Blade assembly for a gas turbine engine

Abstract

A vane assembly for a gas turbine engine includes a rotor, a stator, a seal plate, and a seal member. The rotor includes a rotor vane and a rotor disk. The rotor disk defines a vane slot that receives cooling fluid from a first cavity upstream of the rotor. The seal member includes a control arm. The seal member and the rotor define a flow cavity therebetween that is in fluid communication with an aperture of the seal plate. The flow cavity receives the cooling fluid that flows through the vane slot and the aperture. The control arm and the seal plate define a gap therebetween that fluidly communicates the flow cavity with a second cavity between the stator and the rotor. The control arm deflects at least a portion of the cooling fluid entering the flow cavity.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to vane assemblies, and in particular to vane assemblies for gas turbine engines. BACKGROUND

[0002] Turbines of gas turbine engines generally operate at very high temperatures. Thus, it is critical that the components are sufficiently cooled. Turbines include complex cooling arrangements to ensure that the components are sufficiently cooled, but this requires additional cooling air that detracts from the efficiency of the engine. Thus, it is desirable to use cooling air in an efficient manner.

[0003] Stator well cavities are generally cooled by cooling air introduced through discrete cooling holes. Existing cooling methods generally result in non-uniform mixing of the cooling air with ingested gas, which is derived from the main flow of the gas turbine engine. Such non-uniform mixing in the stator well cavities results in undesirable temperature gradients and high temperatures of the components. Further, existing methods can also result in undesirable temperature gradients between the two cavities, forward and aft of the stator seal. SUMMARY

[0004] According to a first aspect, a vane assembly for a gas turbine engine is provided herein. The vane assembly includes a rotor, a stator, a seal plate, and a seal member. The rotor includes a rotor vane and a rotor disk defining a bucket groove that at least partially receives the rotor vane therein. The bucket groove receives cooling fluid from a first cavity upstream of the rotor. The stator is disposed downstream of the rotor. The stator and the rotor define a second cavity therebetween downstream of the first cavity. The seal plate is coupled to the rotor and faces the first cavity. The seal plate defines an aperture therethrough that is in fluid communication with the bucket groove. The seal member includes a control arm extending at least radially toward the rotor. The seal member and the rotor define a flow cavity therebetween that is in fluid communication with the aperture of the seal plate. The control arm and the seal plate define a gap therebetween that fluidly communicates the flow cavity with the second cavity. The flow cavity receives cooling fluid that flows through the bucket groove of the rotor disk and the aperture of the seal plate. The control arm deflects at least a portion of the cooling fluid entering the flow cavity. The gap allows at least a portion of the cooling fluid to exit the flow cavity and enter the second cavity.

[0005] In some embodiments, the control arm is obliquely angled relative to a main rotational axis of the rotor.

[0006] In some embodiments, the seal member further includes a main portion extending from the rotor disk. The control arm extends from the main portion.

[0007] In some embodiments, the primary portion is obliquely angled relative to a primary rotational axis of the rotor.

[0008] In some embodiments, the primary portion is substantially parallel to a primary rotational axis of the rotor.

[0009] In some embodiments, an angle between the primary portion of the seal member and the primary rotational axis of the rotor is from about -60 degrees to about 60 degrees.

[0010] In some embodiments, the control arm further includes a tip distal from the primary portion of the seal member. In some embodiments, a gap is defined between the tip of the control arm and the seal plate.

[0011] In some embodiments, the control arm extends radially outwardly and axially relative to the primary rotational axis of the rotor from the primary portion toward the seal plate.

[0012] In some embodiments, the rotor further includes a blade platform disposed radially outwardly of the blade slot. In some embodiments, the seal plate extends radially from the rotor disk to the blade platform.

[0013] In some embodiments, the control arm extends at least radially toward the blade platform.

[0014] In some embodiments, the seal plate is coupled to the blade platform and the rotor disk.

[0015] In some embodiments, the stator further includes a stator seal.

[0016] In some embodiments, the seal member further includes one or more fins extending toward and cooperating with the stator seal.

[0017] In some embodiments, the seal member is integral with the rotor disk.

[0018] In some embodiments, the rotor further includes a drive arm extending from the rotor disk. The seal member is separate from the drive arm.

[0019] According to a second aspect, there is provided a turbine for a gas turbine engine, the turbine comprising the blade assembly of the first aspect.

[0020] According to a third aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising the turbine of the second aspect.

[0021] The control arm can avoid a circumferential temperature gradient across the stator well cavity (i.e., the second cavity). Further, the control arm can ensure that the surfaces of the rotor disk, the seal plate, and the rotor blade are always wetted by the cooling fluid. Depending on the length of the control arm and the width of the gap between the seal plate and the control arm, the control arm can allow mixing of the cooling fluid and the core airflow at a relatively higher radial position. This can result in a reduction of the temperature of the second cavity. Further, a high temperature gradient between the two cavities in front of and behind the stator seal can also be avoided. Such a high temperature gradient can otherwise adversely affect the performance of the stator seal.

[0022] The control arm can allow for uniform mixing between the cooling fluid and the acquisition flow from the core airflow of the gas turbine engine. Improving the mixing between the cooling fluid and the acquisition flow can reduce the temperature of various portions of the rotor. This can reduce the amount of cooling fluid needed to meet suitable temperatures. In other words, a lower temperature can be achieved with the same amount of flow of cooling fluid. As a result, this can improve the specific fuel consumption (SFC) of the gas turbine engine and / or allow for the use of lower cost materials.

[0023] Except insofar as the context of the description requires otherwise, any feature of the disclosure can be adopted or combined with any other feature of the disclosure, and the disclosure extends to all combinations and sub-combinations of one or more features described herein.

[0024] As noted elsewhere herein, the disclosure can relate to a gas turbine engine. Such a gas turbine engine can include an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine can include a fan positioned upstream of the engine core (having fan blades).

