The supersonic nozzle of Fig. P9.138 is overexpanded (case G of Fig. 9.12) with Ae/At= 3.0 and a stagnation pressure of 350 kPa. If the jet edge makes a 4° angle with the nozzle centerline, what is the back pressure pr in kPa?

The supersonic nozzle of Fig. P9.138 is overexpanded (case G of Fig. 9.12) with Ae/At= 3.0 and a stagnation pressure of 350 kPa. If the jet edge makes a 4° angle with the nozzle centerline, what is the back pressure pr in kPa?

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

Related questions

Q: Air at 550 kPa and 100°C enters a smooth 1-m-long pipeand then passes through a second smooth pipe…

Q: Engineers call the supersonic combustion in a scramjetengine almost miraculous, “like lighting a…

A: Given Data: The area at section 1 and 4 is: A1=1 m2. The area at section 2 and 4 is: A2=0.2 m2. The…

Q: Q2 (a) In designing a nozzle why nozzle angle is not zero. Explain with the help of sketch. Q2 (b)…

A: Nozzle :- Nozzle is an device which are used to increase the kinetic energy and convert in to heat.…

Q: Air flows through a constant cross-section conduit with Mack number Ma = 1 at 200 KPa pressure and…

A: Given that, Pressure, P=200 kPa,Temperature static, T=100 °C=373 K, Sound…

Q: An air tank of volume 1.5 m3 is initially at 800 kPa and20 °C. At t = 0, it begins exhausting…

A: Given data

Q: Compute stagnation enthalpy of air at normal conditions (P=1atm, T=300K) moving with velocity of…

A: Write the equation for the stagnation enthalpy. h0=h+v22 =CpT+v22 Here, The stagnation enthalpy is…

Q: Consider the Gulfstream IV flying at Mach = 0.8 at an altitude with atmospheric pressure of…

A: Given data Mach number of aircraft = 0.8 Weight = 73000 lb S = 950 ft2 AR = 5.92 CD,0 = 0.015 K =…

Q: Air flows at high speed through a Herschel venturi monitoredby a mercury manometer, as shown in Fig.…

A: Given data, Determine the density of fluid by using the ideal gas equation,

Q: A gas, assumed ideal, flows isentropically from point 1,where the velocity is negligible, the…

A: Hello. Since your question has multiple sub-parts, we will solve first three sub-parts for you. If…

Q: Consider the Gulfstream IV flying at Mach = 0.35. Calculate thrust required given an altitude of…

Q: 3.4. Consider a normal shock wave in air The upstream conditions are given by M 3; 1 atm, and p, = 1…

A: Given: Normal shock wave in air. The upstream condition are: M1=3P1=1 atmρ=1.23 kg/m3

Q: Air, at po = 160 lbf/in2 and To = 300 °F, fl ows isentropicallythrough a converging–diverging…

Q: 6.7 A converging-diverging nozzle is designed to provide flow at Mach 2.0. With the nozzle…

Q: Assume an inviscid, incompressible flow. Also, standard sea level density and pressure are 1.23…

A: Given data: Sea level density, ρ = 1.23 kg/m3 Sea level pressure P = 1.01 × 105 N/m2 density of…

Q: Consider the Gulfstream IV flying at Mach = 0.8 at an altitude with atmospheric pressure of…

A: Given data Speed of aircraft = 0.8 Mach Altitude Pressure = 33283.40 Pa Weight = 73000 lb = 33112.24…

Q: Air, at stagnation conditions of 500 K and 200 kPa, flows through a nozzle. At section 1, where the…

A: The Answer is Below

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: Compressible laminar fl ow, f≈ 64/Re, may occur in capillarytubes. Consider air, at stagnation…

A: Given data: *The stagnation temperature of the air is T=100°C. *The stagnation pressure of the air…

Q: Supersonic air takes a 5° compression turn, as inFig. . Compute the downstream pressure and…

A: Write the given information.

