A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is: the ⁢ ⁢ turbine ⁢ ⁢ diameter ⁢ ⁢ at ⁢ ⁢ an ⁢ ⁢ axial location ⁢ ⁢ of ⁢ ⁢ the ⁢ ⁢ lowest ⁢ ⁢ pressure ⁢ ⁢ rotor ⁢ ⁢ stage ⁢ ground ⁢ ⁢ plane ⁢ ⁢ to ⁢ ⁢ wing ⁢ ⁢ distance and a quasi-non-dimensional mass flow rate Q defined as: Q = W ⁢ T ⁢ ⁢ 0 P ⁢ ⁢ 0 · A flow where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; and Aflow is the flow area of the fan face in m2, and wherein a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 to 0.01.