Multi-valve modulated core ventilation

1 . A gas turbine engine, comprising:

a fan section;

a compressor section;

a combustor section;

a turbine section;

an outer casing disposed radially outside of the compressor section, the combustor section, and the turbine section with respect to an axial centerline of the gas turbine engine;

a core gas path that extends through the compressor section, the combustor section, and the turbine section to a core gas path exhaust, the core gas path disposed radially inside of the outer casing;

a fan bypass air duct extending between and defined by an outer radial flow path boundary and an inner radial flow path boundary, the outer radial flow path boundary comprising one or more inner radial panels of a nacelle, the gas turbine engine disposed in the nacelle, the inner radial flow path boundary including one or more inner frame structure panels;

an outside core annular region, the outside core annular region defined by and radially disposed between the outer casing and the inner frame structure panels of the inner radial flow path boundary, wherein the outside core annular region extends to the core gas path exhaust;

an aft fan annular compartment, the aft fan annular compartment defined by a portion of the outer casing, an annular aft inner diameter panel, and a firewall, the firewall engaged with the outer casing and the annular aft inner diameter panel, the firewall extending in a radial plane between the outer casing and the annular inner diameter panel, the annular aft inner diameter panel radially outside of and spaced apart from the outer casing, and the annular aft inner diameter panel forming a part of the inner radial flow path boundary; and

a fan air circulation system, the fan circulation system comprising a plurality of inlet ports, the inlet ports disposed along a circumference of the annular aft inner diameter panel, the fan air circulation system further comprising a plurality of valves and a plurality of exit ports, the exit ports disposed along a circumference of the firewall, wherein the fan air circulation system is configured such that a respective said inlet port is in fluid communication with a respective valve, and said respective valve is in fluid communication with at least one respective said exit port, and each respective exit port is in fluid communication with the outside core annular region;

wherein each respective valve is selectively controllable to control a flow of fan bypass air through the respective valve, into the outside core annular region at the firewall, and thereafter out of the engine from the outside core annular region,

wherein each respective exit port is configured to direct the flow of fan bypass air at an angle relative to the axial centerline to create a swirling airflow within the outside core annular region; and

wherein a valve of the plurality of valves is in fluid communication with more than one of the plurality of exit ports.

2 . The gas turbine engine of claim 1 , wherein each respective valve is selectively controllable to at least a first operating configuration and a second operating configuration, wherein the first operating configuration is associated with a first flow of fan bypass air through the respective valve and the second operating configuration is associated with a second flow of fan bypass air through the respective valve, and the first flow of fan bypass air through the respective valve is different than the second flow of fan bypass air through the respective valve.

3 . The gas turbine engine of claim 1 , wherein each respective valve is selectively controllable to at least a first operating configuration and a second operating configuration, wherein the first operating configuration is a closed valve configuration, and the second operating configuration is an open valve configuration.

4 . The gas turbine engine of claim 2 , wherein the first operating configuration is a first open valve configuration and the second operating configuration is a second open valve configuration, and wherein the first flow of fan bypass air through the respective valve is volumetrically greater than the second flow or a substantially zero flow of fan bypass air through the respective valve.

5 . The gas turbine engine of claim 1 , wherein each respective valve is selectively controllable to a plurality of open operating configurations, including a first open operating configuration and a second open operating configuration, and wherein the flow of fan bypass air through the respective valve in the first open operating configuration is volumetrically greater than the flow of fan bypass air through the respective valve in the second open operating configuration.

6 . The gas turbine engine of claim 1 , wherein the exit ports are spaced equidistantly from one another.

7 . The gas turbine engine of claim 1 , wherein the exit ports are non-uniformly circumferentially spaced.

8 . The gas turbine engine of claim 1 , wherein each of the inlet ports is configured the same as the other said inlet ports.

9 . The gas turbine engine of claim 1 , wherein the firewall separates the aft fan annular compartment from the outside core annular region.

10 . A method of cooling one or more external components of a gas turbine engine, the gas turbine engine including a fan section, a compressor section, a combustor section, a turbine section, an outer casing disposed radially outside of the compressor section, the combustor section, and the turbine section with respect to an axial centerline of the gas turbine engine, a core gas path that extends through the compressor section, the combustor section, and the turbine section to a core gas path exhaust, the core gas path disposed radially inside of the outer casing; a fan bypass air duct extending between and defined by an outer radial flow path boundary and an inner radial flow path boundary, the outer radial flow path boundary comprising one or more inner radial panels of a nacelle, the gas turbine engine disposed in the nacelle, the inner radial flow path boundary including one or more inner frame structure panels; an outside core annular region, the outside core annular region defined by and radially disposed between the outer casing and the inner frame structure panels of the inner radial flow path boundary, wherein the outside core annular region extends to the core gas path exhaust; an aft fan annular compartment, the aft fan annular compartment defined by a portion of the outer casing, an annular aft inner diameter panel, and a firewall, the firewall engaged with the outer casing and the annular aft inner diameter panel, the firewall extending in a radial plane between the outer casing and the annular inner diameter panel, the annular aft inner diameter panel radially outside of and spaced apart from the outer casing, the annular aft inner diameter panel forming a part of the inner radial flow path boundary, and the annular aft inner diameter panel comprising a plurality of inlet ports, and the firewall comprising a plurality of exit ports, the method comprising:

