Abstract: A method facilitates assembling a gas turbine engine including a compressor and a rotor assembly coupled in axial flow communication downstream from the compressor. The method comprises coupling a bypass system in flow communication with the compressor to channel a portion of flow discharged from the compressor towards the rotor assembly is channeled through the bypass system, and coupling a downstream end of the bypass system within the gas turbine engine such that the flow entering the bypass system flows past the rotor assembly and is discharged downstream from the rotor assembly.

Abstract: A gas turbine engine aspirating face seal includes rotatable and non-rotatable engine members and a leakage path therebetween. Annular generally planar rotatable and non-rotatable gas bearing face surfaces are operably associated with the rotatable and non-rotatable engine members respectively and are circumscribed about and generally perpendicular to a centerline axis. A substantially fully annular pull off biasing element is operably disposed for urging the non-rotatable gas bearing face surface axially away from the rotatable gas bearing face surface and circumscribed about the centerline axis. The pull off biasing element may be at least one wave spring or one bellville washer. The non-rotatable gas bearing face surface may be on a face seal ring mounted on a translatable cylindrical piston which is axially movable and supported by the non-rotatable engine member.

Abstract: An aircraft gas turbine engine includes high and low pressure turbines having respective counter rotatable low pressure inner and outer rotors with low pressure inner and outer shafts in part rotatably disposed co-axially within a high pressure rotor and drivingly connected to first and second fan blade rows and first and second boosters respectively. A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the boosters axially located between the first and second fan blade rows. The engine has a high pressure compressor operable to produce an overall pressure ratio in a range of about 40-65 and a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds in a range of 1000 to 2500 feet per second.

Abstract: A gas turbine engine aspirating face seal includes a rotatable engine member and a non-rotatable engine member and a leakage path therebetween. Annular generally planar rotatable and non-rotatable gas bearing face surfaces circumscribed about a centerline are operably associated the rotatable and non-rotatable engine members respectively. Radially inner and outer tooth rings axially extend away from a first one of the rotatable and non-rotatable gas bearing face surfaces across the leakage path and towards a second one of the gas bearing face surfaces. A pull off biasing means is used for urging the inner and outer tooth rings axially away from the second one of the gas bearing face surfaces. The rotatable engine member may be a side plate mounted on a rotor disk and the non-rotatable engine member is mounted on a translatable piston axially movably supported on a stationary face seal support structure.

Abstract: A support arrangement for a brush seal positioned between a stationary and a rotatable component of a turbomachine. The brush seal has a downstream backing plate that includes an axially-extending lip adjacent its radially outermost end to minimize axial deflection of the seal as a result of a gas pressure differential acting across the seal.

Abstract: A gas turbine engine aspirating face seal includes a rotatable engine member and a non-rotatable engine member and a leakage path therebetween. Annular generally planar rotatable and non-rotatable gas bearing face surfaces circumscribed about a centerline are operably associated the rotatable and non-rotatable engine members respectively. Radially inner and outer tooth rings axially extend away from a first one of the rotatable and non-rotatable gas bearing face surfaces across the leakage path and towards a second one of the gas bearing face surfaces. A pull off biasing means is used for urging the inner and outer tooth rings axially away from the second one of the gas bearing face surfaces. The rotatable engine member may be a side plate mounted on a rotor disk and the non-rotatable engine member is mounted on a translatable piston axially movably supported on a stationary face seal support structure.

Abstract: A gas turbine engine aspirating face seal includes rotatable and non-rotatable engine members and a leakage path therebetween. Annular generally planar rotatable and non-rotatable gas bearing face surfaces are operably associated with the rotatable and non-rotatable engine members respectively and are circumscribed about and generally perpendicular to a centerline axis. A substantially fully annular pull off biasing means s operably disposed for urging the non-rotatable gas bearing face surface axially away from the rotatable gas bearing face surface and circumscribed about the centerline axis. The pull off biasing means may be at least one wave spring or one bellville washer. The non-rotatable gas bearing face surface may be on a face seal ring mounted on a translatable cylindrical piston which is axially movable and supported by the non-rotatable engine member.

Abstract: An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Abstract: An aircraft gas turbine engine includes high and low pressure turbines having respective counter rotatable low pressure inner and outer rotors with low pressure inner and outer shafts in part rotatably disposed co-axially within a high pressure rotor and drivingly connected to first and second fan blade rows and first and second boosters respectively. A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the boosters axially located between the first and second fan blade rows. The engine has a high pressure compressor operable to produce an overall pressure ratio in a range of about 40-65 and a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds in a range of 1000 to 2500 feet per second.

Abstract: An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Abstract: Overall performance in enhanced in gas turbine engines by providing a rotating seal including a rotating member arranged to rotate about an axis and having at least one annular projection extending radially outwardly therefrom, and a stator element having a first surface arranged to contact the projection. The stator element includes at least one slot formed in the first surface, the slot axially traversing the projection so as to allow a flow of purge air to pass. More than one such slot can be used, and each slot is preferably angled circumferentially in the direction of rotation of the rotating member.

Abstract: A blocker and swirl inducer hole configuration for use in connection with a high pressure turbine is described. In one embodiment, the blocker holes are oriented to a 45-degree tangential angle with respect to the direction of rotation of the seal, which results in pre-swirling the air before being injected into the swirl cavity. In addition, the number of blocker holes is reduced by as much as 50% of the number of blocker holes used in the known CFM56 turbine. Further, rather than injecting the air into the first swirl cavity as is known, the air is injected into a second swirl cavity. The combined effect of orienting the holes to the 45-degree tangential angle with respect to the direction of rotation of the seal, locating the holes to open into the second swirl cavity, and reducing the flow area by about 50%, results in an increase in blocker hole pressure ratio. Increasing the blocker hole pressure ratio results in a higher hole exit velocity which maximizes the cavity inlet swirl.

Abstract: A turbine rotor includes a rotor disk having a plurality of dovetail slots defining therebetween respective disk posts. Each post has a top land and a plurality of tangs spaced radially therebelow inside adjacent ones of the slots. A plurality of turbine blades each having a dovetail are mounted in respective ones of the dovetail slots for radially retaining the blades. A platform is joined integrally with the dovetail and extends circumferentially to adjoin adjacent ones of the platforms to define a radially inner flowpath. An airfoil extends integrally from the platform for extracting energy from combustion gases flowable thereover. A thermal barrier coating is fixedly bonded to each of the disk post top lands for providing thermal insulation between the platforms and the disk posts for reducing disk post temperature.