An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

A cooling system for a gas turbine engine may comprise a plenum extending circumferentially around an outer engine case structure. The plenum may comprise a supply conduit and a return conduit. The supply conduit and the return conduit may be in fluid communication with a heat exchanger. The heat exchanger may be disposed between the outer engine case structure and an inner engine case structure. The plenum may be configured to provide enhance heat transfer for the cooling system.

A thermal management system includes a first heat source assembly including a first heat source exchanger, a first thermal fluid inlet line extending to the first heat source exchanger, and a first thermal fluid outlet line extending from the first heat source exchanger; a second heat source assembly including a second heat source exchanger, a second thermal fluid inlet line extending to the second heat source exchanger, and second a thermal fluid outlet line extending from the second heat source exchanger; a shared assembly including a thermal fluid line and a heat sink exchanger, the shared assembly defining an upstream junction in fluid communication with the first thermal fluid outlet line and second thermal fluid outlet line and a downstream junction in fluid communication with the first thermal fluid inlet line and second thermal fluid inlet line; and a controller configured to selectively fluidly connect the first heat source assembly or the second heat source assembly to the shared assembly.

Aircraft propulsion systems including a closed loop-supercritical fluid system having a turbine, a cooler heat exchanger, a compressor, and a recovery heat exchanger arranged along a closed-loop flow path of a supercritical fluid. A shaft is operably coupled to the turbine and configured to be rotationally driven by the turbine. A fan is configured to generate thrust, the fan operably coupled to the shaft to be rotationally driven by the shaft. A burner is configured to combust a fuel and air from the fan to generate a combusted gas and supply said combusted gas to the recovery heat exchanger of the closed loop-supercritical fluid system and out an exhaust nozzle.

Aircraft propulsion systems including a closed loop-supercritical fluid system having a turbine, a cooler heat exchanger, a compressor, and a recovery heat exchanger arranged along a closed-loop flow path of a supercritical fluid. A shaft is operably coupled to the turbine and configured to be rotationally driven by the turbine. A fan is configured to generate thrust, the fan operably coupled to the shaft to be rotationally driven by the shaft. A burner is configured to combust a fuel and air from the fan to generate a combusted gas and supply said combusted gas to the recovery heat exchanger of the closed loop-supercritical fluid system and out an exhaust nozzle.

A system for modulating air flow in a gas turbine engine is provided. The system may include a seal wall comprising an opening, a seal door configured to slideably engage the seal wall, and an actuator configured to move the seal door over the opening. In various embodiments, the system may include a surface forward of the seal door. The seal door may be configured to seal a passage through the surface and the opening of the seal wall. A track may be disposed under the seal door. The track may comprise cobalt. Rollers may be coupled to the seal door with the rollers on the track. The seal door may comprise a nickel-chromium alloy. A sync ring may be coupled to the seal door. The actuator may be coupled through the sync ring to the seal door.

A system for modulating air flow in a gas turbine engine is provided. The system may include a seal wall comprising an opening, a seal door configured to slideably engage the seal wall, and an actuator configured to move the seal door over the opening. In various embodiments, the system may include a surface forward of the seal door. The seal door may be configured to seal a passage through the surface and the opening of the seal wall. A track may be disposed under the seal door. The track may comprise cobalt. Rollers may be coupled to the seal door with the rollers on the track. The seal door may comprise a nickel-chromium alloy. A sync ring may be coupled to the seal door. The actuator may be coupled through the sync ring to the seal door.

A heat exchanger (HEX) for cooling air in a gas turbine engine is provided. The HEX may comprise a central manifold comprising an inlet portion, a first outlet portion, and a second outlet portion. The HEX may further comprise a plurality of tubes coupled to the central manifold, the plurality of tubes comprising at least a first tube, a second tube, a third tube, and a fourth tube, a shroud at least partially encasing said plurality of tubes, and a cooling air flow path defined by at least one of the shroud, the plurality of tubes, and an outer surface of the central manifold, wherein the cooling air flow path is orthogonal to said plurality of tubes.

A heat exchanger (HEX) for cooling air in a gas turbine engine is provided. The HEX may comprise a central manifold comprising an inlet portion, a first outlet portion, and a second outlet portion. The HEX may further comprise a plurality of tubes coupled to the central manifold, the plurality of tubes comprising at least a first tube, a second tube, a third tube, and a fourth tube, a shroud at least partially encasing said plurality of tubes, and a cooling air flow path defined by at least one of the shroud, the plurality of tubes, and an outer surface of the central manifold, wherein the cooling air flow path is orthogonal to said plurality of tubes.

An assembly is provided for a turbine engine. This assembly includes a primary duct, a bleed duct, a plurality of secondary ducts, a heat exchanger and a flow regulator. The bleed duct extends from a bleed duct inlet to a bleed duct outlet. The bleed duct inlet is fluidly coupled with the primary duct. The secondary ducts are arranged in parallel between the bleed duct outlet and the primary duct. The secondary ducts include a first duct and a second duct. The heat exchanger is configured with the second duct. The flow regulator is configured to direct at least a majority of fluid flowing through the bleed duct outlet to: (A) the first duct during a first mode of operation; and (B) the second duct during a second mode of operation.