Say we have an airfoil with leading edge (LE) and trailing edge (TE) located at 0 and 1.84 m, respectively. The pressure coefficient distribution in the upper surface of the airfoil is trapezoidal, with -0.32 constant for all values of between and 63% of the chord, and then ranging linearly from -0.32 at 63% of the chord to 0 at . On the other hand, the pressure coefficient distribution in the lower surface fits well a cosenoidal function with period equal to 0.57 times the chord, and amplitude 0.67, crossing the positive part of the ordinate axis in the leading edge with null slope. Please, compute the value of the airfoil’s global lift coefficient (with 3 decimals; and please, use the comma "," as decimal separator).