Patent Application
20210062862
Exemplary embodiments pertain to the art of gas turbine engines, and more particularly to seal assemblies for gas turbine engines.
In a gas turbine engine, a number of components rotate under tight tolerances about an engine central longitudinal axis relative to static components. For example, the compressor and turbine sections of the gas turbine engine include rotating rotors with rotor blades extending radially outward. The rotor rotates relative to a stator with a small annular gap therebetween. To increase efficiency of the gas turbine engine, it is important that such small gaps be maintained to limit leakage through the gap, but to also allow for rotation of the rotor relative to the stator.
Seals are often utilized to manage leakage through the gaps. Such seals are typically fixed to static components and may be contact seals, such as labyrinth or brush seals, while others may be non-contact seals such as hydrostatic seals. Some hydrostatic seals are configured with a shoe having a plurality of beams, the beams having radial travel in response to a pressure differential across the seal.
The seal shoe floats via its outboard beams. The intent is for the shoes to move radially up and down, perpendicular to the rotor surface, as to maintain a tight gap for sealing. In actual practice, the shoes do not move radially up and down but have a slight swing or imbalance in them. Due to the cantilever design of the shoe, each shoe has a “free end” that is radially softer and a “fixed end” which is radially stiffer. The free end has a larger radial swing than the fixed end. If the imbalance is too high, the seal shoe can clash with the rotor and cause rotor damage.
In one embodiment, a hydrostatic seal and rotor assembly includes a rotor having a rotor rotational direction about a rotor central axis, and a hydrostatic seal assembly including a seal support, a seal shoe configured for sealing between the seal shoe and the rotor, and one or more seal beams operably connecting the seal shoe to the seal support. The one or more seal beams are configured as spring elements integral with the seal shoe to allow radial movement of the seal shoe relative to the seal support. The seal shoe is circumferentially cantilevered from a fixed end to a free end opposite the fixed end. The fixed end is located circumferentially upstream of the free end relative to the rotor rotational direction.
Additionally or alternatively, in this or other embodiments the seal beam extends between a radially fixed first beam end and a radially movable second beam end.
Additionally or alternatively, in this or other embodiments the fixed end of the seal shoe is located at the second beam end.
Additionally or alternatively, in this or other embodiments one or more stops are formed in the seal shoe configured to limit radial travel of the seal shoe.
Additionally or alternatively, in this or other embodiments a seal carrier is configured to axially and radially retain the seal support.
Additionally or alternatively, in this or other embodiments the seal shoe is a plurality of circumferentially spaced and segmented seal shoes.
Additionally or alternatively, in this or other embodiments each seal shoe of the plurality of seal shoes is circumferentially cantilevered from a fixed end to a free end opposite the fixed end. The fixed end is located circumferentially upstream of the free end relative to the rotor rotational direction.
In another embodiment, a turbine section of a gas turbine engine includes a turbine stator, a turbine rotor configured to rotate in a rotor rotational direction about an engine central longitudinal axis relative to the turbine stator, and a hydrostatic seal assembly. The hydrostatic seal assembly includes a seal support, a seal shoe configured for sealing between the seal shoe and the rotor, and one or more seal beams operably connecting the seal shoe to the seal support. The one or more seal beams are configured as spring elements integral with the seal shoe to allow radial movement of the seal shoe relative to the seal support. The seal shoe is circumferentially cantilevered from a fixed end to a free end opposite the fixed end. The fixed end is located circumferentially upstream of the free end relative to the rotor rotational direction.
Additionally or alternatively, in this or other embodiments the seal beam extends between a radially fixed first beam end and a radially movable second beam end.
Additionally or alternatively, in this or other embodiments the fixed end of the seal shoe is located at the second beam end.
Additionally or alternatively, in this or other embodiments one or more stops are formed in the seal shoe and are configured to limit radial travel of the seal shoe.
Additionally or alternatively, in this or other embodiments a seal carrier is configured to axially and radially retain the seal support.
