compressed in the compressor and then expanded through the turbine.

Furthermore, in order to minimize the amount of cooling flow that is used, and thus minimize the impact on engine efficiency, the cooling flow must be used as efficiently as possible. For example, the cooling passages used to cool such turbine components are typically intricate, requiring extensive design and complex manufacturing techniques. This significantly increases the cost of the gas turbine engine.

Still further, the cooling system itself adds mass to the engine.

The present inventors have understood that a gas turbine engine can be improved through selective use of ceramic matrix composites (CMCs) in its turbine. In particular, the inventors have understood that whilst the CMCs may be advantageous in some areas of the turbine, their use may not actually be appropriate in all areas. Through this understanding, the inventors have derived the optimum level of CMC use in the turbine to be in the claimed ranges. For example, whilst the thermal capability of CMCs - which is typically higher than their metallic counterparts - lends itself to use in some areas, the present inventors have understood that the reduced thermal conductivity of CMCs (compared to an equivalent metallic component) means that they are not suitable in some other areas. Purely by way of non-limitative example, the very hottest parts of the turbine may experience temperatures that exceed even the capability of CMCs, and thus still require a degree of cooling flow. In such a case, it may be more appropriate to use a metal than a CMC, due to the greater thermal conductivity of metals potentially improving the effectiveness of the cooling flow in removing heat from the component.

The inventors of the present disclosure have found that taking such considerations into account, the optimum mass of ceramic matrix composite in the high pressure (or second) turbine as a percentage of the total mass of the high pressure (or second) turbine may be within the ranges described and/or claimed herein. Additionally or alternatively, the optimum mass of ceramic matrix composite in the high pressure (or second) and low pressure (or first) turbine combined as a percentage of the total mass of the high pressure (or second) and low (or first) pressure turbine may be in the ranges described and/or claimed herein.

Purely by way of example, the CMC may be SiC-SiC (i.e. silicon carbide fibres in a silicon carbide matrix). However, it will be appreciated that any suitable CMC may be used, and indeed the turbine may comprise more than one composition of CMC (for example having different elements). Any suitable manufacturing method may be used for the CMC, such as a vapour deposition process or a vapour infusion process.

The turbine may comprise stator vanes, rotor blades, seal segments (which together may be said to form a generally annular ring radially outside the rotor blades), rotor discs (on which rotor blades are provided), one or more radially inner casing elements and one or more radially outer casing elements. The turbine mass may be the total mass of all such turbine components.

The minimum mass of ceramic matrix composite in the second turbine may be <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>% or <NUM>% of the total mass of the second turbine. The maximum mass of ceramic matrix composite in the second turbine may be <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>% or <NUM>% of the total mass of the second turbine. The mass of ceramic matrix composite in the second turbine as a percentage of the total mass of the second turbine may be in a range having any of the minimum percentages listed above as a lower bound and any compatible maximum percentage listed above as an upper bound.

The second turbine may be said to be axially upstream of the first turbine. The first turbine may comprise at least one ceramic matrix composite component. The mass of ceramic matrix composite in the first and second turbines may be in the range of from <NUM>% to <NUM>%, optionally <NUM>% to <NUM>%, of the total mass of the first and second turbines.

The minimum mass of ceramic matrix composite in the first and second turbines may be <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>% or <NUM>% of the total mass of the first and second turbines. The maximum mass of ceramic matrix composite in the first and second turbines may be <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>%, <NUM>% or <NUM>% of the total mass of the first and second turbine. The mass of ceramic matrix composite in the first and second turbines as a percentage of the total mass of the first and second turbines may be in a range having any of the minimum percentages listed above as a lower bound and any compatible maximum percentage listed above as an upper bound.

As noted above, the percentages of CMCs used in the turbine described and claimed herein are based on the present inventors' insight into the most appropriate components for which to use CMCs, taking into account, inter alia, the temperature variation though the turbine. Non-limitative examples are provided below of metallic and CMC components in the gas turbine engine.

The second turbine comprises at least two rows of stator vanes. The most axially upstream row of stator vanes is metallic. The most axially upstream row of stator vanes may be directly downstream of the combustor. For example, there may be no rotor blades between the combustor and the stator vanes.

The terms "upstream" and "downstream" are used herein in the conventional manner, i.e. with respect to the flow through the engine in normal use. Thus, for example, the compressor and combustor are in the upstream direction relative to the turbine.

The second turbine comprises at least two rows of rotor blades. The most axially upstream row of rotor blades is metallic. The most axially upstream row of rotor blades may be directly downstream of the most axially upstream row of stator vanes. The most axially upstream row of rotor blades and/or the most axially upstream row of stator vanes may comprise one or more internal cooling passages and/or film cooling holes. Such internal cooling passages and/or film cooling holes may be supplied with cooling flow from the compressor that has bypassed the combustor.

A CMC component may or may not be provided with internal cooling passages and/or film cooling holes.

The most axially upstream row of rotor blades in the turbine is a part of the second turbine. The most axially upstream row of stator vanes in the turbine is a part of the second turbine.

The most axially upstream row of rotor blades in the turbine may be radially surrounded by seal segments. Such seal segments may comprise a ceramic matrix composite.

In general, the seal segments may form the radially outer boundary (which may be annular and/or frusto-conical) inside which the turbine blades rotate in use. The radially outer tips of the turbine blades may be adjacent the radially inner surface of the seal segments.

The second turbine comprises at least two rows of stator vanes. The second most axially upstream row of stator vanes (which may be directly axially downstream of the upstream most row of rotor blades) comprises a ceramic matrix composite.

The second turbine comprises at least two rows of rotor blades. The second most axially upstream row of rotor blades comprises a ceramic matrix composite.

The second most axially upstream row of rotor blades in the turbine is a part of the second turbine. The second most axially upstream row of stator vanes in the turbine may be a part of the second turbine.

The second most axially upstream row of rotor blades may be radially surrounded by ceramic matrix composite seal segments.

The second turbine comprises at least two stator vane rows (for example <NUM>, <NUM>, <NUM>, <NUM> or <NUM>), and one or more of which may comprise a ceramic matrix composite. The second turbine comprises at least two rotor blade rows and/or surrounding seal segments (for example <NUM>, <NUM>, <NUM>, <NUM> or <NUM>), and one or more of which may comprise a ceramic matrix composite.

The axially most upstream row of stator vanes in the first turbine (which may be directly downstream of the axially most downstream row of rotor blades in the second turbine) may comprise a ceramic matrix composite.

The axially most upstream row of rotor blades in the first turbine may comprise a ceramic matrix composite. The axially most upstream row of rotor blades in the first turbine may be surrounded by ceramic matrix composite seal segments.

In any aspect of the present disclosure, any one or more rotor blade, stator vane or seal segment (i.e. seal portion that forms at least a part of the radially outer flow path around a row of rotor blades) that experiences a maximum temperature a maximum power condition at which the engine is certified (which may be commonly known as the "red-line" condition) in the range of from <NUM> to <NUM> - for example in a range having a lower bound of <NUM>, <NUM> or <NUM> and an upper bound of <NUM>, <NUM>, <NUM> or <NUM> - may be manufactured using a CMC. In some arrangements, most, or even all, rotor blades experiencing "red-line" temperatures within such ranges may be manufactured using a CMC. In some arrangements, most, or even all, stator vanes experiencing "red-line" temperatures within such ranges may be manufactured using a CMC. In some arrangements, most, or even all, seal segments experiencing "red-line" temperatures within such ranges may be manufactured using a CMC. Rotor blades, stator vanes and seal segments that do not experience "red-line" temperatures in such ranges may be manufactured using a metal, such as a nickel alloy.

As noted elsewhere herein, the present disclosure relates to a gas turbine engine. Such a gas turbine engine may be said to comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor.

As noted elsewhere herein, the gas turbine engine comprises a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

Purely by way of example, the turbine connected to the core shaft that drives the gearbox may be a first turbine, the compressor connected to the core shaft that drives the gearbox may be a first compressor, and the core shaft that drives the gearbox may be a first core shaft.

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 <NUM>, for example in the range of from <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.

