Abstract: Aa gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first compressor having a first overall pressure ratio, and a second compressor having a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. Further, a section of the gas turbine engine includes a thermally isolated area.

Abstract: A section of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a thermally isolated area, and a first rotor disk and a second rotor disk. Each of the first and second rotor disks are provided within the thermally isolated area.

Abstract: A method comprises the steps of modifying an existing engine which includes a high pressure compressor driven by a high pressure turbine and a low pressure turbine to drive a low pressure compressor. The modifying step includes utilizing the high pressure compressor as a low pressure compressor in a modified gas turbine engine, and designing and incorporating a new high pressure compressor downstream of the low pressure compressor section in the modified engine, such that a portion of the design of the existing engine is utilized in the modified engine. A gas turbine engine is also disclosed.

Abstract: The present disclosure provides systems and methods related to thermal management systems for gas turbine engines. For example, a thermal management system comprises a thermally neutral heat transfer fluid circuit, a first heat exchanger disposed on the fluid circuit, and a second heat exchanger disposed on the fluid circuit.

Abstract: A gas turbine engine includes a turbine section that includes a turbine rotor arranged in a plenum. A compressor section includes a compressor rotor assembly that has spaced apart inner and outer portions that provide an axially extending cooling channel. Compressor blades extend radially outward from the outer portion which provides an inner core flow path. A rotor spoke is configured to receive a first cooling flow and fluidly connect the outer portion to the cooling channel. The compressor rotor assembly has a coolant exit that is in fluid communication with the cooling channel. The compressor rotor assembly is configured to communicate the first cooling flow to the turbine rotor. A bleed source is configured to provide a second cooling flow. A combustor section includes an injector in fluid communication with the bleed source. The tangential onboard injector is configured to communicate the second cooling flow to the turbine rotor.

Abstract: An engine comprises a fan configured to produce propulsive thrust from a fan discharge air stream. A bypass duct is located between a cowl and an engine core, the bypass duct is fluidly coupled to the fan. The bypass duct contains the air stream created by the fan. A pump system provides the air to an environmental control system. The pump system comprises an impeller having an inlet for receiving the air from the duct and an outlet for discharging the air to the environmental control system. An intake manifold is configured to receive the air from the bypass duct and to deliver the air to the inlet. A single discharge manifold is configured to deliver the air from the outlet to the environmental control system.

Abstract: A gas turbine engine may comprise a first rotor with a primary flowpath along an outer diameter of the first rotor. A secondary flowpath may be radially inward from the primary flowpath. The secondary flowpath may pass through an opening through the first rotor. A blade may be disposed on a distal end of the first rotor. The blade may extend into the primary flowpath. A bleed tube may be in a wall of the primary flowpath and forward of the blade. The bleed tube may extend radially inward from the primary flowpath. The bleed tube may fluidly connect to the opening through the first rotor. A plenum may be aft of the blade and radially inward from the primary flowpath. The plenum may be fluidly connected to the opening through the first rotor. A second rotor may be aft of the plenum.

Abstract: A gas turbine engine includes a core engine with a compressor section, a combustor and a turbine. The turbine drives an output shaft, and the output shaft drives at least four gears. Each of the at least four gears extends through a drive shaft to drive an associated fan rotor.

Abstract: A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section, and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a low spool including a low pressure compressor section and a low pressure turbine. A high spool includes a high pressure compressor section. A gear arrangement is defined along an engine axis. The low spool is operable to drive the fan section through the gear arrangement. A mount system includes an aft mount configured to react at least a portion of a thrust load at an engine case generally parallel to an engine axis.

Abstract: A gas turbine engine includes a core engine with a compressor section, a combustor and a turbine. The turbine drives an output shaft, and the output shaft drives at least four gears. Each of the at least four gears extends through a drive shaft to drive an associated fan rotor.

Abstract: A turbine section includes a rotor assembly which includes an internal cooling passage. A segmented seal is adjacent the rotor assembly and includes a fluid passage that is in fluid communication with the internal cooling passage.

Abstract: A rotor assembly of a gas turbine engine may be spoked and includes a rotor and a shell. The rotor has a rotor disk and a plurality of blades each having a platform attached to the rotor disk and with a first channel defined radially between the platforms and the rotor disk. The shell projects aft of the rotor and includes inner and outer walls with a passage defined therebetween. The passage is in fluid communication with the first channel and, together, form part of a secondary flowpath for cooling of adjacent components. The rotor assembly may further include a structure located radially inward of the rotor disk and shell. The structure defines a supply conduit for flowing air from the passage and into a rotor bore defined at least in part by adjacent rotor disks. The entering air, being pre-heated when flowing through the channel and passage, warms the bore and reduces thermal gradients, thus thermal fatigue, across the rotor disk.

Abstract: A rotor for a gas turbine engine includes a cold shell, a hot shell, and a spoke. The spoke is connected to and extends radially outward from the cold shell. The hot shell is connected to the cold shell by the spoke and includes an axially extending outboard segment and an axially extending inboard segment. The outboard segment is connected to the inboard segment and the inboard segment is disposed radially inboard of the outboard segment for sealably engaging a stator blade shroud.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a spool along an engine axis which drives a gear train, the spool including a low stage count low pressure turbine.

Abstract: A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed.

Abstract: A gas turbine engine comprises a compressor section, a combustor section downstream of the compressor section, and a turbine section downstream of the combustor section. A mid-turbine frame includes an outer case portion and is configured to support the turbine section. At least one shaft defines an axis of rotation, and the turbine section comprises an inner rotor directly driving the shaft. The inner rotor includes an inner set of blades. An outer rotor is positioned immediately adjacent to the outer case portion and has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system is positioned downstream of the combustor section, is mounted to the mid-turbine frame, and is coupled to the outer rotor to drive the at least one shaft.

Abstract: A rotor for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a rotor disk rotatable about an axis and a gas path wall coupled to and radially outward of the rotor disk. The gas path wall bounds a radially inward portion of a gas path. A plurality of rotor spokes are radially intermediate the rotor disk and the gas path wall. The plurality of rotor spokes is circumferentially spaced to define a plurality of cooling channels intermediate the rotor disk and the gas path wall. A thermal barrier coating is disposed on a surface of at least one of the plurality of cooling channels. A method of cooling a rotor assembly is also disclosed.

Abstract: A gas turbine engine including a core nacelle defined about an engine axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A gear train is defined along an engine axis. The gear train defines a gear reduction ratio of greater than or equal to about 2.3. A spool along the engine axis drives the gear train. The spool includes a downstream turbine having six or fewer stages. A fan is driven through the gear train by the downstream turbine. A pressure ratio across the fan is less than about 1.45. A fan variable area nozzle is axially movable relative to the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.