Gas turbine engine cooling system

Abstract

The present invention provides a gas turbine engine, which includes a nacelle and an engine core received in the nacelle. The engine core defines a main rotation axis along its length. The engine core and the nacelle define a bypass channel between the two. The gas turbine engine also includes a cooling system, which includes a cooling duct. The cooling duct defines an inlet for receiving bypass air from the bypass channel at an upstream position and an outlet for discharging bypass air at a downstream position. The cooling duct extends axially and circumferentially around a segment of the engine core relative to the main rotation axis. The cooling duct includes: a first portion, which extends at least axially relative to the main rotation axis; a second portion, which is located downstream of the first portion and extends circumferentially around the engine core relative to the main rotation axis; and a third portion, which is located downstream of the second portion and extends at least axially relative to the main rotation axis.

Description

Technical Field

[0001] The present disclosure relates to gas turbine engines and, more particularly, to cooling systems for gas turbine engines. Background Art

[0002] Gas turbine engines are used for power generation and vehicle (eg, aircraft) propulsion. A typical gas turbine engine includes a compressor section, a combustor section, and a turbine section that utilize the primary airflow entering the engine to generate power or propel the vehicle.

[0003] Air may be bled from the compressor section of an engine for cooling and sealing purposes. Cooling air may be required to protect the materials of the engine from premature degradation and to prevent the materials from reaching temperatures that would be detrimental to their properties. However, in some cases, cooling air may not provide sufficient cooling. Furthermore, there is a limit to the amount of air that can be bled from the compressor section without affecting the normal operation of the engine. Summary of the invention

[0004] According to a first aspect, a gas turbine engine is provided, the gas turbine engine comprising a nacelle and an engine core received in the nacelle. The engine core defines a main rotation axis along its length. The engine core and the nacelle define a bypass channel therebetween. The gas turbine engine also includes a cooling system, the cooling system comprising a cooling duct. The cooling duct defines an inlet for receiving bypass air from the bypass channel at an upstream position and an outlet for discharging bypass air at a downstream position. The cooling duct extends axially and circumferentially around a segment of the engine core relative to the main axis. The cooling duct comprises: a first portion, the first portion extending at least axially relative to the main rotation axis; a second portion, the second portion being located downstream of the first portion and extending circumferentially around the engine core relative to the main rotation axis; and a third portion, the third portion being located downstream of the second portion and extending at least axially relative to the main rotation axis.

[0005] Since the cooling system uses bypass air flowing through the bypass passage and discharges the bypass air back into the bypass passage, the cooling system can provide cooling to the engine core without affecting the normal operation of the gas turbine engine. Since the cooling duct extends both axially and circumferentially around the segment of the engine core, heat transfer between the engine core and the cooling duct can be enhanced.

[0006] The engine core may include a combustor segment.The cooling duct extends circumferentially relative to the main axis around a segment of the engine core including the combustor segment.

[0007] The gas turbine engine may also include one or more grids disposed within the cooling duct for preventing debris from entering the cooling duct.

[0008] The one or more grids may include a first grid disposed proximate to the inlet and a second grid disposed proximate to the outlet.

[0009] At least one of the one or more grids may be a fractal grid.The one or more grids may be used to further increase the heat exchange rate.

[0010] The gas turbine engine may also include a check valve disposed in the cooling duct to prevent airflow from the outlet to the inlet. The check valve may be disposed near the outlet. The check valve may be a three-pointed valve. The three-pointed valve may be operated without any actuator.

[0011] The cooling system may further include one or more heat transfer enhancement elements disposed between the engine core and the cooling conduit for increasing heat transfer from the engine core to the cooling conduit. The one or more heat transfer enhancement elements may include at least one of fins and heat pipes.

[0012] The cooling system may further include one or more airflow boosters disposed in the cooling duct. The one or more airflow boosters may include at least one of a nozzle ejector, an electric fan, and a micro compressor.

[0013] The cooling conduit may be helical.

[0014] The cooling system may include a plurality of cooling conduits.

[0015] The cooling system may include two cooling ducts that merge with each other at a junction downstream of their inlets such that the two cooling ducts share a common outlet.

[0016] Each of the two cooling ducts may include a first portion extending at least axially relative to the main rotation axis and a second portion located downstream of the first portion. The second portion may extend circumferentially around the engine core relative to the main rotation axis. The two cooling ducts may merge with each other at a junction downstream of the second portion.

[0017] The two cooling ducts may comprise a common third portion located downstream of the junction and extending at least axially relative to the main rotation axis. The third portion may comprise a common outlet.

[0018] The gas turbine engine may further include a check valve disposed downstream of the junction and upstream of the common outlet for preventing air flow from the common outlet from reaching the inlets of the two cooling ducts.

[0019] The gas turbine engine may further include a mesh disposed downstream of the junction and upstream of the common outlet for preventing debris from entering the common outlet of the two cooling ducts.

[0020] The cooling system of the present disclosure can provide cooling to the engine core without requiring any additional power or actuators, thereby enhancing cooling without affecting the efficiency of the gas turbine engine. In addition, the cooling system may require minimal maintenance or repair.

[0021] As described elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may include an engine core including a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades) located upstream of the engine core.

[0022] The arrangement of the present disclosure may be particularly, but not exclusively, beneficial to fans driven via a gearbox. Thus, the gas turbine engine may include a gearbox that receives input from a mandrel and outputs a drive to the fan so as to drive the fan at a rotational speed lower than the mandrel. The input to the gearbox may come directly from the mandrel or indirectly from the mandrel, for example via a spur gear shaft and / or gears. The mandrel may rigidly connect the turbine and the compressor so that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0023] As described and / or claimed herein, a gas turbine engine may have any suitable general architecture. For example, a gas turbine engine may have any desired number of shafts, such as one, two, or three shafts, connecting a turbine and a compressor. By way of example only, the turbine connected to the mandrel may be a first turbine, the compressor connected to the mandrel may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0024] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor.The second compressor may be arranged to receive flow (eg directly, such as via a generally annular conduit) from the first compressor.

