Gas turbine engine

Abstract

The present invention is entitled gas turbine engine. The present disclosure provides a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core, the engine core comprising a turbine, a compressor and a core shaft connecting the turbine to the compressor; and a fan, the fan being located upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of series fan blades. Each of the series fan blades comprises a main fan blade and an auxiliary fan blade, the auxiliary fan blade being positioned at the rear of the main fan blade, so that over substantially the entire radial span of the auxiliary fan blade, the leading edge of the auxiliary fan blade is behind the nearest point on the trailing edge of the main fan blade, and on a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the rearward extension of the arc line of the airfoil chord section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate one after the other. The auxiliary fan blade is movable within a pitch angle range relative to the main fan blade.

Description

Technical Field

[0001] The present disclosure relates to gas turbine engines and, more particularly, to gas turbine engines having tandem fan blades. Background Art

[0002] Modern gas turbine engines typically include a fan positioned upstream of the engine core, the fan having a circumferential row of fan blades mounted to the fan hub. The fan is driven by a turbine in the engine core to generate an airflow. The radially outer portion of the generated airflow enters a bypass duct to provide propulsive thrust. The remainder of the airflow flows toward the fan hub to the engine core to power the Brayton cycle and provide additional propulsive thrust.

[0003] In a typical turbofan engine, the fan is driven directly by a shaft extending from the engine's low-pressure turbine (LPT) to the fan, so the fan rotates at the same rotational speed as the LPT. In some recently developed turbofan engines, the fan and the low-pressure compressor (LPC) at the core of the engine are driven by the LPT, which is directly driven by a shaft extending from the LPT, and the fan is driven through a power gearbox that receives its drive from the shaft. Therefore, the fan rotates at a lower speed than the LPT. This arrangement can improve the efficiency of the gas turbine engine because for a given specific thrust, the fan tip load increases as the fan reduces its rotational speed, thereby improving the bypass efficiency.

[0004] However, due to the increased fan tip loading, the fan may be more susceptible to self-excited vibrations (known as flutter). More specifically, when the fan blades in a row of blades vibrate, they generate unstable aerodynamic forces. Under most conditions, these unstable aerodynamic forces cause the blades to do work on the surrounding air, and the amplitude of the vibrations decays. However, under certain operating conditions (e.g., certain rotational speeds and / or thrusts and / or combinations thereof), the surrounding air can do work on the fan itself. If the work done by the air exceeds the work dissipated (e.g., through mechanical damping), the vibrations will increase. Such flutter is undesirable because it can produce greater stresses in the engine. Flutter is further discussed in European patent application EP 3361050 A1.

[0005] One method of reducing the susceptibility of fan blades to flutter is to decamber the hub and root of the fan blades, for example, by reducing the angular variation of the camber extending between the leading edge and the trailing edge on a given airfoil chord section. Doing so can reduce the torsional component of the vibration response in the first winglet vibration mode, which can result in improved flutter behavior. However, such a reduction in camber at the bottom of the fan blade reduces the pressure ratio generated by the fan at the fan hub. That is, the ratio of the average total pressure of the airflow entering the engine core to the average total pressure at the fan inlet is reduced. Moreover, in combination with mechanical constraints, the combination of hub chord and camber may not be aerodynamically optimal. Therefore, the performance of the engine core and the overall efficiency or gross weight of the engine may be negatively affected.

[0006] It is therefore desirable to provide a gas turbine engine having improved overall performance and efficiency while reducing its susceptibility to flutter. Summary of the invention

[0007] The present disclosure provides a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0008] an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; and

[0009] a fan located upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades;

[0010] wherein each of the tandem fan blades comprises a main fan blade and an auxiliary fan blade, the auxiliary fan blade being positioned at the rear of the main fan blade such that over substantially the entire radial span of the auxiliary fan blade, the leading edge of the auxiliary fan blade is rearward of the nearest point on the trailing edge of the main fan blade, and over a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on a rearward extension of the camber line of the airfoil chord section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate one behind the other; and

[0011] The auxiliary fan blades are capable of moving relative to the main fan blades within a pitch angle range.

