Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Textbook Question
Chapter 4, Problem 4.7P
For the airfoil given in Problem 4.6, calculate
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= 3570
T AC
=144
Example:
1.5 ft
It is desired to determine the drag force at a given
speed on a prototype sailboat hull. A model is
placed in a test channel and three cables are used
to align its bow on the channel centerline. For a
given speed, the tension is 40 lb in cable AB and
60 lb in cable AE.
Flow
4 ft
Determine the drag force exerted on the hull and
the tension in cable AC.
Choc
ug the h-
as the free hod-
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Question 2: Consider the 5000kg aircraft below being in a 1000m radius looping maneuver (Fig.2).
There are 4 forces acting on the aircraft (Weight, Thrust, Lift, Drag). Assuming that they are all
approximately perpendicular to eachother at points A,B,C, calculate the following:
A. Calculate the Lift Forces at points A, B, C if the aircraft was at a constant speed of 350km/h
throughout the maneuver.
B. Calculate the Lift Forces at points A, B, C if the aircraft starts the maneuver at point A at a
speed of 400km/h, but reduces its speed with constant acceleration so that the pilot feels
weightless at point C.
C. A human can withstand
earth's gravity was more than 9 times what it is now, we would not survive. What is the
maximum of 9 times the gravitational acceleration, i.e. if the
maximum speed the pilot can enter this maneuver?
Figure 2: Looping Aircraft
An airfoil has a lift curve slope at 7/rad and angle at zero lift of -3deg. At what angle of attack will the airfoil develop a lift of 120 lbs/ft at 110 mph under SSLC? Assume c = 10ft.
Chapter 4 Solutions
Fundamentals of Aerodynamics
Ch. 4 - Consider the data for the NACA 2412 airfoil given...Ch. 4 - Consider an NACA 2412 airfoil with a 2-m chord in...Ch. 4 - Starting with the definition of circulation,...Ch. 4 - Starting with Equation (4.35), derive Equation...Ch. 4 - Consider a thin, symmetric airfoil at 1.5 angle of...Ch. 4 - The NACA 4412 airfoil has a mean camber line given...Ch. 4 - For the airfoil given in Problem 4.6, calculate...Ch. 4 - Compare the results of Problems 4.6 and 4.7 with...Ch. 4 - Starting with Equations (4.35) and (4.43), derive...Ch. 4 - For the NACA 2412 airfoil, the lift coefficient...
Ch. 4 - Consider again the NACA 2412 airfoil discussed in...Ch. 4 - For the airfoil in Problem 4.11, calculate the...Ch. 4 - In Section 3.15 we studied the case of the lifting...Ch. 4 - The question is often asked: Can an airfoil fly...Ch. 4 - The airfoil section of the wing of the British...Ch. 4 - For the conditions given in Problem 4.15, a more...
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- Problem 4 – Consider the isentropic flow over an airfoil. The freestream conditions correspond to a standard altitude of 10,000 ft and M, = 0.82. At a given point on the airfoil, M = 1.0. Calculate P and T at this point.arrow_forwardB2 is flying at an altitude of 43,000 ft and a speed of 400 KTAS under standard atmospheric conditions. If the average aerodynamic chord is 39,6 ft, calculate the Reynolds number and the average % aerodynamic chord present on the wingarrow_forwardConsider a standard NACA Five digit aerofoil with h = 0.29 and k_1 = 6.643; x_h = 0.2 Find the zero lift Angle of Anglearrow_forward
- Aerodynamic Model: - Circular Disk Inside Dia.= 40 mm Outside Dia.= 80 mm Angle of Attack= 15 Degree Lift Force= 0.38 N Drag Force=0.40 N Wind Velocity= 20 m/sec Air Temperature= 30 oC (Note: Take Air Density at 30 oC) Required: - Find Lift and Drag Coefficient at 15 Degree Attack Angle.arrow_forward3. The airfoil has an angle of attack of 4 deg. What is the lift curve slope if the angle of attack measured from zero lift is -2 deg., lift/span is 1000 N/m with a speed of 200 m/s at 12,000 m. Assume c= 3 m.arrow_forwardFor the thin aerofoil of mean camber line given by: z = f (1+x/b)(1−x/b) with f the maximum camber, find the lift coefficient, the zero lift AOA, the moment coefficient about the leading edge, the moment coefficient about the aerodynamic centre and the location of the aerodynamic centre Consider a thin symmetric aerofoil at 2° angle of attack. Calculate the lift and the moment about the leading edge. Consider an NACA 2412 aerofoil with a 2-m chord in an air- stream with a velocity of 50m/s at standard sea level conditions. If the lift per unit span is 1353 N, what is the AOA?arrow_forward
- For Gulfstream IV at the conditions stated. Calculate the minimum thrust required and the velocity at which it occurs. The data given are W=73000 lb, S=950 ft2, ρ=8.9068*104 slug/ft3 , Cd,o=0.015, k=0.08 This is a basic Flight Mechanics Question.arrow_forwardThe Drag Polar Equation of an General Aviation Aircraft in clean configuration can be written as: C2 Cp = 0.0270+ %3D 32.787 Its weight is 2300 lb and the wing area is 155ft?, Calculate the (L/D)max and the Minimum Thrust Required (Unaccelerated Level flight and Sea level Condition)arrow_forwardChoose a specific airfoil for an airplane and determine the lift and drag forces after that calculate the power required to overcome drag. You must determine the following: The cruises speed and altitude. The aspect ratio.arrow_forward
- 9. The thin ( 3 mm ) thick plastic panel shown in the figure is lowered from a ship to a construction site on the осean Поог. The plastic panel weighs ( 200 N) in air and is lowered at a rate of ( 2 m/s ). Assuming the ship Cble panel remains vertically oriented, calculate the tension force in the cable. Take ( p = 1000 10 m/s ), and use; kg/m', v 1.31 water Plustic Punel Cplam = 1.328/ Re, Cpturb = 0.072/Re 1/5arrow_forward2. The table below shows experimental data for the shape and pressure distribution on the upper and lower surface of an airfoil at zero angle of attack. X 0 0.25 0.5 0.75 1 Yu 0 0.0952 0.0922 0.0588 0 Fx = dYu dx 5 Y₁ 0 -0.0254 -0.0144 -0.0052 0 PL Pd) d: dx Y = Yu (x) Calculate the drag and lift force on the airfoil by numerically evaluating the integrals = S₁² ( P₂₁ = ['(P₁ - F (P₁ - Pu)dx 0 y = Y₂ (x) Pu 1.000 -1.640 -0.786 -0.212 0 Fy x Pi 1.000 0.589 0.426 0.322 0 Use finite difference approximations of the derivatives and the composite Simpson's rule to evaluate the integrals. Hint: Use central differences to estimate the derivatives wherever possible since they are more accurate than forward or backward differences.arrow_forwardAn aircraft is required to climb with 10° of climb angle. The aircraft has a mass of 30,000 kg and produces 50,000 N of drag. Assume zero angle of attack and zero thrust setting angle. How much thrust the aircraft engine must produce?arrow_forward
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