[0025] The gearbox may be arranged to be driven by the spindle configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first spindle in the above example). For example, the gearbox may be arranged to be driven only by the spindle configured to rotate (e.g. in use) at the lowest rotational speed (e.g. in the above example, only the first spindle, not the second spindle). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first and / or second shafts in the above example.

[0026] The gearbox can be a reduction gearbox (because the output to the fan is at a lower rotational rate than the input from the mandrel). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example, about or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. For example, the gear ratio can be between any two values ​​in the previous sentence. By way of example only, the gearbox can be a "stellar" gearbox having a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0027] In any gas turbine engine as described and / or claimed herein, the burner may be disposed axially downstream of the fan and one or more compressors. For example, where a second compressor is provided, the burner may be located directly downstream of the second compressor (e.g., at its outlet). By way of another example, where a second turbine is provided, the flow at the burner outlet may be provided to the inlet of the second turbine. The burner may be disposed upstream of one or more turbines.

[0028] The compressor or each compressor (e.g., the first and second compressors described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, which may be variable stator vanes (because the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0029] The turbine or each turbine (e.g., the first turbine and the second turbine as described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0030] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas wash position or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or approximately) any of the following: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.32. These ratios may generally be referred to as hub-tip ratios. Both the radius at the hub and the radius at the tip may be measured at the leading edge (or axially forward) portion of the blade. Of course, the hub-to-tip ratio refers to the gas-washed portion of the fan blade, ie, the portion radially outside of any platform.

[0031] The radius of the fan may be measured between the engine centerline and the tips at the leading edges of the fan blades. The fan diameter (which may be just twice the fan radius) may be greater than (or approximately) any of the following: 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The fan diameter may be within an inclusive range defined by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within the range of 240 cm to 280 cm or 330 cm to 380 cm.

[0032] The rotational speed of the fan can vary in use. Generally speaking, for a fan with a larger diameter, the rotational speed is lower. By way of non-limiting example only, the rotational speed of the fan under cruising conditions may be less than 2500rpm, for example, less than 2300rpm. By way of another non-limiting example only, for an engine with a fan diameter in the range of 220cm to 300cm (e.g., 240cm to 280cm or 250cm to 270cm), the rotational speed of the fan under cruising conditions may be in the range of 1700rpm to 2500rpm, for example, in the range of 1800rpm to 2300rpm, for example, in the range of 1900rpm to 2100rpm. By way of another non-limiting example only, for an engine with a fan diameter in the range of 330cm to 380cm, the rotational speed of the fan under cruising conditions may be in the range of 1200rpm to 2000rpm, for example, in the range of 1300rpm to 1800rpm, for example, in the range of 1400rpm to 1800rpm.

[0033] When a gas turbine engine is in use, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to move at a speed U 尖端 The work done by the fan blades on the flow causes an enthalpy rise dH in the flow. The fan tip load can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端 is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (may be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip load at cruise conditions may be greater than (or approximately) any of the following: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all units in this paragraph are Jkg -1 K -1 / (ms -1 ) 2 ). The fan tip load may be within an inclusive range bounded by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0034] A gas turbine engine according to the present disclosure may have any desired bypass ratio, wherein the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass passage to the mass flow rate of the flow through the core under cruise conditions. In some arrangements, the bypass ratio may be greater than (or approximately) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within a range of 12 to 16, or a range of 13 to 15, or a range of 13 to 14. The bypass passage may be substantially annular. The bypass passage may be located radially outside the engine core. The radially outer surface of the bypass passage may be defined by a nacelle and / or a fan case.

[0035] The total pressure ratio of the gas turbine engine 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 outlet of the highest pressure compressor (before entering the combustor). By way of non-limiting example, the total pressure ratio of the gas turbine engine as described and / or claimed herein at cruise may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), such as within a range of 50 to 70.

[0036] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. Under cruise conditions, the specific thrust of the engine described and / or claimed herein may be less than (or approximately) any of the following: 110 Nkg -1 s, 105Nkg -1 s、100Nkg -1 s, 95Nkg -1 s、90Nkg -1 s, 85Nkg -1 s or 80Nkg -1 s. The specific thrust may be within an inclusive range defined by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within 80 Nkg -1 s to 100Nkg -1 s, or 85Nkg -1 s to 95Nkg -1 Such engines are likely to be particularly efficient compared to conventional gas turbine engines.

[0037] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein may produce a maximum thrust of at least (or approximately) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit). By way of example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, such as 350 kN to 400 kN. The thrust mentioned above may be the maximum net thrust under standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), when the engine is stationary.

[0038] In use, the temperature of the flow at the inlet of the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustor, for example just upstream of the first turbine vane, which itself may be referred to as the nozzle guide vane. At cruise, the TET may be at least (or approximately) any one of the following: 1400K, 1450K, 1500K, 1550K, 1600K, or 1650K. The TET at cruise may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit). The maximum TET of the engine when in use may be, for example, at least (or approximately) any one of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within a range of 1800K to 1950K. Maximum TET may occur, for example, at high thrust conditions, such as at maximum takeoff (MTO) conditions.

[0039] The fan blades and / or airfoil portions of the fan blades described and / or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils may be manufactured at least in part from a composite material, such as a metal matrix composite material and / or an organic matrix composite material, such as carbon fiber. In another exemplary manner, at least a portion of the fan blades and / or airfoils may be manufactured at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions manufactured from different materials. For example, the fan blade may have a protective leading edge, which may be manufactured from a material that resists impact (e.g., from birds, ice, or other materials) better than the rest of the blade. Such a leading edge may be manufactured, for example, using titanium or a titanium-based alloy. Therefore, by way of example only, the fan blade may have carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0040] A fan as described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixing that may engage with a corresponding slot in a hub (or disc). By way of example only, such fixings may be in the form of a dovetail that may be inserted into and / or engage with a corresponding slot in the hub / disc so as to fix the fan blade to the hub / disc. In another example, the fan blade may be formed integrally with the central portion. Such an arrangement may be referred to as a blade disk or blade ring. Any suitable method may be used to manufacture such a blade disk or blade ring. For example, at least a portion of the fan blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0041] 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 outlet area of ​​the bypass passage to vary during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0042] A fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades.

[0043] As used herein, cruise condition has a conventional meaning and will be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person will immediately recognize that cruise condition refers to the operating point of the engine for which the gas turbine engine is designed to be attached to the aircraft at mid-cruise for a given mission (which may be referred to in the industry as an "economy mission"). In this regard, mid-cruise is a critical point in the flight cycle of the aircraft at which 50% of the total fuel burned between the peak of ascent and the beginning of descent has been burned (which may be approximated in terms of time and / or distance to the midpoint between the peak of ascent and the beginning of descent. Thus, cruise condition defines the operating point of the gas turbine engine that provides a thrust that will ensure steady-state operation of the aircraft at mid-cruise (i.e., maintaining a constant altitude and a constant Mach number) taking into account the number of engines provided to the aircraft to which the gas turbine engine is designed to be attached. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, then under cruise condition, the engine provides half of the total thrust required for steady-state operation of the aircraft at mid-cruise.

[0044] In other words, for a given gas turbine engine of an aircraft, the cruise condition is defined as the operating point of the engine that provides a specified thrust at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at mid-cruise altitude) (necessary to provide steady-state operation of the aircraft to which the gas turbine engine is designed to be attached, in combination with any other engines on the aircraft, at a given mid-cruise Mach number). For any given gas turbine engine of an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, so the operating point of the engine at cruise conditions is well defined.

[0045] By way of example only, the forward speed at cruise conditions may be anywhere within the range of from Mach 0.7 to Mach 0.9, such as 0.75 to 0.85, such as 0.76 to 0.84, such as 0.77 to 0.83, such as 0.78 to 0.82, such as 0.79 to 0.81, such as about Mach 0.8, about Mach 0.85, or 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise condition may be outside these ranges, such as below Mach 0.7 or above Mach 0.9.

[0046] By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere ISA) at an altitude within the following range: 10000 m to 15000 m, such as within the range of 10000 m to 12000 m, such as within the range of 10400 m to 11600 m (about 38000 feet), such as within the range of 10500 m to 11500 m, such as within the range of 10600 m to 11400 m, such as within the range of 10700 m (about 35000 feet) to 11300 m, such as within the range of 10800 m to 11200 m, such as within the range of 10900 m to 11100 m, such as approximately 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0047] By way of example only, the cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 ft (11,582 m). By way of another example only, the cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 ft (10,668 m).

[0048] In use, the gas turbine engines described and / or claimed herein may be operated at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by a cruise condition (e.g., an intermediate cruise condition) of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0049] According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is an aircraft to which the gas turbine engine has been designed for attachment. Thus, the cruise condition according to this aspect corresponds to an intermediate cruise of the aircraft, as defined elsewhere herein.

[0050] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be carried out at cruise conditions (eg, in terms of thrust, atmospheric conditions and Mach number) as defined elsewhere herein.

[0051] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may include (or may be) operation at an intermediate cruise of the aircraft, as defined elsewhere herein.

[0052] The skilled person will appreciate that, unless mutually exclusive, features or parameters described in any one of the above aspects may be applied to any other aspect. In addition, unless mutually exclusive, any features or parameters described herein may be applied to any aspect and / or combined with any other features or parameters described herein. BRIEF DESCRIPTION OF THE DRAWINGS

[0053] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0054] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0055] Figure 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0056] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0057] Figure 4 is a cross-sectional side view of a gas turbine engine including a cooling system having cooling ducts;

[0058] FIG. 5A to FIG. 5C yes Figure 4 A perspective view, a top view and a front view of a cooling duct of a cooling system;

[0059] Figure 6 is a cross-sectional top view of a gas turbine engine including a cooling system having two cooling ducts sharing a common outlet;

[0060] FIG. 7A to FIG. 7D yes Figure 6 A perspective view, a front view, a side view and a top view of a cooling duct of a cooling system;

[0061] Figure 8 is a schematic diagram of a fractal grid used in a cooling system of a gas turbine engine;

[0062] Fig.9A is a schematic front view of a check valve in a closed position for use in a cooling system of a gas turbine engine;

[0063] Fig. 9B is in the open position Fig.9A A schematic front view of a check valve;

[0064] FIG. 10A to FIG. 10C In closed, partially open and open positions FIG. 9A to FIG. 9B Schematic top view of a check valve;

[0065] Fig.11 is a schematic diagram of one or more heat transfer enhancement elements of a cooling system of a gas turbine engine; and

[0066] Fig.12 A schematic diagram of an airflow booster of a gas turbine engine's cooling system. DETAILED DESCRIPTION

[0067] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying drawings. Additional aspects and embodiments will be apparent to those skilled in the art.

[0068] As used herein, a component extends "axially" relative to an axis if the component extends along the axis. A component extends "circumferentially" relative to the axis if the component extends in a circumferential direction defined around the axis. A component extends "radially" relative to the axis if the component extends radially inward or outward relative to the axis.

[0069] Figure 1 A gas turbine engine 10 having a main axis of rotation 9 is shown. The engine 10 includes an air inlet 12 and a pusher fan 23, which produces two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 includes a core 11 that receives the core air flow A. The engine core 11 includes a low-pressure compressor 14, a high-pressure compressor 15, a compressor section 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20 in an axial flow series manner. The nacelle 21 surrounds the gas turbine engine 10 and defines a bypass channel 22 and a bypass exhaust nozzle 18. The bypass air flow B flows through the bypass channel 22. The pusher fan 23 is attached to the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30 and is driven by the low-pressure turbine.

[0070] In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustor section 16, where the compressed air is mixed with fuel and the mixture is burned. The resulting hot combustion products are then expanded through the high-pressure turbine 17 and the low-pressure turbine 19 before being discharged through the core exhaust nozzle 20, thereby driving the high-pressure turbine 17 and the low-pressure turbine 19 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 through a suitable interconnecting shaft 27. The propulsion fan 23 generally provides most of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0071] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown in FIG. Figure 1 ) drive shaft 26, which is coupled to a sun wheel or sun gear 28 of an epicyclic gear arrangement structure 30. Radially outward of the sun gear 28 and intermeshing with the sun gear is a plurality of planet gears 32, which are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously around the sun gear 28 while causing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the pusher fan 23 via a connecting rod 36 so as to drive the fan to rotate about the engine axis 9. Radially outward of the planet gears 32 and intermeshing with the planet gears is a ring gear or annular gear 38, which is coupled to the fixed support structure 24 via a connecting rod 40.

[0072] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" used in this article may refer to the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the propulsion fan 23), respectively, and / or the turbine stage and the compressor stage connected together by the interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the propulsion fan 23). In some literature, the "low-pressure turbine" and "low-pressure compressor" mentioned in this article may be alternatively referred to as "intermediate-pressure turbine" and "intermediate-pressure compressor". In the case of using such alternative nomenclature, the propulsion fan 23 may be referred to as the first or lowest pressure compression stage.

[0073] exist Figure 3 The epicyclic gearbox 30 is shown in more detail by way of example in FIG. Each of the sun gear 28, the planetary gears 32, and the ring gear 38 includes teeth around its periphery for intermeshing with the other gears. However, for clarity, Figure 3 Only an exemplary portion of the teeth are shown. Four planetary gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of the planetary epicyclic gearbox 30 typically include at least three planetary gears 32.

[0074] exist Figure 2 and Figure 3The epicyclic gearbox 30 shown by way of example in FIG. 1 is a planetary type, in which the planet carrier 34 is coupled to the output shaft via a connecting rod 36, in which the ring gear 38 is fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of another example, the epicyclic gearbox 30 may be a star arrangement in which the planet carrier 34 remains fixed, allowing the ring gear (or ring gear) 38 to rotate. In such an arrangement, the pusher fan 23 is driven by the ring gear 38. By way of another alternative example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0075] It should be understood that Figure 2 and Figure 3 The arrangement shown in is exemplary only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to position the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of another example, the connections (such as the gearbox 30) between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may be Figure 2 The connecting rods 36, 40 in the example may have any desired degree of stiffness or flexibility. By way of another example, any suitable arrangement of bearings between rotating and stationary components of the engine (e.g., between the input and output shafts from the gearbox and a stationary structure such as a gearbox housing) may be used, and the present disclosure is not limited to Figure 2 For example, where the gearbox 30 has a star arrangement (as described above), the skilled person will readily appreciate that the arrangement of the output and support links and bearing locations is typically different than Figure 2 The arrangement structure is shown by way of example in FIG.

[0076] Thus, the present disclosure extends to gas turbine engines having any arrangement in gearbox type (eg sun or planetary), support structure, input and output shaft arrangement and bearing locations.

[0077] Optionally, the gearbox may drive additional and / or alternative components (eg, an intermediate pressure compressor and / or a booster compressor).

[0078] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have alternative numbers of compressors and / or turbines and / or alternative numbers of interconnecting shafts. By way of further example, Figure 1The gas turbine engine shown in has a splitter nozzle 18, 20, which means that the flow passing through the bypass channel 22 has its own nozzle 18, which is separated from the core exhaust nozzle 20 and radially located outside the core exhaust nozzle. However, this is not restrictive, and any aspect of the present disclosure may also be applied to an engine in which the flow passing through the bypass channel 22 and the flow passing through the core 11 are mixed or combined before (or upstream) a single nozzle that may be referred to as a mixed flow nozzle. One or two nozzles (whether mixing or splitting) may have a fixed or variable area. Although the examples described relate to turbofan engines, the present disclosure may be applied to, for example, any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0079] The geometry of the gas turbine engine 10 and its components is defined by conventional shafting, including an axial direction (aligned with the main rotation axis 9), a radial direction (in Figure 1 from bottom to top) and circumferential direction (perpendicular to Figure 1 view). The axial, radial, and circumferential directions are perpendicular to each other.

[0080] Furthermore, the present disclosure is equally applicable to aviation gas turbine engines, marine gas turbine engines, and land-based gas turbine engines.

[0081] Figure 4 FIG. 5 is a cross-sectional side view of a gas turbine engine 50. The gas turbine engine 50 may be used for propulsion of an aircraft. The gas turbine engine 50 may be similar to Figure 1 The gas turbine engine 10 of the present invention. The gas turbine engine 50 includes a nacelle 52, an engine core 54 and a pusher fan 55. The engine core 54 is received in the nacelle 52. The engine core 54 defines a main rotation axis 56 along its length. In other words, the engine core 54 and the nacelle 52 are circumferentially arranged around the main rotation axis 56. The engine core 54 also includes a low-pressure compressor 62, a high-pressure compressor 64, a combustor segment 60, a high-pressure turbine 66, a low-pressure turbine 68 and a core exhaust nozzle 70 in an axial flow series manner.

