1) What is the value of static temperature, stagnation temperature and stagnation pressure in the test section 2) What is the value of stagnation pressure in the storage thank? 3) What do you conclude about state of air? 4) What is the solution to rectify this problem. i.e. 5)What is the minimum temperature of air should you take in reservoir tank? At this condition, what is mass fl ate in test section? Low pressure air chamber

1) What is the value of static temperature, stagnation temperature and stagnation pressure in the test section 2) What is the value of stagnation pressure in the storage thank? 3) What do you conclude about state of air? 4) What is the solution to rectify this problem. i.e. 5)What is the minimum temperature of air should you take in reservoir tank? At this condition, what is mass fl ate in test section? Low pressure air chamber

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
What is the Mach number of an airflow with a velocity of 2000m/s, at 2,000 ft altitude?
= 1.2 bar, and M1 = 3.0. An 2- Conditions before a shock are T1 = 40°C, pl oblique shock is observed at 45° to the approaching air flow. (a) Determine the Mach number and flow direction after the shock. (b) What are the temperature and pressure after the shock?
3:34 ull Q3.pdf Q1) Air at mach number 2 , pressure 0.5* 10° N/m² and a temperature of 0 °C , over a wedge and generation of an oblique shock wave 40° . determine : 1. Static pressure and temperature downstream the shock. 2. The mach number of the flow downstream the shock. 3. The wedge angle. *********** Q2) As shown in figure (1) A mach number 2.4 air stream at oblique shock angle of (B1-2 =30°) of incident shock with turning angle (0=10°). determine : 1. the mach number of the flow upstream the reflected shock. 2. the mach number of the flow downstream the reflected shock. M2 M3 Figure (1)
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
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
An aircraft is flying at supersonic speed. At a component of an aircraft where the flow is perpendicular (normalshock), the density ratio is 5. Solve for: a.Mach Number Downstream b. Pressure Ratio c. Temperature Ratio d. Mach Number upstream
Air at 30 lbf/in2 absolute and 200F flows from a reservoir into a duct. The flow is steady, adiabatic, frictionless. The flow rate is 10lbm/s. What are the cross-sectional area, temperature, pressure, and Mach number at the point in the duct where the velocity is 1,400 ft/s?
For supersonic and hypersonic wind tunnels, a diffuser efficiency, ηD,can be defined as the ratio of the total pressures at the diffuser exit andnozzle reservoir, divided by the total pressure ratio across a normal shockat the test-section Mach number. This is a measure of the efficiency ofthe diffuser relative to normal shock pressure recovery. Consider asupersonic wind tunnel designed for a test-section Mach number of 3.0which exhausts directly to the atmosphere. The diffuser efficiency is 1.2.Calculate the minimum reservoir pressure necessary for running thetunnel.
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…
Q1- A converging nozzle with a throat area of 0.001² m is operated with air at aback pressure of 5.91 bars. The nozzle is fed from a large plenum chamber where the stagnation pressure and the temperature are 10 bars and 60 °C, respectively. The exit Mach number and mass flow rate are to be determined.
Show that for isentropic flow of a perfect gas if a pitot- static probe measures Po, p, and Tg, the gas velocity can be calculated from (k-1Vk v - 2,7, 1 - What would be a source of error if a shock wave were formed in front of the probe?