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Lunar Lander

Lunar Lander. Preliminary propulsion system selection and design analysis Thursday, January 22, 2009. Propulsion system critical design requirements Variable Thrust (throttle for soft landing & trajectory) Must be able to control mass flow rate Eliminates solid propellant engines

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Lunar Lander

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  1. Lunar Lander Preliminary propulsion system selection and design analysis Thursday, January 22, 2009 Thaddaeus Halsmer, Propulsion

  2. Propulsion system critical design requirements Variable Thrust (throttle for soft landing & trajectory) Must be able to control mass flow rate Eliminates solid propellant engines Mission Delta V 1950 m/s preliminary value from mission ops. Dependant on trajectory Payload Currently assumed to be 85 kg Max and Min Thrust Requirements Design features to optimize Reliability – proven design concepts and/or existing hardware Cost Ex: SpaceX Falcon 9, 1925 kg to TLI for $46.8 million  minimum of $24,312/kg Minimize mass and volume of Lunar Lander  Minimize cost of launch vehicle and OTV Thaddaeus Halsmer, Propulsion

  3. Lunar Lander Propulsion system preliminary design tool Xo Results: Hydrazine Mono-Prop System Propellant Mass 140-160 kg Total System Initial Mass 255-280 kg Work with Mission Ops. to derive the thrust profile from the landing trajectory Finish similar model for system size and volume Choose optimum engine and design propulsion system Figure 1: Lunar Lander mass vs. Isp, (Eq. 1.27) Space Propulsion Analysis and Design Payload mass 85 kg, Delta V 1950 m/s, Historical values for inert mass fraction Thaddaeus Halsmer, Propulsion

  4. Figure 2: Propellant mass vs. Isp, (Eq. 1.27) Space Propulsion Analysis and Design Thaddaeus Halsmer, Propulsion

  5. Combine the Ideal Rocket Equation with the given mass definitions to obtain Eq. 1.27 from Space Propulsion Analysis and Design Thaddaeus Halsmer, Propulsion

  6. function [M_prop, M_lander, M_fuel, M_ox] = Prop(delta_V, M_pay, f_inert, Isp)%, f, fuel_dens, ox_dens go = 9.807; %m/s^2 M_prop = M_pay*(exp(delta_V/(Isp*go))-1)*(1-f_inert)/(1-f_inert*exp(delta_V/(Isp*go))); M_inert= f_inert/(1-f_inert)*M_prop; %M_fuel = M_prop*(1/f)/(1+(1/f)); %M_ox = M_prop*f/(1+f); %V_fuel = 1/fuel_dens*M_fuel; %V_ox = 1/ox_dens*M_ox; %V_prop = V_fuel + V_ox; M_lander = M_pay + M_prop + M_inert; Thaddaeus Halsmer, Propulsion

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