[0012] Since the leading edge of the auxiliary fan blade is behind the nearest point on the trailing edge of the main fan blade, the auxiliary fan blade will not produce a significant harmful aerodynamic effect on the airflow passing through the main fan blade. In addition, since the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the extension of the camber of the airfoil chord section of the main fan blade, and since the main fan blade and the auxiliary fan blade are arranged to rotate one after the other (i.e., at the same rotational speed), each auxiliary fan blade can be advantageously used as an extension of its corresponding main fan blade, wherein the airflow leaves the main fan blade along a smooth flow path leading to the auxiliary fan blade. This can allow the main fan blade to have greater design freedom. Specifically, a shallower camber (i.e., a reduced camber) can be implemented on the bottom of the main fan blade to reduce the susceptibility to flutter by reducing the torsional component of the vibration response in the first wing vibration mode, while the associated fan root pressure ratio drop can be at least partially alleviated by the auxiliary fan blade. For example, including an auxiliary fan blade can effectively extend the camber of the main fan blade, which can then increase the total camber angle of its combination and increase the fan hub pressure ratio accordingly.

[0013] On any given airfoil, the camber line extends from the leading edge position of the section to the trailing edge position. Along the circumferential direction in the plane of the airfoil section, the camber line is equidistantly spaced from the pressure and suction surfaces of the airfoil. Throughout this application, the phrase "total camber angle" refers to the combined camber angle of a pair of main fan blade and auxiliary fan blade airfoil chord sections, and the chord sections are paired when the leading edge position of the auxiliary fan blade airfoil chord section is located on the rearward extension of the camber line of the main fan blade airfoil chord section. That is, the total camber angle is the angle between the direction of the camber line at the leading edge position of the main fan blade airfoil chord section and the direction of the camber line at the trailing edge position of its paired auxiliary fan blade airfoil chord section.

[0014] The present disclosure also provides a fan of the aforementioned gas turbine engine. For example, the fan may include a circumferential row of serial fan blades, wherein each of the serial fan blades includes a main fan blade and an auxiliary fan blade, the auxiliary fan blade is positioned at the rear of the main fan blade, so that on a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the rearward extension of the camber of the airfoil chord section of the main fan blade, and wherein the main fan blade and the auxiliary fan blade are arranged to rotate one after the other.

[0015] The present disclosure also provides a series fan blade of the aforementioned gas turbine engine. For example, the series fan blade may include a main fan blade and an auxiliary fan blade, the auxiliary fan blade being positioned at the rear of the main fan blade, so that on a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the rearward extension of the camber line of the airfoil chord section of the main fan blade, and wherein the main fan blade and the auxiliary fan blade are arranged to rotate one in front and one behind.

[0016] Optional features of the present disclosure will now be described. These features may be applied alone or in any combination with any aspect of the present disclosure.

[0017] The auxiliary fan blade may be positioned so that over substantially the entire radial span thereof, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the rearward extension of the camber line of the corresponding airfoil chord section of the main fan blade. For example, the auxiliary fan blade may be positioned so that over at least 90%, 95% or 100% of its radial span, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on the extension of the camber line of the corresponding airfoil chord section of the main fan blade.

[0018] The gas turbine engine may further include an annular splitter located behind the fan and configured to split the airflow generated by the fan into a core airflow flowing through the engine core and a bypass airflow flowing through a bypass duct surrounding the engine core, wherein the auxiliary fan blades have radially outer tips that are inboard of or flush with the radial position of the annular splitter. In this way, the auxiliary fan blades can sweep only the inlet of the engine core and thus can act on the core flow without interfering with the bypass flow.

[0019] In the side elevation view, the projected area of ​​the auxiliary fan blade of each tandem fan blade may be smaller than the projected area of ​​the corresponding main fan blade. For example, the projected area of ​​the auxiliary fan blade of each tandem fan blade may be less than 25% of the projected area of ​​the corresponding main fan blade, and preferably less than 20%, 15%, 10%, 5% or 3%.

