The lift curve of the airfoil of an aircraft is as shown. The weight of the aircraft is 1500 kg. If the stall speed of the aircraft is 60 m/s, determine the wing area.

To determine the wing area of the aircraft, we need to use the lift equation. The lift equation states that the lift force acting on an aircraft is equal to the coefficient of lift (CL) multiplied by the dynamic pressure of the air (q) and the wing area (A).

Lift (L) = CL * q * A

First, let's find the lift force using the weight of the aircraft. The lift force must be equal to the weight of the aircraft for level flight.

Weight (W) = Mass (m) * gravity (g)

Given that the weight of the aircraft is 1500 kg, and assuming the acceleration due to gravity (g) is approximately 9.8 m/s^2:

W = 1500 kg * 9.8 m/s^2 = 14700 N

Next, we need to determine the dynamic pressure (q). Dynamic pressure is given by the formula:

q = 0.5 * ρ * V^2

Where:
ρ (rho) is the air density
V is the velocity of the aircraft

To calculate the dynamic pressure, we need to know the air density, which is not provided in the given information. However, we can estimate the air density to be approximately 1.2 kg/m^3 at sea level.

Now we can calculate the dynamic pressure at the stall speed (V) of 60 m/s:

q = 0.5 * 1.2 kg/m^3 * (60 m/s)^2 = 2160 Pa

Finally, rearranging the lift equation, we can solve for the wing area (A):

A = L / (CL * q)

Substituting the known values:

A = 14700 N / (CL * 2160 Pa)

Since the lift coefficient (CL) is not provided in the given information, we cannot directly calculate the wing area without it. The lift coefficient depends on the specific airfoil design and the angle of attack.