Reference no: EM131394731

Problem 1: Delta-V vs. Payload Mass Trade

You are a spacecraft integrator seeking to visually communicate the trade-space between delta-V capability and payload mass provision. For an ESPA class satellite with a total wet mass of 180.0 kg, create a plot that compares the amount of conferred Total ΔV vs. Payload Mass allocation (zero to max), as a function of the following prevailing propulsion types (specified I_sp): cold gas (75s), monoprop (220s), bi-prop (316s), and Hall Effect Thruster (1380s). The total space vehicle mass is defined as follows:

M_Total = M_Spacecraft + M_Propellant + M_Payload

where the combined payload-propellant mass-fraction ([M_Payload + M_Propellant] / M_Spacecraft) is 50.0%.

Problem 2: Tank Sizing

Using an allocation for propellant that is 27% of the not to exceed (NTE) ESPA space vehicle wet mass, for each propulsion type in Problem #1, use their associated densities and other pertinent data (see table below), to compute the required volume and radius dimensions, as well as empirically-derived mass estimate of the corresponding single spherical propellant tank employing a propellant management device (PMD). Ensure that adequate margin is included in the determined tank volume allocation to account for design growth (20%) and temperature/density-driven propellant expansion (5%). When calculating the required tank size of the Bi-prop, the oxidizer component can be ignored. Create a summary table for the results that specify all utilized input parameters.

* Values at critical point     

Now, assuming the monoprop system utilizes regulated Helium pressurant (R_gas = 2080 J/kg-K) with a maximum beginning of life operating pressure of 27.6 MPa and temperature of 323K, and corresponding EOL limits of 689 kPa and 293K, respectively, determine the following parameters:

a. Mass of pressurant

b. Volume of pressurant tank (without margin)

c. Mass of pressurant tank with specification of whether it is of all-metallic or composite overwrapped construction

Problem 3: LEO →GEO Transfer System Sizing

Suppose a spacecraft in permanent orbit around the earth is to be used for delivering payloads from low earth orbit (LEO) to geostationary equatorial orbit (GEO). Before each flight from LEO, the spacecraft is refueled with propellant, which it uses up in its round trip to GEO. The outbound leg requires four times as much propellant as the inbound return leg (for the same total delta-v). The delta-v for transfer from LEO to GEO is 4.22 km/s. The specific impulse of the propulsion system is 430 sec. If the payload mass is 3500 kg, calculate the empty mass of the vehicle.

Problem 4: Launch Vehicle Selection

Explore the features provided at the NASA Launch Services Program (LSP) Performance Web Site: https://elvperf.ksc.nasa.gov. Utilizing the Performance Query Tool (with default NLS-II contract), create and provide the following (Note: Ensure you review all options in the table and not just the results for the limitation of five selections to plot):

1. A plot of Payload Mass vs. C3 (High Energy option). What is the maximum C3 provided for a 500 kg payload and by which vehicle? Note: only consider LV options for which there is a valid performance result (i.e., do not extrapolate).

2. A plot of Payload Mass vs. Altitude (LEO option), choosing from five different launch vehicle families. What is the maximum mass that can be delivered to a 600 km sun synchronous orbit?

Assignment File - https://www.dropbox.com/s/zhmzfe65cbcmycf/Assignment%20File.rar?dl=0