[0025] The arrangements of the disclosure can be particularly, but not exclusively, beneficial for a fan that is driven via a gearbox. Thus, the gas turbine engine can include a gearbox that receives an input from the core shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox can be directly from the core shaft, or indirectly from the core shaft, e.g., via a spur shaft and / or a gear. The core shaft can rigidly connect the turbine and the compressor, such that the turbine and the compressor rotate at the same speed (whereas the fan rotates at a lower speed).

[0026] A gas turbine engine as described and / or claimed herein can have any suitable overall configuration. For example, the gas turbine engine can have any desired number of shafts connecting the turbine and compressor, such as one, two or three shafts. By way of example only, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor, and the core shaft can be a first core shaft. The engine core can further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft can be arranged to rotate at a higher rotational speed than the first core shaft.

[0027] In such an arrangement, the second compressor can be axially positioned downstream of the first compressor. The second compressor can be arranged to receive (e.g. directly receive, such as via a generally annular passage) the flow from the first compressor.

[0028] The gearbox can be arranged to be driven by the core shaft (e.g. the first core shaft in the above example) configured (e.g. in use) to rotate at the lowest rotational speed. For example, the gearbox can be arranged to be driven by only the core shaft (e.g. only the first core shaft in the above example, and not the second core shaft) configured (e.g. in use) to rotate at the lowest rotational speed. Alternatively, the gearbox can be arranged to be driven by any one or more shafts (e.g. the first and / or second shaft in the above example).

[0029] The gearbox can be a reduction gearbox (where the output to the fan is a lower rotational speed than the input from the core shaft). Any type of gearbox can be used. For example, the gearbox can be a “planetary” or “star” 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), such as greater than 2.5, such as in the range from 3 to 4.2 or 3.2 to 3.8, such as approximately 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. The gear ratio can be, for example, between any two of the values in the preceding sentence. By way of example only, the gearbox can be a “star” gearbox having a ratio in the range from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0030] In any gas turbine engine as described and / or claimed herein, the combustion chamber can be provided axially downstream of the fan and the compressor(s). For example, in the case where a second compressor is provided, the combustion chamber can be directly downstream of the second compressor (e.g. at the outlet of the second compressor). By way of further example, in the case where a second turbine is provided, the flow at the outlet of the combustion chamber can be provided to the inlet of the second turbine. The combustion chamber can be provided upstream of the turbine(s).

[0031] The compressor or each compressor (e.g. the first and second compressors as described above) can comprise any number of stages, e.g. a plurality of stages. Each stage can comprise a row of rotor blades and a row of stator vanes, which can be variable stator vanes (where their angle of incidence can be variable). The row of rotor blades and the row of stator vanes can be axially offset from one another.

[0032] The turbine or each turbine (e.g. the first and second turbines as described above) can comprise any number of stages, e.g. a plurality of stages. Each stage can comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes can be axially offset from one another.

[0033] Each fan blade can be defined as having a radial span extending from a root (or hub) at a radially inner gas wash location or 0% span location to a tip at a 100% span location. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be less than (or about) any of: 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 can be within a closed- ended range bounded by any two of the values in the preceding sentence (i.e. the values can form an upper or lower boundary), e.g. within a range of from 0.28 to 0.32. These ratios can be referred to generally as hub to tip ratios. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward-most) portion of the blade. Of course the hub to tip ratio relates to the gas washing portion of the fan blade, i.e. the portion radially outside of any platforms.

[0034] The radius of the fan can be measured between the engine centerline and the tip of the fan blade at its leading edge. The fan diameter, which can simply be twice the radius of the fan, can be greater than (or about) any of: 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 can be in a range bounded by (i.e., the values can form upper or lower bounds) any two of the values in the preceding sentence, e.g., in a range from 240 cm to 280 cm or 330 cm to 380 cm.

[0035] The rotational speed of the fan can vary in use. Generally, the rotational speed is lower for fans having a higher diameter. By way of non-limiting example only, the rotational speed of the fan can be less than 2500 rpm, e.g., less than 2300 rpm, under cruise conditions. By way of further non-limiting example only, for an engine having a fan diameter in a range from 220 cm to 300 cm, e.g., 240 cm to 280 cm or 250 cm to 270 cm, the rotational speed of the fan can be in a range from 1700 rpm to 2500 rpm, e.g., in a range from 1800 rpm to 2300 rpm, e.g., in a range from 1900 rpm to 2100 rpm, under cruise conditions. By way of further non-limiting example only, for an engine having a fan diameter in a range from 330 cm to 380 cm, the rotational speed of the fan can be in a range from 1200 rpm to 2000 rpm, e.g., in a range from 1300 rpm to 1800 rpm, e.g., in a range from 1400 rpm to 1800 rpm, under cruise conditions.

[0036] In use of a gas turbine engine, 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 velocity Utip. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. The fan tip loading can be defined as dH / Utip. 尖端 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., the (translational) velocity of the fan tip at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge times the angular velocity). The fan tip load can be greater than (or about) any of: 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 J kg -1 K -1 / (ms -1 ) 2 ). The fan tip load can be in a range bounded by inclusive endpoints of any two of the values in the preceding sentence (i.e., the values can form upper or lower bounds), e.g., in a range from 0.28 to 0.31 or 0.29 to 0.3.

[0037] A gas turbine engine according to the present disclosure can have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass passage to the mass flow rate of the flow through the core at cruise conditions. In some arrangements, the bypass ratio can be greater than (or about) any of: 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 can be in a range bounded by inclusive endpoints of any two of the values in the preceding sentence (i.e., the values can form upper or lower bounds), e.g., in a range from 12 to 16, 13 to 15, or 13 to 14. The bypass passage can be substantially annular. The bypass passage can be radially outward of the engine core. A radially outer surface of the bypass passage can be defined by the nacelle and / or the fan casing.

