Lunar Lander Preliminary propulsion system selection and design

Lunar Lander Preliminary propulsion system selection and design analysis Thursday, January 22, 2009 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

1. Propulsion system critical design requirements • Variable Thrust (throttle for soft landing & trajectory) • • • Mission Delta V • • • 1950 m/s preliminary value from mission ops. Dependant on trajectory Payload • • Must be able to control mass flow rate Eliminates solid propellant engines Currently assumed to be 85 kg Max and Min Thrust Requirements 2. Design features to optimize • • Reliability – proven design concepts and/or existing hardware Cost • • Ex: Space. X 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 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

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 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 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

Figure 2: Propellant mass vs. Isp, (Eq. 1. 27) Space Propulsion Analysis and Design AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

Combine the Ideal Rocket Equation with the given mass definitions to obtain Eq. 1. 27 from Space Propulsion Analysis and Design AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

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; AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

BI- PROP Oxidizer Fuel Isp range O/F range Oxidizer Storability Fuel Storability Oxidizer density [kg/m^3] LOX H 2 360 -435 1. 0 -3. 8 LOX RP-1 265 -322 1. 3 -2. 2 N 2 O 4 RP-1 254 -297 2. 0 -3. 5 N 2 O 4 Unsymmetrical dimethylhydrazine (UDMH) N 2 O 4 Fuel density [kg/m^3] Cryo. 90 K Boiling pt. 20. 4 K 1149 71 Cryo. 90 K Boiling pt. 460 -540 K 1149 580 Boiling pt. 294 K Boiling pt. 460 -540 K 1449 580 323 Boiling pt. 294 K Boiling pt. 336 K 1449 784 Aerozine 50 (50% UDMH+ 50% N 2 H 4) 341. 7 Boiling pt. 294 K Boiling pt. 336 K ? ? 1449 894. 5 Monomethylhydrazine (MMH) 341. 5 Boiling pt. 294 K Boiling pt. 360. 6 K 1449 878. 8 H 2 O 2 RP-1 250 -297 3. 0 - 2. 7 Boiling pt. 423 Boiling pt. 460 -540 K 1463 580 F 2 LH 2 388 -442 1. 5 -4. 5 Boiling pt. 85. 02 K Boiling pt. 20. 4 K 1636 71 References: Huzel, Dieter K. : Huang, David H. Modern engineering for Design of liquid-Propellant Rocket engines. AIAA. 1992 SPAD Appendix B RPE: pp. 244 Table 7 -1 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

HYBRID Oxidizer Fuel Isp range O/F range Oxidizer Storability Oxidizer density [kg/m^3] Fuel density [kg/m^3] H 2 O 2 HTPB 220 -300 2. 0 -6. 5 Boiling pt. 423 K 1463 930 LOX HTPB 270 -315 1. 2 -2. 0 Cryo. 90 K 1149 930 N 2 O 4 HTPB 268 -297 2. 0 -3. 1 Boiling pt. 294 K 1449 930 N 2 O 4 PMMA 230 - 280 Boiling pt. 294 K 1449 1683 Notes: 8: 1 throttling (pp. 581 RPE) PVC PE N 2 O (Nitrous Oxide) Cryo. -88 C 1226 HNO 3 (Nitric Acid f_inert = 0. 16 - 0. 20 Isp = 220 - 315 s reference: RPE ~pp. 581 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

MONO-PROP Propellant Isp range [S] Storability density [kg/m^3] N 2 H 4 (Hydrazine) 225 Tf = 274 K, Tb=386 K 1010 H 2 O 2 155 Tf = 267. 4 K, Tb=419 K 1414 Notes: Reference: SPAD Table B. 1 Astronautix. com, Aerojet MR-104 AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion