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 9, Problem 9.18P
(The purpose of this problem is to calculate a two-dimensional expanding supersonic flow and compare it with the analogous quasi-one-dimensional flow in Problem 10.15.) Consider a two-dimensional duct with a straight horizontal lower wall, and a straight upper wall inclined upward through the angle
Expert Solution & Answer
Want to see the full answer?
Check out a sample textbook solution
Students have asked these similar questions
Air flows through a constant cross-section conduit with Mack number Ma = 1 at 200 KPa pressure and 100°C temperature. State the necessary assumptions and calculate:
d) With the specified inlet velocity a gas is flowing isentropically in the converging duct. At the throat if we assume that the velocity is supersonic, how does the mass flow rate be affected compared to sonic velocity at the throat. How could supersonic velocity can be achieved in this case?
Note :
(The properties of air at room temperature are R = 0.287 kPa.m3/kg.K and k = 1.4.)
please answer my question , please help me , don't use hand write , I need by Ms word
Air flows through a constant cross-section conduit with Mack number Ma = 1 at 200 KPa pressure and 100°C temperature. State the necessary assumptions and calculate, (a) stagnation temperature (b) stagnation pressure, and (c) stagnation density.(d) With the specified inlet velocity a gas is flowing isentropically in the converging duct. At thethroat if we assume that the velocity is supersonic, how does the mass flow rate be affectedcompared to sonic velocity at the throat. How could supersonic velocity can be achieved in thiscase?
A supersonic wind tunnel is in the design stage. It is to be driven by a large upstream reservoir of compressed air and discharges to atmospheric conditions downstream. The test section has a constant cross-sectional area and lies downstream of a throat, which is a converging-diverging section that serves to accelerate the flow to supersonic conditions. For the duration of any given experiment, the reservoir can be considered to have constant stagnation conditions that are To = 313K and po = 6x105 Pa. The specific gas constant R = 287 J kg-¹ K-1 and the specific heat ratio is y = 1.4. The wind tunnel test section is designed to run with a cross-sectional area A (test section ) = 1.2 m² and Mach number M (test section ) = 4. Find the area of the throat that lies between the reservoir and the test section. Give your answer in m² to two decimal places.
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...
Additional Engineering Textbook Solutions
Find more solutions based on key concepts
6–1C A mechanic claims to have developed a car engine that runs on water instead of gasoline. What is your resp...
Thermodynamics: An Engineering Approach
What is the importance of modeling in engineering? How are the mathematical models for engineering processes pr...
HEAT+MASS TRANSFER:FUND.+APPL.
Find the change in length of side AB.
Mechanics of Materials, 7th Edition
19.8 Calculate the allowable tensile load for the connection shown. The plates are ASTM A36 steel and the weld ...
Applied Statics and Strength of Materials (6th Edition)
A windowmounted air conditioner removes 3.5kJ from the inside of a home using 1.75 kJ work input. How much ener...
EBK FUNDAMENTALS OF THERMODYNAMICS, ENH
17–1C A high-speed aircraft is cruising in still air. How does the temperature of air at the nose of the aircra...
Thermodynamics: An Engineering Approach
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
- Consider 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_forwardA supersonic wind tunnel is in the design stage. It is to be driven by a large upstream reservoir of compressed air and discharges to atmospheric conditions downstream. The test section has a constant cross-sectional area and lies downstream of a throat, which is a converging-diverging section that serves to accelerate the flow to supersonic conditions. For the duration of any given experiment, the reservoir can be considered to have constant stagnation conditions that are To = 313K and po = 6x105 Pa. The specific gas constant R = 287J kg-1 K-1 and the specific heat ratio is y = 1.4. The wind tunnel test section is designed to run with a cross-sectional area A (test section ) = 1.2 m² and Mach number M (test section ) = 3.5. Find the area of the throat that lies between the reservoir and the test section. Give your answer in m² to two decimal places.arrow_forwardThe stagnation chamber of a wind tunnel is connected to a high pressure air bottle farm which is outside the laboratory building. The two are connected by a long pipe which has a inside diameter of 4 inches. If the static pressure ratio between the bottle farm and the stagnation chamber is 10 and the bottle farm static pressure is 100 atm, how long can the pipe be without choking and what is the change in entropy? Assume adiabatic,subsonic, one-dimensional flow with a friction coefficient of 0.005.arrow_forward
