You must have Credits on your Balance to download this sample
Principles of Aerodynamics
Finance & Accounting
Pages 8 (2008 words)
Task 1: Question No. 1: Area of the aircraft = S = 29 m2 Coefficient of Lift = CL = 0.67 Altitude = A = 29000 ft Velocity = V? = 385 km/hr = 106.94 m/s Lift = L = ? Solution: Density at given altitude ‘A’ = A = 0.4671 kg/m3 (Anderson Introduction 763) Lift on an airfoil is given as L = (?
2: Coefficient of Drag = CD = 0.054 Area = S = 15 m2 Thrust = T= 1500 N Density= A = 0.5 kg/m3 For a steady and level flight, drag force is equal to the thrust produced by engines, D = T = 1500 N D = (? A V?2) S CD = (? (0.5) (V?2)) (15) (0.054) = 1500 => V? = 86.06 m/s = 8.6 E +1 m/s Question No. 3: Question No. 4: The sketches shown below illustrate the trend of variation in CL, CD, and L/D ratios with increasing angle of attack. Question No. 5: Critical Mach number corresponds to that value of Mach number for free stream flow for which a localized mach number of ‘1’ is obtained at any point around the airfoil. When this condition arises, a shock wave is created at the point where the flow reaches the sonic speed. As the speed increases, regions of very low pressure are created. This causes the flow to separate from the airfoil thereby substantially increasing the drag forces on it. The figure illustrates this phenomenon. Some of the important design features incorporated in the aircrafts in order to contain the effects of this situation are using thin airfoil and / or super critical airfoil (Anderson Introduction 763). Making an airfoil thinner increases the value of Critical Mach Number and hence the airplane can fly at very high speeds without a significant increase in drag forces on it. ...
Not exactly what you need?