Abstract: Air mixing systems for gas turbine engines include a heat exchanger, a first extraction conduit fluidly coupled to an inlet of the heat exchanger, a second extraction conduit fluidly coupled to an outlet of the heat exchanger, an injection conduit fluidly coupled to the second extraction conduit, an onboard injector supply chamber fluidly coupled to the injection conduit, and an onboard injector fluidly coupled to the onboard injector supply chamber.

Abstract: A gas turbine engine according to the present disclosure includes a first compressor and a first turbine for driving the first compressor. A core section includes a second compressor and a second turbine for driving the second compressor. A third turbine is arranged fluidly downstream of the first turbine and the second turbine and configured to drive a power take-off. A first duct system is arranged fluidly between the low-pressure compressor and the core section. The first duct system is arranged to reverse fluid flow before entry into the core section.

Abstract: A recuperated gas turbine engine includes an engine core that has a compressor section, a combustor section, and a turbine section. An exhaust duct is located downstream of the turbine section for receiving a hot turbine exhaust stream from the turbine section. The exhaust duct includes a heat exchanger and a temperature-control module upstream of the heat exchanger. A compressor bleed line leads from the compressor section into the heat exchanger and a compressor return line leads from the heat exchanger into the engine core upstream of the combustor section. The compressor bleed line is operable to selectively feed compressed air to the heat exchanger, and the temperature-control module is operable to selectively modulate at least one of temperature and flow of the hot turbine exhaust stream with respect to the heat exchanger.

Abstract: An inlet manifold for a multi-tube pulse detonation engine includes a vaneless diffuser disposed in a first zone to collect an air discharged from a compressor; a vaned diffuser including a plurality of guide vanes disposed in a second zone to slow the air from the compressor; a plenum disposed in a third zone located next to second zone to provide the air from the compressor to chambers; and a splitter disposed in a fourth zone to split the air from the compressor into an airflow required by each pulse detonation tube for detonation.

Abstract: A system includes a stator assembly including at least one stator airfoil. The system also includes a rotor assembly including at least one rotor airfoil configured to rotate about an axis. The system also includes an actuator coupled to the stator assembly and configured to actuate the stator assembly in an axial direction relative to the rotor assembly, creating an axial movement such that a clearance between the at least one rotor airfoil and the stator assembly varies based on an axial position of the stator assembly.

Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine. The engine includes a first rotor disk, a second rotor disk, and a circumferentially segmented seal. The segmented seal engages the first rotor disk and the second rotor disk. The segmented seal further includes a fore surface contacting the first disk, an aft surface contacting the second disk, and a radially outer surface. Further, (1) the aft surface and (2) one of the fore surface and the radially outer surface include perforations to allow fluid to flow through the interior of the segmented seal.

Abstract: An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.

Abstract: A mounting system for a gas turbine engine includes a compressor case portion, an inlet frame, an outlet frame, and a mounting structure. The compressor case portion houses rotatable compressor blades. The inlet frame connects to an inlet end of the compressor case. The outlet frame connects to an outlet end of the compressor case portion at an end opposite the compressor case inlet end. An axially fore mounting structure of the mounting structure connects to the inlet frame. An axially aft mounting structure of the mounting structure connects to the outlet frame. A bridging structure of the mounting structure is offset from the compressor case and connects the fore and aft mounting structures, thereby bridging engine loads across the inlet and outlet frames to reduce load induced distortion of the compressor case portion.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first compressor having a first overall pressure ratio, and a second compressor having a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. Further, a section of the gas turbine engine includes a thermally isolated area.

Abstract: A section of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a thermally isolated area, and a first rotor disk and a second rotor disk. Each of the first and second rotor disks are provided within the thermally isolated area.

Abstract: The present disclosure provides systems and methods related to thermal management systems for gas turbine engines. For example, a thermal management system comprises a thermally neutral heat transfer fluid circuit, a first heat exchanger disposed on the fluid circuit, and a second heat exchanger disposed on the fluid circuit.

Abstract: A cooling system is provided. The cooling system may comprise a heat exchanger and a first conduit fluidly coupled to an outlet of the heat exchanger. An annular passage may be fluidly coupled to the first conduit. A tangential onboard injector (TOBI) may be fluidly coupled to the annular passage. A gas turbine engine is also provided and may comprise a compressor, a combustor in fluid communication with the compressor, and a diffuser around the combustor and a turbine. A heat exchanger may have an inlet fluidly coupled to the diffuser. A second conduit may be fluidly coupled to an outlet of the heat exchanger. An annular passage may be fluidly coupled to the second conduit. A tangential onboard injector (TOBI) may be fluidly coupled to the annular passage.

Abstract: Components include a low pressure turbine having a plurality of rotor assemblies including a first gamma TiAl intermetallic blade having a maximum operating temperature over 1180° F. (638° C.). At least two of the rotor assemblies include gamma TiAl intermetallic alloy blades. In an embodiment, a method of making a turbine having a plurality of rotor assemblies includes attaching a first gamma TiAl intermetallic alloy blade to an upstream stage of the plurality of rotor assemblies.

Abstract: A gas turbine engine may comprise a first rotor with a primary flowpath along an outer diameter of the first rotor. A secondary flowpath may be radially inward from the primary flowpath. The secondary flowpath may pass through an opening through the first rotor. A blade may be disposed on a distal end of the first rotor. The blade may extend into the primary flowpath. A bleed tube may be in a wall of the primary flowpath and forward of the blade. The bleed tube may extend radially inward from the primary flowpath. The bleed tube may fluidly connect to the opening through the first rotor. A plenum may be aft of the blade and radially inward from the primary flowpath. The plenum may be fluidly connected to the opening through the first rotor. A second rotor may be aft of the plenum.

Abstract: A hybrid component for a turbine engine having a casing includes a first part of a gamma TiAl intermetallic alloy and a second part of a material of at least one of nickel, a nickel base, a cobalt base, an iron base superalloy or mixtures thereof. The second part is coupled to and configured to attach the first part to the casing of the engine. The first and second parts are attached to each other by transient liquid phase (TLP) bonding.

Abstract: A turbine section for a gas turbine engine includes a first rotor assembly with a first rotor assembly bleed air source and an aft cavity that is in fluid communication with the first rotor assembly bleed air source. A second rotor assembly includes a forward cavity. A vane bleed air source is in fluid communication with the forward cavity. A seal extends between the first rotor assembly and the second rotor assembly.

Abstract: A gas turbine engine comprises a lower pressure compressor and a higher pressure compressor. A single turbine drives both the lower pressure compressor and the higher pressure compressor through a gear reduction. The gear reduction includes an actuator and at least two available speeds, such that the lower pressure compressor can selectively be operated at either of at least two speeds relative to the higher pressure compressor. A method of operating a gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a propulsor with a power turbine, a power turbine shaft extending forward therefrom defining a centerline axis, and a fan driven by the power turbine shaft. The fan is aligned with the centerline axis forward of the power turbine and is operatively connected to be driven by the power turbine through the power turbine shaft. A gas generator operatively connected to the propulsor is included downstream from the fan and forward of the power turbine, wherein the gas generator defines a generator axis offset from the centerline axis. The gas generator is operatively connected to the power turbine to supply combustion products for driving the power turbine.

Abstract: A pulse detonation combustor may include a valve and a tubular combustor wall, which forms an airflow inlet and a combustion chamber. The valve may be configured to selectively fluidly couple the airflow inlet with the combustion chamber. The valve may include a center body and an annular projection. The center body and the projection may be configured to sealingly engage with one another.

Abstract: A turbine section includes a rotor assembly which includes an internal cooling passage. A segmented seal is adjacent the rotor assembly and includes a fluid passage that is in fluid communication with the internal cooling passage.