- Consider a cone at zero angle of attack in a hypersonic flow. (Hypersonic flow is very high-speed flow, generally defined as any flow above a Mach number of 5.) The half-angle of the cone is θc, as shown inthe figure. An approximate expression for the pressure coefficient on the surface of ahypersonic body is given by the newtonian sine-squared law : Cp = 2 sin2 θcNote that Cp, hence, p, is constant along the inclined surface of the cone. Along the base of the body, we assume that p = p∞. Neglecting the effect of friction, obtain an expression for the drag coefficient of the cone, where CD is based on the area of the base Sb.arrow_forwardOne type of supersonic wind tunnel is a blow-down tunnel, where air is stored in a high-pressure reservoir, and then, upon the opening of a valve, exhausted through the tunnel into a vacuum tank or simply into the open atmosphere at the downstream end of the tunnel. For this example, weconsider just the high-pressure reservoir as a storage tank that is being charged with air by a high-pressure pump. As air is being pumped into the constant-volume reservoir, the air pressure inside the reservoir increases. The pump continues to charge the reservoir until the desired pressure is achieved.Consider a reservoir with an internal volume of 30 m3. As air is pumped into the reservoir, the air pressure inside the reservoir continually increases with time. Consider the instant during the charging process when the reservoir pressure is 10 atm. Assume the air temperature inside the reservoir is held constant at 300 K by means of a heat exchanger.Air is pumped into the reservoir at the rate of 1…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
- Q.2 Air which enters a diverging duct is slowed by a normal shock wave, as shown in figure. The airstream enters the duct at a Mach number of 3 and leaves it at a Mach number of 0.4. the exit cross-sectional area of the duct is twice the inlet cross- sectional area. Determine the pressure ratio across the normal shock wave, and the ratio of the exit pressure to the inlet pressure. M. = 0.4 M; = 3arrow_forwardA Pitot tube is inserted into an airflow where the static pressure is 1 atm. Calculate the flow Mach number when the Pitot tube measures (a) 1.276 atm, (b) 2.714 atm, (c) 12.06 atm.arrow_forwardA piston moves along a tube containing air at an initial sound speed of 330 m/s. When the piston velocity is 250 m/s, it drives a shock wave which propagates at a velocity of 500 m/s. When the piston velocity is 100 m/s, it drives a shock at 400 m/s. Use the hypersonic equivalence principle to calculate the shock angles (in degrees) on a flat plate: At an incidence of 6 degrees and a Mach number of 7.2arrow_forward
- For Air Assume: y =1.4 and R = 287 J/kg K Question B1 a) Air is flowing at a Mach number of 0.6 in a two-dimensional duct at a location where the area is of 0.75 m². At this location the static pressure is 50 kPa and the static temperature is 300 K. 1- Calculate the mass flow rate through the duct; [2 marks] 2- What percentage in area change would be necessary to reach a Mach number of 0.75? [2 marks] 3- What percentage in area change would be necessary to reach a Mach number of 1.00? [2 marks] b) A normal shock wave occurs in a gas with an unknown specific heat ratio. The static pressure ratio across the normal shock wave is 10.6. The Mach number downstream of the shock wave is equal to 0.495, find the specific heat ratio of the gas and the velocity ratio across the shock wave. [8 marks] c) Explain why successive infinitesimal compression waves tend to reinforce and form a shock wave. Use sketches to illustrate your answer. [5 marks] d) Show that for a choked convergent nozzle further…arrow_forwardAir at 290°K and 105 P a approaches a normal shock. The temperature downstream of the shock is 540°K. Find: (a) The velocity downstream of the shock. (b) The pressure change across the shock, and compare it with that calculated for an isentropic flow with the same deceleration.arrow_forwardUnder low-speed incompressible flow conditions, the pressure coefficientat a given point on an airfoil is −0.54. Calculate Cp at this point when thefreestream Mach number is 0.58, usinga. The Prandtl-Glauert ruleb. The Karman-Tsien rulec. Laitone’s rulearrow_forward
arrow_back_ios
SEE MORE QUESTIONS
arrow_forward_ios
Recommended textbooks for you
Principles of Heat Transfer (Activate Learning wi...Mechanical EngineeringISBN:9781305387102Author:Kreith, Frank; Manglik, Raj M.Publisher:Cengage Learning
Principles of Heat Transfer (Activate Learning wi...
Mechanical Engineering
ISBN:9781305387102
Author:Kreith, Frank; Manglik, Raj M.
Publisher:Cengage Learning
Intro to Compressible Flows — Lesson 1; Author: Ansys Learning;https://www.youtube.com/watch?v=OgR6j8TzA5Y;License: Standard Youtube License