High temperature gas turbine engine

Abstract

There is provided a gas turbine engine comprising: a turbine comprising a compressor section, a combustion section and a turbine section in a series flow sequence, the turbine section comprising a turbine housing and a turbine stage, the turbine stage comprising a plurality of turbine stator vanes, the stage being a stage count N downstream of the combustion section, the turbine stator vanes comprising a first vane assembly, the first vane assembly includes a nozzle formed of a first material defining a first coefficient of thermal expansion; and a hanger radially outboard of the nozzle and coupling the nozzle to the turbine housing, the hanger formed of a second material defining a second coefficient of thermal expansion, the gas turbine engine defining a nozzle efficiency offset NEO greater than 1 and less than or equal to 5, where NEO denotes: # imgabs0 #

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine designed to operate at elevated exhaust gas temperatures. Background Art

[0002] A gas turbine engine generally includes a turbine and a rotor assembly. A gas turbine engine (eg, a turbofan engine) may be used for aircraft propulsion. For a turbofan engine, the rotor assembly may be configured as a fan assembly.

[0003] The turbomachine may include a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine in a series flow relationship. The high-pressure turbine and the low-pressure turbine each include an annular array ("row") of stationary vanes or nozzles for directing gas leaving the combustor or a corresponding upstream rotating blade row into a corresponding downstream rotating blade row. A row of nozzles and a row of blades together constitute a "stage". These components operate in a high temperature environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] This specification sets forth a complete and enabling disclosure of the present technology, including the best mode thereof, to one of ordinary skill in the art with reference to the accompanying drawings, in which:

[0005] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0006] Figure 2 yes Figure 1 A schematic cross-sectional view of a portion of a turbine of an exemplary gas turbine engine.

[0007] Figure 3 is a perspective view of a turbine stator vane assembly according to an example embodiment of the present disclosure.

[0008] Figure 4 is a partial exploded view of a nozzle segment assembly viewed from front to rear according to an exemplary embodiment of the present disclosure.

[0009] Figure 5 is another partial exploded view of the nozzle segment assembly also looking from front to rear.

[0010] Figure 6 is a diagram of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0011] Figure 7 is a diagram of a gas turbine engine according to one or more exemplary embodiments of the present disclosure. DETAILED DESCRIPTION

[0012] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. Numerical and alphabetical symbols are used in the detailed description to refer to features in the drawings. Similar or similar symbols in the drawings and description are used to refer to similar or similar parts of the present disclosure.

[0013] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0014] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0015] For example, in the context of “at least one of A, B, and C,” the term “at least one” means only A, only B, only C, or any combination of A, B, and C.

[0016] The term "cooled cooling air system" as used herein refers to a system configured to provide a cooling air flow to one or more components exposed to a working gas flow path of a turbine of a gas turbine engine at a location downstream of a combustor of the turbine and upstream of an exhaust nozzle of the turbine, the cooling air flow being thermally connected to a heat exchanger to reduce the temperature of the cooling air flow at a location upstream of the one or more components.

[0017] The cooled cooling air system contemplated by the present disclosure may include a hot bus cooled cooling air system or a dedicated heat exchanger cooled cooling air system (i.e., a cooled cooling air system including a radiator heat exchanger dedicated to the cooling cooling air system; see Figure 2 ); a bypass heat exchanger cooled cooling air system having a radiator heat exchanger thermally coupled to an airflow through a bypass passage; an air-to-air cooled cooling air system (i.e., a cooled cooling air system having a radiator heat exchanger configured to transfer heat to an airflow); an oil-to-air cooled cooling air system (i.e., a cooled cooling air system having a radiator heat exchanger configured to transfer heat to an oil flow); a fuel-to-air cooled cooling air system (i.e., a cooled cooling air system having a radiator heat exchanger configured to transfer heat to a fuel flow, such as a JetA fuel flow, a liquid hydrogen or hydrogen fuel flow, etc.; see e.g. Figure 2 ); or a combination thereof.

[0018] In one or more exemplary cooled cooling air systems described herein, the cooled cooling air system may receive cooling air from a downstream end of the high pressure compressor (i.e., a position closer to the last stage of the high pressure compressor than the first stage of the high pressure compressor), an upstream end of the high pressure compressor (i.e., a position closer to the first stage of the high pressure compressor than the last stage of the high pressure compressor), a downstream end of the low pressure compressor (i.e., a position closer to the last stage of the low pressure compressor), an upstream end of the low pressure compressor (i.e., a position closer to the first stage of the low pressure compressor), a position between compressors, a bypass passage, a combination thereof, or any other suitable airflow source.

[0019] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0020] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, while "downstream" refers to the direction to which the fluid is flowing.

[0021] The phrases "from X to Y" and "between X and Y" each refer to a numerical range that includes the endpoints (ie, refers to a numerical range that includes both X and Y).

[0022] The term "redline exhaust gas temperature" refers to the maximum rated exhaust gas temperature of a gas turbine engine.

[0023] The term "hanger" as used herein is a load-bearing structural member that connects one or more stator blades to the engine casing of a gas turbine engine. The hanger is radially disposed in the gas turbine engine (see Figure 1 , radial direction R) is positioned at the outer end of one or more stator blades. The pylon can be used in conjunction with stator blades in a high pressure turbine of a gas turbine engine, a low pressure turbine of a gas turbine engine, and / or other sections of a gas turbine engine.