Additionally or alternatively, in this or other embodiments the seal shoe is a plurality of circumferentially spaced and segmented seal shoes.
Additionally or alternatively, in this or other embodiments each seal shoe of the plurality of seal shoes is circumferentially cantilevered from a fixed end to a free end opposite the fixed end. The fixed end is located circumferentially upstream of the free end relative to the rotor rotational direction.
In yet another embodiment, a gas turbine engine includes a combustor, and a turbine section in fluid communication with the combustor. The turbine section includes a turbine stator, a turbine rotor configured to rotate about an engine central longitudinal axis relative to the turbine stator, and a hydrostatic seal assembly. The hydrostatic seal assembly includes a seal support, a seal shoe configured for sealing between the seal shoe and the rotor, and one or more seal beams operably connecting the seal shoe to the seal support. The one or more seal beams are configured as spring elements integral with the seal shoe to allow radial movement of the seal shoe relative to the seal support. The seal shoe is circumferentially cantilevered from a fixed end to a free end opposite the fixed end. The fixed end is located circumferentially upstream of the free end relative to the rotor rotational direction.
Additionally or alternatively, in this or other embodiments the seal beam extends between a radially fixed first beam end and a radially movable second beam end.
Additionally or alternatively, in this or other embodiments the fixed end of the seal shoe is located at the second beam end.
Additionally or alternatively, in this or other embodiments one or more stops are formed in the seal shoe configured to limit radial travel of the seal shoe.
Additionally or alternatively, in this or other embodiments a seal carrier is configured to axially and radially retain the seal support.
Additionally or alternatively, in this or other embodiments the seal shoe is a plurality of circumferentially spaced and segmented seal shoes.
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).
Referring now to
The seal 60 is fixed to the turbine stator 62 via a seal carrier 66 and includes a primary seal 68 and one or more secondary seals 70. The primary seal 68 includes a seal support 104, which supports a seal shoe 72 via one or more seal beams 74 which, as shown best in
In operation, an airflow 78 flows through the air gap 76 from a high pressure area 80 upstream of the seal 60 toward a low pressure area 82 downstream of the seal 60. Further, airflow enters a seal cavity 84 radially outboard of the seal shoe 72 via one or more plate openings 86 in an aft plate 88 of the seal carrier 66, which is downstream of the seal shoe 72 and in some embodiments axially abuts the seal shoe 72. The secondary seals 70 are located upstream of the seal shoe 72 and in some embodiments abut the seal shoe 72. The secondary seal 70 prevents airflow from entering the seal cavity 84 from the high pressure area 80 and/or prevents airflow from exiting the seal cavity 84 via an upstream side of the seal 60. In some embodiments, the secondary seals 70 are axially retained at the seal shoe 72 by a secondary seal cover 106 upstream of the secondary seals 70. Further, a radial and axial position of the secondary seal 70 may be maintained by a spacer 90. The seal shoe 72 moves radially until a pressure equilibrium between the air gap 76 and the seal cavity 84 is reached.
Referring again to
The seal beams 74 extend between a radially fixed first beam end 108 and a radially movable second beam end 110. The seal shoe 72 is cantilevered from the second beam end 110 at a fixed end 112 of the seal shoe 72 to a free end 114 of the seal shoe 72 opposite the fixed end 112. As stated above, the travel of the seal shoes 72 is limited by the features of the primary seal 68. During operation of the seal 60, the free end 114 moves a first radial distance 116 greater than a second radial distance 118 that the fixed end 112 moves. Thus, there is incidental contact between the turbine rotor 64 and the seal shoe 72, such contact is likeliest to occur at the free end 114. To prevent or minimize damage to the turbine rotor 64 during such incidental contact with the seal shoe 72, the seal shoe 72 is particularly oriented and positioned relative to the turbine rotor 64.
The turbine rotor 64 has a rotor rotation direction 120 about a rotor central axis which may be engine central longitudinal axis A, as shown in
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.
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.
This invention was made with Government support awarded by the United States. The Government has certain rights in the invention.
