Lunar Lander Propulsion System 100g, 10kg and Large Payload cases Thaddaeus Halsmer Thursday, April 9, 2009 Lunar Lander propulsion system (final presentation slide) Lunar Lander Propulsion backup slides Thaddaeus Halsmer, Propulsion

Radial Flow Hybrid Engine H2O2 Tank Helium Tank Radial Flow Hybrid Engine 3 ft. Thaddaeus Halsmer, Propulsion

Payload case/Description Lunar Lander Propulsion – Engine Specifications Table 1 Engine performance parameters Engine No. Payload case/Description F_max/min [N] tb [s] 1 10 kg/hop engine 2x 192 (avg.) 134.5 2 100 g/main engine 1100/110 198.6 3 10 kg/main engine 1650/165 190.4 4 Arbitrary/main engine 27000/2700 250.2 (4) (3) (1) (2) Stick is 6.5 feet high, same as a standard doorway Thaddaeus Halsmer, Propulsion

Thaddaeus Halsmer, Propulsion Lunar Lander Propulsion –fluid system diagrams High Pressure Helium Tank HV01 High Pressure Helium Tank HV01 SV01 SV01 REG REG SV02 SV02 CK01 CK01 RV01 H2O2 Tank HV02 RV01 H2O2 Tank HV02 F01 F01 MOV CK02 MOV CK02 SV04 SV03 SV05 100g and Large payload cases 10kg payload case Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion - Propellant/Propulsion system selection Selection Criteria: Thrust min/max throttling Dimensions Short and fat Mass – minimize Propellant storability Purchase/development costs High Reliability Figure X: Propellant mass vs. Isp trade Thaddaeus Halsmer, Propulsion

Thaddaeus Halsmer, Propulsion Lunar Lander Propulsion - Nozzle area ratio and mass optimization Used CEA to compute Isp for given nozzle area ratio All other inputs constant Empirical nozzle mass equation As area ratio, ε, increases Mnozzle increases, but Isp increases also As Isp increases Mprop decreases for a given thrust and burn time Wrote Matlab script that used Matlab CEA interface to compute multiple Isp’s for different area ratio’s and the corresponding Mprop and Mnozzle for a given thrust, and burn time Results: Area ratio for minimum mass occurred at ~150, however this nozzle would be very large and little is gained above ~100 Thaddaeus Halsmer, Propulsion

Fuel grain dimension definitions Lunar Lander Propulsion – Isp analysis approach Fuel grain dimension definitions Thaddaeus Halsmer, Propulsion

Thaddaeus Halsmer, Propulsion Lunar Lander Propulsion – fuel grain and chamber sizing approach Choose Empirical value for initial fuel regression rate Initial O/F ratio for optimum Isp Initial propellant mass flow rate Compute required burn surface area Dimensions of fuel grains Diameter is derived from burn surface area found from values in step #1 and chosen fuel grain geometry Thickness is function of burn time and regression rate Compute Chamber dimensions a. Chamber dimensions approximated from fuel grain size and additional room for insulating materials Thaddaeus Halsmer, Propulsion