[0038] The overall pressure ratio of a gas turbine engine as described and / or claimed herein can be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entering into the combustion chamber). By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein can be greater than (or about) any of: 35, 40, 45, 50, 55, 60, 65, 70, 75 at cruise. The overall pressure ratio can be in a range bounded by inclusive endpoints of any two of the values in the preceding sentence (i.e., the values can form upper or lower bounds), e.g., in a range from 50 to 70.

[0039] The specific thrust of an engine can 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 engines described and / or claimed herein can be less than (or on the order of) any of: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s, or 80 Nkg -1 s. The specific thrust can be in a range bounded by inclusive endpoints defined by any two of the values in the preceding sentence (i.e., the values can form upper or lower bounds), such as in a range from 80 Nkg -1 s to 100 Nkg -1 s, or 85 Nkg -1 s to 95 Nkg -1 s. Such engines can be particularly efficient compared to conventional gas turbine engines.

[0040] Gas turbine engines as described and / or claimed herein can have any desired maximum thrust. By way of non-limiting example only, gas turbines as described and / or claimed herein can have the ability to produce a maximum thrust of at least (or on the order of) any of: 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 can be in a range bounded by inclusive endpoints defined by any two of the values in the preceding sentence (i.e., the values can form upper or lower bounds). By way of example only, gas turbines as described and / or claimed herein can have the ability to produce a maximum thrust in a range from 330 kN to 420 kN (e.g., 350 kN to 400 kN). The thrusts mentioned above can be maximum net thrust at sea level plus 15 degrees Celsius (ambient pressure 101.3 kPa, temperature 30 degrees Celsius) under standard atmospheric conditions, with the engine static.

[0041] In use, the temperature of the flow at the inlet to the high pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine vanes (which can themselves be referred to as nozzle guide vanes). At cruise, the TET can be at least (or approximately) any of: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. At cruise, the TET can be within an inclusive range bounded by any two of the values in the preceding sentence (i.e. the values can form an upper or lower boundary). The maximum TET in use of the engine can be, for example, at least (or approximately) any of: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. The maximum TET can be within an inclusive range bounded by any two of the values in the preceding sentence (i.e. the values can form an upper or lower boundary), for example within a range from 1800 K to 1950 K. The maximum TET can occur, for example, under high thrust conditions, for example under maximum take-off (MTO) conditions.

[0042] The fan blade and / or the airfoil portion of the fan blade described and / or claimed herein can be manufactured from any suitable material or combination of materials. For example, at least portions of the fan blade and / or the airfoil can be at least partially manufactured from a composite, for example a metal matrix composite and / or an organic matrix composite, such as carbon fibre. By way of further example, at least portions of the fan blade and / or the airfoil can be at least partially manufactured from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium lithium alloy) or a steel based material. The fan blade can comprise at least two regions manufactured using different materials. For example, the fan blade can have a guarded leading edge, which can be manufactured using a material that is better able to resist impacts (for example from birds, ice or other materials) than the remainder of the blade. Such a leading edge can be manufactured, for example, using titanium or a titanium based alloy. Thus, by way of example only, the fan blade can have a carbon fibre or aluminium based body (such as an aluminium lithium alloy) with a leading edge of titanium.

[0043] A fan as described and / or claimed herein can include a central portion from which fan blades can extend, e.g. in a radial direction. The fan blades can be attached to the central portion in any desired manner. For example, each fan blade can include a fixing which can engage a corresponding slot in a hub (or disc). By way of example only, such a fixing can be in the form of a dovetail which can slot into and / or engage a corresponding slot in the hub / disc in order to secure the fan blade to the hub / disc. By way of further example, the fan blades can be integrally formed with the central portion. Such an arrangement can be referred to as a bladed disc or bladed ring. Any suitable method can be used to manufacture such a bladed disc or bladed ring. For example, at least part of the fan blades can be machined from a block and / or at least part of the fan blades can be attached to the hub / disc by welding, such as linear friction welding.

[0044] A gas turbine engine as described and / or claimed herein can or can not be provided with a variable area nozzle (VAN). Such a variable area nozzle can allow the exit area of the bypass passage to vary in use. The general principles of the present disclosure can apply to engines with or without a VAN.

[0045] A fan of a gas turbine as described and / or claimed herein can have any desired number of fan blades, e.g. 14, 16, 18, 20, 22, 24 or 26 fan blades.

[0046] As used herein, cruise conditions have the usual 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 appreciate that cruise conditions refer to the operating point of the engine at the mid-cruise of a given mission of the aircraft to which the gas turbine engine is designed to be attached (which can be referred to in the industry as an “economic mission”). In this regard, the mid-cruise is the point in the flight cycle of the aircraft at which 50% of the total fuel burned between the top of climb and the start of descent has been burned (which can approximate to the midpoint in terms of time and / or distance between the top of climb and the start of descent. Cruise conditions thus define an operating point of the gas turbine engine for which, taking into account the number of engines provided to the aircraft, the thrust provided will guarantee steady state operation at the mid-cruise of the aircraft to which the gas turbine engine is designed to be attached (i.e. to maintain a constant altitude and constant Mach number). For example, in the case where the engine is designed to be attached to an aircraft having two engines of the same type, the engine at cruise conditions provides half of the total thrust required for steady state operation of the aircraft at the mid-cruise.

[0047] In other words, for a given gas turbine engine for an aircraft, the cruise condition is defined as the operating point of the engine that provides (in combination with any other engines on the aircraft to provide what is needed for the aircraft to which it is designed to be attached to operate steadily at a given Mach number of the cruise segment) unit thrust at cruise mid-segment atmospheric conditions (defined by the International Standard Atmosphere at the cruise mid-segment altitude according to ISO 2533). For any given gas turbine engine for an aircraft, the cruise mid-segment thrust, atmospheric conditions, and Mach number are known, and thus the operating point of the engine at the cruise condition is clearly defined.