[0020] The main fan blade and the auxiliary fan blade may be configured such that in a side elevation view, the trailing edge of the main fan blade has the appearance of a cutout region at the rear of the base of the main fan blade, which surrounds the auxiliary fan blade. More specifically, the main fan blade radially outward of the auxiliary blade may extend over the tip of the auxiliary fan blade. This may allow the trailing edge of the main fan blade radially outward of the auxiliary blade to form an extension of the line of the trailing edge of the auxiliary fan blade, thereby reducing undesirable flow disturbances at the tip of the auxiliary blade.

[0021] Thus, in side elevation view, the cutout region of the primary fan blade may completely surround the leading edge and tip of the secondary fan blade, so that the total axial extent of the tandem fan blade (and therefore the axial extent of the fan rotor disk for carrying the tandem fan blade) may be the same as the axial extent of a fan having a corresponding single fan blade. This may facilitate retrofitting the tandem fan blade and its corresponding fan rotor disk to existing gas turbine engines, as well as reducing the amount of engine redesign that may be required to accommodate the tandem fan blade.

[0022] The auxiliary fan blade can move within a pitch angle range relative to the main fan blade. For example, the auxiliary fan blade may be able to rotate or pivot around its leading edge, for example, by a gear-based variable pitch mechanism. The variable pitch mechanism may be provided on the fan rotor disk. The variable pitch mechanism may be configured to simultaneously change the pitch angle of all auxiliary fan blades mounted on the fan rotor disk. Advantageously, a change in the pitch angle of the auxiliary fan blades changes the combined camber toward the bottom of the tandem fan blades, thereby changing the total camber angle there. Therefore, the fan hub pressure ratio can be controlled by changing the pitch angle of the auxiliary fan blades. This is particularly beneficial because the fan hub pressure ratio of the turbine engine can be selectively changed according to flight conditions (e.g., take-off and cruise conditions).

[0023] The auxiliary fan blades may be capable of moving within a pitch angle range of at least 10° and preferably at least 20°, whereby increasing the pitch angle increases the total camber angle of a given airfoil chord section of the main and auxiliary fan blades by a corresponding amount. For example, during aircraft takeoff, when the gas turbine engine provides the maximum amount of thrust during flight, the auxiliary fan blades may move to a pitch angle of 0° to reduce the total camber angle. Conversely, during aircraft cruise, the auxiliary fan blades may move to a larger pitch angle (e.g., at least 10°) to increase the total camber angle, thereby increasing the fan hub pressure ratio. For example, under cruise operating conditions, the fan hub pressure ratio may increase from 1.2 to 1.3 or above 1.3.

[0024] The auxiliary fan blades may be configured to contribute at least 20% of the total work done by the tandem fan blades under cruise operating conditions of the engine. Here and hereinafter, the term "total work done" refers to C p .ΔTAT, where C p is the specific heat at constant pressure, and ΔTAT is the rise in total air temperature.

[0025] The auxiliary fan blades may be configured such that moving the auxiliary fan blades through the total pitch angle range changes the total work done by the fan by at least 20%.

[0026] However, another option is that the pitch angle of the auxiliary fan blades is fixed, i.e., non-adjustable. This can reduce the complexity of the fan and associated controls, and can reduce the overall weight of the fan. The fixed pitch angle auxiliary fan blades can provide a total camber angle of about 10° at the base of the tandem fan blades. This can optimize the fan hub pressure ratio for cruise operating conditions.

[0027] The fan may also include a fan rotor disk to which the circumferential row of tandem fan blades is attached, wherein each auxiliary fan blade has a root portion that engages a complementary formation formed in the fan rotor disk to removably mount the auxiliary fan blade to the fan rotor disk. This mounting arrangement may be particularly convenient when the pitch angle of the auxiliary fan blades is fixed. For example, the root portion may be a dovetail and the complementary formation is a matching groove into which the dovetail can slide. Typically, the main fan blade has a mounting arrangement similar to the rotor disk.

