Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
6th Edition
ISBN: 9781259683268
Author: Anderson, John
Publisher: Mcgraw-hill Higher Education (us)
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Chapter 14, Problem 14.3P
Consider a hypersonic vehicle with a spherical nose flying at Mach 20 at a standard altitude of 150,000 ft. where the ambient temperature and pressure are
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Consider a hypersonic vehicle with a spherical nose flying at Mach 20at a standard altitude of 150,000 ft, where the ambient temperature andpressure are 500◦R and 3.06 lb/ft2, respectively. At the point on thesurface of the nose located 20◦ away from the stagnation point, estimatethe: (a) pressure, (b) temperature, (c) Mach number, and (d) velocity ofthe flow.
A normal shock occurs in a stream of oxygen. The oxygen flows at Ma=1.8 and the upstream pressure and temperature are 40 psia and 85 degrees Fahrenheit. a) Calculate the following on the downstream side of the shock: static pressure, stagnation pressure, static temperature, stagnation temperature, static density, and velocity. b)If the Mach number is doubled to 3.6, what will be the resulting values of the parameters listed in part (a)?
What is the Mach number of an airflow with a velocity of 2000m/s, at 2,000 ft altitude?
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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?arrow_forward5. 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 thisarrow_forwardNitrogen at 700 kPa pressure and 300K temperature enters a convergent divergent nozzle at negligible speed and passes through a normal shock wave at the Mach number of 3. Calculate the pressure, temperature, velocity, Mach number and stop pressure in the flow behind the shock wave. Compare the values you find with the values found with the air passing through the normal shockwave under the same conditions.arrow_forward
- Calculate the ratio of kinetic energy to internal energy at a point in an airflow where the Mach number is: (a) M = 2, and (b) M = 20.arrow_forwardThe entropy increase across a normal shock wave is 199.5 J/(kg · K). What is the upstream Mach number?PLease show step by step solns for better understanding thank you!arrow_forwardWhat is the Mach of airflow with 150 m/s at a pressure of 0.95 atm and a Density of 1.15 kg/m^3?arrow_forward
- A uniform supersonic flow of air at Mach 3.8, with a stagnation pressure of 5.0 MPa and stagnationtemperature of 1100 K, expands around a 23o convex corner. Determine the downstream Macharrow_forwardA uniform supersonic airstream travelling at a Mach number of 9.0 passes over a concave corner, as shown in Figure 4. An oblique shockwave, which makes an angle of 30° with the flow direction, is attached to the corner under the given conditions. If the pressure and temperature in the uniform flow are 45 kPa and -30 °C respectively, determine the Mach number and deflection angle behind the wave.arrow_forward2. An air tank with a nozzle has a pressure of 196.32 KPa and density of 1.9 Kg/m³. Outside the converging- diverging nozzle, the pressure is atmospheric and designed to have a Mach No. of 1.0 and 1.5 at the throat and exit respectively. The area at the throat is 0.11m2. Calculate the following: (a) Temperature and speed of sound at the tank. (b) Pressure, density, temperature and speed of sound at the throat. (c) Mass flow at the exit.arrow_forward
- Consider the flow of air through a supersonic nozzle. The reservoir pressure and temperature are 5 atm and 500 K, respectively. If the Mach number at the nozzle exit is 3, calculate the exit pressure, temperature, and density.arrow_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_forwardThe Pitot tube on a supersonic aircraft (see the Video) cruising at an altitude of 25000 ft senses a stagnation pressure of 15.0 psia. If the atmosphere is considered standard, determine (a) the Mach number of the aircraft, (b) the airspeed. A shock wave is present just upstream of the probe impact hole. (a) Ma = (b) V= ft/sarrow_forward
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