Gas turbine engine

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 including:
• 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 blades of the plurality of blades have a characteristic selected from the group consisting of:
• i) a blade camber at 30% span is at least 90% of a maximum blade camber between 50% span and 100% span, ii) the blade camber at 0% span is at least 110% of the maximum blade camber between 50% span and 100% span, and iii) combinations thereof.