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 gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to about 8 at cruise power. A gear arrangement is configured to drive the fan section. A compressor section includes both a first compressor section and a second compressor section. A turbine section is configured to drive the gear arrangement, and may have a low pressure turbine with four stages and a low pressure turbine pressure ratio greater than about 5:1, and a high pressure turbine with two stages. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor section and a pressure ratio across the second compressor section, and greater than about 40, measured at sea level and at a static, full-rated takeoff power.

Abstract: A gas turbine engine includes a main compressor. A tap is fluidly connected downstream of the main compressor. A heat exchanger is fluidly connected downstream of the tap. An auxiliary compressor unit is fluidly connected downstream of the heat exchanger. The auxiliary compressor unit is configured to compress air cooled by the heat exchanger with an overall auxiliary compressor unit pressure ratio between 1.1 and 6.0. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A 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. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.

Abstract: A gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row.

Abstract: A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing.

Abstract: A turbine injection system for a gas turbine engine includes a first end operable to receive air from a heat exchanger, a second end operable to distribute mixed cooling air to a turbine stage, an opening downstream of said first end and a mixing plenum downstream of said first end and said opening. The opening provides a direct fluid pathway into said turbine injection system.

Abstract: A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a high spool, and a low spool including a low pressure turbine that drives the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.

Abstract: A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk at an interface, where the interface is defined along a spoke. A spool for a gas turbine engine includes the rotor disk, the plurality of blades with the interface defined along the spoke radially inboard of a blade platform, a rotor ring axially adjacent to the rotor disk, and a plurality of core gas path seals which extend from the rotor ring. Each of the plurality of core gas path seals extends from the rotor ring at a seal interface, with the seal interface defined along a spoke and the plurality of core gas path seals being axially adjacent to the blade platform.

Abstract: A gas turbine engine includes a plurality of rotatable components housed within a main compressor section and a turbine section. A cooling system is connected to tap air from said main compressor section. A first tap is connected to a first heat exchanger. The first heat exchanger is connected to a cooling compressor for raising a pressure of the tapped air downstream of the first heat exchanger. A second heat exchanger is downstream of the cooling compressor, and a connection is downstream of the second heat exchanger for delivering air to a bearing compartment. A connection intermediate the cooling compressor and the second heat exchanger delivers cooling air to at least one of the rotatable components.

Abstract: A gas turbine engine comprises a main compressor section having a downstream most end, 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 cooling air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are 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 ratio of an outer diameter of a fan hub at a leading edge of the blades to an outer tip diameter of the blades at the leading edge is greater than or equal to about 0.24 and less than or equal to about 0.38. The fan tip diameter is greater than or equal to about 84 inches (213.36 centimeters) and a fan tip speed is less than or equal to about 1050 ft/second (320.04 meters/second). The fan drive turbine has between three and six stages.

Abstract: A disclosed bleed air system utilizes high pressure air from a high pressure compressor to drive the turbo compressor to increase a pressure of bleed air drawn from the low pressure compressor. Air drawn from the low pressure compressor is at a lower temperature and pressure than that encountered from the high pressure compressor. The turbo compressor increases the pressure of airflow and provides that airflow into the main bleed air passage to be communicated to systems utilizing the bleed air.

Abstract: A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section.

Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing.

Abstract: A gas turbine engine includes a main compressor section. A booster compressor includes an inlet and an outlet. The inlet receives airflow from the main compressor section and the outlet provides airflow to a pneumatic system. A recirculation passage is between the inlet and the outlet. A flow splitter valve controls airflow between the outlet and the inlet through the recirculation passage for controlling airflow to the pneumatic system based on airflow output from the booster compressor. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

Abstract: A gas turbine engine comprises an engine having a compressor section, and a turbine section. A firewall and accessory pumps are mounted on a downstream side of the firewall. The accessory pumps are driven by electric motors mounted on the firewall on an upstream side of the firewall.