Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
expand_more
expand_more
format_list_bulleted
Videos
Textbook Question
Chapter 8, Problem 8.17P
When the Apollo command module returned to earth from the moon, it entered the earth’s atmosphere at a Mach number of 36. Using the results from the present chapter for a calorically perfect gas with the ratio of specific heats equal to 1.4, predict the gas temperature at the stagnation point of the Apollo at Mach 36 at an altitude where the freestream temperature is 300 K. Comment on the validity of your answer.
Expert Solution & Answer
Want to see the full answer?
Check out a sample textbook solution
Students have asked these similar questions
What is the Mach number of an airflow with a velocity of 2000m/s, at 2,000 ft altitude?
1. Using the Second and Third-Order Approximations, find the pressure acting on the upper and lower surfaces
of a flat plate inclined at an angle of 10° toward the airflow at initial Mach number of 2 and freestream
pressure of 1 atm.
When the Apollo command module returned to earth from the moon, itentered the earth’s atmosphere at a Mach number of 36. Using the resultsfrom the present chapter for a calorically perfect gas with the ratio ofspecific heats equal to 1.4, predict the gas temperature at the stagnationpoint of the Apollo at Mach 36 at an altitude where the freestreamtemperature is 300 K. Comment on the validity of your answer.
Chapter 8 Solutions
Fundamentals of Aerodynamics
Ch. 8 - Consider air at a temperature of 230 K. Calculate...Ch. 8 - The temperature in the reservoir of a supersonic...Ch. 8 - At a given point in a flow, T=300K,p=1.2atm, and...Ch. 8 - At a given point in a flow, T=700R,p=1.6atm, and...Ch. 8 - Consider the isentropic flow through a supersonic...Ch. 8 - Consider the isentropic flow over an airfoil. The...Ch. 8 - The flow just upstream of a normal shock wave is...Ch. 8 - The pressure upstream of a normal shock wave is 1...Ch. 8 - The entropy increase across a normal shock wave is...Ch. 8 - The how just upstream of a normal shock wave is...
Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Repeat Problems 8.11 and 8.12 using (incorrectly)...Ch. 8 - Derive the Rayleigh Pitot tube formula, Equation...Ch. 8 - On March 16, 1990, an Air Force SR-71 set a new...Ch. 8 - In the test section of a supersonic wind tunnel, a...Ch. 8 - When the Apollo command module returned to earth...Ch. 8 - The stagnation temperature on the Apollo vehicle...Ch. 8 - Prove that the total pressure is constant...
Knowledge Booster
Learn more about
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
- QUESTION 14 The fan of a turbofan engine is tested at high speed. The flow approaches the fan tip at a relative Mach number M1rel = 1.19. An oblique shock forms at the fan tip leading edge, with a shock angle B = 87 degrees. Use cold-air- standard assumptions and the two-dimensional oblique shock formulae to determine the relative Mach number M2rel after the shock. State your answer to two decimal places. Partial credit is awarded for a reasonable approximation to the correct numerical answer. %3Darrow_forwardWhat is the mass flow rate of air at the exit if the inlet has a Mach number of 0.45, an area of 3ft^2, and has standard atmospheric conditions? Your answerarrow_forwardQUESTION 3 Dry air at 270C flows with a velocity of 360 kmph in an adiabatic wind tunnel. The static pressure is measured as 110kPa. Determine the Mach number, Stagnation temperature stagnation pressure and the stagnation density of air at the stagnation point on the nose tip of an aircraft model installed in the tunnel. Assume the cross-section area of the tunnel is 1.0 sq.m, Determine the mass flow rate of the air. For air take R = 287 J/kg; γ = 1.4;arrow_forward
- A long pipe of 0.0254 m diameter has a mean coefficient of friction of 0.003. Air enters the pipe at a mach number of 2.5, stagnation temperature 310 K and static pressure 0.507 bar. Determine for a section at which the mach number reaches 1.2: i) Static pressure and temperature, ii) Stagnation pressure and temperature, iii) Velocity of air, iv) Distance of this section from the inlet and v) mass flow rate of air.arrow_forwardFor supersonic and hypersonic wind tunnels, a diffuser efficiency, ηD,can be defined as the ratio of the total pressures at the diffuser exit andnozzle reservoir, divided by the total pressure ratio across a normal shockat the test-section Mach number. This is a measure of the efficiency ofthe diffuser relative to normal shock pressure recovery. Consider asupersonic wind tunnel designed for a test-section Mach number of 3.0which exhausts directly to the atmosphere. The diffuser efficiency is 1.2.Calculate the minimum reservoir pressure necessary for running thetunnel.arrow_forwardOn October 3, 1967, pilot William J. "Pete" Knight set a world record that still stands to this day for the fastest flight by a crewed, powered aircraft by flying the North American Aviation X-15 as fast as Mach 6.7 at an altitude of 102,000 ft above sea level. Mach number is the ratio of an object’s velocity to the local speed of sound. Using the table for atmospheric properties with respect to altitude, calculate the speed of sound at an altitude of 100,000 ft. What must have been the velocity of the X-15?arrow_forward