[0024] As used herein, ceramic matrix composites or "CMCs" refer to a class of materials that include reinforcement materials (e.g., reinforcing fibers) surrounded by a ceramic matrix phase. In general, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of CMC matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) may also be included in the CMC matrix.

[0025] Some examples of reinforcing fibers for CMCs may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates such as mullite, or mixtures thereof), or mixtures thereof.

[0026] In general, a particular CMC can be represented by its fiber type / matrix type combination. For example, C / SiC represents carbon fiber reinforced silicon carbide; SiC / SiC represents silicon carbide fiber reinforced silicon carbide, SiC / SiN represents silicon carbide fiber reinforced silicon nitride; SiC / SiC-SiN represents silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, the CMC can include a matrix and reinforcing fibers that include oxide-based materials such as aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al2O3·2SiO2), as well as glassy aluminosilicates.

[0027] In certain embodiments, the reinforcing fibers may be bundled and / or coated before being added to the matrix. For example, the fiber bundle may be formed into a reinforcement tape, such as a unidirectional reinforcement tape. Multiple tapes may be laid together to form a preform component. The fiber bundle may be impregnated with a slurry composition before or after forming the preform. Then, the preform may be subjected to heat treatment and subsequent chemical treatment to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo curing or burnout to produce a high carbon residue in the preform, followed by melt infiltration with silicon, or undergo curing or pyrolysis to produce a silicon carbide matrix in the preform, followed by chemical vapor infiltration with silicon carbide. Additional steps may be taken before or after chemical vapor infiltration to improve the densification of the preform, by injecting a liquid resin or polymer into the preform, followed by a heat treatment step to fill the void with silicon carbide. The CMC material used herein may be formed using any known or future developed method, including but not limited to melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP) or any combination thereof.

[0028] Such materials are particularly suitable for high-temperature applications. In addition, these ceramic materials are lightweight compared to high-temperature alloys, but still provide strength and durability to the parts made from them.

[0029] Typically, a gas turbine engine (e.g., a turbofan gas turbine engine) includes a fan and a turbine, wherein the turbine rotates the fan to generate thrust. The turbine includes a compressor section, a combustion section, a turbine section, and an exhaust section, and defines a working gas flow path therethrough. Traditionally, the design practice of a gas turbine engine limits the compressor pressure ratio based at least in part on the gas temperature at the high-pressure compressor outlet stage. Once the heat generated by the combustion process in the combustion section is added to the gas flow, the higher gas temperature at the high-pressure compressor outlet stage may cause the temperature through the turbine section to be too high.

[0030] To achieve new levels of propulsion efficiency, the inventors set out to design a gas turbine engine with a redline exhaust temperature of at least 1,100 degrees Celsius. In designing a gas turbine engine with a redline exhaust temperature of at least 1,100 degrees Celsius, the inventors recognized that using various stages of turbine stator blades made of advanced high temperature materials that can withstand high temperatures can provide significant advantages and overcome the challenges of combining such advanced high temperature materials with other turbine structures. In addition, the inventors recognized that providing cooling air to various stages of turbine stator blades can also provide significant benefits to such a gas turbine engine. Specifically, these benefits may be reflected in allowing the gas turbine engine to operate at the higher temperatures required.

[0031] However, as the temperature within the working gas flow path increases, using cooling air to maintain the turbine components below a certain temperature threshold requires bleeding more cooling air from the compressor section than previously calculated, which means that more work used to compress this air will not drive the turbine. Instead, this bleed air will be diverted out of the core flow path and used to cool the turbine section.

[0032] Furthermore, while high temperature advanced materials are able to withstand elevated temperatures with little or no cooling, they generally have material properties that limit their ability to withstand forces in certain directions, such as in tension or directions where there is a tensile component, such as, but not limited to, torsion or bending. When integrating components formed from these high temperature materials, such as when mounting the high temperature materials to an engine housing using a support structure (referred to herein as a "hanger"), it is necessary to establish a load path between materials having mismatched coefficients of thermal expansion, such as in the case of a turbine stator bucket formed from a ceramic matrix composite (CMC) and a hanger formed at least in part from a metal or metal alloy.

[0033] Others have concluded that if high temperature materials are used in the turbine to achieve elevated operating temperatures, the mismatch in the thermal expansion coefficients of the materials will result in the joints (i.e., where two materials join together) not remaining sufficiently airtight to achieve higher efficiencies at elevated operating temperatures. Thus, while components formed from these high temperature materials would require less cooling and be able to operate at higher temperatures, it is believed that existing engines operating at lower operating temperatures and requiring more cooling air would indeed achieve higher efficiencies because leakage through the working gas flow path would be significantly reduced. Furthermore, the use of high temperature materials is not recommended because it is believed that the joints are susceptible to increased wear and premature failure due to the large amount of temperature-induced, associated motion between components during each operating cycle of a gas turbine engine. Thus, it is concluded that despite the potential efficiency advantages of gas turbine engines operating at redline exhaust temperatures of at least 1,100 degrees Celsius and using advanced high temperature materials in the turbine, the result is a net loss in propulsive efficiency.

[0034] However, the inventors have found that higher engine efficiency can indeed be achieved when operating at high temperatures using advanced materials. The inventors have found that the combination of advanced high temperature materials in the turbine stage and a certain level of cooling airflow supplied to the turbine stage, coupled with increased redline exhaust gas temperatures and overall thrust output, produces the desired results. This combination can be adjusted based on the position of the low pressure or power turbine stator vanes relative to the combustion section (e.g., the position of the first stage, the second stage relative to the combustion section).

[0035] In particular, the inventors discovered unexpectedly during the design of a gas turbine engine having a redline exhaust temperature of at least 1,100 degrees Celsius that, contrary to previous thinking and expectations, the costs involved in including turbine stator vanes formed of advanced high temperature materials (and optionally cooled in the manner described above) can be overcome to achieve a net propulsive efficiency gain. In particular, the inventors discovered during the design of several gas turbine engines having different thrust ratings and propulsive efficiency requirements (including the configurations shown and described in detail herein) that there is a relationship between the thermal expansion coefficient mismatch between the turbine stator vanes and the corresponding pylons, the amount of cooling flow provided to the turbine stator vanes, and the stage count of the turbine stator vanes, such that operating a gas turbine engine at a redline exhaust temperature of at least 1,100 degrees Celsius can result in a net propulsive gain for the overall engine design in accordance with one or more exemplary aspects described herein. As noted above, previous thinking was that for gas turbine engines with redline exhaust gas temperatures of at least 1,100 degrees Celsius, the inefficiencies introduced by having turbine stator blades of a different material than the hanger to which they are mounted and the cooling flow to the turbine stator blades would outweigh the potential propulsion efficiency benefits.

[0036] In order to achieve an improved gas turbine engine that is capable of providing a net propulsion efficiency benefit at elevated redline exhaust gas temperatures, the inventors proceeded as follows: designed a gas turbine engine having turbine stator blades as described herein, wherein the turbine stator blades have various thermal expansion coefficient mismatches with corresponding pylons, various cooling flows are provided to the turbine stator blades, and the turbine stator blades have various stage counts; examined the operability and net propulsion efficiency benefit of the designed gas turbine engine; redesigned the gas turbine engine to change the parameters based on the impact on other aspects of the gas turbine engine; re-examined the operability and net propulsion efficiency benefit characteristics of the redesigned gas turbine engine; and so on, in the process of designing several different types of gas turbine engines, including the gas turbine engine described herein, which will be described in more detail below.

[0037] Referring now to the drawings, wherein like numerals represent like elements throughout the several views, Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In an embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine, sometimes also referred to as a "turbofan engine". Figure 1 As shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline 12. Generally, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

[0038] The exemplary turbine 16 shown generally includes a generally tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 surrounds a compressor section, which includes a supercharger or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24, a combustion section 26, a turbine section, which includes a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30, and an injection exhaust nozzle section 32 in series flow relationship. A high pressure (HP) shaft 34 (which may additionally or alternatively be a spool) drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft 36 (which may additionally or alternatively be a spool) drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and injection exhaust nozzle section 32 collectively define a working gas flow path 37.

[0039] For the illustrated embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 is rotatable about a pitch axis P relative to the disk 42 by means of the fan blades 40 being operably coupled to an appropriate pitch change mechanism 44, the pitch change mechanism being configured to jointly change the pitch of the fan blades 40, e.g., in unison. The gas turbine engine 10 also includes a power gearbox 46, and the fan blades 40, the disk 42, and the pitch change mechanism 44 are rotatable together about the longitudinal centerline 12 through the LP shaft 36 across the power gearbox 46. The power gearbox 46 includes a plurality of gears for adjusting the rotational speed of the fan 38 relative to the rotational speed of the LP shaft 36, so that the fan 38 can rotate at a more efficient fan speed.

[0040] Still refer to Figure 1 In the exemplary embodiment of the fan blade 40, the disk 42 is covered by a rotatable front hub 48 (sometimes also referred to as a “spinner”) of the fan section 14. The front hub 48 has an aerodynamic profile to facilitate airflow through the plurality of fan blades 40.

[0041] Additionally, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the turbine 16. It should be appreciated that in the illustrated embodiment, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Furthermore, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween.

[0042] During operation of the gas turbine engine 10, a quantity of air 58 enters the gas turbine engine 10 through the nacelle 50 and the associated inlet 60 of the fan section 14. As the quantity of air 58 passes through the fan blades 40, a first portion of air 62 is directed or delivered into the bypass airflow passage 56, while a second portion of air 64 is directed or delivered into the working gas flow path 37, or more specifically, into the LP compressor 22, as indicated by arrow 64. The ratio between the first portion of air 62 and the second portion of air 64 is generally referred to as a bypass ratio. The pressure of the second portion of air 64 is then increased as it is delivered through the HP compressor 24 and into the combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66.

[0043] The combustion gases 66 are delivered through the HP turbine 28, where a portion of the heat and / or kinetic energy from the combustion gases 66 is extracted via one or more stages of HP turbine stator blades 90 coupled to the outer casing 18 and HP turbine rotor blades 88 coupled to the HP shaft 34, thereby rotating the HP shaft 34 to support the operation of the HP compressor 24. The HP turbine 28 may include a single stage of HP turbine stator blades 90 and HP turbine rotor blades 88. Alternatively, the HP turbine 28 may include two stages of HP turbine stator blades 90 and HP turbine rotor blades 88 (i.e., a first stage and a second stage). It should be understood that the term "stage" as used herein refers to a set of stator blades and rotor blades.

[0044] The combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36 , thereby rotating the LP shaft 36 to support operation of the LP compressor 22 and rotation of the fan 38 .

[0045] The combustion gases 66 are then routed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased as the first portion of air 62 is routed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the gas turbine engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbine 16.

[0046] However, it should be understood that Figure 1The exemplary gas turbine engine 10 shown in FIG. 1 is merely an example, and in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, although the illustrated gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., including an outer nacelle 50), in other embodiments, the gas turbine engine 10 may be an unducted gas turbine engine (such that the fan 38 is an unducted fan and the outlet guide vanes 52 are cantilevered from the outer casing 18). Additionally or alternatively, although the illustrated gas turbine engine 10 is configured as a geared gas turbine engine (i.e., including a power gearbox 46) and a variable pitch gas turbine engine (i.e., including a fan 38 configured as a variable pitch fan), in other embodiments, the gas turbine engine 10 may alternatively be configured as a direct drive gas turbine engine (such that the LP shaft 36 rotates at the same speed as the fan 38), a fixed pitch gas turbine engine (such that the fan 38 includes fan blades 40 that cannot rotate about the pitch axis P), or both.

[0047] Reference now Figure 2 , provides Figure 1 Schematic cross-sectional view of a portion of a turbine 16 of an exemplary gas turbine engine 10 .

[0048] As described above, during operation of the gas turbine engine 10, the gas flow through the working gas flow path 37 of the gas turbine engine 10 is directed to the compressor section, or more specifically, to the LP compressor 22 (see FIG. Figure 1 ) and the HP compressor 24, are compressed in sequence. The compressed air from the compressor section is then provided to the combustion section 26, where at least a portion of the compressed air is mixed with fuel and combusted to produce combustion gases 66. The combustion gases 66 flow from the combustion section 26 to the turbine section, more specifically, through the HP turbine 28 and the LP turbine 30, which, for the illustrated embodiment, drive the HP turbine 28 and the LP turbine 30 (see Figure 1 The HP shaft 34 is drivingly coupled to the HP turbine 28 and the HP compressor 24 .

[0049] Special References Figure 2 , the HP compressor 24 includes a plurality of compressor stages 80, wherein each compressor stage 80 includes, for example, a plurality of HP compressor rotor blades 82 and a rotor 84. Each of the respective compressor stages 80 is drivingly coupled to the HP shaft 34, such that the HP turbine 28 can drive the HP compressor 24 via the HP shaft 34. Among the plurality of compressor stages 80 of the HP compressor 24, the last compressor stage 80A is located at the rear end of the HP compressor 24. The compressor rotor blades 82 of the last compressor stage 80A are attached to the corresponding rotor 84 of the last compressor stage 80A.

[0050] The HP turbine 28 and the LP turbine 30 similarly include a plurality of turbine stages 86, wherein each turbine stage 86 includes, for example, a plurality of turbine rotor blades 88, a plurality of turbine stator vanes 90, and a rotor 92. Among the plurality of turbine stages 86, there is a first turbine stage 86A located at the frontmost / upstream end of the HP turbine 28, and a second turbine stage 86B. The first and second turbine stages 86A, 86B are the first and second stages of the HP turbine 28. The plurality of turbine stages 86 also include a third turbine stage 86C and a fourth turbine stage 86D. The first to fourth turbine stages 86A-86D are arranged in sequence. The third and fourth turbine stages 86C, 86D are part of the LP turbine 30. In this manner, it should be understood that each set of a plurality of turbine rotor blades 88 and turbine stator vanes 90 defines a stage count N, and is a stage count N downstream of the combustion section 26 (i.e., the first turbine stage 86A defines a stage count N, which is equal to 1, and the fourth turbine stage 86D defines a stage count N, which is equal to 4).

[0051] Briefly, it should be understood that each of the plurality of turbine stator blades 90 is mounted at a radially outer end along a radial direction R using a structural member 94, referred to herein as a hanger. The structural member 94, in turn, may be coupled to a turbine casing 95 of the engine 10. It should be understood that, as used herein, the term "casing" generally refers to a support structure in an engine, such as for turbine stator blades (e.g., a turbine casing, a turbine frame, etc.).

[0052] As described above, the gas turbine engine 10 may be designed to have an elevated redline exhaust temperature, such as at least 1,100 degrees Celsius. To accommodate the elevated temperatures, the turbine stator vanes 90, particularly the turbine stator vanes 90 of at least the first, second, and third turbine stages 86A, 86B, 86C, are made of advanced high temperature materials. For example, the turbine stator vanes 90 of at least the first, second, and third turbine stages 86A, 86B, 86C may be formed of a ceramic matrix composite (CMC) material.

[0053] Additionally, for the illustrated embodiment, gas turbine engine 10 includes cooling system 100. Cooling system 100 generally includes inlet duct 102 configured to receive cooling airflow 106 and distribution duct 104 (or distribution duct network) for providing cooling airflow 106 to one or more components of turbine 16.

[0054] Specifically, for the illustrated embodiment, the inlet duct 102 is in airflow communication with the HP compressor 24 for receiving the bleed airflow from the HP compressor 24 as the cooling airflow 106. Specifically, for the illustrated embodiment, the inlet duct 102 includes a first inlet duct 102A in airflow communication with the HP compressor 24 at a first location and a second inlet duct 102B in airflow communication with the HP compressor 24 at a second location (upstream of the first location). In this way, the first and second inlet ducts 102A, 102B can be configured to extract air from the HP compressor 24 at different pressures and temperatures.

[0055] It should be appreciated that while inlet duct 102 is depicted as having two sections, in other exemplary embodiments, inlet duct 102 may include any other suitable number of sections to extract additional or alternative air flows from HP compressor 24 , LP compressor 22 , or both.

[0056] Additionally, for the illustrated embodiment, the turbine 16 includes a fuel line 108 extending to a fuel nozzle 110 of the combustion section 26. The cooling system 100 includes a heat exchanger 112 to thermally couple the cooling air flow 106 through the inlet duct 102 with the fuel flow through the fuel line 108. In this manner, the cooling system 100 may cool the cooling air flow 106 before providing it to the turbine stage 86. In this manner, the cooling system 100 may be referred to as a cooled cooling air (CCA) system.

[0057] It is noted that in other exemplary embodiments, cooling system 100 may not include heat exchanger 112 or any other similar heat exchanger such that cooling system 100 does not provide a cooling cooling air flow therethrough.

[0058] The distribution duct 104 extends to the turbine stator blades 90 of the first turbine stage 86A, the turbine stator blades 90 of the second turbine stage 86B, and optionally the turbine stator blades 90 of the third turbine stage 86C, and provides a cooling airflow 106 thereto.

[0059] In brief summary, the cooling system 100 also includes a valve 114 to control the mass flow rate W of the cooling airflow 106 provided to a given stage of turbine stator blades 90. N Specifically, the cooling system 100 includes a first valve 114A for controlling the mass flow rate W of the cooling air flow 106 provided to the turbine stator vanes 90 of the first turbine stage 86A. N1 A second valve 114B for controlling the mass flow rate W of the cooling air flow 106 provided to the turbine stator blades 90 of the second turbine stage 86B N2and optionally a third valve 114C for controlling the mass flow rate W of the cooling airflow 106 supplied to the turbine stator vanes 90 of the third turbine stage 86C N3 .

[0060] It will further be appreciated that the mass flow rate W of the cooling airflow 106 provided to a given stage of turbine stator blades 90 is N The mass flow rate W of the gas flow through the working gas flow path 37 at the inlet of the HP compressor 24 may be referred to as 25 The percentage of N Divide by W 25 The ratio of may indicate how much cooling airflow is provided to a given stage of turbine stator buckets 90 .

[0061] It should be appreciated that while the cooling system 100 includes inlet ducts 102 that are configured to converge into a single stream prior to providing the cooling airflow 106 to the distribution ducts 104, in other exemplary embodiments, the cooling system 100 may include multiple dedicated networks of cooling airflows 106 such that, for example, the cooling airflow 106 from the first inlet duct 102A is provided to a location in the turbine section without mixing with the cooling airflow 106 from the second inlet duct 102B.

[0062] Reference now Figure 3 , Figure 3 2 is a perspective view of a stator vane assembly 200 according to an example embodiment of the present disclosure. In an example embodiment, the stator vane assembly 200 may be located within the HP turbine 28 and / or the LP turbine 30 (e.g., as referenced above). Figure 2 The stator vane assembly 200 is formed of one or more stator vane segment assemblies 202. The stator vane segment assemblies 202 direct the combustion gases downstream through a subsequent row of rotor blades (not shown). The stator vane assembly 200 and the plurality of stator vane segment assemblies 202 defining the stator vane assembly 200 facilitate passage of the combustion gases through the immediately downstream rotor blades (e.g., Figure 2 As shown; for example, turbine stator vanes 90 and downstream turbine rotor blades 88) extract energy.

[0063] The stator blade segment assembly 202 includes an inner band 204, an outer band 216, and a plurality of struts 208 ( Figure 3 Each strut extends through one of a plurality of nozzle airfoils 210 (or simply "nozzles"). The inner band 204 and the outer band 216 extend 360 degrees circumferentially around the longitudinal centerline 215.

[0064] The stator vane assembly 200 is formed of a plurality of stator vane segment assemblies 202, each stator vane segment assembly 202 including an internal support structure 212, at least one nozzle airfoil 210, and a hanger 214. The hanger 214 is a load-bearing structural member that holds one or more stator vanes ( Figure 3 At least one nozzle airfoil 210 in the embodiment of the present invention is connected to an engine case (not shown) of the gas turbine engine. A cradle 214 is positioned at the outer end of one or more stator blades in a radial direction R. The cradle 214 is depicted as being separate from the outer band 216, but in other embodiments, the outer band 216 may be formed together with the cradle 214. In addition, it should be understood that the cradle 214 may be directly coupled to the at least one nozzle airfoil 210, or in alternative embodiments, may be coupled to the at least one nozzle airfoil 210 through one or more intermediate components, for example, in an embodiment where the outer band 216 is separate from the cradle 214, it may be coupled to the at least one nozzle airfoil 210 through the outer band 216.

