Abstract: Systems, apparatus, articles of manufacture, and methods to provide for cross-diffuser bleed are provided herein. An example gas turbine engine includes a frame defining a cavity in a forward side of a diffuser of the gas turbine engine; a compressor including the diffuser, the diffuser defining a primary flow path to provide air flow to a combustor and including at least one conduit, the at least one conduit fluidly coupled to the cavity; and a downstream sink fluidly coupled to the cavity via the at least one conduit defining at least a portion of a bleed air path.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: Systems, apparatus, articles of manufacture, and methods to provide for cross-diffuser bleed are provided herein. An example gas turbine engine includes a frame defining a cavity in a forward side of a diffuser of the gas turbine engine; a compressor including the diffuser, the diffuser defining a primary flowpath to provide air flow to a combustor and including at least one conduit, the at least one conduit fluidly coupled to the cavity; and a downstream sink fluidly coupled to the cavity via the at least one conduit defining at least a portion of a bleed air path.

Abstract: Systems, apparatus, articles of manufacture, and methods to provide for cross-diffuser bleed are provided herein. An example gas turbine engine includes a frame defining a cavity in a forward side of a diffuser of the gas turbine engine; a compressor including the diffuser, the diffuser defining a primary flow path to provide air flow to a combustor and including at least one conduit, the at least one conduit fluidly coupled to the cavity; and a downstream sink fluidly coupled to the cavity via the at least one conduit defining at least a portion of a bleed air path.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T O ? U ? T A D ? T ? E ? x ? i ? t ) 2 * EGT A HPCExit * 10 - 11 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine includes a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a controller configured to autonomously regulate performance of the gas turbine engine in response to at least one of: a thrust demand, an energy efficiency target, or a flight profile condition.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 10 - 11 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine has a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a nozzle vane and an engine casing defining a fluid supply plenum. The engine casing is configured to route cooling fluid from the fluid supply plenum to the nozzle vane. The routing of the cooling fluid facilitates cooling of the nozzle vane and increases an overall pressure ratio of the gas turbine engine. Integration of the cooling fluid routing within the engine casing reduces parasitic bleed air losses, enables tighter turbine stage spacing, and improves the thermal efficiency of the engine, thereby enhancing performance at elevated pressure ratios.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: (TOUT/?{square root over (ADTExit)})2*EGT/AHPCExit*10?11; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 10 - 11 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T OUT A DTExit ) 2 * EGT A HPCExit * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.14 and less than 0.0003294×EGT2?0.306×EGT+77.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T O ? U ? T A D ? T ? E ? x ? i ? t ) 2 * E ? G ? T A H ? P ? C ? E ? x ? i ? t * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.

Abstract: A stator structure including a plurality of stator blades and redirection structures including a first portion and a second portion, the first portion disposed on a front edge surface of a stator hub and a second portion disposed on a facing of the stator hub is provided. The stator hub includes the facing and the front edge surface, the facing being disposed generally perpendicular to the casing, and the front edge surface is disposed generally perpendicular to the facing. During operation of a turbine engine a core air flow moves along the longitudinal axis and past the plurality of stator blades, and a leakage air flow moves in a direction different to the core air flow, the redirection structures are effective to redirect the leakage air flow to merge into the core air flow.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T O ? U ? T A D ? T ? E ? x ? i ? t ) 2 * E ? G ? T A H ? P ? C ? E ? x ? i ? t * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.

Abstract: A turbomachine for an aircraft is provided. The turbomachine includes a plurality of radially-extending blades and an annular endwall opposite the radially-extending blades. The endwall includes an endwall treatment recessed into the endwall. The endwall treatment is characterized by a casing treatment volume compressibility factor (CTVCF), and a casing treatment normalized volume (CTNV), and a blade tip Mach number (Mtip).

Abstract: A turbomachine for an aircraft is provided. The turbomachine includes a plurality of radially-extending blades and an annular endwall opposite the radially-extending blades. The endwall includes an endwall treatment recessed into the endwall. The endwall treatment is characterized by a casing treatment volume compressibility factor (CTVCF), and a casing treatment normalized volume (CTNV), and a blade tip Mach number (Mtip).

Abstract: A stator structure including a plurality of stator blades and redirection structures including a first portion and a second portion, the first portion disposed on a front edge surface of a stator hub and a second portion disposed on a facing of the stator hub is provided. The stator hub includes the facing and the front edge surface, the facing being disposed generally perpendicular to the casing, and the front edge surface is disposed generally perpendicular to the facing. During operation of a turbine engine a core air flow moves along the longitudinal axis and past the plurality of stator blades, and a leakage air flow moves in a direction different to the core air flow, the redirection structures are effective to redirect the leakage air flow to merge into the core air flow.