Note: Show your work completely with calculation steps, not just the result(s). Problems (25 points each) 1. An airplane

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Note Show Your Work Completely With Calculation Steps Not Just The Result S Problems 25 Points Each 1 An Airplane 1 (107.04 KiB) Viewed 44 times

Note Show Your Work Completely With Calculation Steps Not Just The Result S Problems 25 Points Each 1 An Airplane 2 (63.58 KiB) Viewed 44 times

Note Show Your Work Completely With Calculation Steps Not Just The Result S Problems 25 Points Each 1 An Airplane 3 (157.79 KiB) Viewed 44 times

Note Show Your Work Completely With Calculation Steps Not Just The Result S Problems 25 Points Each 1 An Airplane 4 (83.61 KiB) Viewed 44 times

Note: Show your work completely with calculation steps, not just the result(s). Problems (25 points each) 1. An airplane flying at Mach 0.8 and 10 km altitude (speed of sound a = 230 m/s) has following uninstalled engine measurement: Air mass flow rate mo = 50 kg/s Fuel mass flow rate my= 2.55 kg/s Exhaust velocity Ve= 1250 m/s Calculate the uninstalled thrust F and the net power out of engine Wout. 2. The compressor of a gas turbine has 16 stages, ac = 30, ns = 0.92. Knowing that each stage has the same pressure ratio and air specific heat ratio is 1.4, find the polytropic efficiency ec, isentropic efficiency ne, and temperature ratio te of the compressor. Tc Po 3. A real convergent-divergent nozzle (Cn=0.975) which has an h-s diagram shown expands gas (constant cp = 0.257 Btu/lbm.R, gas constant R= 0.05 Btu/lbm.R) from 400 Psia/1,320 R. The temperature at throat is measured 600°F and diameter of the throat is 2”. Determine the mass flow (lbm/s) of the nozzle without considering friction loss. entrance pressure P = 58P throat pressure outlet pressure be - diagram for a convergent divergent no
Po entrance pressure P, = 0 58P throat pressure outlet pressure do C his diagram for a convergent divergent nozzle Diagram of Problem 3
Merlin ID 4. The rocket engine Merlin ID has some performance data as shown. Estimate (in SI units) the propellant mass flow rate and effective exhaust velocity at the vacuum condition. United States Country of origin Manufacturer SpaceX Application Falcon 9. Falcon Heavy Liquid ſuch engine Propellant LOX/RP-1 (rocket grade kerosene) Cycle gas-generator Performance Thrust (vac.) w 690 KN (155.000 lbf) 650 kN (147.000 lbf) Thrust (SL) Thrust-to-welght ratio 150 Chamber pressure 9.7 MPa (1.410 psi) Ip (vac.) 310 s (3.0 km/s)
For graduate students (20 points) 5. Referring to the frictionless-steady air flows going through the converging nozzles and diverging nozzles as shown, Ma <1 Ma > 1 Ma<1 Ma> A B С D circle all correct answers in the following questions: 1) possibly have subsonic flows at outlet in 2) possibly have sonic flows at outlet in (A, B, 3) possibly have supersonic flows at outlet in (A, B, C, D) C, D) B, C, D)