[0028] The main fan blade and the auxiliary fan blade of each tandem fan blade in the tandem fan blade can be removed independently for upgrading or inspection and maintenance. Specifically, only the auxiliary fan blade can be replaced with a fan blade with a different configuration to achieve, for example, a different total camber angle.

[0029] The main fan blade may have a radial span extending from the hub at a 0% span position to the tip at a 100% span position, wherein the average camber angle of the radially innermost 10% of the radial span of the main fan blade is less than 75% of the average camber angle of the 10% portion of the radial span of the main fan blade having the largest average camber angle. Such a configuration may be referred to as reduced camber. Advantageously, reducing the camber of the hub portion of the blade may reduce the susceptibility to flutter of the main fan blade.

[0030] The leading edge of the auxiliary fan blade may abut the trailing edge of the main fan blade. Advantageously, this may reduce turbulence at the transition between the main fan blade and the auxiliary fan blade. Alternatively, the leading edge of the auxiliary fan blade may be spaced apart from the trailing edge of the main fan blade, i.e. to provide a smaller gap between the auxiliary fan blade and the main fan blade. Advantageously, this may reduce the likelihood of interference between the main fan blade and the auxiliary fan blade when the auxiliary fan blade changes pitch angle. The gap may provide a gap between the leading edge of the auxiliary fan blade and the trailing edge of the main fan blade, the width of the gap measured in the axial direction of the engine being no greater than 10% of the maximum axial spacing between the leading edge and the trailing edge of the main fan blade, and preferably no greater than 5%. In this way, the airflow leaving the trailing edge of the main fan blade can flow effectively and uninterruptedly to the leading edge of the auxiliary fan blade, i.e. static components (such as stator blades) are generally not required in the gap between the main fan blade and the auxiliary fan blade.

[0031] 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, the fan being connected to the LPT via a shaft, or connected to the LPT / IPT via a transmission mechanism.

[0032] 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).

[0033] 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.

[0034] 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.

[0035] 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.

[0036] 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.

[0037] 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.

[0038] 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.

[0039] Each main 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 main fan blade at the hub to the radius of the main 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 main 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). 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.

[0040] The radius of the fan can be measured between the engine centerline and the tip at the leading edge of the main fan blades. The fan diameter (which may be just twice the fan radius) may be greater than (or approximately) any of the following: 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), or 390 cm (about 155 inches). The fan diameter 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).

[0041] 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 250cm to 300cm (e.g., 250cm to 280cm), 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 320cm 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 1600rpm.

[0042] 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: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all units in this paragraph are Jkg-1 K -1 / (ms -1 ) 2 ). The fan tip load 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).

[0043] 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 duct 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, or 17. 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). The bypass duct may be substantially annular. The bypass duct may be located radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0044] 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).

[0045] 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 bounded by any two values ​​in the preceding sentence (ie, these values ​​may form an upper or lower limit). Such engines may be particularly efficient compared to conventional gas turbine engines.

[0046] 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 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). The thrust mentioned above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), when the engine is stationary.

[0047] 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). Maximum TET may occur, for example, at high thrust conditions, such as at maximum takeoff (MTO) conditions.

[0048] The fan blades (e.g., tandem fan blades) and / or the 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.

[0049] A fan as described and / or claimed herein may include a central portion from which the tandem fan blades may extend, for example, in a radial direction. The tandem fan blades may be attached to the central portion in any desired manner. For example, each tandem fan blade (e.g., a primary fan blade and an auxiliary fan blade) may include a fixing that can engage a corresponding slot in a hub (or disc). By way of example only, such fixings may be in the form of a dovetail that can be inserted and / or engaged with a corresponding slot in the hub / disc so as to fix the fan blade to the hub / disc. In another example, the tandem fan blades may be formed integrally with the central portion. Such an arrangement may be referred to as an integral blade disk or integral blade ring. Any suitable method may be used to manufacture such an integral blade disk or integral 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).

