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.

Air flows from a tank through a nozzle into the standardatmosphere, as in Fig.  A normal shock stands in theexit of the nozzle, as shown. Estimate (a) the pressure inthe tank and (b) the mass flow.

M4

(a) A pitot tube indicates a pressure of 265 kPa in an airstream in which the temperature is 10°C and the local Mach number is 1.5. Find the static pressure in the airstream. (b) In another problem, a normal shock wave occurs in air at a point where the stagnation temperature is 300°C and the velocity is 700 m/s. The stagnation pressure is 700 kPa. Under this situation do the following: (i) Draw a diagram and label all flow conditions. (ii) Evaluate the Mach numbers, static and stagnation pressures, static and stagnation temperatures before and after the normal shock. Give reasons to justify your answers.

MA The Mach number at the maximum velocity point on the upper surface of an airfoil is 0.6 for the freestream conditions of M = 0.5. Calculate the Mach MA = 0.6 number at the same point for the freestream conditions of M = 0.7. Use the convenient similarity rule. М,- 0.7 M= 0.5

AERODYNAMICS 1. A uniform supersonic air flow at Mach 2.0 passes over a wedge. An oblique shock, making an angle of 40° with the flow direction, is attached to the wedge. If the static pressure and temperature in the freestream are 0.5 Pa and 0°C, determine: a. the deflection angle in degrees b. the pressure ratio c. the Mach number in the downstream region

None

Compressible Flow and Propulsion 1. Determine the speed of sound in Argon at 120 °C. Molecular weight of Argon = 40 kg/kmol and y= 1.667. (Ro = 8314.5 J/kmol K) %3D [Ans. 369 m/s] 2. Calculate the Mach number corresponding to the following speeds at an am temperature of 15 °C. Rair = 287 J kg*K, 1 mph = 0.447 ms?. ent (a) National speed limit of 70 mph [Ans. 0.09] (b) A top speed of 230 mph attained by a Formula driver. [Ans. 0.3] 3. Calculate the speed of an aircraft cruising in 20°C air where the angle of the zone of silence is a. 10° to the direction of travel. [1946ms) b. 20° to the direction of travel. [943ms] c. 80° to the direction of travel. [60ms] d. What is the maximum expected angle for M = 1? 4. Calculate the speed of sound during the flight of an aircraft for the two conditions (a) when it is at a cruising altitude of 10,000 m and the outside temperature is-60 °C. (b) when it lands in a desert city at a temperature of 45 °C (c) If the Mach number during cruising condition…

i need the answer quickly

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)