[0065] Reference now Figure 4 and 5 , Figure 4 is a partial exploded view of stator vane segment assembly 202 looking from front to rear in accordance with an example embodiment. Figure 5 Another partial exploded view of stator bucket segment assembly 202 also looking from front to rear.

[0066] The struts 208 carry loads from the radially inner side of the stator bucket segment assembly 202 at the inner support structure 212 to the radially outer side at the pylon 214, where the loads are transferred to the structure of the engine, such as, but not limited to, the engine case, and mechanically support the nozzle airfoil 210. The struts 208 may be connected to the inner support structure 212 and the pylon 214 by, for example, bolting, fastening, fixing, combinations thereof, and integrally formed with one or both of 212 and 214.

[0067] In an exemplary embodiment, the stator blade segment assembly 202 includes an internal support structure 212 formed of a first metallic material and a hanger 214 formed of a second metallic material (which may be the same as the first metallic material in some embodiments). The struts 208 are connected to or integrally formed with the internal support structure 212. The struts 208 may be hollow and may each have at least one thickened inner wall sufficient to provide the desired stiffness and strength. The struts 208 include a first mating end 206 (at Figure 3The invention also includes a first metal material 206 and a second metal material 207. The first metal material 206 includes a first metal material 208 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 209 and a second metal material 209. The first metal material 208 includes a first metal material 208 ...

[0068] The nozzle airfoil 210 , including the body of the nozzle airfoil 210 , is formed of an advanced high temperature material, such as a material having a low coefficient of thermal expansion, such as a ceramic matrix composite (CMC). The nozzle airfoil 210 extends between the inner band 204 and the outer band 216 .

[0069] It should be understood that the nozzle airfoil 210 is formed of a material having a lower coefficient of thermal expansion than the material forming the internal support structure 212, the pylon 214, the struts 208, and other metal components. In this way, when the gas turbine engine is operated at elevated temperatures, the thermal growth of the internal support structure 212, the pylon 214, the struts 208, and other metal components relative to the nozzle airfoil 210 may create gaps, causing gas flow to leak from the working gas flow path, which may cause gas turbine engine inefficiencies. It is worth noting that the higher the temperature to which the components are exposed, the greater the relative growth of the components.

[0070] As previously described, the inventors, in the process of designing a gas turbine engine with an elevated redline exhaust gas temperature—that is, designing a gas turbine engine with an elevated redline exhaust gas temperature with various thermal expansion coefficient mismatches between the turbine stator blades and the corresponding pylons, providing various cooling flows to the turbine stator blades, and having various stage counts of the turbine stator blades, and evaluating overall engine and net propulsion efficiency increases—surprisingly discovered that there is a significant relationship between the thermal expansion coefficient mismatch between the turbine stator blades and the corresponding pylons, the cooling flow provided to the turbine stator blades, and the stage count of the turbine stator blades. This relationship can be viewed as an indicator of the ability of a gas turbine with an elevated redline exhaust gas temperature to achieve a net propulsion efficiency increase by avoiding excessive cooling flows and flow path inefficiencies introduced by the relative motion of these components, and taking into account the reliability of these components.

[0071] This relationship is applicable to gas turbine engines having elevated redline exhaust temperatures. This relationship relates to the mismatch in thermal expansion coefficients between the turbine stator vanes and the corresponding pylons for engines having redline exhaust temperatures of 1,100 degrees Celsius or above, the cooling flow provided to the turbine stator vanes, and the stage count of the turbine stator vanes, as described in more detail herein.

[0072] Specifically, the inventors have discovered that in engines with elevated redline exhaust gas temperatures, an increase in the mismatch in coefficients of thermal expansion between the turbine stator vanes and the corresponding hangers generally results in a corresponding increase in relative thermal expansion between the turbine stator vanes and the corresponding support structure (exacerbated by the elevated redline exhaust gas temperatures of the engine), which increases the chance of leakage flow in the working gas flow path escaping through the joints. Furthermore, with the increase in relative thermal expansion, the likelihood of premature wear of the components increases. Furthermore, the inventors have discovered that when a gas turbine engine with elevated redline exhaust gas temperatures is operated, an increase in cooling flow can generally be provided to maintain the components within a desired operating range.

[0073] However, the inventors discovered in the process of designing an engine in accordance with the present disclosure that the thermal expansion mismatch between the turbine stator blades and the corresponding pylons can be balanced by the amount of cooling provided to the turbine stator blades at a given stage count of the turbine section. When balanced in accordance with the present disclosure, the above-mentioned inefficiencies associated with operating at elevated redline exhaust gas temperatures can be overcome, thereby providing a net propulsive efficiency benefit. Specifically, the inventors discovered that by targeted reductions in cooling flow to offset potential inefficiencies introduced by higher levels of relative thermal expansion (which may be caused by a combination of thermal expansion mismatch and elevated operating temperatures), the engine can be allowed to avoid the above-mentioned inefficiencies and obtain the propulsive efficiency benefits that can be achieved by engines with elevated redline exhaust gas temperatures.