[0050] 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 duct to vary during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0051] A fan of a gas turbine as described and / or claimed herein may have any desired number of tandem fan blades, such as 16, 18, 20, or 22 tandem fan blades, each tandem fan blade comprising a primary fan blade and a secondary fan blade.

[0052] As used herein, cruise conditions may refer to cruise conditions of an aircraft to which a gas turbine engine is attached. Such cruise conditions may generally be defined as conditions in the middle of a cruise, such as conditions experienced by an aircraft and / or engine at a midpoint (in terms of time and / or distance) between the top of a climb and the start of a descent.

[0053] By way of example only, the forward speed at a cruise condition may be anywhere within the range of 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 a cruise condition. For some aircraft, the cruise condition may be outside these ranges, such as below Mach 0.7 or above Mach 0.9.

[0054] By way of example only, cruise conditions may correspond to standard atmospheric conditions 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. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0055] By way of example only, cruise conditions may correspond to: forward Mach number of 0.8; pressure of 23000 Pa; and temperature of -55°C.

[0056] As used anywhere herein, "cruise" or "cruise conditions" may refer to an aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (including, for example, one or more of Mach number, environmental conditions, and thrust requirements) at which a fan is designed to operate. For example, this may refer to the conditions at which a fan (or gas turbine engine) is designed to have optimal efficiency.

[0057] 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.

[0058] Those skilled in the art 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

[0059] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0060] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0061] Figure 2 yes Figure 1 a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0062] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0063] Figure 4 is with Figure 1 and Figure 2 A schematic side elevation view of a tandem fan blade for use with a gas turbine engine;

[0064] Figure 5A yes Figure 4 Series fan blades online P A -P A A schematic chord cross section in a first configuration;

[0065] Figure 5B yes Figure 4 Series fan blades online P B -P B Schematic chord cross section on ;

[0066] Figure 6 yes Figure 4 Series fan blades online P A -P A a schematic chord cross section in a second configuration; and

[0067] Figure 7 is a schematic side elevation view of a variant tandem fan blade. DETAILED DESCRIPTION

[0068] 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.

[0069] Figure 1A gas turbine engine 10 having a main axis of rotation 9 is shown. The gas turbine engine 10 comprises an air inlet 12 and a pusher fan having a circumferential row of tandem fan blades 23, which produces two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 comprises a core 11 receiving the core air flow A. The engine core 11 comprises a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20 in axial series. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 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 air flow A is accelerated and compressed by the low-pressure compressor 14 and is directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion device 16, in which 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 series fan blades 23 generally provide 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 ) drives the shaft 26, which is coupled to a sun wheel or sun gear 28 of an epicyclic gearbox 30. Radially outward of and intermeshing with the sun gear 28 are 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 fan via a connecting rod 36 so as to drive the fan in rotation about the engine axis 9. Radially outward of and intermeshing with the planet gears 32 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" as used herein may refer to the lowest pressure turbine stage and the lowest pressure compressor stage, respectively (i.e., excluding the fan), and / or the turbine stage and the compressor stage connected together by an interconnecting shaft having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the fan). In some documents, the "low-pressure turbine" and "low-pressure compressor" mentioned herein may be alternatively referred to as "intermediate-pressure turbine" and "intermediate-pressure compressor". In the case of using such alternative nomenclature, the fan may be referred to as the first compression stage or the 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 less 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 3 The 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 connecting rods 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 fan is driven by the ring gear 38. By way of another alternative example, the epicyclic 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 epicyclic gearbox 30 in the gas turbine engine 10 and / or for connecting the epicyclic gearbox 30 to the gas turbine engine 10. By way of another example, the connections (such as the couplings) between the epicyclic gearbox 30 and other components of the gas turbine engine 10 (such as the input shaft, the output shaft, and the fixed support structure 24) may be made. Figure 2The 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 epicyclic 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 will typically be 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 1 The gas turbine engine shown in has a split nozzle, which means that the flow passing through the bypass duct 22 has its own nozzle, which is separated from the core exhaust nozzle 20 and radially 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 duct 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 an epicyclic 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 axis of rotation 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] like Figure 1 and Figure 2As shown, the fan includes a plurality of tandem fan blades 23, which are attached to the fan rotor disk of the fan hub 50 and are arranged in a circumferential row around the fan hub 50. Each of the tandem fan blades includes a main fan blade 23a and an auxiliary fan blade 23b, which is positioned behind the main fan blade 23a, so that on a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade 23b is located on the rearward extension of the arc line of the airfoil chord section of the main fan blade 23a. In addition, the main fan blade 23a and the auxiliary fan blade 23b in each of the tandem fan blades 23 are arranged to rotate one after the other, so that they maintain their angular positions relative to each other.

