Air flows in the channel in the figure below. The pressure at station 1 is 59 psia. Determine the Mach number, static and stagnation pressures at station 3. Assume isentropic flow except for the normal shock wave. 3 1 P03 = i Shock wave A₁ = 100 in.² Ma₁ = 2.1 Ma3 = i P3= i 2 A₂ = 75 in.2 psia psia A3 = 120 in.²

Air flows in the channel in the figure below. The pressure at station 1 is 59 psia. Determine the Mach number, static and stagnation pressures at station 3. Assume isentropic flow except for the normal shock wave. 3 1 P03 = i Shock wave A₁ = 100 in.² Ma₁ = 2.1 Ma3 = i P3= i 2 A₂ = 75 in.2 psia psia A3 = 120 in.²

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

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
Q11: Air enters the converging section shown in Figure and a normal shock occurs at the exit. The entering Mach number is 2.8 and the area ratio 1.7. Compute the overall static A1/A2 = temperature ratio T3/T1. Ans.: T3/T1 = 2.43 A,|A, = 1.7 M = 2.8
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)
Q1/2. Air flow isentropically through a supersonic convergent divergent nozzle with 0.6 kg/s. At the inlet, the pressure is 700 kPa and the temperature is 300 K, and the area is 7 cm. If the exit area is 14 cm? calculate: a- The stagnation pressure and temperature. b- The exit Mach number. c- The exit pressure and temperature. d- The area and velocity at the throat. e- What will be the maximum rate of flow and the corresponding exit Mach number if the flow completely subsonic in the nozzle?
QUESTION 2 A compressor blade design tested in a cascade is found to choke with an inlet Mach number of 0.9 when the inlet flow angle is 52 deg. If the ratio of the throat area to the frontal area, A*/H1s, for the cascade is).625, calculate the loss of stagnation pressure between the far upstream and the throat and express this as a loss coefficient. Give your comments on what could cause this loss.
b) In a converging-diverging duct the flow conditions are such that a normal shock is formed at the exit of the diverging section. Upstream of the shock the flow conditions are as below. Mach number= '5.198', Pressure= '311886.9' Pa and Temperature= '228.72' K. Obtain the Mach number, pressure, and temperature at the downstream of the shock. Assume standard air properties (y=1.4).
An industrial nozzle as shown in Figure 3 works with an inert gas rush through the nozzle where the speed of the gas at the nozzle throat is recorded equal to the speed of sound. The inlet velocity, temperature, and pressure are 98 m/s, 600 oC, and 88 psi, respectively.Take the property of inert gas for k = 1.66 and cp = 5.192 kJ/kg. K. Calculate:a) Stagnation pressureb) Mach number at the nozzle inletc) Temperature of the gas at the throatd) Pressure of the gas at the throat
Q1. A turbojet aircraft flying at speed of 1645 km/h and altitude of 9000 m. A supersonic intake was designed with a conical center body, which generated one oblique shock wave and normal shock wave at the inlet of the aircraft. The deflection of the conical center body is a 10°. Calculate the fluid flow properties and velocity at the aircraft inlet of the intake. Assume a weak solution. Take: R= 0.287 kJ/kg K, and Cp=0.9 kJ/kg K.
A turbine nozzle vane at high speed exhibits a normal shock. The flow is modelled as analogous to that through a convergent-divergent de Laval nozzle with a normal shock at the nozzle exit. Just upstream of the shock, the exit Mach number Me = 1.09, the absolute static pressure pe = 280 kPa, the static temperature Te = 586 K, the ratio of specific heats of the combustion productsk = 1.35, and the constant pressure specific heat capacity Cp = 1027 J-kg "K-1. Assuming the combustion products are constant property ideal gasses, determine the density of the combustion products pp after the shock. State your answer in kg m to two decimal places and enter the numerical value only.
A turbojet engine with a convergent divergent nozzle expands the flow such that Po = 0.8 P, The engine has the following data at a flight Mach number of 1.5 : Te = 7.5 ,Tt4=1400 K, e. = et = 0.9 , y = 1.34 ,CP; = 1135 J/kg.K If the jet velocity is 1025 m/s, determine the flight altitude.
Q1: A convergent-divergent nozzle has an exit area to throat area ratio of underline 2 Air enters the nozzles with a stagnation pressure of 6.5 bar and a stagnation temperature of 93 degrees * C The throat area is 6.25 cm ^ 2 If there is a normal shock wave standing at a point where M= 1.5, determine the pressure, temperature on either side of the plane of shock and the mach number on the downstream side of the plane. Find also the exit mach number of the nozzle. p 2:44 /
A subsonic inlet diffuser of area ratio 3 is employed in a Mach-2 jet engine. A normal shock is formed just upstream of the inlet. The free stream pressure and temperature are 10 kPa and 300 K. Determine the Mach number, pressure and temperature at the diffuser exit. [Ans. M. = 0.16, P. = 55.4 kPa, T. = 537.8 K]