[0048] By way of example only, the forward speed at the cruise condition can be at any point in the range from Mach 0.7 to 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 approximately Mach 0.8, approximately Mach 0.85, or in the range from 0.8 to 0.85. Any individual speed in these ranges can be part of the cruise condition. For some aircraft, the cruise condition can be outside these ranges, such as below Mach 0.7 or above Mach 0.9.

[0049] By way of example only, the cruise condition can correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude in the range from 10,000 m to 15,000 m, such as in the range from 10,000 m to 12,000 m, such as in the range from 10,400 m to 11,600 m (approximately 38,000 ft), such as in the range from 10,500 m to 11,500 m, such as in the range from 10,600 m to 11,400 m, such as in the range from 10,700 m (approximately 35,000 ft) to 11,300 m, such as in the range from 10,800 m to 11,200 m, such as in the range from 10,900 m to 11,100 m, such as approximately 11,000 m. The cruise condition can correspond to standard atmospheric conditions at any given altitude in these ranges.

[0050] By way of example only, the cruise conditions can correspond to an operating point of the engine in which a known required thrust level (e.g. a value in the range from 30 kN to 35 kN) is provided at a forward Mach number of 0.8 and at standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m). By way of further example only, the cruise conditions can correspond to an operating point of the engine in which a known required thrust level (e.g. a value in the range from 50 kN to 65 kN) is provided at a forward Mach number of 0.85 and at standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m).

[0051] In use, the gas turbine engine described and / or claimed herein can be operated at cruise conditions as defined elsewhere herein. Such cruise conditions can be determined by cruise conditions (e.g. cruise mid-section conditions) of an aircraft to which at least one (e.g. 2 or 4) gas turbine engine can be mounted so as to provide propulsive thrust.

[0052] According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is an aircraft to which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the cruise mid-section of the aircraft as defined elsewhere herein.

[0053] According to an aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation can be at cruise conditions (e.g. in terms of thrust, atmospheric conditions and Mach number) as defined elsewhere herein.

[0054] According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. The operation according to this aspect can comprise (or can be) operation at the cruise mid-section of the aircraft as defined elsewhere herein.

[0055] The skilled person will appreciate that features or parameters described in relation to any one of the above aspects can be applied to any other aspect, except where mutually exclusive, and further that any feature or parameter described herein can be applied to any aspect and / or in combination with any other feature or parameter described herein. BRIEF DESCRIPTION OF DRAWINGS

[0056] Embodiments will now be described, by way of example only, with reference to the accompanying drawings in which:

[0057] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0058] Figure 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0059] Figure 3 is a partial cutaway view of a gear box for a gas turbine engine;

[0060] Figure 4A is a schematic cross-sectional side view of a portion of a vane assembly for a gas turbine engine;

[0061] Figure 4B is an enlarged cross-sectional side view of the vane assembly of Figure 4A

[0062] Figure 5A is a schematic cross-sectional side view of a portion of another vane assembly for a gas turbine engine; and

[0063] Figure 5B is an enlarged cross-sectional side view of the vane assembly of Figure 5A DETAILED DESCRIPTION

[0064] Aspects and embodiments of the present disclosure will now be discussed with reference to the drawings. Further aspects and embodiments will be apparent to those of ordinary skill in the art.

[0065] As used herein, if a member extends along an axis, the member extends "axially" relative to the axis. If a member extends in a circumferential direction about the axis, the member extends "circumferentially" relative to the axis. If a member extends radially inward or outward relative to the axis, the member extends "radially" relative to the axis. If a first member is disposed "radially outward" of a second member, the first member is disposed at a greater radial distance from the axis than the second member. If a first member is disposed "radially inward" of a second member, the first member is disposed at a lesser radial distance from the axis than the second member.

[0066] Figure 1 A gas turbine engine 10 having a main rotational axis 9 is illustrated in FIG. 1. The engine 10 includes an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial flow series, 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. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass passage 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass passage 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.​​

[0067] In use, core airflow A is accelerated and compressed by low pressure compressor 14 and is directed into high pressure compressor 15 where further compression takes place. Compressed air discharged from high pressure compressor 15 is directed into combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resulting hot combustion products then expand through high and low pressure turbines 17, 19 before being discharged through core discharge nozzle 20 and in the process drive high and low pressure turbines 17, 19 to provide some of the propulsive thrust. High pressure turbine 17 drives high pressure compressor 15 through a suitable interconnecting shaft 27. Fan 23 provides substantially all of the propulsive thrust. Planetary gearbox 30 is a reduction gearbox.

[0068] In Figure 2 an arrangement for an example of a geared fan gas turbine engine 10 is shown. Low pressure turbine 19 (see Figure 1 ) drives a shaft 26 which is coupled to a sun or sun gear 28 of planetary gear arrangement 30. Radially outward of, and intermeshing with, sun gear 28 are a plurality of planet gears 32 which are coupled together by a carrier 34. Carrier 34 constrains planet gears 32 to travel synchronously around sun gear 28 whilst simultaneously allowing each planet gear 32 to rotate about its own axis. Carrier 34 is coupled via a link 36 to fan 23 so as to drive it to rotate about engine axis 9. Radially outward of, and intermeshing with, planet gears 32 is an annulus or ring gear 38 which is coupled via a link 40 to fixed support structure 24.

[0069] Note that the terms "low pressure turbine" and "low pressure compressor" as used herein can be taken to mean the lowest pressure turbine and compressor stages respectively (i.e. excluding fan 23) and / or the turbine and compressor stages which are connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e. excluding the gearbox output shaft which drives fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein can alternatively be referred to as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, fan 23 can be taken to be the first or lowest pressure compression stage.