[0081] Each auxiliary fan blade 23b is pivotable about its leading edge so that it can move within a range of pitch angles relative to the main fan blade 23a. The pivotal movement of the auxiliary fan blades is actuated by a gear-based variable pitch mechanism 52 provided in the fan hub 50.

[0082] Figure 4 A schematic side elevation view of the tandem fan blades 23 relative to the main rotation axis 9 is shown. In this view, the main fan blades 23a project a larger area than the auxiliary fan blades 23b. The main fan blades 23a also have a larger span in the radial direction R of the engine than the auxiliary fan blades 23b. Specifically, the main fan blades 23a extend from the hub 62 to the tip 64 adjacent to the inner surface of the nacelle 21, while the auxiliary fan blades 23b extend from the hub 72 to the tip 74 at a radial position or radially inside the leading edge of the annular splitter 51 located behind the fan, which divides the airflow generated by the fan into a core airflow A and a bypass airflow B. Therefore, the auxiliary fan blades 23b have little or no effect on the bypass airflow B.

[0083] The primary and secondary fan blades 23a, 23b have respective leading edges 66, 76 and trailing edges 68, 78. The primary fan blade 23a has a root portion, such as a dovetail (not shown), which engages a complementary formation formed in the fan rotor disk to removably mount the blade to the disk.

[0084] As described above, the auxiliary fan blade 23b can be pivoted about its leading edge 76 by the variable pitch mechanism 52. More specifically, the auxiliary fan blade 23b can be pivoted about the rotation axis 56 (eg, Figure 4The auxiliary fan blade 23b pivots as shown by the dashed line in FIG. 2 . There is a small gap between the trailing edge 68 of the main fan blade 23a and the leading edge 76 of the auxiliary fan blade 23b. This ensures that the leading edge 76 of the auxiliary fan blade 23b does not interfere with the trailing edge 68 of the main fan blade 23a when the auxiliary fan blade pivots. The gap may provide a clearance between the leading edge of the auxiliary fan blade and the trailing edge of the main fan blade, the width of which measured in the axial direction of the engine is no greater than 10%, and preferably no greater than 5%, of the maximum axial spacing between the leading and trailing edges of the main fan blades.

[0085] Figure 4 Series fan blades 23 online P A -P A The schematic airfoil chord section on Figure 5A Line P A -P A Within the first 10% of the span of the main fan blade 23a, as measured from its hub 62. In contrast, Figure 4 Main fan blades 23a online P B -P B The schematic airfoil chord section on Figure 5B Line P B -P B Within the first 40% of the span of the primary fan blade 23a, as measured from its hub 62, but above the tip 74 of the secondary fan blade 23b, thus within Figure 5B The airfoil chord section of the auxiliary fan blade is not shown.