Q: In 1976, the SR-71A, flying at 20 km standard altitude, seta jet-powered aircraft speed record of…

A: The standard flying altitude is h = 20 km = 20000 m. The speed of aircraft is v = 3326 km/hr. Since,…

Q: Q2: Air at a stagnation temperature 365K and a stagnation pressure 760 kpa flows through a…

A: Given data as per question Stagnation temperature =365 k stagnation pressure =760 kpa Back pressure…

Q: Air flows from a reservoir where p = 300 kPa and T = 500 K through a throat to in Fig. where there…

Q: The orientation of a hole can make a difference. Considerholes A and B in Fig. which are identical…

A: Write the relation between pressures at entry and exit points of hole B and the Mach number of hole…

Q: c) Prove mathematically that shock wave cannot occur in subsonic flow

A: Solution: Change in entropy across shock is given by, ∆sR=γγ-1ln2(γ+1)M2+γ-1γ+1+1γ-1ln2γγ+1M2-γ-1γ+1…

Q: Carbon dioxide convergent-divergent to a Lule of 60 m / s, 310 °C and 300 kPa in its case, it enters…

A: Nozzle: Nozzle is a device which having varying in cross section, in which pressure is reduced by…

Q: At a section upstream of the throat of a converging-diverging nozzle, the air flow conditions are: T…

Q: Consider the Gulfstream IV flying at Mach = 0.35. Calculate thrust required given an altitude of…

A: Given data Mach number = 0.35 Altitude = 30000 ft Weight = 73000 lb S = 950 ft2 AR = 5.92 CD,0 =…

Q: Air, at stagnation conditions of 450 K and 250 kPa, fl owsthrough a nozzle. At section 1, where the…

A: Given data To determine the Mach number and the stagnation pressure just downstream of the shock.

Q: Q2: Air flows isentropically through converging nozzle which has a throat of 7 cm?. The stagnation…

A: Given data: - The throat area of the nozzle is A = 7 cm2. The stagnation pressure is Po = 130 kPa.…

Q: Air, from a reservoir at 350 K and 500 kPa, fl ows througha converging–diverging nozzle. The throat…

Q: 78. 5.23. Assume that a supersonic nozzle operating isentropically delivers air at an exit Mach…

Q: If there are no shock waves and the pressure at one ductsection in Fig. is 55.5 kPa, what is the…

A: given data as per the question The temperature of tank =400 k The pressure of tank =300 Kpa…

Q: 02: Air flows isentropically through converging nozzle which has a throat of 7 cm?. The stagnation…

Q: An experimental scramjet aircraft is flying at an altitude of 25,000m at Mach 5. Considering it;s…

A: Given: M1 =5 delta = 5 deg P2/P1 = ?

Q: Consider the venturi nozzle of Fig. , with D = 5 cmand d = 3 cm. Stagnation temperature is 300 K,…

A: Given one-dimensional isentropic flow of air. The problem looks sticky—sparse, scattered…

Q: Air, with po = 500 kPa and To = 600 K, flows through aconverging–diverging nozzle. The exit area is…

A: Understand that if the flow is supersonic in the divergent nozzle that the nozzle must be chocked.…

Q: The converging-diverging nozzle shown in Fig. P3.22 expands and accelerates dry air to supersonic…

Q: Air flows isentropically in a converging-divergingnozzle with a throat area of 3 cm2. At section 1,…

A: Given: Throat area A=3 cm2P1=101 kPaT1=300 KV1=868 m/s To determine: Whether the flow is choked or…

Q: Problem (1): A large vacuum tank, held at sokf -absolute. Suck, sea-level standard air through a…

A: Given Data The pressure inside the tank is: P2=60 kPa The nozzle throat diameter is: d=3 cm=0.03 m…

Q: Consider steady, isentropic flow of air through a duct with varying area. At section 1, p1 = 60.0…

Q: Air flows through a converging–diverging nozzle betweentwo large reservoirs, as shown in Fig. A…

A: Given Data: The reading of manometer is, h=15 cm. The pressure at the nozzle inlet is, P0=300 kPa.…

Q: Air flows isentropically through a variable-area duct. Atsection 1, A1= 20 cm2, p1= 300 kPa, ρ1=…

A: If the areas are the same but the velocities are different, there must be a sonic throat in between:…

Q: For low-speed (nearly incompressible) gas flow, the stagna- tion pressure can be computed from…

A: Assumption 1) Non-Viscous 2) Adiabatic 3) Isentropic 4) Steady flow

Q: A converging-diverging nozzle has a throat area of 10 cm2and an exit area of 28.96 cm2. A normal…

A: giventhroat area a=10 cm2exit area A= 28.96 cm2upstream tank temperature T= 400 Knormal shock stands…

Q: A converging–diverging nozzle exits smoothly to sealevelstandard atmosphere. It is supplied by a…

A: Given:Volume of tank,V=40m3Initial pressure of tank,Pi=800KPaInitial temperature of…

Q: Air flows at high speed through a Herschel venturi monitored by a mercury manometer, as shown in…

Q: Supersonic airflow takes a 5° expansion turn, as inFig. Compute the downstream Mach number…

Question

I need the answer as soon as possible

The supersonic nozzle of Fig. P9.138 is
overexpanded (case G of Fig. 9.12) with
Ae/At= 3.0 and a stagnation pressure of 350
kPa. If the jet edge makes a 4° angle with the
nozzle centerline, what is the back pressure
pr in kPa?
P,?
4°
Air
Jet
edge
Р9.138

Expert Solution

This question has been solved!

