Design a rocket nozzle, knowing that you have a chamber pressure of 6.28 MPa, mass flow rate of 1.23, and temperature of 825K.

5. For the subsonic axial-flow air expander specified, calculate the stagnation and static pressures and temperatures and the Mach number at the rotor-inlet plane (figure P2.5). Also find the rotor rotational speed and the nozzle-inlet blade height for constant-outer-diameter blading. Sketch the form of the complete turbine expansion on an enthalpy-entropy chart. The axial velocity will be constant at this design point. Mass flow, m = 2 kg/s Nozzle-inlet stagnation temperature, Tuni = 400 °C. Absolute nozzle-inlet stagnation pressure, Po,ni = 3 bars(= 3 x 105 N/m²) Mean diameter, dm = 0.25 m. Blade height at rotor entry, 1 = 0.1dm(= 1/2[ds - dn]). Flow angle at nozzle exit, a₁ = 70° to axial direction. Drop in stagnation pressure in nozzle, Apo = 0.05 bar Rotor peripheral speed at mean diameter, um = 0.5x (component of nozzle outlet velocity, Ce,1).

a turbojet on a test stand air at inlet has density of 1.1 kg/m^3 at the exit gas has higher velocity and higher temp, but pressure is the same. inlet area is 0.5m^2 and free stream velocty is 200m/s. fuel air mix is 1:15 at exit, engine area is 0.3m^2, gas velocity 800m/s Determine the density of gas at exit and evaluate reaction forces to hold test stand Turbojet Engine V1 V2 P2 A1 A2 Test Stand

Q4) The nozzles receive steam at 1.75 MPa, 300 °C, and exit pressure of steam is 1.05 MPa. If there are 16 nozzles, find the cross-sectional area of the exit of each nozzle for a total discharge to be 280 kg/min. Assume nozzle efficiency of 90%. If the steam has velocity of 120 m/s at the entry to the nozzles, by how much would the discharge be increased?

Calculate the thrust of a turbojet engine with an inlet diameter of 4 ft. If an airplane is flying at altitude of 15km at Mach 0.8 with an exit velocity of 900ft/s.  The turbojet engine thrust at 15-km. (in Lbf)