[0082] The engine core 54 and the nacelle 52 define a bypass passage 58 therebetween. The gas turbine engine 50 also includes an air inlet 72 and a bypass exhaust nozzle 74. The pusher fan 55 generates two airflows: a core airflow C and a bypass airflow D. The engine core 54 receives the core airflow C. The bypass airflow D flows through the bypass passage 58. The bypass airflow D may be interchangeably referred to as "bypass air D".

[0083] The gas turbine engine 50 also includes a cooling system 100. The cooling system 100 includes a cooling duct 102. Figure 4In the example of FIG. 1 , the cooling system 100 includes a single cooling conduit 102 . However, in some other examples, the cooling system 100 may include a plurality of cooling conduits 102 similar to the cooling conduit 102 .

[0084] The cooling conduits 102 extend both axially and circumferentially around a segment of the engine core 54 relative to the primary rotation axis 56. In one example, the cooling conduits 102 extend circumferentially around a segment of the engine core 54 including the combustor segment 60 relative to the primary rotation axis 56. The segment of the engine core 54 may be a high pressure segment.

[0085] The cooling duct 102 is a hollow tubular member having an inlet 104 and an outlet 106. The cooling duct 102 defines an inlet 104 for receiving bypass air D from the bypass passage 58 at an upstream position 122 and an outlet 106 for discharging bypass air D at a downstream position 124. The inlet 104 and the outlet 106 may be flush with the outer skin of the core. The cooling duct 102 guides a portion of the bypass air D from the upstream position 122 to the downstream position 124, which is located upstream of the bypass discharge nozzle. The cooling duct 102 also includes a first portion 108, a second portion 110, and a third portion 112. The first portion 108 extends at least axially relative to the main rotation axis 56. The second portion 110 is downstream of the first portion 108. The second portion 110 extends circumferentially around the engine core 54 relative to the main rotation axis 56. The third portion 112 is located downstream of the second portion 110 and extends at least axially relative to the main rotation axis 56.

[0086] The cooling duct 102 surrounds the engine core 54. The cooling duct 102 can guide the bypass air D around the engine core 54 from the upstream position 122 of the bypass passage 58 to the downstream position 124 thereof by means of the Venturi effect. In one example, the cooling duct 102 can be arranged along the outer surface ( Figure 4 In some examples, the cooling conduit 102 contacts the outer surface of the segment of the engine core 54 including the combustor segment 60. In some examples, the cooling conduit 102 may contact and extend along the outer surface of the high pressure segment of the engine core 54 adjacent to the combustor segment 60. The high pressure segment may include a high pressure compressor 64 and a high pressure turbine 66.

[0087] The bypass air D is guided around the engine core 54 by the cooling duct 102. To minimize pressure losses, the Mach number may be maintained substantially constant downstream of the first portion 108 of the cooling duct 102 until a point close to the outlet 106. The cross-sectional area of ​​the cooling duct 102 may be reduced adjacent to the outlet 106 to increase the Mach number and provide the desired thrust. The increase in the Mach number may create a suction force to draw the bypass air D into the inlet 104. In some examples, the difference in Mach number between two points in the cooling duct 102 may be used to create a cross flow around the engine core 54, thereby guiding the bypass air D from the inlet 104 through to the outlet 106.

[0088] In this example, the cooling conduit 102 is spiral. The cooling conduit 102 is spiraled about the engine core 54 relative to the main rotation axis 56 for half a circle (i.e., about 180 degrees). In other examples, the cooling conduit 102 may spiral less than half a circle (i.e., less than 180 degrees) around the engine core 54 or spiral more than half a circle (i.e., greater than 180 degrees) around the engine core 54. In some examples, the cooling conduit 102 may spiral one full circle or more circles (i.e., greater than or equal to 360 degrees) around the engine core 54. Increasing the number of turns of the cooling conduit 102 (or the angular range of the circle, if there is less than one circle or a fraction of a circle) can achieve higher heat exchange. In addition, the smaller cross-section of the cooling conduit 102 can increase the flow rate of the bypass air D, which in turn can further increase the heat exchange. The number of turns in the cooling conduit 102 and the cross-section of the cooling conduit 102 can be changed according to the desired cooling characteristics and the space requirements of the application.

[0089] The cooling system 100 of the gas turbine engine 50 may also include one or more grids 116, 118 disposed within the cooling conduit 102 for preventing debris from entering the cooling conduit 102. Each of the one or more grids may have a rectangular grid pattern. In other examples, each of the one or more grids may have a square, triangular, polygonal, circular, or irregular grid pattern. In some examples, at least one of the one or more grids may be a fractal grid, as described below with reference to Figure 8 Described in detail. Figure 4 In the example shown, two grids are disposed in the cooling conduit 102 to prevent debris from entering the cooling conduit 102. Specifically, a first grid 116 and a second grid 118 are disposed in the cooling conduit 102. The first grid 116 is disposed near the inlet 104. The second grid 118 is disposed near the outlet 106.

[0090] The cooling system 100 may further include one or more check valves disposed within the cooling conduit 102 to prevent air flow from the outlet 106 to the inlet 104. Figure 4In the example shown, the cooling system 100 of the gas turbine engine 50 includes a check valve 120 disposed within the cooling conduit 102. The check valve 120 is disposed proximate the outlet 106. The check valve 120 may be a three-pointed valve. The check valve 120 may prevent backflow in the cooling conduit 102.

[0091] In some examples, the cooling system 100 also includes one or more heat transfer enhancement elements (e.g., heat transfer enhancement elements) disposed between the engine core 54 and the cooling conduit 102. Figure 4 102) for increasing heat transfer from the engine core 54 to the cooling duct 102. The one or more heat transfer enhancement elements may include at least one of fins and heat pipes, as described below with reference to Fig.11 Explain in more detail.

