Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 8, Problem 8.6P
Consider the isentropic flow over an airfoil. The freestream conditions correspond to a standard altitude of 10.000 ft and
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Chapter 8 Solutions
Fundamentals of Aerodynamics
Ch. 8 - Consider air at a temperature of 230 K. Calculate...Ch. 8 - The temperature in the reservoir of a supersonic...Ch. 8 - At a given point in a flow, T=300K,p=1.2atm, and...Ch. 8 - At a given point in a flow, T=700R,p=1.6atm, and...Ch. 8 - Consider the isentropic flow through a supersonic...Ch. 8 - Consider the isentropic flow over an airfoil. The...Ch. 8 - The flow just upstream of a normal shock wave is...Ch. 8 - The pressure upstream of a normal shock wave is 1...Ch. 8 - The entropy increase across a normal shock wave is...Ch. 8 - The how just upstream of a normal shock wave is...
Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Repeat Problems 8.11 and 8.12 using (incorrectly)...Ch. 8 - Derive the Rayleigh Pitot tube formula, Equation...Ch. 8 - On March 16, 1990, an Air Force SR-71 set a new...Ch. 8 - In the test section of a supersonic wind tunnel, a...Ch. 8 - When the Apollo command module returned to earth...Ch. 8 - The stagnation temperature on the Apollo vehicle...Ch. 8 - Prove that the total pressure is constant...
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- 1 atm = 2116 lb/ft2 = 1.01 × 105 N/m2. Calculate cp, cv, e, and h fora. The stagnation point of a high-speed missile are 934 ◦R and 7.8 atm, respectively.b. Air at standard sea level conditionsarrow_forwardMA The Mach number at the maximum velocity point on the upper surface of an airfoil is 0.6 for the freestream conditions of M = 0.5. Calculate the Mach MA = 0.6 number at the same point for the freestream conditions of M = 0.7. Use the convenient similarity rule. М,- 0.7 M= 0.5arrow_forwardSR-71 Blackbird cruised at Mach 3.2 at 80,000 ft altitude. However, it was designed to leak a finite amount of fuel during takeoff because the surface temperature of the aircraft increases (which leads to expansion of the surfaces) due to aerodynamic heating during flight. Estimate the flow temperature around the aircraft during cruise condition. Assume specific heat ratio is 1.4.arrow_forward
- The shock waves on a vehicle in supersonic flight cause a component ofdrag called supersonic wave drag Dw. Define the wave-drag coefficient asCD,w = Dw/q∞S, where S is a suitable reference area for the body. Insupersonic flight, the flow is governed in part by its thermodynamicproperties, given by the specific heats at constant pressure cp and atconstant volume cv. Define the ratio cp/cv ≡ γ . Using Buckingham’spi theorem, show that CD,w = f (M∞, γ ). Neglect the influence of friction.arrow_forwardThe upstream flow properties are T1= 280 K and V= 2000 mph. What is downstream Mach number?arrow_forwardUsing Epsilon and Angle of Attach to solve for the answers.arrow_forward
- The instrument fairing on an aircraft is shown below, consisting of a forward ramp at30◦, a horizontal section, and a rear ramp at 25◦. During a flight in air at Mach 5.5,the static pressure is 42.6 kPa and the static temperature is 250 K. An oblique shockforms at the turn of the forward ramp, two expansion fans form at the front and rear ofthe horizontal section, and a second oblique shock forms at the turn of the rear ramp. (a) Calculate the Mach numbers and pressures at regions 2, 3, 4, and 5. (b) Determine stagnation pressure ratio p05/p01. (c) Tabulate the angles of all oblique shocks waves leading/trailing expansion waves relative to horizontal (not relative to flow angle). Gamma = 1.4, R=287 J/kgK PLEASE SHOW ALL WORKarrow_forwardAn airfoil is in a freestream where p.(assigned altitude)atm, p(density) (assigned altitude) kg/m3, and V 980 m/s. At a point on the airfoil surface, the pressure is 0.5 atm. Assuming isentropic flow, calculate the velocity at that point. assigned altitude 36.22kmarrow_forward4. For the 2D symmetric airfoil with a diamond profile A = 5° as shown in figure, compute the lift and drag coefficients in the supersonic flow through air (y = 1.4) and free stream mach number M = 2 %3D with AOA a = 10°. (2 (3arrow_forward
- Consider a Lear jet flying at a velocity of 250 m/s at an altitude of 10 km,where the density and temperature are 0.414 kg/m3 and 223 K,respectively. Consider also a one-fifth scale model of the Lear jet beingtested in a wind tunnel in the laboratory. The pressure in the test section ofthe wind tunnel is 1 atm = 1.01 × 105 N/m2. Calculate the necessaryvelocity, temperature, and density of the airflow in the wind-tunnel testsection such that the lift and drag coefficients are the same for thewind-tunnel model and the actual airplane in flight.arrow_forwardThe lift curve for the 4 digit NACA 2421 airfoil is shown in the Figure. Consider a wing with AR = 8, sweep ? = 15 ???, and airplane flies at the Mach number equal to 0.3. 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_forwardConsider the isentropic flow over an airfoil. The freestream conditionscorrespond to a standard altitude of 10,000 ft and M∞ = 0.82. At a givenpoint on the airfoil, M = 1.0. Calculate p and T at this point.arrow_forward
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