[0086] Series fan blades 23 online P A -P A The combined arc 80 on Figure 5A and only the main fan blade 23a is on line P B -P B The arc 82 on Figure 5B The combined arc 80 is the main fan blade 23a and the auxiliary fan blade 23b on the line P A -P A More specifically, Figure 5A In the embodiment, the camber of the airfoil chord section of the auxiliary fan blade 23b is the rearward continuation of the camber of the airfoil chord section of the main fan blade 23a. Therefore, the airflow over the main fan blade 23a continues with almost no disturbance over the auxiliary fan blade 23b. Figure 5A A combined camber of the type shown (i.e. one in which the leading edge of the airfoil chord section of the auxiliary fan blade 23b is located on the rearward extension of the camber of the airfoil chord section of the main fan blade 23a) is provided over substantially the entire radial span of the auxiliary fan blade 23b.

[0087] exist Figure 5A, the angle between the direction of the combined camber line 80 at the leading edge of the main fan blade airfoil chord section and the axial direction is α1, and the angle between the direction of the combined camber line at the trailing edge of the auxiliary fan blade airfoil chord section and the axial direction is α2. Therefore, the total camber angle of the airfoil chord section is (α1–α2).

[0088] exist Figure 5B , the angle between the direction of the camber line 82 at the leading edge of the main fan blade airfoil chord section and the axial direction is β1, and the angle between the direction of the camber line at the trailing edge of the main fan blade airfoil chord section and the axial direction is β2. Therefore, the camber angle is (β1-β2).

[0089] like Figure 5A and Figure 5B The main fan blade 23a shown is a reduced camber fan blade. For example, the average camber angle of the radially innermost 10% of the radial span of the main fan blade is less than 75% of the average camber angle of the 10% portion of the radial span of the main fan blade with the largest average camber angle. Therefore, the fan hub pressure ratio at the bottom of the main fan blade can be less than 1.2.

[0090] During aircraft takeoff, the gas turbine engine provides the greatest amount of thrust during flight. This is an operating condition that is generally limited by the T30 delivery temperature of the high pressure compressor 15 to the combustion device 16. Therefore, it is desirable to limit the fan hub pressure ratio in order to reduce this effect on the T30 temperature. For example, Figure 5A As shown, the auxiliary fan blades 23b can be pivoted to their lowest pitch angle, so that the total camber angle (α1-α2) is low. Therefore, the auxiliary fan blades 23b do not significantly increase the fan hub pressure ratio.

[0091] However, during, for example, cruise operating conditions (which are not limited by the T30 delivery temperature), the auxiliary fan blades 23b may be pivoted to a pitch angle such that the total camber angle (α1 - α2) is higher, such as Figure 6 As shown, the figure is Figure 5A The same schematic airfoil chord section, but in which the auxiliary fan blade 23b is pivoted about its leading edge 76. In this way, the auxiliary fan blade 23b supplements the flow steering capability at the bottom of the main fan blade 23a, thereby increasing the fan hub pressure ratio from 1.2 to 1.3 or greater. This increases the T30 temperature and the total pressure ratio, and accordingly improves the efficiency of the gas turbine engine. The auxiliary fan blade 23b can be configured to contribute at least 20% of the total work done by the series fan blade 23 under the cruise operating conditions of the engine. Calculations show that the use of auxiliary fan blades during cruise can increase the T30 temperature by about 20°C and significantly improve the specific fuel consumption (SFC).

[0092] Typically, the auxiliary fan blades 23b can pivot through a pitch angle range of about 15°, with a pitch angle of 0° corresponding to the lowest total camber angle (e.g., about 10° of (α1-α2)) and a pitch angle of 15° corresponding to the highest total camber angle (e.g., about 25° of (α1-α2)). At a pitch angle of 15°, the auxiliary fan blades 23b essentially act as a supercharger for the low-pressure compressor 14. Moving the auxiliary fan blades within the total pitch angle range can vary the work done by the fan by about 20%.

[0093] In a variant tandem fan blade (not shown), the pitch angle of the auxiliary fan blades is fixed, e.g., non-adjustable. In this variant tandem fan blade, the pitch angle of the auxiliary fan blades can be set to provide an overall camber angle that balances the need to increase the fan base pressure ratio under cruise operating conditions and appropriately limit T30 temperatures at takeoff. Since the variant fan blade does not require a variable pitch mechanism, it can help reduce the overall weight of the fan.

