An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions.

An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions.

A turbine engine including a fan, a nacelle circumscribing at least the fan, a compressor section downstream of the fan, and a conduit defined, at least in part, by the nacelle. The conduit includes a first opening at the compressor section, a second opening downstream of the fan and upstream of the compressor section, and a third opening upstream of the fan. Pressure sensors coupled to the nacelle are communicatively coupled to at least one actuator. The at least one actuator can adjust airflow between the first opening and the second opening, or between the first opening and the third opening. The pressure sensors can provide outputs for generating commands that control the at least one actuator.

A gas turbine engine includes a propulsor. A speed reduction device is drivingly connected to the propulsor. A compressor section with a high pressure compressor has between 8 and 13 stages and a pressure ratio of at least 16:1 and less than 35:1. A turbine section including a transition duct is located between a high pressure turbine and a low pressure turbine including fewer support struts than vanes in a first vane row of the low pressure turbine. The first vane row of the low pressure turbine is located downstream of the transition duct and downstream of a first row of blades in the low pressure turbine. The first row of blades in the low pressure turbine are immediately downstream of the support struts.

A gas turbine engine includes a propulsor. A speed reduction device is drivingly connected to the propulsor. A compressor section with a high pressure compressor has between 8 and 13 stages and a pressure ratio of at least 16:1 and less than 35:1. A turbine section including a transition duct is located between a high pressure turbine and a low pressure turbine including fewer support struts than vanes in a first vane row of the low pressure turbine. The first vane row of the low pressure turbine is located downstream of the transition duct and downstream of a first row of blades in the low pressure turbine. The first row of blades in the low pressure turbine are immediately downstream of the support struts.

A gas turbine engine has a fan rotor including at least one stage, with the at least one stage delivering a portion of air into a low pressure duct, and another portion of air into a compressor. The compressor is driven by a turbine rotor, and the fan rotor is driven by a fan drive turbine. A channel selectively communicates air from the low pressure duct across a boost compressor.

A gas turbine engine has a fan rotor including at least one stage, with the at least one stage delivering a portion of air into a low pressure duct, and another portion of air into a compressor. The compressor is driven by a turbine rotor, and the fan rotor is driven by a fan drive turbine. A channel selectively communicates air from the low pressure duct across a boost compressor.

An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions.