A propulsion system is provided including an unducted rotating fan defining a fan axis; and a turbomachine disposed downstream from the unducted rotating fan, wherein the turbomachine defines a working gas flowpath flowing therethrough; wherein the propulsion system defines a third stream flowpath and an inlet passage having an inlet that is offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flowpath, and wherein the third stream flowpath bypasses at least a portion of the turbomachine.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R.sub.1 and a primary fan hub radius R.sub.2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R.sub.3 and a secondary fan hub radius R.sub.4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to %Fn.sub.3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ratio of R.sub.1 to R.sub.3 equals

[00001]$\sqrt{\left(EFP\right)\frac{\left(1-Rq{R}_{sec.-Fan}{}^2\right)}{\left(1-Rq{R}_{Prim.-Fan}{}^2\right)}\left(\frac{1}{\%F{n}_{3s}}-1\right)};$

wherein EFP is between 1.5 and 11, wherein RqRp.sub.rim.-Fan is a ratio of R.sub.1 to R.sub.2, and wherein RqR.sub.Sec.-Fan is a ratio of R.sub.3 to R.sub.4.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R.sub.1 and a primary fan hub radius R.sub.2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R.sub.3 and a secondary fan hub radius R.sub.4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to %Fn.sub.3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ratio of R.sub.1 to R.sub.3 equals

[00001]$\sqrt{\left(EFP\right)\frac{\left(1-Rq{R}_{sec.-Fan}{}^2\right)}{\left(1-Rq{R}_{Prim.-Fan}{}^2\right)}\left(\frac{1}{\%F{n}_{3s}}-1\right)};$

wherein EFP is between 1.5 and 11, wherein RqRp.sub.rim.-Fan is a ratio of R.sub.1 to R.sub.2, and wherein RqR.sub.Sec.-Fan is a ratio of R.sub.3 to R.sub.4.

A gas turbine engine which includes an outer casing; a central longitudinal hollow shaft with a forward air inlet; a three stage rotating superconducting electric bypass fan with front and rear fan blades and a diffuser blade interposed between said front and rear fan blades wherein the diffuser blade rotates in an opposite direction to the front and rear fan blades; a multiple stage superconducting axial compressor positioned aft of the three stage rotating superconducting electric bypass fan; a multiple stage superconducting electric turbine core positioned aft of the multiple stage variable speed superconducting axial compressor, whereby the electric power from the multiple stage superconducting electric turbine core powers the three stage superconducting electric bypass fan and the multiple stage superconducting axial compressor.

A three-stream gas turbine engine with an embedded electric machine and methods of operating the same are disclosed. In one aspect, a three-stream engine includes an electric machine operatively coupled with a shaft of the engine. The three-stream engine also includes a core engine and a primary fan and a mid-fan positioned upstream of the core engine. The primary fan and the mid-fan are operatively coupled with the shaft. During operation, the three-stream engine defines a tip speed ratio being defined by a tip speed of a rotor of the electric machine to a tip speed of a mid-fan blade of the mid-fan. The tip speed ratio is defined as being equal to or greater than 0.2 and less than or equal to 1.0.

A three-stream gas turbine engine with an embedded electric machine and methods of operating the same are disclosed. In one aspect, a three-stream engine includes an electric machine operatively coupled with a shaft of the engine. The three-stream engine also includes a core engine and a primary fan and a mid-fan positioned upstream of the core engine. The primary fan and the mid-fan are operatively coupled with the shaft. During operation, the three-stream engine defines a tip speed ratio being defined by a tip speed of a rotor of the electric machine to a tip speed of a mid-fan blade of the mid-fan. The tip speed ratio is defined as being equal to or greater than 0.2 and less than or equal to 1.0.

A disclosed gas turbine engine includes a first fan section including a plurality of fan blades rotatable about an axis, a compressor in fluid communication with the first fan section, a combustor in fluid communication with the compressor and a first turbine section in fluid communication with the combustor. The first turbine section includes a low pressure turbine that drives the first fan section. A second fan section is supported between the first fan section and the compressor and is driven by a second turbine section disposed between the second fan section and the compressor for driving the second fan section.

A disclosed gas turbine engine includes a first fan section including a plurality of fan blades rotatable about an axis, a compressor in fluid communication with the first fan section, a combustor in fluid communication with the compressor and a first turbine section in fluid communication with the combustor. The first turbine section includes a low pressure turbine that drives the first fan section. A second fan section is supported between the first fan section and the compressor and is driven by a second turbine section disposed between the second fan section and the compressor for driving the second fan section.

A gas turbine engine has a fan rotor including at least one stage, with the at least one stage delivering a portion of air into a low pressure duct, and another portion of air into a compressor. The compressor is driven by a turbine rotor, and the fan rotor is driven by a fan drive turbine. A channel selectively communicates air from the low pressure duct across a boost compressor.

A gas turbine engine has a fan rotor including at least one stage, with the at least one stage delivering a portion of air into a low pressure duct, and another portion of air into a compressor. The compressor is driven by a turbine rotor, and the fan rotor is driven by a fan drive turbine. A channel selectively communicates air from the low pressure duct across a boost compressor.