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 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: 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 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: A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil includes a first portion near the airfoil base with a first density and includes a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction.

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: A flow-management system may comprise a center body impermeable to air. A conical surface of the center body may face forward. A blocking surface of the center body may be coaxial with the conical surface and may comprise an annular recess. An annular ring may be aft of the center body and fluidly coupled with the blocking surface. A tube may encase the center body and annular ring. The annular ring may comprise an air-foil shape to direct a pulse to the blocking surface. The blocking surface may comprise a central peak and a circular ridge separated by the annular recess.

Abstract: A process for manufacturing a turbine engine component includes the steps of: providing a powder containing gamma titanium aluminide; and forming a turbine engine component from said powder using a direct metal laser sintering technique.

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 gas turbine engine includes a turbine section fluidly connected to a combustor by a primary flow path. The turbine section includes a first portion at a high pressure relative to a second portion. A thermally isolated cooling plenum is positioned radially inward of the primary flow path. The cooling plenum is defined by a forward wall, a shaft structure, an aft wall, and an inner diameter wall of the primary flow path. Air in the thermally isolated cooling plenum is thermally isolated from air in the primary flow path. At least one cooling air system is operable to provide cooling air to the thermally isolated cooling plenum.

Abstract: A flow-management system may comprise a center body impermeable to air. A conical surface of the center body may face forward. A blocking surface of the center body may be coaxial with the conical surface and may comprise an annular recess. An annular ring may be aft of the center body and fluidly coupled with the blocking surface. A tube may encase the center body and annular ring. The annular ring may comprise an air-foil shape to direct a pulse to the blocking surface. The blocking surface may comprise a central peak and a circular ridge separated by the annular recess.

Abstract: A turbine injection system for a gas turbine engine includes a first end operable to receive air from a heat exchanger, a second end operable to distribute mixed cooling air to a turbine stage, an opening downstream of said first end and a mixing plenum downstream of said first end and said opening. The opening provides a direct fluid pathway into said turbine injection system.

Abstract: A seal segment according to an exemplary aspect of the present disclosure includes, among other things, a first axial wall, a second axial wall radially spaced from the first axial wall and a radially outer wall that interconnects the first axial wall and the second axial wall. At least one curved member is radially inwardly offset from the radially outer wall and extending between the first and second axial walls.

Abstract: A turbine casing may comprise a casing body a heat pipe disposed in the casing body. The heat pipe may include a vaporization section and a condensation section. The vaporization section may be located forward the condensation section. The vaporization section may be located in a high pressure turbine region of the casing body. The condensation section may be located in a low pressure turbine region of the casing body.

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: An engine includes an elongated pulse detonation combustor tube having an arcuate combustion path over a majority of an entire length of the combustor tube, and an elongated portion of the combustor tube being oriented transverse to a central axis of the engine.

Abstract: A gas turbine engine comprises a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface defining a diffuser chamber radially outwardly of the combustor, and a cooling air chamber wall positioned outwardly of the diffuser chamber and the combustor, and radially inwardly of a second wall to define a cooling air chamber. The turbine section includes a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap taps air having been compressed by the compressor and is passed through a heat exchanger. Air downstream of the heat exchanger passes into the cooling chamber, and then to the blade outer air seal.

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: 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 rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk, adjacent ones of the plurality of blades are joined by a flexible web.