Patent Application
20240360772
Exemplary embodiments of the present disclosure pertain to the art of gas turbine engines, and in particular to blade outer air seal (BOAS) configurations of turbine sections or compressor sections of gas turbine engines.
During a typical rapid acceleration of a gas turbine engine, the rotors of the turbine and/or compressor expand radially outwardly more rapidly than the corresponding blade outer airseals (BOAS). This results in a pinch condition and excessive rub of the BOAS, resulting in an increased radial clearance between the rotor blade tip and the BOAS when the engine returns to a cruise operating condition. This increased clearance reduces performance of the gas turbine engine.
The art would welcome a configuration allowing for rapid radially outward movement of the BOAS during such rapid accelerations.
In one embodiment, a blade outer airseal assembly of a gas turbine engine includes a blade outer airseal (BOAS) segment, and a BOAS carrier positioned radially outboard of the BOAS segment relative to an engine central longitudinal axis. The BOAS segment is secured to the BOAS carrier, and an adjustment lever is operably connected to the BOAS carrier. Rotation of the adjustment lever about a pivot axis at a pivot urges movement of the BOAS segment in a radial direction.
Additionally or alternatively, in this or other embodiments the adjustment lever includes a first lever arm and a second lever arm. The first lever arm is engaged with a complimentary engagement feature of the BOAS carrier.
Additionally or alternatively, in this or other embodiments a pin rotably secures the adjustment lever to a casing element of the gas turbine engine. The pin defines the pivot axis.
Additionally or alternatively, in this or other embodiments an actuator is operably connected to the second lever arm. The actuator is configured to urge rotation of the adjustment lever about the pivot axis.
Additionally or alternatively, in this or other embodiments the engagement feature is a slot in the BOAS carrier. The first lever arm extends at least partially through the slot.
Additionally or alternatively, in this or other embodiments the BOAS carrier, the adjustment lever and the pivot are all disposed in a common pressure cavity of the gas turbine engine.
In another embodiment, a turbine section of a gas turbine engine includes a turbine rotor positioned at an engine central longitudinal axis, and a blade outer airseal (BOAS) assembly located radially outboard of the turbine rotor. The BOAS assembly includes a plurality of BOAS segments arrayed circumferentially about the engine central longitudinal axis, and a plurality of BOAS carriers located radially outboard of the plurality of BOAS segments. Each BOAS carrier is supportive of at least one BOAS segment. The BOAS assembly further includes a plurality of adjustment levers. Each adjustment lever is operably connected to at least one BOAS carrier of the plurality of BOAS carriers. Rotation of each adjustment lever about a respective pivot axis at a respective pivot of a plurality of pivots urges movement of the plurality of BOAS segments in a radial direction thereby adjusting a radial gap between the turbine rotor and the plurality of BOAS segments.
Additionally or alternatively, in this or other embodiments each adjustment lever includes a first lever arm and an opposing second lever arm extending from the pivot axis.
Additionally or alternatively, in this or other embodiments the first lever arm is engaged with a complimentary engagement feature of at least one BOAS carrier of the plurality of BOAS carriers.
Additionally or alternatively, in this or other embodiments the first lever arm is engaged with complimentary engagement features of two circumferentially adjacent BOAS carriers of the plurality of BOAS carriers.
Additionally or alternatively, in this or other embodiments the engagement feature is a slot in the BOAS carrier. The first lever arm extends at least partially through the slot.
Additionally or alternatively, in this or other embodiments an actuator is operably connected to the second lever arm. The actuator is configured to urge rotation of the adjustment lever about the pivot axis.
Additionally or alternatively, in this or other embodiments a pin rotably secures the adjustment lever to a casing element of the gas turbine engine. The pin defines the pivot axis.
Additionally or alternatively, in this or other embodiments the plurality of BOAS carriers, the plurality of adjustment levers and the plurality of pivots are all positioned in a common pressure cavity of the gas turbine engine.
In yet another embodiment, a gas turbine engine includes a combustor, and a turbine section driven by combustion products from the combustor. The turbine section includes a turbine rotor positioned at an engine central longitudinal axis, and a blade outer airseal (BOAS) assembly located radially outboard of the turbine rotor. The BOAS assembly includes a plurality of BOAS segments arrayed circumferentially about the engine central longitudinal axis, and a plurality of BOAS carriers located radially outboard of the plurality of BOAS segments. Each BOAS carrier is supportive of at least one BOAS segment. The BOAS assembly further includes a plurality of adjustment levers. Each adjustment lever is operably connected to at least one BOAS carrier of the plurality of BOAS carriers. Rotation of each adjustment lever about a respective pivot axis at a respective pivot of a plurality of pivots urges movement of the plurality of BOAS segments in a radial direction thereby adjusting a radial gap between the turbine rotor and the plurality of BOAS segments.
Additionally or alternatively, in this or other embodiments each adjustment lever includes a first lever arm and an opposing second lever arm extending from the pivot axis, and the first lever arm is engaged with a complimentary engagement feature of at least one BOAS carrier of the plurality of BOAS carriers.
Additionally or alternatively, in this or other embodiments the first lever arm is engaged with complimentary engagement features of two circumferentially adjacent BOAS carriers of the plurality of BOAS carriers.
Additionally or alternatively, in this or other embodiments the engagement feature is a slot in the BOAS carrier, the first lever arm extending at least partially through the slot.
Additionally or alternatively, in this or other embodiments an actuator is operably connected to the second lever arm. The actuator is configured to urge rotation of the adjustment lever about the pivot axis.
Additionally or alternatively, in this or other embodiments the plurality of BOAS carriers, the plurality of adjustment levers and the plurality of pivots are all disposed in a common pressure cavity of the gas turbine engine.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
Illustrated in
Referring now to
To effect movement of the BOAS segments 70, the BOAS carriers 72 are operably connected to a plurality of adjustment levers 76 secured to the case member 74. The adjustment levers 76 are each retained at the case member 74 via a pin 78 extending through a lever pivot 80 and a casing flange 82, best shown in
The rotation of the adjustment lever 76 is driven and controlled by an actuator 98 operably connected to the adjustment lever 76. In one embodiment, illustrated in
Referring again to
Referring now to
Referring again to
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