[0070] In Figure 3 planetary gearbox 30 is shown in more detail by way of example. Each of sun gear 28, planet gears 32 and ring gear 38 comprise teeth around their outer periphery to intermesh with the other gears. However, for clarity, only an example portion of the teeth is shown in Figure 3The figure is a schematic. Four planetary gears 32 are illustrated in this figure, however it will be apparent to the skilled person that more or fewer planetary gears 32 can be provided within the scope of the claimed invention. Practical applications of planetary gearboxes 30 are generally comprised of at least three planetary gears 32.

[0071] In Figure 2 and Figure 3 The planetary gearbox 30 illustrated by way of example in

[0072] It will be appreciated that the arrangements shown in Figure 2 and Figure 3 are by way of example only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement can be used for positioning the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of further example, the connections between the gearbox 30 and other parts of the engine 10, such as the input shaft 26, the output shaft and the fixed structure 24 (such as the gearbox housing) can have any desired degree of rigidity or flexibility. By way of further example, any suitable arrangement of bearings between the rotating and fixed parts of the engine (for example between the input and output shafts from the gearbox and the fixed structure, such as the gearbox housing) can be used, and the present disclosure is not limited to the arrangements of the examples of Figure 2 For example, in the case where the gearbox 30 has a star arrangement (described above), the skilled person will readily appreciate that the arrangement of the output and support links and bearing positions will generally be different to that illustrated by way of example in Figure 2 Figure 2

[0073] The present disclosure therefore extends to gas turbine engines having gearboxes of any arrangement (for example star or planetary), support structures, input and output shaft arrangements and bearing positions.

[0074] Optionally, the gearbox can drive additional and / or alternative components (for example an intermediate pressure compressor and / or a booster compressor).

[0075] ​​Other gas turbine engines to which the present disclosure can be applied can have alternative configurations. For example, such engines can have alternative numbers of compressors and / or turbines and / or alternative numbers of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 the core discharge nozzle 20 radially outwardly. However, this is not limiting and any aspect of the present disclosure can also be applied to engines in which the flow through the bypass passage 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle, which can be referred to as a mixed flow nozzle. One or both (whether mixed or split flow) nozzles can have fixed or variable area. Although the described examples relate to a turbofan engine, the present disclosure can 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 can not include a gearbox 30.

[0076] The geometry of the gas turbine engine 10, and its components, is defined by a conventional axis system, which includes an axial direction (which is aligned with the main rotational axis 9), a radial direction (which is aligned with the radius of the engine 10), and a circumferential direction (which is perpendicular to the page in Figure 1 the bottom-to-top direction in Figure 1 the bottom-to-top direction in the bottom-to-top direction in

[0077] Furthermore, the present invention is equally applicable to an aeronautical gas turbine engine, a marine gas turbine engine, and a land-based gas turbine engine.

[0078] Each of the high pressure turbine 17 and the low pressure turbine 19 of the gas turbine engine 10 (shown in Figure 1 FIG. 1) can include one or more rows of stator vanes (not shown in Figure 1 FIG. 1) alternating with one or more rows of rotor blades (not shown in Figure 1 FIG. 1). The stator and the rotor immediately downstream of the stator can form a stage. The stator can include an annular array of radially extending stator vanes. The rotor can include an annular array of radially extending rotor blades mounted to a rotor disk. The combusted gases impinge on the stator vanes, which present the gases at an appropriate angle to efficiently drive the rotor blades. The one or more rows of stator vanes and the one or more rows of rotor blades can form a bladed assembly for the gas turbine engine 10.

[0079] Figure 4A and Figure 4B are respectively illustrated for the gas turbine engine 10 (shown in Figure 1FIG. 1 is a perspective view of a blade assembly 100 of a gas turbine engine 10 (shown in phantom) in accordance with an embodiment of the present disclosure.

[0080] Referring to Figure 1 , Figure 4A and Figure 4B , the blade assembly 100 includes a rotor 102 and a stator 104. The rotor 102 includes rotor blades 106 and a rotor disk 108. The rotor blades 106 and the rotor disk 108 can be joined together. The rotor disk 108 defines a blade slot 110 that at least partially receives the rotor blades 106 therein. The blade slot 110 can further define a space or passage between the rotor blades 106 and the rotor disk 108. In some cases, the blade slot 110 can be a dovetail slot. The rotor blades 106 and the rotor disk 108 can form a dovetail joint. The blade slot 110 has a width 112. The width 112 is defined below the rotor blades 106. The width 112 can be defined by a gap between the rotor blades 106 and the blade slot 110. The rotor 102 can have a ring-like array of such rotor blades 106 supported on the rotor disk 108 via the blade slots 110.

[0081] The rotor 102 further includes a blade platform 114 disposed radially outward of the blade slot 110. The blade slot 110 receives the cooling fluid 70 from a first cavity 116 upstream of the rotor 102 as denoted by arrow 74. The cooling fluid 70 can be cooling air bled from a flow from a compressor (e.g., a low pressure compressor 14) of the gas turbine engine 10.

[0082] The stator 104 is disposed downstream of the rotor 102. The stator 104 includes stator blades 107. The rotor 102 and the stator 104 define a second cavity 118 therebetween downstream of the first cavity 116. The stator 104 further includes a stator seal 120. The stator seal 120 seals the second cavity 118 from a third cavity 119. The rotating portion and the static portion of the blade assembly 100 form the stator seal 120 therebetween. In some cases, the stator seal 120 can include a labyrinth seal. The stator seal 120 can be intended to minimize performance losses from cooling fluid 70 leakage across the stator 104 and across the cavities 116, 118, and 119. In some cases, the stator seal 120 can form an inter-stage seal between two adjacent stages of the gas turbine engine 10.

[0083] The vane assembly 100 further includes a seal plate 122 and a seal member 124. In the illustrated embodiment, the seal plate 122 is coupled to the rotor 102 and faces the first cavity 116. In some embodiments, the seal plate 122 is also coupled to the vane platform 114 and the rotor disk 108. Further, the seal plate 122 extends radially from the rotor disk 108 to the vane platform 114. The seal plate 122 can have a substantially annular and flat structure with curved portions at both radial ends. The curved portions can be received in corresponding notches of the vane platform 114 and the rotor disk 108 to attach the seal plate 122 to the vane platform 114 and the rotor disk 108.

[0084] The seal plate 122 defines a hole 126 therethrough that is in fluid communication with the vane slot 110. The hole 126 can be one of an array of circumferentially spaced holes defined through the seal plate 122. The hole 126 can have any suitable shape, such as circular, elliptical, oval, polygonal, etc. The hole 126 has a width 128. In the illustrated embodiment, the width 128 of the hole 126 of the seal plate 122 is greater than or equal to the width 112 of the vane slot 110 of the rotor disk 108. However, in some other cases, the width 128 of the hole 126 of the seal plate 122 can be less than or equal to the width 112 of the vane slot 110 of the rotor disk 108.

[0085] In some embodiments, the number of holes 126 can be different than the number of rotor vanes 106. In other words, one hole 126 can be provided for a plurality of rotor vanes 106. For example, one hole 126 can be provided for a group of three adjacent rotor vanes 106. In some embodiments, the number of holes 126 can depend on the amount of cooling fluid 70 required, manufacturing considerations, and the minimum allowable area available in the vane assembly 100.

[0086] In the illustrated embodiment, the seal member 124 extends from the rotor disk 108. In some embodiments, the seal member 124 can be integral with the rotor disk 108. In some other embodiments, the seal member 124 can be formed separately from the rotor disk 108 and then connected to the rotor disk 108. The seal member 124 includes a main portion 130 that extends from the rotor disk 108. The main portion 130 is oriented substantially perpendicular to the main rotational axis 9 of the rotor 102 (or in the direction of the main rotational axis 9 in the case of a radial seal member 124). The main portion 130 can have a substantially annular and flat structure with curved portions at both radial ends. The curved portions can be received in corresponding notches of the rotor disk 108 and the seal plate 122 to attach the seal member 124 to the rotor disk 108 and the seal plate 122. Figure 4BThe main portion 130 of the seal member 124 is obliquely angled and subtends an angle Al with the main rotational axis 9 (or parallel lines 9' shown in FIG. 1). In the illustrated embodiment, the angle Al is positive. However, in some other cases, the angle Al can be zero or negative, based on the direction of the oblique of the main portion 130 of the seal member 124. In some embodiments, the angle Al between the main portion 130 of the seal member 124 and the main rotational axis 9 of the rotor 102 is from about -60 degrees to about 60 degrees. The main portion 130 thus extends both axially and radially relative to the rotor disk 108. The angle Al can depend on the geometry of the gas turbine engine 10.

[0087] The seal member 124 and the rotor 102 define a flow cavity 134 therebetween that is in fluid communication with the hole 126 of the seal plate 122. The flow cavity 134 receives the cooling fluid 70 flowing through the blade slot 110 of the rotor disk 108 and the hole 126 of the seal plate 122, as indicated by arrow 76.

[0088] The seal member 124 further includes a control arm 132 extending at least radially toward the rotor 102. The control arm 132 is part of the seal member 124 of the blade assembly 100. In some embodiments, the control arm 132 extends at least radially toward the blade platform 114. In other words, the control arm 132 can need to extend as close as possible to the blade platform 114. In some embodiments, the control arm 132 extends from the main portion 130. The control arm 132 can extend as high as possible to a radial position 146, and if desired, the control arm 132 can act as an axial retainer by contacting the rotor blade 106. The high radial position 146 of the control arm 132 can cause a greater surface area of the rotor blade 106 to contact the cooling fluid 70, in turn providing improved mixing of the cooling fluid 70 and resulting in better cooling of the flow cavity 134. The control arm 132 extends relative to the main rotational axis 9 (or parallel lines 9" shown in FIG. 1) of the rotor 102 both radially and axially. Figure 4B The control arm 132 includes an angled surface 133 upon which the cooling fluid 70 impinges. The angled surface 133 can be arcuate. The angled surface 133 can also be obliquely angled relative to the main rotational axis 9 at an angle A2. The impingement of the cooling fluid on the angled surface 133 can cause the cooling fluid to spread apart, particularly in the circumferential direction, thereby evening out the pressure distribution across the second cavity 118. The control arm 132 further includes a tip 131 distal from the main portion 130 of the seal member 124.

[0089] The control arm 132 includes an angled surface 133 upon which the cooling fluid 70 impinges. The angled surface 133 can be arcuate. The angled surface 133 can also be obliquely angled relative to the main rotational axis 9 at an angle A2. The impingement of the cooling fluid on the angled surface 133 can cause the cooling fluid to spread apart, particularly in the circumferential direction, thereby evening out the pressure distribution across the second cavity 118. The control arm 132 further includes a tip 131 distal from the main portion 130 of the seal member 124.

[0090] The control arm 132 deflects at least some of the cooling fluid 70 entering the flow cavity 134. As indicated by arrow 80, the deflected cooling fluid can swirl around the flow cavity 134, thereby reducing the temperature of the flow cavity 134 and then exit the flow cavity 134 to enter the second cavity 118.

[0091] In the illustrated embodiment, the control arm 132 extends radially outwardly and axially from the main portion 130 toward the seal plate 122 relative to the main rotational axis 9 of the rotor 102. The control arm 132 and the seal plate 122 define a gap 136 therebetween that fluidly communicates the flow cavity 134 with the second cavity 118. In some embodiments, the gap 136 is defined between the tip 131 of the control arm 132 and the seal plate 122. As indicated by arrow 82, the gap 136 allows at least some of the cooling fluid 70 to exit the flow cavity 134 and enter the second cavity 118. The gap 136 has a width 138. The width 138 is the distance between the seal plate 122 and the tip 131 of the control arm 132. The control arm 132 can have a converging configuration such that the width 138 is as small as possible. In other words, the control arm 132 can converge at a high radial location 146. In the illustrated embodiment, the control arm 132 can be curved near the tip 131 such that the tip 131 extends substantially parallel to the seal plate 122.

