Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 9, Problem 9.7P
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The flow just upstream of a normal shock wave is given by p1 = 1 atm,T1 = 288 K, and M1 = 2.6. Calculate the following properties justdownstream of the shock: p2, T2, ρ2, M2, p0,2, T0,2, and the change inentropy across the shock.
A normal shock wave was formed on the surface of a supersonic aircraft at a velocity of 1,600 m/s into still atmospheric air at standard seal level conditions. Calculate:
a) M₁
b) M₂
c) P₂
d) T₂
e) V₂
An aircraft is flying at supersonic speed. At a component of an aircraft where the flow is perpendicular, the density ratio is 5. Solve for:
a.Mach Number Downstream
b. Pressure Ratio
c. Temperature Ratio
d. Mach Number upstream
Chapter 9 Solutions
Fundamentals of Aerodynamics
Ch. 9 - A slender missile is flying at Mach 1.5 at low...Ch. 9 - Consider an oblique shock wave with a wave angle...Ch. 9 - Equation (8.80) does not hold for an oblique shock...Ch. 9 - Consider an oblique shock wave with a wave angle...Ch. 9 - Consider the flow over a 22.2 half-angle wedge. If...Ch. 9 - Consider a flat plate at an angle of attack a to a...Ch. 9 - A 30.2 half-angle wedge is inserted into a...Ch. 9 - Consider a Mach 4 airflow at a pressure of 1 atm....Ch. 9 - Consider an oblique shock generated at a...Ch. 9 - Consider the supersonic flow over an expansion...
Ch. 9 - A supersonic flow at M1=1.58 and p1=1atm expands...Ch. 9 - A supersonic flow at M1=3,T1=285K, and p1=1atm is...Ch. 9 - Consider an infinitely thin flat plate at an angle...Ch. 9 - Consider a diamond-wedge airfoil such as shown in...Ch. 9 - Consider sonic flow. Calculate the maximum...Ch. 9 - Consider a circular cylinder (oriented with its...Ch. 9 - Consider the supersonic flow over a flat plate at...Ch. 9 - (The purpose of this problem is to calculate a...Ch. 9 - Repeat Problem 9.18, except with =30. Again, we...Ch. 9 - Consider a Mach 3 flow at 1 atm pressure initially...Ch. 9 - The purpose of this problem is to explain what...
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- Air flowing isentropically through a converging nozzle discharges to the atmosphere at standard sea level conditions. At the section where the absolute pressure is 179 kPa, the temperature is 39˚C and the air speed is 177 m/s. Determine the nozzle throat pressure.arrow_forwardAn aircraft is flying at supersonic speed. At a component of an aircraft where the flow is perpendicular (normalshock), the density ratio is 5. Solve for: a.Mach Number Downstream b. Pressure Ratio c. Temperature Ratio d. Mach Number upstreamarrow_forwardA normal shock produced by an explosion propagate at a constant velocity of 450 m/s into still air of 100 kPa and 23 °C. The ratio of stagnation pressure upstream the shock to stagnation pressure of the gas flow behind the wave isarrow_forward
- If there are no shock waves and the pressure at one ductsection in Fig. is 55.5 kPa, what is the velocity atthat section?(a) 166 m/s, (b) 232 m/s, (c) 554 m/s, (d ) 706 m/s,(e) 774 m/sarrow_forwardThe velocity and temperature behind a normal shock wave are 329 m/s and 1500 K, respectively. Calculate the velocity in front of the shock wave.arrow_forwardThe converging-diverging flow domain is shown in Figure 1. The inlet diameter is 0.2 m, the throat diameter is 0.15 m, and the outlet diameter is 0.24 m. The axial distance from the inlet to the throat is 0.30 m—the same as the axial distance from the throat to the outlet. At the inlet section, the stagnation pressure P0 is set to 220 kPa (absolute), while the stagnation temperature T0, at the inlet is set to 300 K.arrow_forward
- A large reservoir at 20 °C and 800 kPa is used to fill a smallinsulated tank through a converging–diverging nozzle with1-cm2 throat area and 1.66-cm2 exit area. The small tankhas a volume of 1 m3 and is initially at 20 °C and 100 kPa. Estimate the elapsed time when (a) shock waves begin toappear inside the nozzle and (b) the mass flow begins todrop below its maximum value.arrow_forwardThe pressure ratio across a normal shock wave in air is 4.5. What are the Mach numbers in front of and behind the wave? What are the density and temperature ratios across the wave?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_forward
- Air at 20°C and 1 atm flow past the flat plate in Fig.under laminar conditions. There are two equally spacedpitot stagnation tubes, each placed 2 mm from the wall.The manometer fluid is water at 20°C. If U = 15 m/s andL = 50 cm, determine the values of the manometer readingsh1 and h2, in mm.arrow_forwardAt an altitude of 7 km, air is travelling at a supersonic speed and decelerated by a normal shock wave which causes a density ratio of 0.325 across the wave. Compute: a) M₁ b) V₁ c) V d) M₂ e) V₂ f) P2 g) p₂ h) Tarrow_forwardThe reservoir pressure and temperature for a convergent-divergent nozzleare 5 atm and 520◦R, respectively. The flow is expanded isentropically tosupersonic speed at the nozzle exit. If the exit-to-throat area ratio is 2.193,calculate the following properties at the exit: Me, pe, Te, ρe, ue, p0,e, T0,e.arrow_forward
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