Explore an expertly crafted, step-by-step solution for a thorough understanding of key concepts.

This is a popular solution!

SEE SOLUTION Check out a sample Q&A here

Given

VIEW

Solution

VIEW

Page 2

VIEW

Trending now

This is a popular solution!

Step by step

Solved in 3 steps with 2 images

SEE SOLUTION Check out a sample Q&A here

Knowledge Booster

Learn more about

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

Consider the venturi nozzle of Fig. , with D = 5 cmand d = 3 cm. Stagnation temperature is 300 K, and theupstream velocity V1 = 72 m/s. If the throat pressure is124 kPa, estimate, with isentropic flow theory, (a) p1,(b) Ma2, and (c) the mass flow.
Air, at stagnation conditions of 450 K and 250 kPa, fl owsthrough a nozzle. At section 1, where the area is 15 cm2,there is a normal shock wave. If the mass fl ow is 0.4 kg/s,estimate (a) the Mach number and (b) the stagnation pressurejust downstream of the shock.
Helium, at T0 = 400 K, enters a nozzle isentropically. Atsection 1, where A1 = 0.1 m2, a pitot-static arrangement(see Fig. P9.25) measures stagnation pressure of 150 kPaand static pressure of 123 kPa. Estimate (a) Ma1, (b) massfl ow m # , (c) T1, and (d ) A*.
Compute stagnation enthalpy of air at normal conditions (P=1atm, T=300K) moving with velocity of 1398 m/s. What temperature of air we will observe?
A converging-diverging nozzle has a throat area of 10 cm2and an exit area of 28.96 cm2. A normal shock stands in theexit when the back pressure is sea-level standard. If theupstream tank temperature is 400 K, estimate (a) the tankpressure and (b) the mass fl ow.
Air flows through a converging–diverging nozzle betweentwo large reservoirs, as shown in Fig. A mercurymanometer between the throat and the downstream reservoirreads h = 15 cm. Estimate the downstream reservoirpressure. Is there a normal shock in the flow? If so, does itstand in the exit plane or farther upstream?
Air flows isentropically in a channel. Properties at section1 are V1 = 250 m/s, T1 = 330 K, and p1 = 80 kPa. Atsection 2 downstream, the temperature has dropped to 0°C.Find (a) the pressure, (b) velocity, and (c) Mach number atsection 2.
Air, at stagnation conditions of 450 K and 250 kPa, flows through a nozzle. At section 1, where the area is 15 cm², there is a normal shock wave. If the mass flow is 0.4 kg/s, estimate (a) the Mach number and (b) the stagnation pres sure just downstream of the shock.
Air, at stagnation conditions of 500 K and 200 kPa, flows through a nozzle. At section 1, where the area is 12 cm², the density is 0.32 kg/m³. Assuming isentropic flow, (a) find the mass flow. (b) Is the flow choked? If so, estimate A*. Also estimate (c) p₁ and (d) Ma₁.
When a pitot tube such as in Fig. is placed in asupersonic flow, a normal shock will stand in front ofthe probe. Suppose the probe reads p0 = 190 kPa andp = 150 kPa. If the stagnation temperature is 400 K,estimate the (supersonic) Mach number and velocityupstream of the shock.
A uniform supersonic flow of air at Mach 3.8, with a stagnation pressure of 5.0 MPa and stagnationtemperature of 1100 K, expands around a 23o convex corner. Determine the downstream Mach
he reservoir pressure and temperature for a Laval nozzle are 5 atm [abs] and 520 K,respectively. The flow is isentropically expanded to supersonic speed at the nozzle exitwith area Ae = 4in2. If the exit-to-throat area ratio is 2.295, calculate the followingexit properties: g) ̇mass flow rate gamma = 1.4, R = 287 J/kgK Please show all work