Question 3 ULO2 A Laval nozzle has a Mach number of 0.4 (located near the inlet at section 1) where the area is 150 cm², the pressure is 300 kPa and temperature 70°C. A normal shock wave is formed near the nozzle exit where the area is 120 cm?. The exit of the nozzle has an area of 140 cm?. Assume the fluid is air, determine the Mach number, static pressure, stagnation pressure, and mass flow rate at station 2. Shock wave

Question 3 ULO2 A Laval nozzle has a Mach number of 0.4 (located near the inlet at section 1) where the area is 150 cm², the pressure is 300 kPa and temperature 70°C. A normal shock wave is formed near the nozzle exit where the area is 120 cm?. The exit of the nozzle has an area of 140 cm?. Assume the fluid is air, determine the Mach number, static pressure, stagnation pressure, and mass flow rate at station 2. Shock wave

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

Problem 1.1MA

See similar textbooks

Similar questions

A convergent divergent diffuser is exposed to air at pressure of 80kPa and temperature of 350K and Mach number Mi=2.5 as shown in fig. The inlet area Ai = 20 cm2, throat area = 10 cm2, and exit area = 25 cm2. 1- If the back pressure is 650kPa the shock wave is take place at the divergent part find: Throat Mach number , mass flow rate, Shock wave location ,exit Mach number. 2- If the incoming Mach number is decrease to 1.8 State the flow condition through the duct and find, the mass flow rate , exit pressure and Mach number M=2.5, P=80, T=350 Flow condition through the duct for CASE 1)
A normoa shock wave exists in air flow with up stream M=2 and a pressure of 20kpa and temperature of 15 c find the Mach number pressure stagntion pressure temperature stagntion temperature and air velocity down stream of the shock Wave
An air tank with a nozzle, has a pressure of twice the standard sea level pressure and density of 2.45 kg/m³. Outside the converging-diverging nozzle, the pressure corresponds to an altitude of 3km and designed to have a mach number of 1.0 and 1.8 at the throat and exit, respectively. The area at the throat is 0.15 square meters. Calculate: a) the temperature and speed of sound in the tank. b) pressure and density at the throat c) mass flow rate at the exit.
Q6: Air following through constant area duct encounters stationary shock as shown in figure. Find the Mach number, temperature, pressure, stagnation pressure, stagnation temperature and velocity after the shock. Ans.: M2 = 0.6746, T2 = 344 K, p2 = 137 kPa, To = 375 K, po = 186 kPa. V1=500 m/s P;=50 kPa 1-250 K → (1) | (2)
(Q) A converging-diverging nozzle discharges an over-expanded jet into a room. Calculate the deflection angle in degrees caused by the oblique shock forming at the nozzle exit? Ratio between the room pressure and the pressure at the nozzle exit is 5.4. Mach number at the nozzle exit is 1.5. Assume the fluid to be air as a perfect gas with k=1.4
Air at P1 = 1 MPa, T1 = 600 °C and V1 = 150 m/s enters a converging nozzle. The exit area is A2 = 50 cm2 and the back pressure is P2 = 0.7 MPa. What is the mass flow rate (kg/s) through the nozzle? What the is Mach number at the nozzle exit? What would be the flow rate you calculate if you assumed isothermal flow?
Q1) Air flows over a plane wall at a Mach number 3.5, the pressure in the flow being 100kpa. The wall turns through an the pressure downstream of the that angle leading to generation of an oblique shock wave whose strength is such corner is 548kpa. Find the turning angle of corner.
The large compressed - air tank shown in the figure bellow exhausts from a nozzle at an exit velocity of Ve = 235 m / s. Assuming isentropic flow, compute: a) the exit Mach number. b) the pressure in the tank. (take; k = 1.396, Cp = 917 J / Kg.k and R = 260 J / Kg.k) air at 30 ° C tank conditions remain constant Pate = 101 kPa
Q2: Air (y = 1.4) enters a converging-diverging diffuser with a Mach number of 2.8, static pressure pi of 100 kPa, and a static temperature of 20°C. For the flow situation shown in Figure, find the exit velocity, exit static pressure, and exit stagnation pressure. Ans: Ve = 55.093 m/s; Pe = 2236.678 kPa; Poe = 2252.44 kPa i M₁ = 2.8 A₁ = 0.10 m² A₂ = 0.50 m² A₁ = 0.25 m²
Design the 2D steady MOC contour for a super sonic nozzle assuming that the expansion section is a simple corner of some fixed angle, with the straightening section being designed to produce uniform outflow. The inflow for your nozzle is to be M = 1, gamma = 1.4, P = 101325, T = 300K, incoming flow area of 1m, with 1m depth into the page. The exit Mach number of your design is to be M = 3. Note to execute this properly you will need to compute v (nu) as a function of M, and viceversa. Please use Matlab for results.
A normal shock wave exists in air flow with up stream M=2, and a pressure of 20 kpa and temperature of 15c. find The mach number, pressure, stagnation pressure temperature stagnation temperature and air velocity down stream of the shock wave.
A normal shock occurs in a stream of oxygen. The oxygen flows at Ma=1.8 and the upstream pressure and temperature are 40 psia and 85 degrees Fahrenheit. a) Calculate the following on the downstream side of the shock: static pressure, stagnation pressure, static temperature, stagnation temperature, static density, and velocity. b)If the Mach number is doubled to 3.6, what will be the resulting values of the parameters listed in part (a)?