Lunar Lander Propulsion System 100 g 10 kg

Lunar Lander Propulsion System 100 g, 10 kg and Large Payload cases Thaddaeus Halsmer Thursday, April 9, 2009 1. Lunar Lander propulsion system (final presentation slide) 2. Lunar Lander Propulsion backup slides AAE 450 Spring 2009 Thaddaeus Halsmer, Propulsion

H 2 O 2 Tank Helium Tank Radial Flow Hybrid Engine 3 ft. AAE 450 Spring 2009 1 Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion – Engine Specifications Table 1 Engine performance parameters Engine No. Payload case/Description F_max/min [N] 1 10 kg/hop engine 2 x 192 (avg. ) 2 100 g/main engine 1100/110 3 10 kg/main engine 1650/165 4 Arbitrary/main engine 27000/2700 tb [s] 134. 5 198. 6 190. 4 250. 2 (4) (3) (1) (2) Stick is 6. 5 feet high, same as a standard doorway AAE 450 Spring 2009 2 Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion –fluid system diagrams High Pressure Helium Tank HV 01 SV 01 REG SV 02 CK 01 RV 01 H 2 O 2 Tank F 01 HV 02 MOV CK 02 SV 03 100 g and Large payload cases SV 04 SV 05 10 kg payload case AAE 450 Spring 2009 3 Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion - Propellant/Propulsion system selection Selection Criteria: 1. Thrust a. min/max b. throttling 2. Dimensions a. Short and fat 3. Mass – minimize 4. Propellant storability 5. Purchase/development costs 6. High Reliability Figure X: Propellant mass vs. Isp trade AAE 450 Spring 2009 4 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 AAE 450 Spring 2009 5 Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion – Isp analysis approach Fuel grain dimension definitions AAE 450 Spring 2009 6 Thaddaeus Halsmer, Propulsion

Lunar Lander Propulsion – fuel grain and chamber sizing approach 1. Choose a. Empirical value for initial fuel regression rate b. Initial O/F ratio for optimum Isp c. Initial propellant mass flow rate Compute required burn surface area 2. Dimensions of fuel grains a. Diameter is derived from burn surface area found from values in step #1 and chosen fuel grain geometry b. Thickness is function of burn time and regression rate 3. Compute Chamber dimensions approximated from fuel grain size and additional room for insulating materials AAE 450 Spring 2009 7 Thaddaeus Halsmer, Propulsion