[0092] In some examples, the cooling system 100 further includes one or more airflow boosters ( Figure 4 The one or more airflow boosters include at least one of a nozzle injector, an electric fan, and a micro compressor. Fig.12 The airflow booster is described in more detail.

[0093] The cooling system 100 may be a passive and actuator-free system. In other words, the cooling system 100 may not require any additional components to power the cooling system 100. The cooling system 100 may reduce the temperature of the high-pressure cooling air flowing around the combustor segment 60 of the gas turbine engine 50. The cooling air may flow within the engine core 54. The air may then reduce the temperature of downstream components such as high-pressure and intermediate-pressure turbines, high-pressure and intermediate-pressure blades, high-pressure, intermediate-pressure and / or low-pressure rotating disks, inter-turbine bearing chambers, and other turbine seal flows. The cooling system 100 may improve the cooling effect of the gas turbine engine 50 and then extend the life of its components.

[0094] FIG. 5A to FIG. 5C Different views of the cooling conduit 102 are shown. Figure 5A A perspective view of the cooling conduit 102 is shown. Figure 5B A top view of the cooling conduit 102 is shown. Figure 5C A front view of a cooling conduit 102 is shown. The cooling conduit 102 includes an inlet 104, an outlet 106, a first portion 108, a second portion 110, and a third portion 112. In this example, the cooling conduit 102 has a generally circular cross-section along its length. In some examples, the cooling conduit 102 may have one of an elliptical, rectangular, and polygonal cross-section along its length. FIG. 5A to FIG. 5CIn the example shown, each of the inlet 104 and the outlet 106 has a generally elliptical shape. In this example, the area of ​​the inlet 104 is different from the area of ​​the outlet 106. Specifically, the area of ​​the inlet 104 is smaller than the area of ​​the outlet 106. In some other examples, the area of ​​the inlet 104 may be equal to or greater than the area of ​​the outlet 106.

[0095] The first portion 108 extends at least axially and radially relative to the main rotation axis 56. The first portion 108 may also extend circumferentially relative to the main rotation axis 56. The second portion 110 is located downstream of the first portion 108 and is connected to the first portion 108 via a first bend 126. The second portion 110 extends circumferentially around the engine core 54 relative to the main rotation axis 56. The second portion 110 may be a spiral portion of the cooling duct 102. In this example, the second portion 110 may extend circumferentially about 180 degrees relative to the main rotation axis 56. However, the second portion 110 may extend circumferentially less than 180 degrees or more than 180 degrees. In some examples, the second portion 110 may form one or more turns relative to the main rotation axis 56. The third portion 112 is located downstream of the second portion 110. The second portion 110 is connected to the third portion 112 via a second bend 128. The third portion 112 also extends at least axially and radially relative to the main rotation axis 56.

[0096] The first bend 126 may ensure a smooth transition between the first portion 108 and the second portion 110. Similarly, the second bend 128 may allow a smooth transition between the second portion 110 and the third portion 112. Therefore, the first bend 126 and the second bend 128 may allow the bypass air D to flow smoothly between the first portion 108, the second portion 110, and the third portion 112.

[0097] Figure 6 2 is a cross-sectional top view of a gas turbine engine 50 including a cooling system 200. The cooling system 200 includes two cooling ducts 202A, 202B. The two cooling ducts 202A, 202B merge with each other at a junction 226 downstream of their inlets 204A, 204B, such that the two cooling ducts 202A, 202B share a common outlet 206.

[0098] Each of the cooling ducts 202A, 202B extends both axially and circumferentially around a segment of the engine core 54 relative to the primary rotation axis 56. In one example, each of the cooling ducts 202A, 202B extends circumferentially around a segment of the engine core 54 including the combustor segment 60 relative to the primary rotation axis 56.

[0099] Each of the cooling conduits 202A, 202B is a hollow tubular member and defines a respective inlet 204A, 2024B for receiving bypass air D from the bypass passage 58 at an upstream position 222 and a common outlet 206 for discharging the bypass air D at a downstream position 224.

[0100] In this example, each of the two cooling ducts 202A, 202B includes a first portion 208A, 208B and a second portion 210A, 210B. The first portion 208A of the cooling duct 202A defines the inlet 204A. Similarly, the first portion 208B of the cooling duct 202B defines the inlet 204B. Each of the first portions 208A, 208B extends at least axially relative to the main rotation axis 56. Each of the second portions 210A, 210B is located downstream of the corresponding first portion 208A, 208B. Specifically, the second portion 210A is located downstream of the first portion 208A. The second portion 210B is located downstream of the first portion 208B. Each of the second portions 210A, 210B extends circumferentially around the engine core 54 relative to the main rotation axis 56. Each of the two cooling conduits 202A, 202B merges with each other at a junction 226 downstream of each of the second portions 210A, 210B. In this example, the two cooling conduits 202A, 202B also include a common third portion 212 located downstream of the junction 226. The third portion 212 extends at least axially relative to the main rotation axis 56. The third portion 212 defines a common outlet 206.

[0101] In some examples, cooling system 200 includes one or more grids, one or more check valves, one or more heat transfer enhancement elements ( Figure 6 ) and one or more airflow boosters ( Figure 6 not shown).

[0102] exist Figure 6In the example shown, the cooling system 200 includes a first grid 216, a second grid 218, and a third grid 219. The first grid 216 and the second grid 218 are respectively arranged near the inlet 204A, 204B. Specifically, the first grid 216 is arranged in the cooling conduit 202A near the inlet 204A. The first grid 216 can prevent debris from entering the cooling conduit 202A via the inlet 204A. The second grid 218 is arranged in the cooling conduit 202B near the inlet 204B. The second grid 218 can prevent debris from entering the cooling conduit 202B via the inlet 204B. The cooling system 200 also includes a third grid 219 arranged downstream of the joint 226 and upstream of the common outlet 206 to prevent debris from entering the common outlet 206 of the two cooling conduits 202A, 202B. Therefore, only one grid near the common outlet 206 may be required in the cooling system 200.