[0074] Thus, the relationship discovered below can identify a gas turbine engine with an elevated redline exhaust temperature that is capable of achieving the desired net propulsive efficiency benefit and is suitable for a particular mission requirement, taking into account efficiency, weight, complexity, reliability, and other factors that influence the optimal selection of a gas turbine engine with an elevated redline exhaust temperature.

[0075] In addition to producing an improved gas turbine engine, as explained in detail above, by utilizing this relationship, the inventors have discovered that the number of suitable or feasible gas turbine engine designs capable of having an increased redline exhaust temperature can be significantly reduced, thereby facilitating faster selection of designs to be considered when developing a gas turbine engine. This advantage can provide a deeper understanding of the requirements of a particular gas turbine engine before the specific technology, integration, and system requirements are fully developed. This advantage can avoid late-stage redesigns.

[0076] The inventors have discovered that a desirable relationship that provides an improved gas turbine engine is the nozzle efficiency offset (NEO), which is expressed as:

[0077]

[0078] Where N is the stage count of the turbine stator vane stage in question, which is the turbine stator vane stage located downstream of the combustion section;

[0079] Wherein CTER is the coefficient of thermal expansion ratio, which is equal to the coefficient of thermal expansion of the hanger of the turbine stator blade at the Nth stage divided by the coefficient of thermal expansion of the nozzle of the turbine stator blade at the Nth stage;

[0080] Where W N is the cooling flow rate of the cooling airflow provided to the turbine stator blades at stage N (in kg / s);

[0081] Where W 25 is the compressor inlet gas flow rate in kg / s.

[0082] Table 1 below provides values ​​for the above relationships within the scope of the present disclosure.

[0083]

[0084] Reference now Figure 6 and 7 , various exemplary gas turbine engines are shown according to one or more exemplary embodiments of the present disclosure. Specifically, Figure 6 A form is provided that includes Figure 7 The corresponding values ​​for several gas turbine engines are plotted in . Figure 7 is a graph 300 of a gas turbine engine showing nozzle efficiency offset on a Y-axis 304 and stage count N on an X-axis 302 , according to one or more exemplary embodiments of the present disclosure.

[0085] As shown in the figure, Figure 7 The graph 300 in FIG. 3 depicts a first range 306 where NEO values ​​are greater than 1 and less than or equal to 5, and N values ​​are from 1 to 8. Figure 7 Also depicted is a second range 308 where the NEO value is between 2.5 and 4.5 and the N value is between 1 and 5; and a third range 310 where the NEO value is between 3 and 4 and the N value is from 1 to 3.

[0086] It should be understood that for certain exemplary embodiments, each turbine stator vane in stages 1, 2, and 3 may have a NEO value within one or more of the above ranges. For example, the following are sample values ​​for an engine having a redline exhaust gas temperature greater than or equal to 1,100 degrees Celsius and less than 1,700 degrees Celsius according to an embodiment of the present disclosure, where each turbine stator vane in stages 1, 2, and 3 may have a NEO value within one or more of the above ranges:

[0087] N CTER <![CDATA[W N / IN 25 ]]> NEO 1 4 .1 3.6 2 4 .015 3.55 3 4 0 3.39

[0088] Further aspects are provided by the subject matter of the following clauses:

[0089] A gas turbine engine, comprising: a turbine, the turbine comprising a compressor section, a combustion section, and a turbine section in a series flow sequence, the turbine section comprising a turbine casing and a turbine stage, the turbine stage comprising a plurality of turbine stator blades, the stage being a stage count N, located downstream of the combustion section, the turbine stator blades comprising a first blade assembly, the first blade assembly comprising: a nozzle, the nozzle being formed of a first material defining a first thermal expansion coefficient; and a hanger, the hanger being positioned radially outward of the nozzle and coupling the nozzle to the turbine casing, the hanger being formed of a second material defining a second thermal expansion coefficient, the gas turbine engine defining a thermal expansion coefficient ratio CTER equal to the second thermal expansion coefficient divided by the first thermal expansion coefficient, wherein during high power operation of the gas turbine engine, the compressor section is configured to operate at a compressor inlet mass flow rate W 25 (in kg / s) to provide airflow to the combustion section, the turbine is configured to provide a cooling flow rate W N (in kg / s) to provide cooling airflow to the turbine stator bucket stage, and wherein the redline exhaust temperature of the gas turbine engine is greater than or equal to 1,100 degrees Celsius and less than or equal to 1,700 degrees Celsius, wherein the gas turbine engine defines a nozzle efficiency offset NEO greater than 1 and less than or equal to 5, wherein NEO is expressed as:

[0090]

[0091] A gas turbine engine as in any preceding clause, wherein the stage count is greater than or equal to one and less than or equal to three.

[0092] A gas turbine engine as claimed in any preceding clause, wherein the turbine section comprises a high pressure turbine and a low pressure turbine, and wherein the stage of turbine stator buckets is a first stage of low pressure turbine stator buckets.

[0093] A gas turbine engine according to any preceding clause, wherein W N is 0.

[0094] A gas turbine engine as in any preceding clause, wherein said NEO is greater than 2.5 and less than or equal to 4.5.

[0095] A gas turbine engine as described in any preceding clause, wherein said NEO is greater than or equal to three.