using a plurality of valves to selectively control a flow of fan bypass air received from the fan bypass air duct into the outside core annular region at the firewall and to exit the gas turbine engine from the outside core annular region, wherein each respective valve is in fluid communication with a respective inlet port among the plurality of inlet ports, wherein each respective valve is in fluid communication with at least one respective exit port, wherein each respective exit port is in fluid communication with the outside core annular region, wherein each said exit port is configured to direct the flow of fan bypass air at an angle relative to the axial centerline to create a swirling airflow within the outside core annular region, and wherein each respective valve is controllable to vary the flow of fan bypass air through the respective valve by moving from a first operating configuration to a second operating configuration;

wherein a valve of the plurality of valves is in fluid communication with more than one of the plurality of exit ports.

11 . The method of claim 10 , wherein the flow of fan bypass air through a respective valve in the first operating configuration is volumetrically greater than the flow of fan bypass air through the respective valve in the second operating configuration.

12 . The method of claim 10 , wherein the step of using the plurality of valves to selectively control the flow of fan bypass air includes controlling the plurality of valves in unison to cause each said valve of the plurality of valves to be in the same operating configuration.

13 . A gas turbine engine, comprising:

a fan section;

a compressor section;

a combustor section;

a turbine section;

an outer casing disposed radially outside of the compressor section, the combustor section, and the turbine section with respect to an axial centerline of the gas turbine engine;

a core gas path that extends through the compressor section, the combustor section, and the turbine section to a core gas path exhaust, the core gas path disposed radially inside of the outer casing;

a fan bypass air duct extending between and defined by an outer radial flow path boundary and an inner radial flow path boundary, the outer radial flow path boundary comprising one or more inner radial panels of a nacelle, the gas turbine engine disposed in the nacelle, the inner radial flow path boundary including one or more inner frame structure panels;

an outside core annular region, the outside core annular region defined by and radially disposed between the outer casing and the inner frame structure panels of the inner radial flow path boundary, wherein the outside core annular region extends to the core gas path exhaust;

an aft fan annular compartment, the aft fan annular compartment defined by a portion of the outer casing, an annular aft inner diameter panel, and a firewall, the firewall engaged with the outer casing and the annular aft inner diameter panel, the firewall extending in a radial plane between the outer casing and the annular inner diameter panel, the annular aft inner diameter panel radially outside of and spaced apart from the outer casing, and the annular aft inner diameter panel forming a part of the inner radial flow path boundary; and

a fan air circulation system, the fan air circulation system including a plurality of inlet ports disposed in the annular aft inner diameter panel and spaced apart from each other along a circumference of the annular aft inner diameter panel, a plurality of valves, and a plurality of exit ports, the exit ports disposed along a circumference of the firewall, wherein each said exit port is in fluid communication with the outside core annular region and the fan air circulation system is configured to receive a flow of fan bypass air from the fan bypass air duct and using a respective valve of the plurality of valves to selectively pass the received flow of fan bypass air into the outside core annular region at the firewall and thereafter exhaust from the gas turbine engine from the outside core annular region, and wherein each said exit port is configured to direct the flow of fan bypass air at an angle relative to the axial centerline to create a swirling airflow within the outside core annular region; and

a system controller in communication with the plurality of valves and a non-transitory memory storing instructions, which instructions when executed cause the system controller to:

selectively control each respective valve of the plurality of valves to control the flow of fan bypass air between a respective inlet port of the plurality of inlet ports and at least one respective exit port of the plurality of exit ports and thereafter into the outside core annular region, the selective control including controlling each respective valve to pass a first said flow of fan bypass air during a first engine operating condition, and to pass a second said flow of fan bypass air during a second engine operating condition, wherein the first said flow of fan bypass air is volumetrically greater than the second said flow of fan bypass air;

wherein a valve of the plurality of valves is in fluid communication with more than one of the plurality of exit ports.

14 . The gas turbine engine of claim 13 , wherein the first engine operating condition is a take-off condition and the second engine operating condition is a cruise condition.

15 . The gas turbine engine of claim 13 , wherein the selective control of the flow of fan bypass air into the outside core annular region is based on a thermal environment within the outside core annular region.