The measured lift slope for the NACA 23012 airfoil is 0.1080 degree
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- Consider the NACA 2412 airfoil discussed . The airfoil is flying at a velocity of 60 m/s at a standard altitude of 3 km . The chord length of the airfoil is 2 m. Calculate the lift per unit span when the angle of attack is 4◦.arrow_forwardHelp me pleasearrow_forwardThe test results of a NACA 2518 shows: Angle of attack 0 deg 8 deg CL 0.045 0.75 If the airfoil is used to construct a wing with a span of 60 ft and a wing area of 450 square feet. Determine the wing lift curve slope. (e = 0.88)arrow_forward
- Consider an airfoil at 12◦ angle of attack. The normal and axial forcecoefficients are 1.2 and 0.03, respectively. Calculate the lift and dragcoefficients.arrow_forwardCompute the lift and drag coefficients for a symmetric, diamond-shaped airfoil with a thickness-to-chord ratio t/c equal to 0.10, flying at Mach 3.5 in air (y=1.4) at zero angle of attack. M₂ - 3.5arrow_forwardFor 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.arrow_forward
- The car shown in the figure below moves at a constant speed on a highway and has a drag coefficient Cpc of 0.32 with the windows and roof closed. What is the percent increase of horsepower needed to maintain the speed if the windows and roof are then opened? With the windows and roof open, the drag coefficient increases to Cpo = 0.43. Assume the frontal area remains the same. Windows and roof closed: CD=CDc Windows open; roof open: C₂=CDoarrow_forwardApproximate the value of the lift coefficient corresponding to the stalling angle of attack. Use the lift curve corresponding to NACA 2421 with a Reynolds Number of 6x10^6 with standard roughness.arrow_forwardA flagpole 16 m high has the shape of a cylinder 100 mm in diameter. The air temperature is 30°C and the atmospheric pressure is 101 KPaa. With what speed is the air blowing against the pole if the moment developed at the base is 2.7 KN.m? The drag coefficient is 1.3.arrow_forward
- 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 1000000arrow_forwardThe airfoil section of the wing of the British Spitfire of World War II fame is an NACA 2213 at the wing root, tapering to an NACA 2205 at the wing tip. The root chord is 8.33 ft. The measured profile drag coefficient of the NACA 2213 airfoil is 0.006 at a Reynolds number of 9 × 106. Consider the Spitfire cruising at an altitude of 19000 ft. Assume that μ varies as the square root of temperature. At what velocity is it flying for the root chord Reynolds number to be 9 × 106? (Round the final answer to the nearest whole number.) The velocity at which the spitfire is flying for the root chord Reynolds number to be 9 × 106 is ..........ft/s.arrow_forwardA 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.arrow_forward
- Principles of Heat Transfer (Activate Learning wi...Mechanical EngineeringISBN:9781305387102Author:Kreith, Frank; Manglik, Raj M.Publisher:Cengage Learning
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