Abstract: A case clearance control system comprising a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate the blade outer air seal support structure; a thermal control ring coupled to the blade outer air seal support structure, the thermal control ring including a plurality of receivers configured to couple with the plurality of protrusions; a thermal break formed between the plurality of protrusions and the plurality of receivers, the thermal break configured to control heat transfer between the blade outer air seal support structure and the thermal control ring; and a plurality of flow passages formed between the blade outer air seal support structure, the thermal control ring and the plurality of protrusions, the plurality of flow passages configured to allow cooling air flow to condition the thermal control ring and maintain thermal control ring dimensions.

Abstract: A compressor case to blade tip clearance system comprising a rotor having blades with tips, the case including an inner case comprising at least one surface feature fluidly coupled to a distribution manifold disposed in a cooling air passageway, the at least one surface feature configured to interact with the cooling air, and a tip clearance located between the tips and the inner case; wherein the tip clearance is maintained responsive to a flow of the cooling air over the at least one surface feature.

Abstract: A cooling system includes an inlet, a passageway, and a feed passage. The inlet is defined by a strut of a diffuser case that extends between a first case portion and a second case portion. The inlet extends from a leading surface towards a trailing surface along a first axis. The passageway is defined by the strut and extends from the inlet towards the second case portion along a second axis that is disposed transverse to the first axis. The feed passage is defined between a body of a tangential onboard injection system that extends between and is connected to an inner vane platform of a guide vane and the second case portion. The feed passage extends along an axis that is disposed parallel to the first axis.

Abstract: A case clearance control system comprising a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate the blade outer air seal support structure; a thermal control ring coupled to the blade outer air seal support structure, the thermal control ring including a plurality of receivers configured to couple with the plurality of protrusions; a thermal break formed between the plurality of protrusions and the plurality of receivers, the thermal break configured to control heat transfer between the blade outer air seal support structure and the thermal control ring; and a plurality of flow passages formed between the blade outer air seal support structure, the thermal control ring and the plurality of protrusions, the plurality of flow passages configured to allow cooling air flow to condition the thermal control ring and maintain thermal control ring dimensions.

Abstract: A compressor case to blade tip clearance system comprising a rotor having blades with tips, the case including an inner case comprising at least one surface feature fluidly coupled to a distribution manifold disposed in a cooling air passageway, the at least one surface feature configured to interact with the cooling air, and a tip clearance located between the tips and the inner case; wherein the tip clearance is maintained responsive to a flow of the cooling air over the at least one surface feature.

Abstract: A gas turbine engine is disclosed. The gas turbine engine includes a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades, a first bore cavity between the first rotor and the second rotor, a first fluid passageway configured to provide cooled air to the first bore cavity and a first anti-vortex component positioned proximate the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity.

Abstract: An example method of cooling a compressor of a gas turbine includes, among other things, diverting a flow from a compressor, and directing the flow at the compressor in a direction, the direction having a circumferential component and an axial component.

Abstract: A gas turbine engine rotor includes a rotor that provides a cooling cavity. The cooling cavity has a first chamber and a second chamber that are fluidly connected to one another by a passageway. At least one of the first and second rotor portions is configured to support a blade that is fluidly isolated from the cavity. A phase change material is arranged in the cavity. The phase change material is configured to be arranged in the first chamber in a first state and in the second chamber in the second state. The passageway is configured to carry the phase change material between the second and first chambers once changed between the first and second states.

Abstract: A gas turbine engine is disclosed. The gas turbine engine includes a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades, a first bore cavity between the first rotor and the second rotor, a first fluid passageway configured to provide cooled air to the first bore cavity and a first anti-vortex component positioned proximate the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity.

Abstract: An example method of cooling a compressor section of a gas turbine engine includes diverting a flow from a compressor through a heat exchanger, the flow moving from the compressor in a first direction, and moving the flow from the heat exchanger back to the compressor in a second direction. An example spacer for a compressor of a gas turbine engine includes a first side portion, a second side portion spaced apart from the first side portion, and a middle web arranged between the first and second side portions. At least one of the first and second side portions and the middle web include at least one orifice to communicate flow in a direction that is different from a core flowpath flow direction. An example compressor including the spacer is also disclosed.

Abstract: A gas turbine engine rotor includes a rotor that provides a cooling cavity. The cooling cavity has a first chamber and a second chamber that are fluidly connected to one another by a passageway. At least one of the first and second rotor portions is configured to support a blade that is fluidly isolated from the cavity. A phase change material is arranged in the cavity. The phase change material is configured to be arranged in the first chamber in a first state and in the second chamber in the second state. The passageway is configured to carry the phase change material between the second and first chambers once changed between the first and second states.

Abstract: An example method of cooling a compressor of a gas turbine includes, among other things, diverting a flow from a compressor, and directing the flow at the compressor in a direction, the direction having a circumferential component and an axial component.

Abstract: A gas turbine engine includes a rotor having at least one rotor bore cavity and a heat exchanger. The heat exchanger is fluidically connected to a conditioning air source and to the rotor bore cavity for flowing air from the conditioning air source to the rotor bore cavity. The heat exchanger is also fluidically connected to a hot fluid source and a cool fluid source for selectively heating and cooling conditioning air flowing to the rotor bore cavity.