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 10, Problem 10.3P
A Pitot tube inserted at the exit of a supersonic nozzle reads
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1. Calculate the density of air entering the nozzle, if the air leaves a nozzle at a P=1.5atm and
T=600K at a velocity of 400m/s. The air enters at 150 m/s and the inlet to exit diameter ratio
is 3:1.5.
The choked nozzle that converges and diverges. Find the nozzle areas ratio that divides by the throat area for a match integer of 0.30. When k = 1.40 and the isentropic that moves for a gas.
The jet engine on a test stand, shown in the figure, admitsair at 20oC and 1 atm at section 1, where the area is 0.5 m2and the velocity is 250 m/s. The fuel-to-air ratio is 1:30.The air leaves at section 2 at atmospheric pressure and ahigher temperature. The exit velocity is 900 m/s and thearea is 0.4 m2. Compute the horizontal test stand reactionRx needed to hold the engine fixed.
Chapter 10 Solutions
Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
Ch. 10 - The reservoir pressure and temperature for a...Ch. 10 - A flow is isentropically expanded to supersonic...Ch. 10 - A Pitot tube inserted at the exit of a supersonic...Ch. 10 - For the nozzle flow given in Problem 10.1, the...Ch. 10 - A closed-form expression for the mass flow through...Ch. 10 - Prob. 10.6PCh. 10 - A convergent-divergent nozzle with an...Ch. 10 - For the flow in Problem 10.7, calculate the mass...Ch. 10 - Consider a convergent-divergent nozzle with an...Ch. 10 - A 20 half-angle wedge is mounted at 0 angle of...
Ch. 10 - The nozzle of a supersonic wind tunnel has an...Ch. 10 - We wish to design a supersonic wind tunnel that...Ch. 10 - Consider a rocket engine burning hydrogen and...Ch. 10 - For supersonic and hypersonic wind tunnels, a...Ch. 10 - Return to Problem 9.18. where the average Mach...Ch. 10 - Return to Problem 9.19, where the average Mach...Ch. 10 - A horizontal flow initially at Mach I flows over a...Ch. 10 - Consider a centered expansion wave where M1=1.0...
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Need a deep-dive on the concept behind this application? Look no further. Learn more about this topic, mechanical-engineering and related others by exploring similar questions and additional content below.Similar questions
- Air enters a constant duct at M1= 0.3 , p 1 =1atm. and T1 = 273 K. Inside the duct the heat added is per unit mass is q = 300 kJ/kg. a. Calculate M2, p2, T2, b. Repeat when q is increased to 800 kJ/kg. Do the inflow conditions change? If yes evaluate the new conditions.arrow_forwardGood day. Here is my question (10) Consider a low-speed subsonic wind tunnel with a 12/1 contraction area ratio for the nozzle. If the flow in the test section is at a standard sea level conditions with a velocity of 50 m /s, calculate the height difference in a U-tube mercury manometer with one side connected to the nozzle inlet and the other to the test section. ρHg = 13.6 x 103 kg/ m^3.arrow_forwardAn industrial nozzle as shown in Figure 3 works with an inert gas rush through the nozzle where the speed of the gas at the nozzle throat is recorded equal to the speed of sound. The inlet velocity, temperature, and pressure are 98 m/s, 600 oC, and 88 psi, respectively.Take the property of inert gas for k = 1.66 and cp = 5.192 kJ/kg. K. Calculate:a) Stagnation pressureb) Mach number at the nozzle inletc) Temperature of the gas at the throatd) Pressure of the gas at the throatarrow_forward
- Air flows into a converging duct, and a normal shock stands at the exit of the duct. Downstream of the shock, the Mach number is 0.54. If p2/p1 = 2, compute the Mach number at the entrance of the duct and the area ratio A1/A2.arrow_forwardAir flows at 250 m/s through the pipe. The temperature is 400 K and the absolute stagnation pressure is 280 kPa. Assume isentropic flow. For air R = 286.9 J/[kg. K] and k = 1.40. (Figure 1) Figure 0.3 m 1 of 1 Part A Determine the pressure within the flow. Express your answer using three significant figures. Submit ΜΑ Provide Feedback h P= 220072.60 Pa Review your calculations and make sure you round to 3 significant figures in the last step. No credit lost. Try again. ? Previous Answers Request Answerarrow_forwardFollowing manufacture, it is planned to use the biodiesel to power a rocket which uses a de Laval nozzle. The nozzle throat has a cross-sectional area of 5.5 x 10-3 m2. The hot air enters the nozzle at 300 bar(atm) and 77 °C and exits nozzle at 1 bar(atm). If we have choked flow at the throat, what is the pressure at the throat? Considering we have choked flow at the throat, what is the flow rate at the exit of the nozzle?arrow_forward
- Nozzle is assuming steady one-dimensional flow. M = 2.731. This is the the supersonic flow of air through a convergent-divergent nozzle. The stagnation temperature = 300K, stagnation pressure at the inlet = 107500Pa, static pressure at the exit=4400Pa, C1 is a constant = 0.1097 for calculating circular cross-sectional area of a convergent-divergent nozzle: A = C1 + x^2 and x (axial distance from the throat) =1m. γ = 1.4 and R=287. Calculate the mass flow rate of air through the nozzle. Thank You.arrow_forwardConsider a low-speed subsonic wind tunnel with a 12/1 contraction area ratio for the nozzle. If the flow in the test section is at a standard sea level conditions with a velocity of 50 m/s, calculate the height difference in a U-tube mercury manometer with one side connected to the nozzle inlet and the other to the test section. pHg = 13.6 x 103 kg/m3.arrow_forwardSuppose an aircraft is flying at standard sea-level at M = 0.8 and using a pitot-static tube for airspeed measurement. Determine the actual difference in total and static pressures as would be measured by the pitot-static system. Compare the speed of the aircraft as determined using the actual pressure difference and incompressible flow versus that knowing the actual Mach number. Use k = 1.4.arrow_forward
- 2) flow is exhausted from the nozzle exit at standard sea level conditions with a velocity of 50 m/s, calculate the height difference in a U-tube mercury manometer with one side connected Consider the nozzle of a subsonic turbojet with contraction ratio of 2:1. If the jet to the nozzle inlet and the nozzle exit.arrow_forwardThe nozzle of a supersonic wind tunnel has an exit-to-throat area ratioof 6.79. When the tunnel is running, a Pitot tube mounted in the testsection measures 1.448 atm. What is the reservoir pressure for thetunnel?arrow_forwardCalculate the dimesions of a rocket nozzle(Area of throat, Area of exit, Length of converging side, Length of Diverging side), knowing that CO2 is the combustion gas with a chamber pressure of 2000000 Pa, T = 850 K and k = 1.2.arrow_forward
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