Question

Consider a flying wing aircraft made using a NACA 2412 airfoil with a wing area of 250 ft2, a wing span of 50 ft, and a span efficiency factor of 0.9. The aircraft is flying at a 6° angle of attack and a Reynolds number of approximately 9 x 106. (a) Determine 2D (airfoil) and 3D (wing) lift curves and plot them together. (b) Discuss stall APA for the lift curves in (a), noting that wing stall AOA is only an estimate. (c) Determine the drag polar and plot drag coefficient vs lift coefficient. (c) If the aircraft is flying at sea level at v. = 280 ft/s, what are lift and drag values? Problem 2. (Individual Submission) An F-15 is flying straight and level (lift equals weight) in 25,000 ft standard day conditions at a velocity of 850 ft/s. Given the following data for the F-15, answer the questions below: parasite drag coefficient (Coo) = 0.018 wing span (b) = 42.8 ft wing area (S) = 610 ft? weight (W) = 50,000 lb Oswald efficiency factor (e.) = 0.915 a. What is the wing's aspect ratio (AR)? b. What is the lift force (L) generated by the aircraft? C. What is the aircraft's drag coefficient (Co)? d. What is the drag force (D) generated by the aircraft? Problem 3. (Individual Submission) An SR-71 is traveling at Mach 2.5 in standard day conditions at an altitude of 10 km. (a) What is the Mach cone angle of the airplane? (b) What is its velocity in m/s? (c) If the airplane descends to 5 km altitude while maintaining the same velocity as in (b), what would be its Mach cone angle?

Answer 1

To determine the speed at which the air must move over the upper surface of the wing to create the ideal lift, we can use Bernoulli's principle and the lift equation. The same principle and equation can be used to calculate the air speed over the upper surface of the wing at a cruising speed and altitude with a different air density.

Step-by-step explanation:

To find out how fast the air must move over the upper surface of the wing to create the ideal lift, we can use Bernoulli's principle.

Bernoulli's principle states that the sum of the pressure and kinetic energy per unit volume of a fluid is constant along a streamline.

The ideal lift can be given by the equation:

Lift = 0.5 * density * velocity^2 * area * lift coefficient

Given that the lift coefficient is 1000 N/m^2 and the wing area is 250 ft^2, we can calculate the air speed:

1000 N/m^2 = 0.5 * 1.29 kg/m^3 * (60.0 m/s + v)^2 * (250 ft^2 * 0.092903 m^2/ft^2) * lift coefficient

where v is the air speed over the upper surface of the wing.

(b) Calculation of Cruising Speed:

Using the same equation, we can calculate the air speed over the upper surface of the wing at a cruising speed of 245 m/s and at an altitude where air density is one-fourth that at sea level:

1000 N/m^2 = 0.5 * (1.29 kg/m^3 /4) * (245 m/s + v)^2 * (250 ft^2 * 0.092903 m^2/ft^2) * lift coefficient