Abstract: A turbine exhaust case is provided for a gas turbine engine. This turbine exhaust case includes a first gas turbine engine exhaust case, a second gas turbine engine exhaust case, a strut and a fastener. The second gas turbine engine exhaust case includes a flange and a bracket. The strut is coupled to the first case and the second case. The fastener fastens the flange and the strut. The second gas turbine engine exhaust case is co-axial with the first gas turbine engine exhaust case. The strut contacts the bracket when the engine is powered-on.

Abstract: A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.

Abstract: A seal system for a gas turbine engine includes a cover plate with a radial flange; a seal carrier adjacent the cover plate; an outer compliant seal supported by the seal carrier; an inner compliant seal supported by the seal carrier, the inner compliant seal engaged with the radial flange; and a spring between the cover plate and the seal carrier.

Abstract: A turbine exhaust case is provided for a gas turbine engine. This turbine exhaust case includes a first gas turbine engine exhaust case, a second gas turbine engine exhaust case, a strut and a fastener. The second gas turbine engine exhaust case includes a flange and a bracket. The strut is coupled to the first case and the second case. The fastener fastens the flange and the strut. The second gas turbine engine exhaust case is co-axial with the first gas turbine engine exhaust case. The strut contacts the bracket when the engine is powered-on.

Abstract: A clearance control system for a gas turbine engine includes a full hoop blade outer air seal carrier ring, a full hoop thermal control ring at least partially contained within the full hoop blade outer air seal blade outer air seal carrier ring to form an interference fit therebetween, and a multiple of blade outer air seals mounted to the full hoop blade outer air seal carrier ring.

Abstract: A clearance control system for a gas turbine engine includes a full hoop blade outer air seal carrier ring, a full hoop thermal control ring at least partially contained within the full hoop blade outer air seal blade outer air seal carrier ring to form an interference fit therebetween, and a multiple of blade outer air seals mounted to the full hoop blade outer air seal carrier ring.

Abstract: A seal system for a gas turbine engine includes a cover plate with a radial flange; a seal carrier adjacent the cover plate; an outer compliant seal supported by the seal carrier; an inner compliant seal supported by the seal carrier, the inner compliant seal engaged with the radial flange; and a spring between the cover plate and the seal carrier.

Abstract: A rotating blade tip clearance system includes a control ring carrier defining a centerline axis. The control ring carrier has a connecting portion, a retaining portion radially inward of the connecting portion, and a flange connecting radially between the connecting portion and the retaining portion. The flange isolates the retaining portion of the control ring carrier from the thermal deflection of a case and assists in keeping the control ring carrier aligned about centerline axis A during thermal deflection. The retaining portion includes radially inner and outer diameter sides defining a retaining cavity therebetween. The system includes a control ring within the retaining cavity. The control ring has a different thermal response rate from the control ring carrier so that the control ring thermally deflects slower than the control ring carrier, thereby controlling the rate and/or extent of thermal deflection of the control ring carrier.

Abstract: A seal system for a gas turbine engine includes an annular seal carrier and an annular compliant seal mounted around the annular seal carrier, the compliant seal includes a multiple of legs in cross-section.

Abstract: A gas turbine engine includes a core assembly with a core case having a containment section for containing liberated compressor and turbine blades and blade fragments. The containment section includes first and second containment layers and the containment section is configured to have a non-linear rate of energy dissipation across the first and second containment layers, thereby to improve containment of blades and blade fragments.

Abstract: A gas turbine engine includes a containment ring to contain liberated compressor and turbine blades and blade fragments within a core assembly. The combination of a containment gap between the core case and the containment ring and a containment layer disposed in the gap helps dissipate the energy generated by loose body impacts on the core assembly. The containment layer deforms, deflects, and/or redirects the impact energy acting in a radial direction, thereby to improve containment.

Abstract: An assembly for use with a gas turbine engine includes a tie rod and a connector. The tie rod, which is for extending radially outwardly from a latitudinal axis of the gas turbine engine, has a hollow length having a longitudinal axis and a base having a width in parallel to the latitudinal axis. The base has a counterbore disposed therein and is wider than a width of the length. The connector, for attaching the base of the tie rod to a bearing assembly of the gas turbine engine, has a hollow body having a shaft removably attaching, at a first end portion thereof, to the counterbore. The shaft and the counterbore are disposed in parallel with the longitudinal axis.

Abstract: A rotating blade tip clearance system includes a control ring carrier defining a centerline axis. The control ring carrier has a connecting portion, a retaining portion radially inward of the connecting portion, and a flange connecting radially between the connecting portion and the retaining portion. The flange isolates the retaining portion of the control ring carrier from the thermal deflection of a case and assists in keeping the control ring carrier aligned about centerline axis A during thermal deflection. The retaining portion includes radially inner and outer diameter sides defining a retaining cavity therebetween. The system includes a control ring within the retaining cavity. The control ring has a different thermal response rate from the control ring carrier so that the control ring thermally deflects slower than the control ring carrier, thereby controlling the rate and/or extent of thermal deflection of the control ring carrier.

Abstract: A gas turbine engine includes a seal assembly that is supported by a member at a joint. The seal assembly includes a seal support having a radial flange secured to the joint. A first bend adjoins the radial flange to a first leg, which is oriented generally in an axial direction. A second bend adjoins the first leg to a second leg, which is conical in shape. A seal is supported by the second leg.

Abstract: A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.

Abstract: A gas turbine engine includes high and low pressure turbines. A mid turbine frame is arranged axially between the high and low pressure turbines. A bearing is operatively supported by a support structure. An inner case is secured to the support structure and includes a first conical member and a bearing support to which the bearing is mounted. The bearing support includes a second conical member that is secured to the first conical member at a joint. The first and second conical members are arranged radially inward of the joint.

Abstract: A seal system for a gas turbine engine includes an annular seal carrier and an annular compliant seal mounted around the annular seal carrier, the compliant seal includes a multiple of legs in cross-section.

Abstract: A seal system for a gas turbine engine includes a cover plate with a radial flange; a seal carrier adjacent the cover plate; an outer compliant seal supported by the seal carrier; an inner compliant seal supported by the seal carrier, the inner compliant seal engaged with the radial flange; and a spring between the cover plate and the seal carrier.

Abstract: A gas turbine engine includes bound assemblies with an inner diameter ring, struts, and an outer diameter ring. The strut is connected to the inner diameter ring and extends radially outward therefrom to connect to the outer diameter ring. A strain relief feature is disposed adjacent to or at the connection between the strut and the inner diameter ring and/or the outer diameter ring. The strain relief feature lengthens the arc segment of fillet curvature. For a constant thermal punch load, the lengthened arc segment of fillet curvature results in a decreased maximum strain in the bound assembly.

Abstract: A gas turbine engine includes high and low pressure turbines. A mid turbine frame is arranged axially between the high and low pressure turbines. A bearing is operatively supported by a support structure. An inner case is secured to the support structure and includes a first conical member and a bearing support to which the bearing is mounted. The bearing support includes a second conical member that is secured to the first conical member at a joint. The first and second conical members are arranged radially inward of the joint.