The SR-71 Blackbird is the fastest production plane capable of reaching speeds of more than Mach 3, roughly 3000 kph. If a Blackbird is cruising at a speed of 585 m/s an altitude of 33,435 ft and assumin it produces a Normal Shockwave in the process, what would be its downstream Mach number? Note: Specific Heat of Air, R= 287 J/kg"K

The SR-71 Blackbird is the fastest production plane capable of reaching speeds of more than Mach 3, roughly 3000 kph. If a Blackbird is cruising at a speed of 585 m/s an altitude of 33,435 ft and assumin it produces a Normal Shockwave in the process, what would be its downstream Mach number? Note: Specific Heat of Air, R= 287 J/kg"K

Principles of Heat Transfer (Activate Learning with these NEW titles from Engineering!)

8th Edition

ISBN:9781305387102

Author:Kreith, Frank; Manglik, Raj M.

Publisher:Kreith, Frank; Manglik, Raj M.

Chapter5: Analysis Of Convection Heat Transfer

Section: Chapter Questions

Problem 5.29P: Air at 20C flows at 1 m/s between two parallel flat plates spaced 5 cm apart. Estimate the distance...

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Air at 20C flows at 1 m/s between two parallel flat plates spaced 5 cm apart. Estimate the distance from the entrance to the point at which the hydrodynamic boundary layers meet.
A large reservoir at 20 °C and 800 kPa is used to fill a smallinsulated tank through a converging–diverging nozzle with1-cm2 throat area and 1.66-cm2 exit area. The small tankhas a volume of 1 m3 and is initially at 20 °C and 100 kPa. Estimate the elapsed time when (a) shock waves begin toappear inside the nozzle and (b) the mass flow begins todrop below its maximum value.
On March 16, 1990, an Air Force SR-71 set a new continental speedrecord, averaging a velocity of 2112 mi/h at an altitude of 80,000 ft.Calculate the temperature (in degrees Fahrenheit) at a stagnation point onthe vehicle.
(a) The maximum speed achieved by the X-15A-2 rocket-powered research craft was 6630 ft/s at 99,000 ft. Assuming standard day conditions, whatwas its Mach number?(b) An SR-71 is traveling at Mach 3.2 in standard day 80,000-ft conditions.What is its velocity in nautical miles per hour?
The scramjet of Fig. operates with supersonic fl owthroughout. Assume that the heat addition of 500 kJ/kg,between sections 2 and 3, is frictionless and at constantarea of 0.2 m2. Given Ma2 = 4.0, p2 = 260 kPa, and T2 =420 K. Assume airflow at k = 1.40. At the combustionsection exit, find (a) Ma3, (b) p3, and (c) T3.
The 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: a) Me; b) pe; c) Te; gamma = 1.4, R=287 J/kgK
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, from a reservoir at 350 K and 500 kPa, fl ows througha converging–diverging nozzle. The throat area is 3 cm2. Anormal shock appears, for which the downstream Machnumber is 0.6405. (a) What is the area where the shockappears? Calculate (b) the pressure and (c) the temperaturedownstream of the shock.
1. Consider a flat plate at an angle of attack α to a M∞ = 4.2 airflow at p∞ = 1 atm[abs].• What is the maximum static pressure that can occur on the plate surface andwhile the shock wave is still attached at the leading edge?• At what value of α does this occur? gamma = 1.4 and R = 287 J/kgK. THERE IS NO PICTURE FOR THIS QUESTION
A large tank, at 500 K and 200 kPa, supplies isentropicairfl ow to a nozzle. At section 1, the pressure is only 120kPa. What is the Mach number at this section?(a) 0.63, (b) 0.78, (c) 0.89, (d ) 1.00, (e) 1.83 assume one-dimensional fl ow of ideal air, R 5287 J/(kg ? K) and k 5 1.4.
Engineers call the supersonic combustion in a scramjetengine almost miraculous, “like lighting a match in a hurricane.”Figure C9.8 is a crude idealization of the engine.Air enters, burns fuel in the narrow section, then exits, all at supersonic speeds. There are no shock waves. Assume areasof 1 m2 at sections 1 and 4 and 0.2 m2 at sections 2 and3. Let the entrance conditions be Ma1 = 6, at 10,000 mstandard altitude. Assume isentropic fl ow from 1 to 2,frictionless heat transfer from 2 to 3 with Q = 500 kJ/kg,and isentropic fl ow from 3 to 4. Calculate the exit conditionsand the thrust produced.
When operating at design conditions (smooth exit to sealevelpressure), a rocket engine has a thrust of 1 million lbf.The chamber pressure and temperature are 600 lbf/in2absolute and 40008R, respectively. The exhaust gasesapproximate k = 1.38 with a molecular weight of 26. Estimate(a) the exit Mach number and (b) the throat diameter.