[0092] The cooling fluid 70 flows from the first cavity 116 through the blade slot 110 of each rotor blade 106. The cooling fluid 70 exits through the holes 126 into the flow cavity 134. In some cases, the cooling fluid 70 exits through a plurality of such holes 126. Further, the cooling fluid 70 can exit uniformly through the gap 136. The cooling fluid 70 can further exit to the second cavity 118 at the high radial location 146. As indicated by arrow 84, the cooling fluid further flows through the gap 136 out of the second cavity 118.

[0093] In some embodiments, the seal member 124 further includes one or more fins 142 that extend toward and cooperate with the stator seal 120. The fins 142 seal against a static seal portion 144 of the stator seal 120. The stator seal 120 can create a barrier to the flow of cooling fluid 70 by forcing the cooling fluid 70 laterally through the fins 142. During operation, a relatively small amount of cooling fluid 70 can pass through the seal member 124 from the second cavity 118 to the third cavity 119 to provide cooling thereto.

[0094] The control arm 132 can avoid a circumferential temperature gradient across the stator well cavity (i.e., the second cavity 118). Further, the control arm 132 can ensure that the surfaces of the rotor disk 108, the seal plate 122, and the rotor blade 106 are always wetted by the cooling fluid 70. Depending on the length 148 of the control arm 132 and the width 138 of the gap 136 (i.e., the distance between the seal plate 122 and the tip 131 of the control arm 132), the control arm 132 can allow mixing of the cooling fluid 70 and the core airflow A at a high radial position 146. This can result in a reduction of the temperature of the second cavity 118. Further, a high temperature gradient between the two cavities (i.e., the second and third cavities 118, 119) in front of and behind the stator seal 120 can also be avoided. Such a high temperature gradient can otherwise adversely affect the performance of the stator seal 120.

[0095] The control arm 132 can allow for an even mixing between the cooling fluid 70 and the extracted flow from the core airflow A. Improving the mixing between the cooling fluid 70 and the extracted flow can reduce the temperature of the blade platform 114 and the aft or downstream portions of the rotor blade 106 and the rotor disk 108. This can reduce the amount of cooling fluid 70 required to meet the appropriate temperature. In other words, a lower temperature can be achieved with the same amount of flow of cooling fluid. Thus, this can improve the specific fuel consumption (SFC) of the gas turbine engine 10 and / or allow for the use of lower cost materials. The control arm 132 can also reduce the weight, manufacturing complexity, and / or cost of the blade assembly 100. The estimated temperature reduction achieved by the flow including the control arm 132 can be between 60 K and 100 K.

[0096] Figure 5A A cross-sectional side view of a portion of a blade assembly 200 for a gas turbine engine 10 is illustrated. Figure 5B An enlarged view of the blade assembly 200 is illustrated. Reference is made to Figure 5A and Figure 5BThe vane assembly 200 is similar to the vane assembly 100. The vane assembly 200 includes a rotor 202, a stator 204, rotor vanes 206, stator vanes 207, a rotor disk 208, a vane slot 210, a vane platform 214, a first cavity 216, a second cavity 218, a third cavity 219, a stator seal 220, a seal plate 222, a seal member 224, a bore 226, a flow cavity 234, a gap 236, a fin 242, and a static seal portion 244, which are respectively equivalent to the rotor 102, the stator 104, the rotor vanes 106, the stator vanes 107, the rotor disk 108, the vane slot 110, the vane platform 114, the first cavity 116, the second cavity 118, the third cavity 119, the stator seal 120, the seal plate 122, the seal member 124, the bore 126, the flow cavity 134, the gap 136, the fin 142, and the static seal portion 144 of the vane assembly 100.

[0097] The rotor 202 further includes a drive arm 254 extending from the rotor disk 208. The drive arm 254 can be annular. The drive arm 254 can engage another drive arm 260. The drive arm 260 is part of the seal member 224. In the illustrated embodiment, the seal member 224 is separate from the drive arm 254. However, in some other cases, the seal member 224 can be integral with the drive arm 254. Further, the seal member 224 includes a main portion 230. The main portion 230 is generally parallel to the main rotational axis 9 of the rotor 202.

[0098] The vane assembly 200 further includes a control arm 232. The control arm 232 includes a tip 231 distal from the main portion 230 of the seal member 224. The control arm 232 extends radially outwardly and axially from an end 256 of the main portion 230 toward the seal plate 222 relative to the main rotational axis 9. A distance 258 between the end 256 of the main portion 230 and the seal plate 222 is greater than a width 238 of the gap 236 (i.e., a distance between the tip 231 of the control arm 232 and the seal plate 222). In other words, the control arm 232 converges at a high radial location 246 near the tip 231 of the control arm 232. It can be desirable for the gap 236 to be as small as possible at the high radial location 246.

