Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Textbook Question
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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Students have asked these similar questions
If the pressure difference in the wind tunnel experiment is 50 mmH,0, the density of air is 1.2 kg/m in wind
tunnel experiment, and the airfoil span is 204 mm, and the cord length is 98 mm, what is the lift coefficient if
the lift force is L=13.4 N
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.)
4. 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;
= ?
da
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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- 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).arrow_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_forwardA fl at-plate airfoil with C = 1.2 m is to have a lift of30 kN/m when fl ying at 5000-m standard altitude withU∞= 641 m/s. Using Ackeret theory, estimate (a) the angleof attack and (b) the drag force in N/m.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_forwardThe 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_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
- a)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_forwardTheory of Flight A- Airbus 320 fly in steady level flight with altitude 11000m, find the lift force required to steady level flight. At steady level flight aircraft has weight as (65000 kg), wing area is (124 m?), aircraft speed is (147 m/s) and lift coefficient is (1.3) Atmosphere data for 11000m shown in table below Туре of atmosphere layer Stratosphere Rang (m) Temperature gradient (a) k/m Temperature (k) 0-11000 -0.0065 288-216.6 Given data (R=287, g=9.81 m/s?, and po=1.225 kg/m³)arrow_forward
- An airfoil generates a 1200 N/m sectional lift when traveling at 75 m/s at sea level. What is the circulation generated by this airfoil?arrow_forwardThe drag coefficient for a car with a frontal area of 26 ft? is being measured in an 8 ft x 8 ft wind tunnel. The density of air under the test conditions is 2.4 x 10 3 slugs/ft, When the air flow rate is, 500,000 ft/min, the drag force on the car was measured to be 170 lb. The drag coefficient under the test conditions is most nearly O 0.30 O 0.32 O 0.28 O 0.34arrow_forwardA 5,000 lb airplane with wing area of 250 sq. ft and power available of 450 hp is at sea level flight standard condition. The airplane drag-to-lift relation is close to Co = 0.025 + 0.048 CL². What is the rate of climb (in feet per minute) at a flight speed of 180 miles per hour?arrow_forward
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