-8D Static Free-stream pressure pressure Ps Stagnation pressure 4 to 8 holes +10% Static pressure Po Ps Stagnation pressure -10% 0° Differential 10° 20 Yaw angle 0 pressure transducer JO.Lg

Q: Air at 7.20 MPa (abs.) and a temperature 1100K flows isentropically through a converging-diverging…

A: According to the given data as we know that  Po = 7.2Mpa  To =1100k  A*  = 0.01m2  for mach number 4…

Q: A large tank supplies helium through a converging-diverging nozzle to the atmosphere. The Pressure…

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Q: An aircraft with nose M1 P1 T1 properties of the upstream unifrom air M2 angle 0 = 15° as shown in…

A: nose angle θ = 15° Mach no M1 = 2.5 Upstream air pressure = P1= 60 kPa Upstream temperature = T1 =…

Q: Air enters a converging- diverging nozzle at a pressure of 830 kPa abs and a temperature of 32°C.…

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Q: Calculate : 1. Upstream velocity Vi and deflection angle 0. 2. Downstream Mach number M2

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Q: Q5: air velocity decrease from 480m/s chrough the shock. If the pressure and density upstream of the…

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Q: We are given the pressure on one side of a normal shock wave is 1 atmosphere, and the pressure and…

A: given Px = 1atm Py = 10 atm , Ty = 775K

Q: The Mach no. of a fixed geometry inlet at overspeeded for starting mode (Ai= 0.13 m2, ,At= 0.095 m2)

A: Given data: Area at inlet, Ai = 0.13 m2 Area at throat, At = 0095 m2

Q: The stagnation density after the shock (px= 55 kPa, Tx= 335 K, vx= 525 ms- ‚1) 0.26 kgm-3 O 0.77…

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Q: The large compressed - air tank shown in the figure bellow exhausts from a nozzle at an exit…

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Q: Design the 2D steady MOC contour for a super sonic nozzle assuming that the expansion section is a…

A: To design the 2D steady Method of Characteristics (MOC) contour for a supersonic nozzle, we need to…

Q: If the exit Mach number in Fig. is 2.2, what is theexit area?(a) 0.10 m2, (b) 0.12 m2, (c) 0.15 m2,…

A: Write the given data.

Q: An air formed a shockwave with a Mach of 2.3. What is the pressure downstream when the pressure…

A: Write the expression for the static pressure before the normal shock. Here, ρ represents the…

Q: Shock waves are treated as discontinuities here, but theyactually have a very small fi nite…

A: The variation of the velocity, temperature, pressure and entropy of the before and after the shock…

Q: A normal shock wave exists in air flow with up stream M=2, and a pressure of 20 kpa and temperature…

A: M1=2  , and P1=20 kPa T1=15°C M1=V1γRT1V1=M1×γRT1V1=2×1.4×287×288V1=680.34 msec Now…

Q: When operating at design conditions (smooth exit to sealevelpressure), a rocket engine has a thrust…

A: Write the given data.

Q: -Converging nozzle is supplied with Po=10 bars and To=1000 K and discharges to atmosphere with…

A: Given: Stagnation Pressure P0=10barsStagnation Temperature T0=1000KExit pressure P=9.8bars…

Q: ir at P1 = 1 MPa, T1 = 600 °C and V1 = 150 m/s enters a converging nozzle. The exit area is A2 = 50…

A: A convergent nozzle is one that starts large and gradually gets smaller, resulting in a reduction in…

Q: impact tube

A: Given,  Static pressure=3psia or 20.6843kPa Total pressure=116kPa k=1.4 Temperature of…

Q: 1) What is the value of static temperature, stagnation temperature and stagnation pressure in the…

A: To calculatethe static temperature, stagnation temperature and stagnation pressure in the test…

Q: If the Mach number behind a normal shock is equal to 0.4, calculate the ratio of the readings of…

A: Calculate the Mach number shift across a normal shock to determine the ratio of readings from Pitot…

Q: A Rayleigh flow at Mach = 0.3, P1 = 150KPa and T1 = 80oC, enters a non-adiabatic pipeline. It adds…

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Q: he entropy change across a normal shock wave (Tox= 505 K, vx= 490 ms-1)

A: Given data: Tox = 505 K Vx = 490 m/s For air, specific heat, cp = 1.005 kJ/kg.K = 1005 /kg.K Gas…

Q: A compressed Air enters a convergent-divergent nozzle, at 4.0 atm, 300 K and at Mach 2.0. The nozzle…

A: Note: Since you have posted a question with multiple sub-parts, we will solve first three sub- parts…

Q: Q4: : The pressure ratio across a normal shock wave that occurs in air is 1.4. A head of the shock…

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Q: the upstream Mach number, an oblique shock wave with W = 4. Determine considering 32° and a pressure…

A: A dimensionless quantity which represents the ratio berween the local flow velocity of the object in…

Q: The Pitot tube on a supersonic aircraft (see the Video) cruising at an altitude of 25000 ft senses a…

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Q: The stagnation density after the shock (px= 55 kPa, Tx= 335 K, vx= 525 ms-1),

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Q: Air flows through a duct as in Fig, where A1 =24 cm2, A2 = 18 cm2, and A3 = 32 cm2. A normalshock…

A: Given: Mach number at section 1,Ma1=2.5 The pressure at section 1,p1=40 kPa The temperature at…

Q: es with a speed of Two oblique sh window edges as and window edge eam the first shock per upstream…

A: Given: v1= 2750 km/hr= 763.89 m/s T1= 27 c= 300k P1=100 kpa Solution: M1= V1r×R T1      =…

Q: Example 2. A high-speed AC 130 gunship is flying at a pressure altitude of 10 km. A Pitot tube on…

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Q: Consider the following problem. We wish to design a convergent-divergentnozzle to expand a gas from…

A: A nozzle is a relatively simple device, just a specially shaped tube through which hot gases flow.…

Q: Q1: Air expands from a storage tank through a converging diverging nozzle having a hroat area of (70…

A: Given:The throat area is .The shock area is .The exit area is .The pressure and temperature is  and…

Q: A uniform supersonic airstream travelling at a Mach number of 9.0 passes over a concave corner, as…

A: Mach number,M1=9.0shock angle,Assume,Deviation angle be

Q: A uniform supersonic flow of air at Mach 3.8, with a stagnation pressure of 5.0 MPa and stagnation…

A: Given Data: M1 = 3.8 P = 5Mpa T = 1100K v2-v1 = 23o To find: 1) Downstream Mach 2) Downstream…

Q: Fuel is to be burned in a stream of air moving with a Mach No. 2.0, a temperature (-160 F), and a…

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