Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 11, Problem 11.5P
For a given airfoil, the critical Mach number is 0.8. Calculate the value of
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A jet transport is flying at a standard altitude of 30,000 feet with a velocity of 550 miles per hour. What is the Mach number?
MA
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= 0.5
5. At 30000ft, estimate the magnitude of transonic drag rise. Using this estimate,
calculate the maximum velocity of the airplane at this altitude. Assume drag-divergence
Mach number of 0.82 and d(D/D₁)/dM = 14.3 where D₁=1750lb is drag at Mach number
0.9 and D₂ = 4250lb at Mach number 1.
6 Estimate maximum range at 30000€ Also calculate the flight speed to obtain this
Chapter 11 Solutions
Fundamentals of Aerodynamics
Ch. 11 - Consider a subsonic compressible flow in cartesian...Ch. 11 - Using the Prandtl-Glauert rule, calculate the lift...Ch. 11 - Under low-speed incompressible flow conditions,...Ch. 11 - In low-speed incompressible flow, the peak...Ch. 11 - For a given airfoil, the critical Mach number is...Ch. 11 - Consider an airfoil in a Mach 0.5 freestream. At a...Ch. 11 - Prob. 11.7PCh. 11 - Consider the flow over a circular cylinder; the...Ch. 11 - In Problem 11.8, the critical Mach number for a...
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- Consider an airfoil at a given angle of attack, say α1 . At low speeds, the minimum pressure coefficient on the top surface of the airfoil is −0.90. What is the critical Mach number of the airfoil? detailed solution plsarrow_forwardThe 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_forwardIf the Mach number behind a normal shock is equal to 0.4, calculate the ratio of the readings of pitot tubes located before and after the shock.arrow_forward
- Under low-speed incompressible flow conditions, the pressure coefficientat a given point on an airfoil is −0.54. Calculate Cp at this point when thefreestream Mach number is 0.58, usinga. The Prandtl-Glauert ruleb. The Karman-Tsien rulec. Laitone’s rulearrow_forwardProblem 5. A Boeing 747 cruises at a Mach number of Ma = 0.87 at an altitude of z = 13 km on a standard day. A window in the cockpit is located where the external flow outside the window is at a Mach number of Ma = 0.2 relative to the plane surface (just outside the boundary layer). The cabin is pressurized to an equivalent altitude of z = 2.5 km for a standard atmosphere. (a) Estimate the pressure difference across the window and specify the direction of the net pressure force. (b) Sketch the stagnation pressure, static pressures, and critical pressure on a T-s diagram.arrow_forwardThe upstream flow properties are T1= 280 K and V= 2000 mph. What is downstream Mach number?arrow_forward
- An aircraft is flying at an altitude of 10000ft and at a certain part of its airframe a dynamic pressure is measured to be about 150 psf. Calculate for the corresponding Mach number of the aircraf* * 0.4658 0.5958 0.5195 0.5787 O 0.4785arrow_forwardA normal shock wave propagates into atmosphere where the atmospheric temperature is equal 336K and pressure is 0.8 atm. For a given pressure ratio of p2/p₁ 41 to T₁ = a) Assuming stationary atmosphere, calculate the shock wave velocity and the velocity induced behind the shock wave. Also calculate the temperature, total pressure and total temperature of the fluid particles behind the shock wave using equations for moving normal shock. =arrow_forwardConsider a cone at zero angle of attack in a hypersonic flow. (Hypersonic flow is very high-speed flow, generally defined as any flow above a Mach number of 5.) The half-angle of the cone is θc, as shown inthe figure. An approximate expression for the pressure coefficient on the surface of ahypersonic body is given by the newtonian sine-squared law : Cp = 2 sin2 θcNote that Cp, hence, p, is constant along the inclined surface of the cone. Along the base of the body, we assume that p = p∞. Neglecting the effect of friction, obtain an expression for the drag coefficient of the cone, where CD is based on the area of the base Sb.arrow_forward
- Consider the Gulfstream IV flying at Mach = 0.8 at an altitude with atmospheric pressure of 33,283.40-Pa. Calculate thrust required assuming a weight of 73,000-lb. Airplane data: S = 950 ft2, AR = 5.92, CDo = 0.015, and K = 0.08. The thrust required. (in lbf)arrow_forwardThe shock wave photograph of a wedge structure is given. By using this photograph determine the Mach number.arrow_forwardA nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.2, with a velocity of 2500 m/s, and a density of 1.0 kg/ m³ in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium. Te = i Pe= To= Po= Mi K kPa K kPaarrow_forward
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