[0099] In the illustrated embodiment, the control arm 232 includes two portions, a first portion 250 and a second portion 252. However, in some cases, the control arm 232 can include more than two portions. In some embodiments, each of the first and second portions 250, 252 of the control arm 232 can have the same shape and size. In some other embodiments, each of the first portion 250 and the second portion 252 of the control arm 232 can have different shapes and sizes. The first portion 250 of the control arm 232 extends radially outwardly and axially from the end 256 of the main portion 230 toward the seal plate 222 relative to the main rotational axis 9 (or in the direction of the main rotational axis 9 in the case of the second portion 252).Figure 5B The parallel lines 8 shown in FIG. 6 are obliquely inclined and subtend an angle A3 with the main rotational axis 9. In the illustrated embodiment, the angle A3 is positive. However, in some other cases, the angle A3 can be negative based on the direction in which the control arm 232 is inclined. In other words, the first portion 250 of the control arm 232 can be inclined in a downward direction relative to the main rotational axis 9. The second portion 252 of the control arm 232 is substantially perpendicular to the main rotational axis 9 (or subtends an angle A4 of 0° with the main rotational axis 9) at the high radial location 246. In other words, the second portion 252 of the control arm 232 is substantially parallel to the seal plate 222. However, in other cases, the first portion 250 and the second portion 252 of the control arm 232 can have different configurations so long as the control arm 232 converges at the high radial location 246. For example, the first portion 250 and the second portion 252 of the control arm 232 can be inclined at the same angle relative to the main rotational axis 9. In some other examples, the control arm 232 can have a single straight portion that is inclined at an angle. In some other examples, the control arm 232 can have a single curved portion that converges at the high radial location 246. Figure 5B

[0100] As indicated by arrow 86, the cooling fluid 70 is received within the vane slot 210 from the first cavity 216. Further, as indicated by arrow 88, the cooling fluid 70 passes through the aperture 226 and enters the flow cavity 234. As indicated by arrow 90, the cooling fluid 70 swirls around the flow cavity 234. The cooling fluid 70 can provide cooling to the rotor 202 within the flow cavity 234. As indicated by arrow 92, the cooling fluid 70 passes through the slit 236 at the high radial location 246, exits the flow cavity 234, and enters the second cavity 218. Further, as indicated by arrow 94, the cooling fluid 70 exits the second cavity 218 and mixes with the core gas flow A. At least a relatively small portion of the cooling fluid 70 passes from the second cavity 218 to the third cavity 219 through the fins 242 of the seal member 224 to provide cooling thereto. The control arm 232 can allow for uniform mixing of the cooling fluid 70 with the extraction flow from the core gas flow A.

[0101] It will be understood that the present application is not limited to the embodiments described above and various modifications and improvements can be made without departing from the concepts described herein.​

Claims

1. A vane assembly (100, 200) for a gas turbine engine (10), the vane assembly (100, 200) comprising: a rotor (102, 202) including a rotor blade (106, 206), a vane platform (114), and a rotor disk (108, 208), the rotor blade and the rotor disk defining a vane slot (110, 210) that at least partially receives the rotor blade (106, 206) therein, wherein the vane slot (110, 210) receives cooling fluid (70) from a first cavity (116, 216) upstream of the rotor (102, 202); a stator (104, 204) disposed downstream of the rotor (102, 202), wherein the stator (104, 204) and the rotor (102, 202) define a second cavity (118, 218) therebetween downstream of the first cavity (116, 216); and a seal plate (122, 222) extending radially between and coupled to the vane platform (114) and the rotor disk (108, 208) of the rotor (102, 202) and facing the second cavity (118, 218), wherein the seal plate (122, 222) defines an aperture (126, 226) therethrough that is in fluid communication with the vane slot (110, 210); the vane assembly characterized by, the seal plate (122, 222) having an annular and planar configuration and being received in corresponding slots of the vane platform (114) and the rotor disk (108, 208) so as to axially retain the rotor blade (106, 206) in the vane slot (110, 210) of the rotor disk; and the vane assembly having a seal member (124, 224) extending from, integral with, or separately formed from and then connected to the rotor disk (108, 208) and including a main portion (130) extending from the rotor disk and a control arm (132, 232) extending obliquely in radial and axial directions relative to a main rotational axis of the rotor toward the seal plate (122, 222); wherein the seal member (124, 224) and the rotor (102, 202) define a flow cavity (134, 234) therebetween that is in fluid communication with the aperture (126, 226) of the seal plate (122, 222), and the control arm (132, 232) and the seal plate (122, 222) define a gap (136, 236) therebetween that fluidly communicates the flow cavity (134, 234) with the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, 216) through the aperture (126, 226) of the seal plate (122, 222) and into the flow cavity (134, 234) and then through the gap (136, 236) and into the second cavity (118, 218); and the vane assembly further characterized by, the control arm (132, 232) being configured to direct cooling fluid (70) from the first cavity (116, wherein the flow cavity (134, 234) receives the cooling fluid (70) flowing through the blade slot (110, 210) and the hole (126, 226) of the seal plate (122, 222) of the rotor disk (108, 208), the control arm (132, 232) deflects at least part of the cooling fluid (70) entering the flow cavity (134, 234), and the gap (136, 236) allows at least part of the cooling fluid (70) to exit the flow cavity (134, 234) and enter the second cavity (118, 218).

2. The blade assembly (100) of claim 1, wherein a major portion (130) of the seal member (124) is obliquely inclined with respect to the main rotational axis (9) of the rotor (102).

3. The blade assembly (200) of claim 1, wherein a major portion (230) of the seal member (224) is substantially parallel to the main rotational axis (9) of the rotor (202).

4. The blade assembly (100) of claim 1, wherein an angle (Al) between the major portion (130) of the seal member (124) and the main rotational axis (9) of the rotor (102) is from -60 degrees to 60 degrees.

5. The blade assembly (100) of any of the preceding claims, wherein the control arm (132) further comprises a tip (131) distal from the major portion (130) of the seal member (124), and wherein the gap (136) is defined between the tip (131) of the control arm (132) and the seal plate (122).

6. The blade assembly (100) of any of the preceding claims, wherein the stator (104) further comprises a stator seal (120).

7. The blade assembly (100) of claim 6, wherein the seal member (124) further comprises one or more fins (142) extending toward and cooperating with the stator seal (120).

8. A turbine (17, 19) for a gas turbine engine (10), the turbine (17, 19) comprising a blade assembly (100) according to any of the preceding claims.

9. A gas turbine engine (10) for an aircraft, the gas turbine engine (10) comprising a turbine (17, 19) according to claim 8.

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