In any gas turbine engine as described and/or claimed herein, a combustor is provided axially downstream of the fan and compressor(s). The flow at the exit to the combustor is provided to the inlet of the second turbine, where a second turbine is provided. The combustor is provided" upstream of the turbine(s).

The first turbine may comprise any number of stages, for example multiple stages.

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). 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> (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) cm, <NUM> (around <NUM> inches) or <NUM>. 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 <NUM> rpm, for example 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>) 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. 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 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 units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). 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: <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). 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).

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). 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: 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). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> deg C (ambient pressure <NUM>. 3kPa, temperature <NUM> deg C), with the engine static.

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 conditions. 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 an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (<NUM>). 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 50kN to 65kN) at a forward Mach number of <NUM> and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (<NUM>).

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.

The engine core <NUM> comprises, in axial flow series, a low pressure compressor <NUM> (which may be referred to herein as a first compressor <NUM>), a high-pressure compressor <NUM> (which may be referred to herein as a second compressor), combustion equipment <NUM>, a high-pressure turbine <NUM> (which may be referred to herein as a second turbine), a low pressure turbine <NUM> (which may be referred to herein as a first turbine) and a core exhaust nozzle <NUM>.

The planet carrier <NUM> constrains the planet gears <NUM> to process around the sun gear <NUM> in synchronicity whilst enabling each planet gear <NUM> to rotate about its own axis.

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, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

<FIG> shows a part of the turbine in greater detail. In particular, <FIG> shows a downstream portion of the combustor <NUM>, the second (high pressure) turbine <NUM>, and an upstream portion of the first (low pressure) turbine <NUM>. The high pressure turbine <NUM> is connected to the second core shaft <NUM>. The low pressure turbine <NUM> is connected to the first core shaft <NUM>.

In the illustrated example, the high pressure turbine <NUM> comprises, in axial-flow series, a first (most axially upstream) stator vane row <NUM>, a first (most axially upstream) rotor blade row <NUM>, a second (second most axially upstream) stator vane row <NUM>, and a second (second most axially upstream) rotor blade row <NUM>.

The first rotor blade row <NUM> is connected to a rotor disc <NUM>. The second rotor blade row <NUM> is connected to a rotor disc <NUM>. The two rotor discs <NUM>, <NUM> are rigidly connected together by a link member <NUM>. At least one of the rotor discs (in the illustrated example the first rotor disc <NUM>) is connected to the second core shaft <NUM> via an arm <NUM>. Accordingly, in use, the second core shaft <NUM>, rotor discs <NUM>, <NUM> and rotor blades <NUM>, <NUM> all rotate together, at the same rotational speed.

The gas turbine engine <NUM> also comprises seal segments <NUM> provided radially outside the first rotor blade row <NUM>. The gas turbine engine <NUM> also comprises seal segments <NUM> provided radially outside the second rotor blade row <NUM>. The seal segments <NUM>, <NUM> form the radially outer flow boundary (which may be referred to as the radially outer annulus line) in the region of the respective rotor blade row <NUM>, <NUM>, for example over the axial extent of the tips of the rotor blades <NUM>, <NUM>. The seal segments <NUM>, <NUM> may form a seal with the tips of the rotor blades to prevent - or at least restrict - flow passing over or past the tips of the rotor blades. The seal segments <NUM>, <NUM> may be abradable by the rotor blades. Thus, for example, the seal segments <NUM>, <NUM> may be abraded by the rotor blades in use so as to form an optimal seal therebetween. Each segment may form an annular segment or a frusto-conical segment.

In the illustrated example, the high pressure turbine <NUM> is a two-stage high pressure turbine, in that it comprises two stages of vanes and blades, each stage comprising a stator vane row followed by a rotor blade row. However, it will be appreciated that gas turbine engines <NUM> in accordance with the present disclosure may comprise a high pressure turbine with any number of stages, for example <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or more than <NUM> stages of stator vanes and rotor blades.

