Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 9, Problem 9.3P
Equation
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Example 2. A high-speed AC 130 gunship is flying at a pressure altitude of 10 km. A
Pitot tube on the wingtip measures a pressure of 4.24 x 10ª N/m2. Calculate the
Mach number at which the aircraft is flying.
Solution: Solving for P1 at an altitude of 10000 m, we get 2.65 x 104 N/m2
k-1
1.4-1
Po
k
4.24 x 104)
1.4
M?
k – 1
- 1
- 1
1.4 – 1
2.65 x 104
M? = 0.719
M1 = 0.848
I need the answer as soon as possible
3:34
ull
Q3.pdf
Q1) Air at mach number 2 , pressure 0.5* 10° N/m² and a temperature of 0 °C ,
over a wedge and generation of an oblique shock wave 40° . determine :
1. Static pressure and temperature downstream the shock.
2. The mach number of the flow downstream the shock.
3. The wedge angle.
***********
Q2) As shown in figure (1) A mach number 2.4 air stream at oblique shock angle of
(B1-2 =30°) of incident shock with turning angle (0=10°). determine :
1. the mach number of the flow upstream the reflected shock.
2. the mach number of the flow downstream the reflected shock.
M2
M3
Figure (1)
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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- 78. For an aircraft flies with the same Mach no. at elevations of 8800m (T = - 56 C°) and 11800 m (T =- 63 C°), if the aircraft flies 13% slower at 11800 m altitude, find the Mach no. of the aircraft.arrow_forwardAt a point in the flow over an F-15 high-performance fighter airplane, the density,temperature and Mach number are 1 kg/m3, 240 Kelvin, and Mach 1.6.(a) At this point, calculate T0, ρ0, p0, a0, T ∗, a∗, p∗ and the flow velocity.(b) Calculate M ∗1 and, using the Prandtl relation, calculate M ∗2 and u2 after a normalshock.arrow_forward4. It is important to remember that the isentropic relations and normal shock relationsare monotonic with Mach number and can be inverted. To illustrate this, consider thefollowing problems, (c) If the temperature jumps from 0 Celsius ahead of a shock to 200 Celsius behinda shock, what is the shock Mach number?(d) If the velocity behind a Mach 5 shock wave is 200 m/s, what is the upstreamvelocity?arrow_forward
- = 1.2 bar, and M1 = 3.0. An 2- Conditions before a shock are T1 = 40°C, pl oblique shock is observed at 45° to the approaching air flow. (a) Determine the Mach number and flow direction after the shock. (b) What are the temperature and pressure after the shock?arrow_forward1. The flow is considered compressible when the value of Mach number is.... (A) greater than unity (B) greater or equal to 0.3 (C) greater than 0.5 (D) less than 0.3 2. In c-d nozzle, when dA/dV is positive, then the flow is.... (A) subsonic (B) sonic (C) supersonic (D) incompressible 3. Along the adiabatic duct, the...... remains constant. (A) stagnation pressure (B) static pressure (C) static temperature (D) stagnation temperature 4. Across the normal shock wave, the gas velocity ...... (A) decreases↓ (B) increases ↑ (C) remains constant (D) varies randomly v2 (C) 2Cp (D) PV² 5. The dynamic temperature can be calculated from..... (A): M² 2Cp (B) -M² B 6. A supersonic flow with 38° oblique shock angle passes over a concave corner of angle 0=20°. Then, the free stream Mach number M₁ is.... (A) 2 (B) 5 7. Across the expansion waves, the ..... remains constant. (A) Mach number (B) stagnation pressure (C) 3 (C) gas velocity (D) 4 (D) temperature. (D) Pe=0 8. The maximal thrust of the…arrow_forwardQ.28 Superheated steam at 1500 kPa, has a specific volume of 2.75 m³/kmol and compressibility factor (Z) of 0.95. The temperature of steam is integer] (a) 249 (c) 522 °C [Round off to the nearest (b) 471 (d) 198arrow_forward
- Need step by step explanationarrow_forwardCompressible Flow and Propulsion 1. Determine the speed of sound in Argon at 120 °C. Molecular weight of Argon = 40 kg/kmol and y= 1.667. (Ro = 8314.5 J/kmol K) %3D [Ans. 369 m/s] 2. Calculate the Mach number corresponding to the following speeds at an am temperature of 15 °C. Rair = 287 J kg*K, 1 mph = 0.447 ms?. ent (a) National speed limit of 70 mph [Ans. 0.09] (b) A top speed of 230 mph attained by a Formula driver. [Ans. 0.3] 3. Calculate the speed of an aircraft cruising in 20°C air where the angle of the zone of silence is a. 10° to the direction of travel. [1946ms) b. 20° to the direction of travel. [943ms] c. 80° to the direction of travel. [60ms] d. What is the maximum expected angle for M = 1? 4. Calculate the speed of sound during the flight of an aircraft for the two conditions (a) when it is at a cruising altitude of 10,000 m and the outside temperature is-60 °C. (b) when it lands in a desert city at a temperature of 45 °C (c) If the Mach number during cruising condition…arrow_forwardLooking for correct solutionarrow_forward
- determine the Mach number, total pressure, and total temperature after the heat interaction for the following gases: a) y= 1.4 and cp 0.24 Btu/(lbm- °R) b) y= 1.325 and cp = 0.28 Btu/(1bm · °R) %3Darrow_forwardWhat is the Mach number of an airflow with a velocity of 2000m/s, at 2,000 ft altitude?arrow_forwardPlease show full solutionsarrow_forward
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