[0096] A gas turbine engine according to any preceding clause, wherein the stage count is 1 or 2, and wherein W N / W 25 Greater than 0 and less than or equal to 0.2.

[0097] A gas turbine engine according to any preceding clause, wherein W N / W 25 Less than or equal to 0.15.

[0098] A gas turbine engine according to any preceding clause, wherein said CTER is greater than 3.

[0099] A gas turbine engine as claimed in any preceding clause, wherein said high power operating condition is a take-off operating condition.

[0100] A method of operating a gas turbine engine, the method comprising: operating the gas turbine engine under a high power operating condition, wherein operating the gas turbine engine under the high power operating condition comprises: 25 (in kg / s) to provide airflow to the combustion section, and at a cooling flow rate W N (in kilograms per second) to provide a cooling airflow to a turbine stator vane stage, wherein the gas turbine engine has a redline exhaust temperature greater than or equal to 1,100 degrees Celsius and less than or equal to 1,700 degrees Celsius, wherein the turbine stator vane stage is stage count N downstream of a combustion section, the turbine stator vane stage includes a first vane assembly, the first vane assembly including: a nozzle, the nozzle formed of a first material defining a first thermal expansion coefficient; and a hanger, the hanger positioned radially outward of the nozzle and coupling the nozzle to the turbine housing, the hanger formed of a second material defining a second thermal expansion coefficient, the gas turbine engine defining a thermal expansion coefficient ratio CTER equal to the second thermal expansion coefficient divided by the first thermal expansion coefficient, wherein the gas turbine engine defines a nozzle efficiency relationship NEO greater than 1 and less than or equal to 5, wherein NEO is expressed as:

[0101]

[0102] A method as in any preceding clause, wherein the stage count is greater than or equal to one and less than or equal to three.

[0103] A method as in any preceding clause, wherein the turbine section comprises a high pressure turbine and a low pressure turbine, and wherein the turbine stator bucket stage is a first stage of low pressure turbine stator buckets.

[0104] A method according to any preceding clause, wherein W N is 0.

[0105] A method as in any preceding clause, wherein the NEO is greater than 2.5 and less than or equal to 4.5.

[0106] A method as described in any preceding clause, wherein the NEO is greater than or equal to 3.

[0107] A method as in any preceding clause, wherein the stage count is 1 or 2, and wherein W N / W 25 Greater than 0 and less than or equal to 0.2.

[0108] A method according to any preceding clause, wherein W N / W 25 Less than or equal to 0.15.

[0109] A method as in any preceding clause, wherein the CTER is greater than 3.

[0110] A method as in any preceding clause, wherein the high power operating condition is a take-off operating condition.

[0111] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable those skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. These other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or include equivalent structural elements that do not differ substantially from the literal language of the claims.

Claims

1. A gas turbine engine, characterized in that: include: A turbomachine comprising a compressor section, a combustion section, and a turbine section in a series flow sequence, the turbine section comprising a turbine housing and a turbine stage, the turbine stage comprising a plurality of turbine stator blades, the turbine stage being a stage count N downstream of the combustion section, the turbine stator blades comprising a first blade assembly, the first blade assembly comprising: a nozzle formed of a first material defining a first coefficient of thermal expansion; as well as a hanger positioned radially outwardly of the nozzle and coupling the nozzle to the turbine housing, the hanger being formed of a second material defining a second coefficient of thermal expansion, the gas turbine engine defining a coefficient of thermal expansion ratio CTER equal to the second coefficient of thermal expansion divided by the first coefficient of thermal expansion, During high power operation of the gas turbine engine, the compressor section is configured to operate at a compressor inlet mass flow rate W 25 (in kg / s) provides airflow to the combustion section, and the turbine is configured to provide a cooling flow rate W N (in kg / s) to provide cooling airflow to the turbine stator blade stage, and wherein the redline exhaust temperature of the gas turbine engine is greater than or equal to 1,100 degrees Celsius (C) and less than or equal to 1,700 degrees Celsius, wherein the gas turbine engine defines a nozzle efficiency offset NEO greater than 1 and less than or equal to 5, wherein NEO is expressed as:

2. The gas turbine engine according to claim 1, characterized in that Wherein the stage count is greater than or equal to 1 and less than or equal to 3.

3. The gas turbine engine according to claim 1, characterized in that Wherein the turbine section comprises a high pressure turbine and a low pressure turbine, and wherein the turbine stator blade stage is a first stage of low pressure turbine stator blades.

4. The gas turbine engine according to claim 3, characterized in that Where W N is 0.

5. The gas turbine engine according to claim 1, characterized in that Wherein the NEO is greater than 2.5 and less than or equal to 4.

5.

6. The gas turbine engine according to claim 1, characterized in that Wherein the NEO is greater than or equal to 3.

7. The gas turbine engine according to claim 1, characterized in that Wherein the stage count is 1 or 2, and wherein W N / W 25 Greater than 0 and less than or equal to 0.

2.

8. The gas turbine engine according to claim 7, characterized in that Where W N / W 25 Less than or equal to 0.

15.

9. The gas turbine engine according to claim 1, characterized in that wherein the CTER is greater than 3.

10. The gas turbine engine according to claim 1, characterized in that Wherein the high power operating condition is a take-off operating condition.

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