A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: $\frac{\begin{array}{c}\mathrm{the}\mathrm{turbine}\mathrm{diameter}\mathrm{at}\mathrm{an}\mathrm{axial}\mathrm{location}\\ \mathrm{of}\mathrm{the}\mathrm{lowest}\mathrm{pressure}\mathrm{rotor}\mathrm{stage}\end{array}}{a\mathrm{distance}f\mathrm{rom}}$

An apparatus and method for assembling an inducer assembly for inducing a rotation on an airflow passing within a turbine engine. The inducer assembly can provide a volume of fluid from a compressor section to a turbine section of the engine. The inducer assembly can include the combination of separate segments to form an annular inducer.

An apparatus and method for assembling an inducer assembly for inducing a rotation on an airflow passing within a turbine engine. The inducer assembly can provide a volume of fluid from a compressor section to a turbine section of the engine. The inducer assembly can include the combination of separate segments to form an annular inducer.

An airfoil for a turbine engine includes an array of features positioned in an interior portion of the airfoil. Each feature extends from a pressure side to a suction side. The array includes multiple radial rows (A-N) of features with the features in each row (A-N) being interspaced radially to define coolant passages therebetween. The radial rows (A-N) are spaced along a forward-to-aft direction toward an airfoil trailing edge. The coolant passages of the array are fluidically interconnected to lead a pressurized coolant toward the trailing edge via a serial impingement on to the rows of features. The coolant passages are geometrically configured to bias a coolant flow therethrough toward a first side in relation to a second side of the outer wall to effect a greater cooling of the first side than the second side.

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: $\frac{\begin{array}{c}\mathrm{the}\mathrm{turbine}\mathrm{diameter}\mathrm{at}\mathrm{an}\mathrm{axial}\\ \mathrm{location}\mathrm{of}\mathrm{the}\mathrm{lowest}\mathrm{pressure}\mathrm{rotor}\mathrm{stage}\end{array}}{a\mathrm{distance}\mathrm{from}a\mathrm{ground}}$

A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is:

[00001]$\frac{\begin{array}{c}\mathrm{the}.Math..Math.\mathrm{turbine}.Math..Math.\mathrm{diameter}.Math..Math.\mathrm{at}.Math..Math.\mathrm{an}.Math..Math.\mathrm{axial}.Math..Math.\mathrm{location}\\ \mathrm{of}.Math..Math.\mathrm{the}.Math..Math.\mathrm{lowest}.Math..Math.\mathrm{pressure}.Math..Math.\mathrm{rotor}.Math..Math.\mathrm{stage}\end{array}}{\mathrm{ground}.Math..Math.\mathrm{plane}.Math..Math.\mathrm{to}.Math..Math.\mathrm{wing}.Math..Math.\mathrm{distance}}$

and a quasi-non-dimensional mass flow rate Q defined as:

[00002]$Q=W.Math.\frac{\sqrt{T.Math.0}}{P.Math..Math.0.Math.A_{\mathrm{flow}}}$

where: W is mass flow rate through the fan in Kg/s; T0 is average stagnati

In accordance with one aspect of the disclosure, a stream diverter for a gas turbine engine is disclosed. The stream diverter may include a first air duct, a second air duct, a third air duct, and a door operatively associated with the second and third air ducts of the gas turbine engine. The door may have at least an open position allowing air from the second air duct to flow into the third air duct and a closed position preventing air from flowing between the ducts.

An apparatus and method for assembling an inducer assembly for inducing a rotation on an airflow passing within a turbine engine. The inducer assembly can provide a volume of fluid from a compressor section to a turbine section of the engine. The inducer assembly can include the combination of separate segments to form an annular inducer.

An airflow deflector is provided that includes at least one liner attachment flange, a deflector panel, and at least one deflector leg. The at least one liner attachment flange has a liner side surface and an opposing outer surface. The at least one deflector leg extends between and is attached to the at least one liner attachment flange. The at least one deflector leg extends a distance between the deflector panel and the liner attachment flange to maintain the deflector panel a separation distance from the liner attachment flange. The airflow deflector is configured for insertion of the at least one deflector leg and the deflector panel within a liner aperture.

An integrated fuel cell and engine combustor assembly includes an engine combustor having a combustion chamber fluidly coupled with a compressor and a turbine. The assembly also includes a fuel cell stack circumferentially extending around the combustion chamber of the combustor. The fuel cell stack includes fuel cells configured to generate electric current. The fuel cell stack is positioned to receive discharged air from the compressor and fuel from a fuel manifold. The fuel cells in the fuel cell stack generate electric current using the discharged air and at least some of the fuel. The fuel cell stack is positioned to radially direct partially oxidized fuel from the fuel cells into the combustion chamber of the combustor. The combustor combusts the partially oxidized fuel into one or more gaseous combustion products that are directed into and drive the downstream turbine.