The low pressure turbine <NUM> is provided downstream of the high pressure turbine <NUM>. An axially most upstream row of stator vanes <NUM> in the low pressure turbine <NUM> is provided immediately downstream of the final row of rotor blades <NUM> of the high pressure turbine <NUM>. An axially most upstream row of rotor blades <NUM> in the low pressure turbine <NUM> is provided immediately downstream of the axially most upstream row of stator vanes <NUM>. The axially most upstream row of rotor blades <NUM> is connected to the first core shaft <NUM> via a rotor disc. In use, the rotor blades <NUM> of the low pressure turbine <NUM> drive the first core shaft <NUM>, which in turn drives the low pressure compressor <NUM>, and also drives - via a gearbox <NUM> - the fan <NUM>.

<FIG> only shows an upstream portion of the low pressure turbine <NUM>. However, it will be appreciated that downstream of the illustrated portion there may be provided further rows of stator vanes and rotor blades. For example, the low pressure turbine <NUM> may comprise <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or more than <NUM> stages of stator vanes and rotor blades. The axially most upstream row of rotor blades <NUM> are connected to one or more (not shown) downstream rotor blade rows via a linkage <NUM> that is connected to the disc <NUM> on which the rotor blades <NUM> are supported.

At least a part of the high pressure turbine <NUM> and/or the low pressure turbine <NUM> comprises a CMC in the illustrated example. Purely by way of example, the CMC material may be silicon carbide fibres and/or a silicon carbide matrix (SiC-SiC), although it will be appreciated that other CMCs may be used, such as an oxide-oxide (Ox-Ox CMC material), a monolithic ceramic, and/or the like.

As noted elsewhere herein, CMCs have different properties to conventional turbine materials, such as nickel alloys. For example, CMCs typically have lower density and are able to withstand higher temperatures than metals such as nickel alloys. The present inventors have understood that these properties can be useful in some areas of the turbine <NUM>, <NUM>, but other properties - such as lower thermal conductivity of CMCs compared to nickel alloys - mean that their use is not appropriate in all areas of the turbine <NUM>, <NUM>.

For example, depending on the type of engine (for example in terms of architecture and/or maximum thrust), any one or more of the first (most axially upstream) stator vane row <NUM>, first (most axially upstream) rotor blade row <NUM>, second (second most axially upstream) stator vane row <NUM>, second (second most axially upstream) rotor blade row <NUM> and first or second set of seal segments <NUM>, <NUM> of the high pressure turbine may be manufactured using CMCs. Components in the above list that are not manufactured using CMCs may be manufactured using a metal, such as a nickel alloy. Optionally, in any aspect or arrangement described and/or claimed herein and regardless of the number of stages in the high pressure turbine <NUM>, the rotor blades of each stage in the high pressure turbine <NUM> may be surrounded by seal segments, and the seal segments surrounding any one or more stage (for example all stages) may be made from a CMC.

According to the invention, in the <FIG> arrangement, the second stator vane row <NUM>, second rotor blade row <NUM> and optionally a first set of seal segments <NUM> and second set of seal segments <NUM> of the high pressure turbine are manufactured using CMCs, whereas the first stator vane row <NUM> and the first rotor blade row <NUM> are manufactured using a nickel alloy. In this particular example, the temperature experienced by the first stator vane row <NUM> and the first rotor blade row <NUM> may be even higher than that which can be tolerated by CMCs. Accordingly, for this particular example, this means that the first stator vane row <NUM> and the first rotor blade row <NUM> - which experience higher temperatures than downstream components due to their proximity to the combustor exit <NUM> - can take advantage of the relatively high thermal conductivity of the nickel alloy so as to be cooled more effectively using cooling air (taken from the compressor, for example) which may be provided to passages running through the components.

The total mass of the high pressure turbine <NUM> may include the masses of the stator vanes <NUM>, <NUM>, rotor blades <NUM>, <NUM>, seal segments <NUM>, <NUM>, rotor discs <NUM>, <NUM>, one or more radially inner casing elements that form the inner flow boundary <NUM> over the axial extent of the high pressure turbine <NUM>, and one or more radially outer casing elements that form the outer flow boundary <NUM> over the axial extent of the high pressure turbine <NUM>.

CMCs may be used in appropriate parts of the low pressure turbine <NUM>, although in some engines <NUM> their use in the low pressure turbine <NUM> may not be appropriate, and thus they may not be used. Purely by way of non-limitative example, in the <FIG> arrangement, the axially most upstream row of stator vanes <NUM> is manufactured using a CMC, whereas the axially most upstream row of rotor blades <NUM> is manufactured using a metal alloy (such as a nickel alloy). In this particular example, the temperature experienced by the axially most upstream row of rotor blades <NUM> may not be sufficiently high to benefit from the use of CMCs, although it will be appreciated that in other engines <NUM> in accordance with the present disclosure, the axially most upstream row of rotor blades <NUM> and/or the associated seal segments <NUM> may be manufactured using CMCs. Indeed, in some engines, components (such as vanes, blades and seals) downstream of the axially most upstream row of rotor blades <NUM> in the low pressure turbine <NUM> may be manufactured using CMCs.

Any component manufactured using CMCs may also be provided with an environmental barrier coating (EBC). Such an EBC may cover at least the gas washed surface of such components. Such an EBC may protect the CMC from environmental deterioration, for example deterioration due to the effects of water vapour. Such an EBC may be, for example ytterbium disilicate (Yb<NUM>Si<NUM>O<NUM>), which may be applied by any suitable method, such as air plasma spray.

As noted elsewhere herein, CMCs have a higher temperature capability than conventional materials, such as metal alloys. This means that selective use of CMCs in the turbine can mean that some components that would need to be cooled if they were to be made from a metal alloy do not need to be cooled because they are made from a CMC and/or some components manufactured using a CMC require less cooling than if they were to be made from a metal alloy. Additionally or alternatively, through use of CMCs it may be possible to expose some components to a higher temperature than would otherwise be possible.

Purely by way of non-limitative example, optimizing the use of CMCs in the engine (for example in one or more components of the turbine <NUM>, <NUM> as described herein) may reduce the cooling flow C requirement, which may result in a more efficient engine core (because less flow is bypassing the combustor), meaning that for a given amount of core power, the mass flow entering the core can be reduced and/or the size and/or mass of the turbine(s) <NUM>, <NUM> can be reduced.

<FIG> and <FIG> schematically show turbine cooling apparatus <NUM>. The turbine cooling apparatus extracts cooling flow C from the compressor <NUM>, <NUM>. The cooling flow C bypasses the combustor <NUM>. The cooling flow C is then delivered to the high pressure turbine <NUM> and optionally the low pressure turbine <NUM>. Although the turbine cooling apparatus <NUM> is shown in <FIG> and <FIG> as extracting cooling flow C from a specific position in the high pressure compressor <NUM> and delivering it to a specific position in the high pressure turbine <NUM>, it will be appreciated that this is merely for ease of schematic representation, and that the cooling flow C may be extracted from any suitable locations (for example multiple locations in the high pressure compressor <NUM> and/or the low pressure compressor <NUM>) and delivered to any desired locations (for example multiple locations in the high pressure turbine <NUM> and/or the low pressure turbine <NUM>).

A reduction in the amount of cooling flow C is desirable, because the cooling flow is not combusted and thus the amount of work that can be extracted from it is lower than for the flow that passes through the combustor <NUM>. With reference to <FIG>, the gas turbine engine <NUM> has a bypass ratio defined as the mass flow rate of the flow B through the bypass duct <NUM> divided by the mass flow rate of the flow A entering the engine core at cruise conditions. As the bypass ratio is increased - for example to increase engine efficiency - proportionally less flow A goes through the core. This means that for a given size of engine and/or to be able to withstand a given turbine entry temperature, a higher proportion of the core flow A may be required to be used as turbine cooling flow C. In this regard, as used herein, turbine entry temperature (T<NUM>turb_in) may be the maximum stagnation temperature measured at a position <NUM> in the gas flow path that is immediately upstream of the most axially upstream rotor blade row <NUM>, i.e. at a so-called "red-line" operating condition of the engine at which the engine is certified.

