Problem 4. High-speed aircraft use "air data computers" to compute air speed from Pitot-static tubemeasurements of the stagnation pressure and static pressure. (a) Plot the percent error in computing the Machnumber assuming incompressibility (thus, using the Bernoulli equation) versus actual Mach number forMa = 0.1 to Ma=0.9. (b) For an aircraft cruising at an altitude of z = 12 km, plot the speed calculatedassuming incompressible flow and accounting for compressibility versus actual Mach number for Ma = 0.1 toMa = 0.9.

Given Data and Assumptions,Specific heat ratio for air, γ = 1.4Gas constant for air, R = 287 J/(kg…

Answered: Problem 4. High-speed aircraft use "air…

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

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derive an expression for the mass flowrate of air for a C-D nozzle arrangement operating in choked flow conditions. Hint: start with the fact that: mmax = p* A*V * show all your steps. upload your answer here.
Example 2. A high-speed AC 130 gunship is flying at a pressure altitude of 10 km. A Pitot tube on the wingtip measures a pressure of 4.24 x 10ª N/m2. Calculate the Mach number at which the aircraft is flying. Solution: Solving for P1 at an altitude of 10000 m, we get 2.65 x 104 N/m2 k-1 1.4-1 Po k 4.24 x 104) 1.4 M? k – 1 - 1 - 1 1.4 – 1 2.65 x 104 M? = 0.719 M1 = 0.848
1.AB
title of the book: Fluid Mechanics and Thermodynamics of Turbomachinery 6th edition by S.L. Dixon and C.A. Hall
The mass flow rate of a calorically perfect gas (C₁ = constant) through a Convergent-Divergent nozzle is given by, m = A Pt VT₁ M (1+ y +1 YM²) 2(3-1) (a) Derive the above relation from fundamental principals. Here, P₁, is total (or Stagnation) pressure and T₁ is total temperature, M is Mach number, A is local area of nozzle. (b) Prove mathematically that the maximum airflow limit occurs when the Mach number is equal to one and obtain below relation. m A Pt √√Tt Note: The limiting of the maximum mass flow rate when M =1, is called choking of the flow or Choked flow. Hint: https://www.grc.nasa.gov/www/k-12/rocket/mflchk.html y+1 2(y-1) √(4¹) R
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
Consider steady, isentropic flow of air through a duct with varying area. At section 1, p1 = 15 kPa (abs), T = 10°C, Vị = 592 m/s, A1 = 0.25 m². At a downstream section 2, T2 = 137°C, Maz = 0.75. Determine: a. Mach number, Mai b. sketch geometry of the duct c. mass flow rate [kg/s] d. pressure, P2 [kPa] e. cross-sectional area, A2 [m2]
A flow with Mach number M1 = 2 and pressure p1 = 1 atm is turned away from itself twice,first through an angle of 10 deg and then through a second angle of 20 deg. Compute the Machnumber and the static pressure downstream of the second turn. Then, suppose that the original flow(M1 = 2 and p1 = 1 atm) is turned away from itself through a single turn of 30 deg. Compute theMach number and static pressure downstream of this turn, and show that the values are the sameas for the first flow with two turns totaling 30 deg.3
3. The NASA X-43 flies at a Mach number of 9.4 at an altitude of 30,000 m, where thepressure is 1171.8 Pa and the temperature is 226 K. If a supersonic wind tunnel isdesigned to reproduce these conditions, calculate the following:(a) The velocity (m/s), total/stagnation temperature (K), and total/stagnation pres-sure (kPa) in the test section.(b) The velocity (m/s), total temperature (K), and total pressure (kPa) behind thenormal shock formed in front of a blunt surface in the test section.(c) The change in entropy across this normal shock.
For Air: y = 1.4 PART B Question B1 a) With the aid of sketches give a clear account of the operation of convergent nozzles as a function of back pressure. In your answer clearly indicate the critical points (pressure ratios) along the pressure ratio axis. b) Determine the required area ratio A part of the expansion of new teaching and research facilities at QMUL, a new two dimensional blow- down supersonic jet facility is going to be designed, manufactured and installed at our Whitehead Aeronautical Laboratory, Figure QB1. The design Mach number for this facility is M₁ = 2.0. A = A₁ A R=287J/kg K A₁ location at which the area ratio ( A₁ c) What should be the supply pressure to the nozzle (P) in order to have a normal shock wave at a Ax =) is 1.5? A Settling Chamber Po To for this design Mach number. d) For part c, what would be the discharge mass flow rate per unit throat area? You may assume P =105 N/m² and To = 300K. atm Convergent-Divergent Nozzle A AX Schematic of a Supersonic Jet…
At a section in a nozzle (air flow), following properties are known. Po = 1000 KPa; To=500k ; V = 200 m/see é) The temperature (T) at this section is (in K), a. 480; b. 500; c. 520; d. 540; e.560 (ti) The Mach number (M) i's, a. 0.26 ; b. 0:36 ; C. 0.46;d.0.56;e.a66 3.
2) Prove that the mass of gases (m) flowing in the convergent-divergent nozzle of an afterburning jet engine is expressed by 7 Po :A 1 im _7+1_ (1+4M²)VR/To Where Po, To are the total pressure and temperature at a nozzle section having an area of A For chocking conditions,, the Mach number is unity at throat section (Ath), and the above relation is reduced to 20-1) Y Po Ath VR/To

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