- The flow inside the intake of a scramjet is found to be supersonic with Mach number of 2.5. From the instrumentations located inside the intake section, the static pressure and temperature were measured as 1 atm and 290 K, respectively. Inside the intake, the flow encounters the first compression ramp at an angle of 22º. As an intake engineer, investigatethe flow properties (Mach number, static pressure and temperature, stagnation pressure and temperature) after the resulting first oblique shock wave, so that the 2ndcompression ramp can be designed according to the requirements. Use the relevant tables and charts to assist your investigation. Commentif more shock waves are required to ensure that subsonic combustion is possible.arrow_forward3.1. At a given point in the high-speed flow over an airplane wing, the local Mach number, pressure and temperature are 07, 09 atm, and 250 K, respectively Calculate the values of p,, T, p*, T*, and a* at this pointarrow_forwardIf a calorically perfect gas flow is steady, isentropic and one dimensional, which of the following quantities do not change? Total enthalpy, total temperature, statics temperature, total pressure, static pressure, total density, static density, critical temperature, critical pressure, critical density, characteristic Mach number, entropy, enthalpy.arrow_forward
- At the entry of a diverging section, the temperature of a gas is 27*C, and the Mach number is 1.5. The exit Mach number is 2.5. Assume isentropic flow of a perfect gas with a ratio of specific heats of 1.3. The temperature of the gas at exit is most nearly?arrow_forward3.1 At a given point in the high-speed flow over an airplane wing, the local Mach number, pressure, and temperature are 0.7, 0.9 atm, and 250 K, respectively. Calculate the values of po, To, p*, T*, and a* at this point.arrow_forward1. The pressure in a reservoir of a supersonic wind tunnel is 178.42 KPa, and the density is 1.7 kg/m3. The pressure outside the nozzle is atmospheric, with a Mach number of 1.8 at the exit. The throat has a surface area of 0.15m2 with a Mach Number of 1.0. Solve for the value of the following: (a) Reservoir temperature and speed of sound. (b) Pressure, density, temperature and speed of sound at the throat. (c) Mass flow at the exit.arrow_forward
arrow_back_ios
SEE MORE QUESTIONS
arrow_forward_ios
Recommended textbooks for you
Elements Of ElectromagneticsMechanical EngineeringISBN:9780190698614Author:Sadiku, Matthew N. O.Publisher:Oxford University Press
Mechanics of Materials (10th Edition)Mechanical EngineeringISBN:9780134319650Author:Russell C. HibbelerPublisher:PEARSON
Thermodynamics: An Engineering ApproachMechanical EngineeringISBN:9781259822674Author:Yunus A. Cengel Dr., Michael A. BolesPublisher:McGraw-Hill Education
Control Systems EngineeringMechanical EngineeringISBN:9781118170519Author:Norman S. NisePublisher:WILEY
Mechanics of Materials (MindTap Course List)Mechanical EngineeringISBN:9781337093347Author:Barry J. Goodno, James M. GerePublisher:Cengage Learning
Engineering Mechanics: StaticsMechanical EngineeringISBN:9781118807330Author:James L. Meriam, L. G. Kraige, J. N. BoltonPublisher:WILEY
Elements Of Electromagnetics
Mechanical Engineering
ISBN:9780190698614
Author:Sadiku, Matthew N. O.
Publisher:Oxford University Press
Mechanics of Materials (10th Edition)
Mechanical Engineering
ISBN:9780134319650
Author:Russell C. Hibbeler
Publisher:PEARSON
Thermodynamics: An Engineering Approach
Mechanical Engineering
ISBN:9781259822674
Author:Yunus A. Cengel Dr., Michael A. Boles
Publisher:McGraw-Hill Education
Control Systems Engineering
Mechanical Engineering
ISBN:9781118170519
Author:Norman S. Nise
Publisher:WILEY
Mechanics of Materials (MindTap Course List)
Mechanical Engineering
ISBN:9781337093347
Author:Barry J. Goodno, James M. Gere
Publisher:Cengage Learning
Engineering Mechanics: Statics
Mechanical Engineering
ISBN:9781118807330
Author:James L. Meriam, L. G. Kraige, J. N. Bolton
Publisher:WILEY
Intro to Compressible Flows — Lesson 1; Author: Ansys Learning;https://www.youtube.com/watch?v=OgR6j8TzA5Y;License: Standard Youtube License