A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having 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 (A.sub.HPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn.sub.Total) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn.sub.Total?EGT/(A.sub.HPCExit.sup.2?1000).

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

A turbine engine assembly is provided. The assembly includes a low-pressure turbine assembly including a first turbine section configured to rotate in a first rotational direction at a first rotational speed, and a second turbine section configured to rotate in a second rotational direction at a second rotational speed. The second rotational direction is opposite the first rotational direction and the second rotational speed is lower than the first rotational speed. The assembly also includes a first drive shaft coupled to the first turbine section, and a fan assembly including a first fan section coupled to the first drive shaft such that the first fan section rotates in the first rotational direction at the first rotational speed, and a second fan section coupled to the second turbine section such that the second fan section rotates in the second rotational direction at the second rotational speed.

A turbine engine assembly is provided. The assembly includes a low-pressure turbine assembly including a first turbine section configured to rotate in a first rotational direction at a first rotational speed, and a second turbine section configured to rotate in a second rotational direction at a second rotational speed. The second rotational direction is opposite the first rotational direction and the second rotational speed is lower than the first rotational speed. The assembly also includes a first drive shaft coupled to the first turbine section, and a fan assembly including a first fan section coupled to the first drive shaft such that the first fan section rotates in the first rotational direction at the first rotational speed, and a second fan section coupled to the second turbine section such that the second fan section rotates in the second rotational direction at the second rotational speed.

The invention relates to an aircraft comprising a fuselage (1), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan (7) and a downstream fan (8), driven by two contra-rotating rotors (5, 6) of a power turbine (3). The two fans (7, 8) and the turbine (3) are integrated into a nacelle (14) which projects downstream from the fuselage (1) and through which air flows. According to the invention, at least one of the fans (7, 8) of the aircraft and, in particular, the downstream fan (8) comprises variable-spacing blades, and at least one stator-forming variable-spacing blade ring (25) in the aircraft is placed upstream of the upstream fan (7). The variable-spacing stator blades (25) and the variable-spacing blades of the downstream fan (8) are mutually configured to direct the air flow in a first mode in which the air flows through the nacelle (14) from upstream to downstream and in a second mode in which the air is pushed back upstream through the nacelle (14).

The invention relates to an aircraft comprising a fuselage (1), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan (7) and a downstream fan (8), driven by two contra-rotating rotors (5, 6) of a power turbine (3). The two fans (7, 8) and the turbine (3) are integrated into a nacelle (14) which projects downstream from the fuselage (1) and through which air flows. According to the invention, at least one of the fans (7, 8) of the aircraft and, in particular, the downstream fan (8) comprises variable-spacing blades, and at least one stator-forming variable-spacing blade ring (25) in the aircraft is placed upstream of the upstream fan (7). The variable-spacing stator blades (25) and the variable-spacing blades of the downstream fan (8) are mutually configured to direct the air flow in a first mode in which the air flows through the nacelle (14) from upstream to downstream and in a second mode in which the air is pushed back upstream through the nacelle (14).

An after-fan system for an engine may comprise an after-fan turbine an electrical generator operationally coupled to the after-fan turbine, and an electric motor electrically coupled to the electrical generator. The electrical generator may be configured to generate an electrical current in response to rotation of the after-fan turbine. The electric motor may be configured to generate torque.