[0103] exist Figure 6 In the example shown, the cooling system 200 further includes a check valve 220, which is disposed downstream of the junction 226 and upstream of the common outlet 206 to prevent air flow from the common outlet 206 to the inlets 204A, 204B of the two cooling ducts 202A, 202B. Therefore, the check valve 220 can prevent backflow in the cooling ducts 202A, 202B.

[0104] Although in this example, the cooling system 200 has two cooling ducts, the cooling system may include more than two cooling ducts. The cooling ducts may be arranged around (e.g., symmetrically around) the engine core 54, which may reduce potential distortion of the flow through the bypass passage 58. Figure 6 In the example shown, the two cooling ducts 202A, 202B have a symmetrical configuration that may allow for installation of other components around the engine core 54. In addition, the inlets 204A, 204B may also be disposed in suitable locations that allow for installation of other components such as stays.

[0105] Figure 6 The inlets 204A, 204B are shown disposed on the upper side of the engine core 54 and the common outlet 206 is disposed on the lower side of the engine core 54. However, in some other examples, the inlets 204A, 204B may be disposed on the lower side of the engine core 54 and the common outlet 206 may be disposed on the upper side of the engine core 54. The inlets may also be disposed on opposite sides of the engine core 54.

[0106] FIG. 7A to FIG. 7D They are shown respectively Figure 6 A perspective view, a front view, a side view, and a top view of two cooling conduits 202A, 202B of the cooling system 200 are shown.

[0107] Each of the two cooling ducts 202A, 202B includes a first portion 208A, 208B and a second portion 210A, 210B. Each of the first portions 208A, 208B extends at least axially relative to the main rotation axis 56. In this example, each of the first portions 208A, 208B extends axially, circumferentially, and radially relative to the main rotation axis 56. Each of the second portions 210A, 210B extends circumferentially around the engine core 54 relative to the main rotation axis 56. The first portion 208A of the cooling duct 202A is connected to the second portion 210A via a bend 226A. The bend 226A can achieve a smooth transition between the first portion 208A and the second portion 210A. In addition, the first portion 208B of the cooling duct 202B is connected to the second portion 210B via a bend 226B. The bend 226B can achieve a smooth transition between the first portion 208B and the second portion 210B. Each of the two cooling conduits 202A, 202B merges with each other at a junction 226 downstream of each of the second portions 210A, 210B. The two cooling conduits 202A, 202B also include a common third portion 212 extending at least axially relative to the main rotation axis 56. Each of the first portions 208A, 208B is radially inclined at an angle relative to the main rotation axis 56. Similarly, the common third portion 212 is also radially inclined at an angle relative to the main rotation axis 56. In some other examples, the first portions 208A, 208B and the common third portion 212 may be substantially parallel to the main rotation axis 56.

[0108] Each of the second portions 210A, 210B may be a spiral portion of the respective cooling conduit 202A, 202B. In this example, each of the second portions 210A, 210B may extend circumferentially about 180 degrees relative to the main axis of rotation 56. However, each of the second portions 210A, 210B may extend circumferentially less than 180 degrees or greater than 180 degrees. In some examples, each of the second portions 210A, 210B may form one or more turns relative to the main axis of rotation 56.

[0109] exist FIG. 7A to FIG. 7D In the example shown, each of the two cooling conduits 202A, 202B has a circular cross-section along its length. In some other examples, each of the two cooling conduits 202A, 202B may have one of an elliptical, rectangular, and polygonal cross-section along its length. The cooling conduit 202A may be a mirror image of the cooling conduit 202B. However, in some other examples, the two cooling conduits 202A, 202B may not be mirror images of each other. FIG. 7A to FIG. 7DIn the example shown, each of the inlets 204A, 204B has an elliptical shape. In addition, the common outlet 206 has an elliptical shape. The area of ​​each of the inlets 204A, 204B is different from the area of ​​the common outlet 206. Specifically, the area of ​​the common outlet 206 is greater than the area of ​​each of the inlets 204A, 204B. The common outlet 206 can discharge the combined flow received at the inlets 204A, 204B. The areas of the inlets 204A, 204B and the common outlet 206 can be changed according to the desired flow characteristics.

[0110] Figure 8 A schematic diagram of a fractal grid 300 is shown. Figure 8 A fractal grid having a particular fractal pattern is shown, but it should be understood that this is for illustration purposes only and that other fractal grids may be used. The grid 300 includes a plurality of microgrids arranged to follow a fractal or recursive pattern. In the grid 300, the microgrids may also include smaller microgrids or "fractal cells". The grid 300 may act as a turbulator and create turbulence in the airflow, thereby increasing heat exchange with the engine core 54. The grid 300 may be arranged in Figure 4 1. The mesh 300 may also prevent debris from entering the cooling conduit 102. In contrast to conventional foreign object debris (FOD) screens, the mesh 300 may further reduce pressure drop to increase airflow in the cooling conduit 102. In some examples, the mesh 300 may be disposed proximate to the inlet 104 and / or outlet 106 of the cooling conduit 102. The mesh 300 may also be disposed at Figure 6 The grid 300 can prevent debris from entering the cooling conduits 202A, 202B. In some examples, the grid 300 can be disposed proximate to the inlets 204A, 204B and / or the common outlet 206.