[0094] Figure 7 A schematic side elevation view of a variant tandem fan blade 123 including a main fan blade 123a and an auxiliary fan blade 123b is shown. In this view, the trailing edge of the main fan blade 123a has the appearance of a cutout area at the rear of the base of the main fan blade 123a, which surrounds the auxiliary fan blade. For example, the size of the cutout area is set so that the trailing edge 168 of the main fan blade radially outside the auxiliary blade forms an extension of the line of the trailing edge 178 of the auxiliary fan blade. This can help reduce undesirable flow disturbances at the tip 174 of the auxiliary blade. In this side elevation view, the variant tandem fan blade 123 exhibits a profile similar to a single fan blade of reduced camber.

[0095] 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 for an aircraft, the gas turbine engine comprising: an engine core, the engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; as well as a fan located upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades; wherein each of the tandem fan blades comprises a main fan blade and an auxiliary fan blade, the auxiliary fan blade being positioned at the rear of the main fan blade such that over substantially the entire radial span of the auxiliary fan blade, the leading edge of the auxiliary fan blade is rearward of the nearest point on the trailing edge of the main fan blade, and over a given airfoil chord section of the main fan blade, the leading edge position of the airfoil chord section of the auxiliary fan blade is located on a rearward extension of the camber line of the airfoil chord section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate one behind the other; and The auxiliary fan blade is movable within a pitch angle range relative to the main fan blade by pivoting about the leading edge of the auxiliary fan blade.

2. A gas turbine engine according to claim 1, wherein the auxiliary fan blade is positioned so that over substantially the entire radial span thereof, the leading edge of the corresponding airfoil chord section of the auxiliary fan blade is located on the rearward extension of the camber line of the corresponding airfoil chord section of the main fan blade.

3. The gas turbine engine of claim 1 , wherein the auxiliary fan blade is positioned such that the leading edge of the auxiliary fan blade is rearward of a nearest portion of the trailing edge of the main fan blade over substantially its entire radial span.

4. The gas turbine engine according to claim 1, further comprising an annular splitter positioned behind the fan and configured to split the airflow generated by the fan into a core airflow flowing through the engine core and a bypass airflow flowing through a bypass duct surrounding the engine core, wherein the auxiliary fan blades have radially outer tips that are inward of or flush with a radial position of the annular splitter.

5. The gas turbine engine of claim 1, wherein in side elevation view, the projected area of ​​the auxiliary fan blade of each tandem fan blade is smaller than the projected area of ​​the corresponding main fan blade.

6. A gas turbine engine according to claim 1, wherein the main fan blade and the auxiliary fan blade are configured so that in a side elevation view, the trailing edge of the main fan blade has the appearance of a cutout area at the rear of the base of the main fan blade, and the cutout area surrounds the auxiliary fan blade.

7. A gas turbine engine according to claim 1, wherein the auxiliary fan blades are capable of moving within a pitch angle range of at least 10°, whereby increasing the pitch angle increases the total camber angle of the given airfoil chord section of the main fan blades and the auxiliary fan blades by a corresponding amount.

8. The gas turbine engine of claim 1, wherein the fan is configured such that moving the auxiliary fan blades through a total pitch angle range varies the total work performed by the fan by at least 20%.

9. The gas turbine engine of claim 1, wherein the auxiliary fan blades are configured to contribute at least 20% of the total work done by the tandem fan blades under cruise operating conditions of the engine.

10. A gas turbine engine according to claim 1, wherein the main fan blade has a radial span extending from a hub at a 0% span position to a tip at a 100% span position, and wherein the average camber angle of the radially innermost 10% of the radial span of the main fan blade is less than 75% of the average camber angle of the 10% portion of the radial span of the main fan blade having the largest average camber angle.

11. The gas turbine engine of claim 1 , further comprising: A gearbox receives input from the spindle and outputs drive to the fan to drive the fan at a slower rotational speed than the spindle.

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