Use 2 stages for Earth to LEO and a single stage for LEO to lunar orbit. Assume ΔU=7,600 m/s
for Earth to LEO and ΔU=3,900 m/s for LEO to Lunar orbit. Estimate the delta-U and altitude at staging. Select the Ae/A* for each stage using figure 11.3 for guidance. Select the liquid propellants for each stage and run CEA. Here's the website. https://cearun.grc.nasa.gov/  Use this to determine the throat area. Find the density of your propellants and determine the size of the propellant tanks. Find the diameter and length of your rocket, for now neglect the volume of the structure and engines. Use a 9 degree half angle for the nozzle cone. Assume the density of your payload is 600 kg/m^3 and can be packaged into any shape. Submit your results using the table provided. Here's figure 11.3 that's referenced. The fuel for each stage is liquid CH4 and the oxidizer is liquid O2 for all 3 stages.  Decide initial masses ( Structural, engine, propellant, etc).

 Heres the table that needs to be filled. Use to determine values to calculate.

 

Transcribed Image Text:

9 J Pony 1.8 1.6 1.4 1.2 1.0 0.8 0.6 0,4 0.2 0 1 Y = 1.2 Line of maximum thrust N 4 0.050 Shock line 10 20 0.030 40 -0.020 0.010 Pa/Po=0 -0.005 0.003 11 -0.002 100 200 400 0.001 1000 9 J Pony 1.8 1.6 1.4 1.2 1.0 0.8 0.6 0,4 0.2 0 1 Y = 1.2 Line of maximum thrust N 4 0.050 Shock line 10 20 0.030 40 -0.020 0.010 Pa/Po=0 -0.005 0.003 11 -0.002 100 200 400 0.001 1000

Answer rating: 100% (QA)

Im sorry but I wont be able to fill out an external table or run simulations using the NASA CEA Chemical Equilibrium with Applications tool as part of my assistance However I can walk you through the ... View the full answer