Consider an infinitely thin flat plate of chord c at an angle of attack
Want to see the full answer?
Check out a sample textbook solutionChapter 1 Solutions
Loose Leaf for Fundamentals of Aerodynamics
Additional Engineering Textbook Solutions
Fundamentals of Heat and Mass Transfer
Fluid Mechanics: Fundamentals and Applications
Statics and Mechanics of Materials (5th Edition)
Thinking Like an Engineer: An Active Learning Approach (4th Edition)
Vector Mechanics For Engineers
Foundations of Materials Science and Engineering
- 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_forwardConsider a circular cylinder in a hypersonic flow, with its axisperpendicular to the flow. Let φ be the angle measured between radiidrawn to the leading edge (the stagnation point) and to any arbitrary pointon the cylinder. The pressure coefficient distribution along the cylindricalsurface is given by Cp = 2 cos2 φ for 0 ≤ φ ≤ π/2 and 3π/2 ≤ φ ≤ 2πand Cp = 0 for π/2 ≤ φ ≤ 3π/2. Calculate the drag coefficient for thecylinder, based on projected frontal area of the cylinder.arrow_forwardConsider a low-speed open-circuit subsonic wind tunnel. The tunnel is turned on, and the pressure difference between the inlet (the settling chamber) and the test section is read as a height difference of 10 cm on a U-tube mercury manometer. (The density of liquid mercury is 1.36 × 104 kg/m3.) Assume that a Pitot tube is inserted into the test-section flow of the wind tunnel. The tunnel test section is completely sealed from the outside ambient pressure. Calculate the total pressure measured by the Pitot tube, assuming the static pressure at the tunnel inlet is atmospheric. Given that A2/A1 = 1/12. (Round the final answer to two decimal places.) The total pressure measured by the Pitot tube is × 105 N/m2.arrow_forward
- 1 atm = 2116 lb/ft2 = 1.01 × 105 N/m2. Consider the isentropic flow over an airfoil. The freestream conditions areT∞ = 245 K and p∞ = 4.35 × 104 N/m2. At a point on the airfoil, thepressure is 3.6 × 104 N/m2. Calculate the density at this point.arrow_forwardConsider the isentropic flow through a supersonic wind-tunnel nozzle. The reservoir properties are T0= 500 K and p0 = 10 atm. If p (corresponds to your assigned altitude) at the nozzle exit, calculate the exit temperature and density.ASSIGNED ALTITUDE = 9522 ftarrow_forwardConsider a flat plate with a chord length (from leading to trailing edge) of 1 m. The free-stream flow properties are M1 = 3, p1 = 1 atm, and T, = 270 K. Using shock-expansion theory, tabulate and plot on graph paper these properties as functions of angle of attack from 0 to 30° (use increments of 5°): a. Pressure on the top surface b. Pressure on the bottom surface c. Temperature on the top surface d. Temperature on the bottom surface e. Lift per unit span f. Drag per unit span g. Lift/drag ratio (Note: The results from this problem will be used for comparison with linear supersonic theory in Chap. 9.)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_forwardConsider an infinitely thin flat plate with a 1 m chord at an angle of attackof 10◦ in a supersonic flow. The pressure and shear stress distributions onthe upper and lower surfaces are given by pu = 4 × 104(x − 1)2 +5.4 × 104, pl = 2 × 104(x − 1)2 + 1.73 × 105, τu = 288x−0.2, andτl = 731x−0.2, respectively, where x is the distance from the leading edgein meters and p and τ are in newtons per square meter. Calculate thenormal and axial forces, the lift and drag, moments about the leadingedge, and moments about the quarter chord, all per unit span. Also,calculate the location of the center of pressure.arrow_forwardConsider a wing mounted in the test-section of a subsonic wind tunnel. The velocity of the airflow is 160 ft/s. If the velocity at a point on the wing is 195 ft/s, what is the pressure coefficient at this point?arrow_forward
- Consider a supersonic flow with M = 2, p = 1 atm, and T = 288 K. This flow isdeflected at a compression corner through 20◦. Calculate M, p, T , p0, and T0 behind the resulting oblique shock wave.arrow_forwardIn your own words, write a summary of the differences between incompressible flow, subsonic flow, and supersonic flow.arrow_forwardConsider a normal shock wave in air where the upstream flow properties are u1 = 660 m/s, T1 = 288 K, and p1=1 atm. Calculate the velocity, temperature, and pressure downstream of the shock.arrow_forward
- Principles of Heat Transfer (Activate Learning wi...Mechanical EngineeringISBN:9781305387102Author:Kreith, Frank; Manglik, Raj M.Publisher:Cengage Learning