Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 8, Problem 8.9P
The entropy increase across a normal shock wave is
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The entropy increase across a normal shock wave is 199.5 J/(kg · K). What is the upstream Mach number?PLease show step by step solns for better understanding thank you!
A supersonic aircraft flies with a velocity
lower by 39 m-s at 7800 m elevation, (T = - 54 C°
at 7800 m) and (T = - 57 C° at 8800 m, M = 2),
determine the difference between Mach no.s.
56. A supersonic aircraft flies with a velocity
greater by 39 m-s1 at 7600 m elevation, (T = - 51
C° at 7600 m) and (T =- 55 C° at 8600 m, M = 1.9),
determine the ratio of Mach no.s.
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...
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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
- For non-isentropic constant-area flow with stagnation temperature change the following relation was determined: Y 1 To _ ²(y + 1)M² (1 + ¹ Z ¹ M²) 2 TO (1+yM²)² It is possible to use the above equation and calculate the downstream Mach number without resorting to iteration for a flow where the upstream Mach number, as well as the upstream and downstream stagnation temperatures, are known. This is a common calculation for flows through engine combustors. Presuming the left side is a known quantity, show that the above equation can be directly solved as a quadratic in M² and which roots correspond to the subsonic/supersonic solution. Rewrite the equation as: aM4 + bM² + c = 0, and then M² = (−b ± √b² - 4ac)/2a. Determine the appropriate expressions for a, b, and c.arrow_forwardThe Mach number of an aircraft that travels with a speed of 260 m/s in air at 25° C while it undergoes the compressibility effect will be: (Use speed of sound in air at 0° C: 331 m/s) Select one: a. M = 0.69 b. M= 0.70 c. M= 0.75arrow_forward4. Determine the upstream Mach number, considering an oblique shock wave with ew = 32° and a pressure ratio, P2/P1 = 3.0.arrow_forward
- Air is flowing in a convergent nozzle. At a particular location within the nozzle the pressure is 280 kPa, the stream temperature is 345 K. and the velocity is 150 m/s. If the cross-sectional area at this location is 9.29 x 103 m², find: (a) The Mach number at this location, (b) The stagnation temperature and pressure. (c) The area, pressure, and temperature at the exit where M-1.0. (d) The mass rate of flow for the nozzle. Indicate any assumptions you may make and the source of data used in the solution.arrow_forward(b) Air flows through a cylindrical duct at a rate of 2.3 kg/s. Friction between air and the duct and friction within air can be neglected. The diameter of the duct is 10cm and the air temperature and pressure at the inlet are T₁ 450 K and P₁ = 200 kPa. If the Mach number at the exit is Ma2 determine the rate of heat transfer and the pressure difference across the duct. The constant pressure specific heat of air is cp = 1.005 kJ/kg-K. The gas constant of air is R = 0.287 kJ/kg-K and assume k = 1.4. -arrow_forwardi need the answer quicklyarrow_forward
- aircraft flies at the same Mach number but 50 m/s slower at 8 km compared to its speed at sea level. Find this Mach number a. 1.45 b. 1.55 c. 2.25 d. 1.65 e. 2.50.arrow_forwardFast please.arrow_forward2. Carbon dioxide gas (CO2) flows adiabatically along a duct. At station 1 the static pressure P₁=120 kPa and the static temperature T1= 120 oC. At station 2 further along the duct the static pressurep2=75 kPa and the velocity C 2= 150 m/s.Determine (1) the Mach number M 2 (2) the stagnation pressure poz (3) stagnationtemperature Toz (4) the Mach number M 1 For CO 2 take R = 188 J/(kg K) and K = 1.30.arrow_forward
- Q1: when aircraft is ffying at subsonic velocity, the pressure at its nose i e. the stagnation point is found to be 160 kpa. if the ambient pressure and temperature are 100 kpa and 298k respectively, find the speed and Mach number at which the aircraft is flying.-arrow_forwardAir flows isentropically at a rate of 1.3 kg/s from a large chamber through a convergent- divergent duct and leave to the outlet at Mach number 2.72. The air velocity, pressure OUTM EXAMINATION SESSION 2020/2021 (a) Sketch the system and label all components with subsonic/supersonic and UTM N 2020/2021 TM FINAL EXAMIN 2020/2021 answer. 2020/2021 st TION STAL EXAMINATION SEMESTAR I, SESSJO , SESSIOVON 3, SERIONON Esto ON b/202 2020/202 and temperature at a location somewhere along the system were found to be 900 m/s, OUTM 150 kPa and 60°C, respectively. FINAL EXAMINATIC SEMESTER IL SESSION 2025/202 RATION 2020/2021 FINAL EXAMINA BEMENTER I SESSION 2020/202 diffuser/nozzle according to the effect of area change. Justify your ALE eSTER R, SESSION 202b/202 (b) Determine the pressure and temperature of the air in the large chamber, the area at throat, and the velocity at outlet. zb/202arrow_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_forward
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What is entropy? - Jeff Phillips; Author: TED-Ed;https://www.youtube.com/watch?v=YM-uykVfq_E;License: Standard youtube license