A cooling to bypass efficiency ratio may be defined as the ratio of the mass flow rate C of the turbine cooling flow to the mass flow rate B of the bypass flow at cruise conditions. Using an understanding of the constraints and/or technologies described by way of example herein, the cooling to bypass efficiency ratio may be optimized to be as described and/or claimed herein. Additionally or alternatively, the mass of the high pressure turbine <NUM> and/or the low pressure turbine <NUM> may be optimized (for example reduced) relative to a conventional engine. In turn, this may reduce the mass of the high pressure turbine <NUM> and/or the low pressure turbine <NUM> as a proportion of the overall mass of the gas turbine engine <NUM>.

Using an understanding of the constraints and/or technologies described by way of example herein, the normalized thrust may be optimized. In this regard, the normalized thrust is defined as the maximum net thrust of the engine <NUM> at sea level divided by the total mass of the turbines <NUM>, <NUM> in the engine <NUM>. The illustrated example has a high pressure turbine <NUM> and a low pressure turbine <NUM>, however, it will be appreciated that where further turbines are included in the engine the total turbine mass includes the mass of all turbines.

As noted elsewhere herein, the optimized use of CMCs may result in a reduced turbine cooling flow requirement. Additionally or alternatively, through use of CMCs it may be possible to expose some components to a higher temperature than would otherwise be possible. This may result in the ability to increase the turbine entry temperatures relative to engines <NUM> that do not include optimized use of CMCs. In this regard, it has been found that higher turbine entry temperatures are desirable from an engine efficiency perspective.

Using an understanding of the constraints and/or technologies described by way of example herein, the cooling efficiency ratio may be optimized. In this regard, the cooling efficiency ratio is defined as the ratio between the turbine entry temperature (as defined elsewhere herein) and the cooling flow requirement. The cooling flow requirement may be defined as the mass flow rate of the turbine cooling flow C divided by the mass flow rate of the flow A entering the engine core at cruise conditions.

A core size CS may be defined for the gas turbine engine <NUM> as: <MAT> where:.

Using an understanding of the constraints and/or technologies described by way of example herein may allow a thrust to core efficiency ratio TC and/or a fan to core efficiency ratio FC to be optimised to be in the ranges described and/or claimed herein, where the thrust to core efficiency ratio TC and the fan to core efficiency ratio FC are as defined below (with T<NUM>turb_in being the turbine entry temperature at position <NUM> shown in <FIG>, as described above). <MAT> <MAT>.

It will be appreciated that the understanding and/or technology described and/or claimed herein results in a particularly efficient gas turbine engine <NUM>. For example, the understanding and/or technology described and/or claimed herein may provide a particularly efficient gas turbine engine <NUM> in which a fan <NUM> that is driven by a gearbox <NUM> is complemented by an optimized engine core.

Claim 1:
A gas turbine engine (<NUM>) for an aircraft comprising:
an engine core (<NUM>) comprising:
a turbine (<NUM>, <NUM>), a combustor (<NUM>), and a compressor (<NUM>, <NUM>), the turbine comprising a first turbine (<NUM>) and a second turbine (<NUM>) and the compressor comprising a first compressor (<NUM>) and a second compressor (<NUM>);
a first core shaft (<NUM>) connecting the first turbine to the first compressor;
a second core shaft (<NUM>) connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising:
a fan (<NUM>) comprising a plurality of fan blades; and
a gearbox (<NUM>) that receives an input from the first core shaft (<NUM>) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein:
the second turbine comprises at least one ceramic matrix composite component;
a bypass duct is defined radially outside the engine;
part of the flow (C) that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine;
the cooling to bypass flow efficiency ratio defined as the ratio of the mass flow rate of the turbine cooling flow to the mass flow rate of the bypass flow at engine cruise conditions is less than <NUM>,
the mass of ceramic matrix composite in the second turbine is in the range of from <NUM>% to <NUM>% of the total mass of the second turbine; and
the turbine entry temperature, defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is at least <NUM>; characterised in that the second turbine (<NUM>) comprises first and second rotor blade and stator vane rows, wherein the most axially upstream row of stator vanes (<NUM>) and rotor blades (<NUM>) of the second turbine (<NUM>) are metallic, whilst the second most axially upstream row of stator vanes (<NUM>) and rotor blades (<NUM>) of the second turbine (<NUM>) comprise Ceramic Matrix Composites (CMCs).