Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 2, Problem 2.2P
Consider an airfoil in a wind tunnel (i.e., a wing that spans the entire test section). Prove that the lift per unit span can be obtained from the pressure distributions on the top and bottom walls of the wind tunnel (i.e., from the pressure distributions on the wails above and below the airfoil).
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Consider an airfoil in a wind tunnel (i.e., a wing that spans the entire testsection). Prove that the lift per unit span can be obtained from the pressuredistributions on the top and bottom walls of the wind tunnel (i.e., from thepressure distributions on the walls above and below the airfoil).
please answer do not image format
Chapter 2 Solutions
Fundamentals of Aerodynamics
Ch. 2 - Consider a body of arbitrary shape. If the...Ch. 2 - Consider an airfoil in a wind tunnel (i.e., a wing...Ch. 2 - Consider a velocity field where the x and y...Ch. 2 - Consider a velocity field where the x and y...Ch. 2 - Consider a velocity field where the radial and...Ch. 2 - Consider a velocity field where the x and y...Ch. 2 - The velocity field given in Problem 2.3 is called...Ch. 2 - The velocity field given in Problem 2.4 is called...Ch. 2 - Is the flow field given in Problem 2.5...Ch. 2 - Consider a flow field in polar coordinates, where...
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- A light combat aircraft weighs 75,000 N and has a wing area of 27 m?. The maximum lift coefficient with high-lift devices is 1.8, and the structural limit load factor is 6.0. While flying at 250 km/h, the aircraft makes a 90 deg turn in 8 s at sea level holding a constant altitude and at an angle of attack such that the lift-to-drag ratio is 8.0. Find (a) the bank angle, (b) load factor, (c) radius of turn, and (d) the thrust required. [Answer: (a) µ = 54.26 deg, (b) n = 1.7120, (c) R %3D 353.7665 m, and (d) T = 16,050 N.)arrow_forward4arrow_forward4. Write down the lift curve equation (lift coefficient versus angle of attack) according to the thin airfoil theory. What is the amount of the lift curve slope for a thin airfoil ? NACA 2412 airfoil Eg. (4,57) C 2.0 - 1.e 1.2 Lift coefficient 0.8 0.4 -0.4 Eg. (4,64) -0.1 Moment coefficient -0.8 -0.2 -1.2 -0.3 Re - 3.1 X 10 O Re- 8.9 x 10 -0.4 16 a, degrees de; = ? daarrow_forward
- The lift curve for the 4 digit NACA 2421 airfoil is shown in the Figure. -Consider a wing with AR = 12, no sweep, and airplane flies at the Mach number equal to 0.65. The Oswald efficiency span number ?1 = 0.95 Calculate the lift slope for the finite wing. Clearly show the formula and explanations in the solution.arrow_forwardAnswer this olearrow_forwardA closed-loop wind tunnel has a test section of cross-section W =3 ft by H =5 ft and length L =4 m. An aerofoil with 21% thickness to chord ratio and a chord length of 0.52 m is mounted vertically in the test section and spans the entire height. If the measured lift and drag coefficients are C; = 0.325 and Ca =0.0071, what is the lift coefficient after correcting for solid blockage, wake blockage and streamline curvature? Give your answer to 3 decimal places.arrow_forward
- A smooth aluminum sphere hangs in air on a 0.5 m-long cable. The sphere has a diameter of 10 cm and weighs 1.5 N (it is hollow). It is observed that a steady wind causes the sphere and cable to be displaced so that = 20°. Assume that the weight of the cable and the drag force on the cable are negligible. (a) What is the magnitude of the (horizontal) drag force on the sphere? (b) What is the wind speed U? Answer: (a) D≈ 0.55 N, (b) U≈ 15 m/s Uarrow_forwardPlease show all workarrow_forward6. Write down the lift curve equation (lift coefficient versus angle of attack) according to the thin airfoil theory. What is the amount of the lift curve slope for a thin airfoil ? NACA 2412 airfoil LEg. (4.57) * 2.0 1.6 1.2 Lift coefficient 0.8 0.4 -0.4 Eg. (4,64) --0.1 Moment coefficient -0.8- -1.2 -0.3 • Re -3.1 X 10 O Re -8.9 X 10 L-0.4 -8 16 24 a, degrees dc =? daarrow_forward
- 1arrow_forwarda)What is the impact of increasing Reynolds number on skin friction and pressure drags over an airfoil? What can be happened for separation in this case? b) What is an adverse pressure gradient and where does it occur on an airfoil (show that on a sketch)? c) Why lift-to-drag ratio is an important parameter for an aircraft? d)How can changing in altitude affect the aircraft power required, PR? Show thatmathematically and graphically?arrow_forwardAn aircraft is equipped with a wing of symmetrical airfoils. The lift curve slope of the total aircraft is estimated to be 0CL/Oa = 5.0265 1/rad. The stall angle of attack (AOA) is 12°. Wing area is 16 m2. Use g = g0 and r = rO. What is the aircraft's mass during a flight on which a stall speed of 50 kt was observed?arrow_forward
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