[0111] Fig.9A and Fig. 9B 4 is a front view of the check valve 400. In this example, the check valve 400 is a tri-point valve. The check valve 400 is also an actuator-less valve. In other words, the check valve 400 may not require a separate actuator. The check valve 400 is a one-way valve that provides one-way flow thereon. Fig.9A The closed position of the check valve 400 is shown. Fig. 9B The open position of the check valve 400 is shown. The check valve 400 includes a flap 402 and a leg 404. The flap 402 may be approximately triangular. The leg 404 may prevent the check valve 400 from opening beyond a desired angle. With the aid of the leg 404, different intermediate positions of the check valve 400 are also possible. The check valve 400 may be arranged in the cooling system 100 ( Figure 4104) to prevent air flow from the outlet 106 to the inlet 104. The check valve 400 may be in a closed position when air flow is from the outlet 106 to the inlet 104, and may be in an open position during normal air flow. The check valve 400 may be disposed near the outlet 106. The check valve 400 may also be disposed at Figure 6 The cooling system 200 is shown to prevent airflow from the common outlet 206 from reaching the inlets 204A, 204B.

[0112] FIG. 10A to FIG. 10C Shows Figure 4 The cross section of the cooling conduit 102 includes Fig.9A and Fig. 9B The check valve 400 shown. In this example, the check valve 400 is disposed in the cooling duct 102. The check valve 400 allows a unidirectional airflow AF in the cooling duct 102. The airflow AF exists in both the top side and the bottom side of the fin 402. The legs 404 can ensure a minimum angle of the respective fin 402 relative to the axis of the cooling duct 102. The legs 404 can hold the fins 402 in an open position. In the event of a reverse airflow, the fins 402 can be hinged to contact each other so that the check valve 400 is in a closed position. See FIG. 10A to FIG. 10C , the check valve 400 is shown in a closed position, a partially open position, and an open position, respectively.

[0113] Fig.11 FIG. 5 is a cross-sectional view of the engine core 54 and the cooling system 200, showing one or more optional heat transfer enhancement elements 502 of the cooling system 200. Note that the combustor can 504 in the combustor segment 60 is Fig.11 The one or more heat transfer enhancement elements 502 are disposed between the engine core 54 and the cooling conduits 202A, 202B, and in this example are arranged around the core 54. The heat transfer enhancement elements 502 may be disposed in an outer wall or casing of the engine core 54, or otherwise disposed at the junction between the cooling conduits 202A, 202B and the engine core 54.

[0114] The one or more heat transfer enhancement elements 502 may include fins and / or heat pipes and may promote or enhance turbulence in the airflow around the core 54 to increase the heat removal rate. The cooling air may flow around the heat pipes and / or fins. The heat transfer enhancement element 502 may have different shapes depending on the desired application characteristics. For example, the fins may be rectangular, triangular, shark fin-shaped or cylindrical geometric shapes or any other suitable geometric shapes. The heat pipes may be rectangular, cylindrical or curved geometric shapes or any other suitable geometric shapes. The heat pipes may be used together with the fins to further increase the heat removal rate.

[0115] Fig.12Schematic diagram of an optional airflow booster 600 of the cooling system 200 of the gas turbine engine 50. The airflow booster 600 may be, for example, one or more of a nozzle ejector, an electric fan, or a micro compressor, and introduces a relatively small amount of high-pressure bleed air to increase the amount of airflow through the cooling duct 202A. This can increase the amount of airflow through the cooling ducts 102, 202A, 202B. Each nozzle ejector in the nozzle ejector may be a converging nozzle design or a converging-dispersing nozzle design. In some examples, each nozzle ejector in the nozzle ejector may have a small diameter, such as less than 5 mm, less than 4 mm, less than 2 mm, or less than 1 mm. The nozzle ejector may be directly welded or bolted to the outer casing of the engine core 54. The airflow booster 600 may be arranged near the outlet 106 or a common outlet 206.

[0116] Although heat transfer enhancement element 502 and airflow enhancer 600 have been generally described in the context of a cooling system 200 having two ducts, it should be understood that they may be equally used in a system 100 having one duct or a system having more than two ducts.

[0117] It should be understood that the present invention is not limited to the above-described embodiments, and various modifications and improvements may be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A gas turbine engine, comprising: Nacelle; an engine core received within the nacelle and defining a primary rotational axis along a length of the engine core, the engine core and the nacelle defining a bypass passage therebetween; and a cooling system including a cooling duct defining an inlet for receiving bypass air from the bypass passage at an upstream location and an outlet for discharging the bypass air back into the bypass passage at a downstream location and extending both axially and circumferentially about a segment of the engine core relative to the primary rotational axis; The cooling duct comprises: a first portion extending at least axially relative to the main rotation axis; a second portion located downstream of the first portion and extending circumferentially around the engine core relative to the main rotation axis; and a third portion located downstream of the second portion and extending at least axially relative to the main rotation axis, The cooling system further includes one or more heat transfer enhancement elements disposed between the engine core and the cooling duct for increasing heat transfer from the engine core to the cooling duct.

2. The gas turbine engine of claim 1, wherein the engine core comprises a combustor segment, and wherein the cooling duct extends circumferentially relative to the main rotation axis around a segment of the engine core comprising the combustor segment. 3 . The gas turbine engine of claim 2 , further comprising one or more grids disposed within the cooling duct for preventing debris from entering the cooling duct. 4 . The gas turbine engine of claim 3 , wherein the one or more grids include a first grid disposed proximate to the inlet and a second grid disposed proximate to the outlet. 5 . The gas turbine engine of claim 3 , wherein at least one of the one or more grids is a fractal grid. 6 . The gas turbine engine of claim 1 , further comprising a check valve disposed within the cooling duct for preventing air flow from the outlet to the inlet.

7. The gas turbine engine of claim 6, wherein the check valve is disposed proximate to the outlet.

8. The gas turbine engine of claim 6, wherein the check valve is a tri-pointed valve.

9. The gas turbine engine of claim 1, wherein the one or more heat transfer enhancement elements include at least one of fins and heat pipes.

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