Gas turbine engine having a bypass passage, a fan duct, and a core duct

A gas turbine engine comprising:

• a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the 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 defining 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, and
• wherein the primary fan includes:
• a plurality of blades, each of the plurality of blades comprising:
• a blade root proximal to an axis of rotation;
• a blade tip remote from the axis;
• and
• a blade span measured between the blade root and the blade tip,
• wherein a circumferential averaged tangential velocity, Cu, and a radius of airflow, R, and a loading of a respective blade of the plurality of blades at R is defined as a change in RCu, ΔRCu, over a row of blades in the plurality of blades, and
• wherein the primary fan has a load distribution such that, at any location between the blade root of a respective blade and 30% span, a value of ΔRCu is greater than or equal to 60% of a peak ΔRCu.