16. Consider a finite wing with aspect ratio A-8 and elliptical planform. The aerofoil section is thin and symmetric. Calculate the lift and induced drag coefficients when angle of attack is five degrees.
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- For the NACA 2412 airfoil, the lift coefficient and moment coefficient about the quarter-chord at -6° angle of attack are -0.39 and -0.045, respectively. At 4° angle of attack, these coefficients are 0.65 and -0.037, respectively. Calculate the location of the aerodynamic center. -6. For a wing of aspect ratio AR, having an elliptical lift distribution, the induced drag coefficient is (where CL is the lift coefficient) (b) TAR CL 2TAR (d) TAR2 (a) TAR 7. Laminar flow airfoil is used to reduce the (a) pressure drag (b) induced drag (c) skin friction drag (d) wave drag 8. Set the aerofoil your chosen angle of incidence is gives a good installed pressure distribution. (a) twelve degree (b) ten degree (c) eight degree 9. The critical Mach number for a thick airfoil will be (a) lesser than a thin airfoil. (b) greater than a thin airfoil. (c) equal to a thin airfoil. (d) cannot be related to thin airfoil 10. What is the dimension for Lift coefficient? (a) N/s (b) kg/N (c) DimensionlessA wing has a planform area S of 200 ft? and a total span b of 40 feet. The airfoils are symmetric all along the span. The airfoil has a 2-D lift curve slope of 27 per radian. The wing has a rectangular planform, and thus has zero taper. The wing is untwisted. a. Compute the lift coefficient C and the drag coefficient Coi at an angle of attack of 4 degrees. Use two terms in the series expansion for circulation. T= 2bV,[4, sin ø + A, sin 3ø] b. Repeat the above calculation, now with just one term T=2bVA1sino. Compare the lift drag coefficient C and Cp values to problem #2 above. c. Compare the results for drag coefficient from part (b) above with that for an elliptically loaded wing at this lift coefficient.
- Consider a thin, symmetric airfoil at 1.5◦ angle of attack. From the resultsof thin airfoil theory, calculate the lift coefficient and the momentcoefficient about the leading edge.For a 10 deg included angle wedge at 0 deg AOA, calculate the lift and drag coefficient (Cl = 0, Cd = 0.082). Then calculate the lift and drag at AOA=4 deg (Cl = 0.16, Cd = 0.093). Let M=2.Consider a finite wing with an aspect ratio of 3. Assume an elliptical liftdistribution. The lift slope for the airfoil section is 0.1/degree. Calculateand compare the lift slopes for (a) a straight wing, and (b) a swept wing,with a half-chord line sweep of 45 degrees.
- Calculate the lift and moment coefficients using thin airfoil theory approximations for a range of angle of attacks for the following airfoils below. 1. NACA 0008 2. NACA 0018 3. AG04 4. Clark-Y 5. NACA 2415 Angle of attacks taken to be between 0 degree to 15 degree. Also calculate value of Cm at quarter chord. (Cm at 0.25c) for each. Reynolds number 1000000Problem 06.025 - Calculate maximum velocity The Predator UAV has the following characteristics: wingspan = 14.85 m, wing area = 11.45 m², maximum weight = 1110 kgf, and fuel weight = 295 kgf. The power plant is a Rotax four- cylinder, four-stroke engine of 100 horsepower driving a two- blade, variable-pitch pusher propeller. Assume: the Oswald efficiency factor is 0.7; the zero-lift drag coefficient is 0.03; the propeller efficiency is 0.9; and the specific fuel consumption is 0.2 kgf of fuel per horsepower per hour. Calculate the maximum velocity of the Predator at sea level. The maximum velocity of the Predator UAV at sea level is m/s.- Study how the lift and drag varies with a change in angle of attack for a flat plate, symmetrical airfoil and asymmetrical airfoil when Re is below 50'000 and compare it to higher Re.
- Estimate the landing ground roll distance at sea level for the CJ-1. No thrust reversal is used; however, spoilers are employed such that L drag coefficient by 10 percent. The fuel tanks are essentially empty, so neglect the weight of any fuel caried by the airplane. The maximum lift coefficient, with flaps fully em- ployed at touchdown, is 2.5. 0. The spoilers increase the zero-lift,An untwisted wing uses the same airfoil along the span and the lift distribution can be considered elliptical. Data: mo = 6.1/rad b = 30 ft; S = 150ft²; a = 4°; aoL = -2.5° Find: CL(a); Cpi; CL; and Induced angle of attack at y = 5 ftConsider a wing with a linear AR of 8.0 and a taper of 0.5, the wing has a wingspan of 196 ft and a thin profile, a CL max of 1.6, estimate the maximum lift coefficient of the wing and the position in the chord of the profile when it enters in loss. Consider an arrow angle (1/4 of the chord) of 30 °