A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: $\frac{\mathrm{bypass}\mathrm{exhaust}\mathrm{nozzle}\mathrm{pressure}\mathrm{ratio}}{\mathrm{core}\mathrm{exhaust}\mathrm{nozzle}\mathrm{pressure}\mathrm{ratio}}$
is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;

wherein a bypass to core ratio of:

[00001]$\frac{\mathrm{bypass}\mathrm{exhaust}\mathrm{nozzle}\mathrm{pressure}\mathrm{ratio}}{\mathrm{core}\mathrm{exhaust}\mathrm{nozzle}\mathrm{pressure}\mathrm{ratio}}$

is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

An apparatus for the control of fluid flow in a gas turbine engine comprises a first plurality of inlet guide vanes disposed upstream of a fan, a compressor, a combustor, and a turbine; at least one airflow splitter adapted to split air admitted through the first plurality of inlet guide vanes into a core airflow which flows through the fan, the compressor, the combustor, and the turbine and a bypass airflow which flows through the fan; wherein the first plurality of inlet guide vanes comprise a radially-inward first portion adapted to direct air admitted through the first plurality of inlet guide vanes to the core airflow and a radially-outward second portion adapted to direct air admitted through the first plurality of inlet guide vanes to the bypass airflow, and wherein the first portion comprises a fixed vane and the second portion comprises a variable vane.

Methods are provided for creating a fluidic barrier between the core stream and the bypass stream in a turbofan engine. A method comprises compressing the bypass and core streams with a fan between an upstream splitter and a downstream splitter which divides the bypass and core streams, and imparting a first momentum into the air stream proximate the fan in a region between the core and bypass streams and the upstream and downstream splitters to form a fluid barrier, wherein the first momentum of the air stream in the region is higher than a second momentum of the air stream adjacent the fluid barrier.

A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of high pressure fluid jets originating from the low pressure side of the fan blades, the jets restricting the migration of the core stream into the bypass stream through the fluid passage.

A gas turbine engine is provided having a compressor section, a combustion section located downstream of the compressor section, and a turbine section located downstream of the combustion section. A structural member extends from the compressor section to the turbine section for strengthening one or more components of the gas turbine engine. The structural member also defines a flowpath extending between an inlet in airflow communication with the compressor section and an outlet in airflow communication with the turbine section. The flowpath is configured to provide bleed air from the compressor section to the turbine section to capture at least a portion of the energy in such bleed air.

A tip turbine engine provides increased efficiency while eliminating or reducing the number of axial compressor stages by moving the core airflow inlet aft of the fan. As a result, the core airflow entering the core airflow inlet is the fan exhaust, which is already at an increased pressure. A portion of the fan exhaust is guided radially inward, then axially forward and then radially outward through compressor chambers in the hollow fan blades for further, centrifugal compression.