1. AAE450 Spring 2009 Lunar Lander Propulsion System 100g, 10kg and Large Payload cases Thaddaeus Halsmer Thursday, April 9, 2009 1.Lunar Lander propulsion system (final presentation slide) 2.Lunar Lander Propulsion backup slides Thaddaeus Halsmer, Propulsion

2. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion 1 H 2 O 2 Tank Radial Flow Hybrid Engine Helium Tank 3 ft.

3. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion 2 (2) (3) (4) (1) Table 1 Engine performance parameters Engine No.Payload case/DescriptionF_max/min [N]tb [s] 110 kg/hop engine 2x192 (avg.)134.5 2100 g/main engine1100/110198.6 310 kg/main engine1650/165190.4 4Arbitrary/main engine27000/2700250.2 Stick is 6.5 feet high, same as a standard doorway Lunar Lander Propulsion – Engine Specifications

4. AAE450 Spring 2009 SV01 SV02 High Pressure Helium Tank HV01 REG CK01 CK02 MOV F01 H 2 O 2 Tank HV02 RV01 Thaddaeus Halsmer, Propulsion 3 Lunar Lander Propulsion –fluid system diagrams SV01 SV02 High Pressure Helium Tank HV01 REG CK01 CK02 MOV F01 H 2 O 2 Tank HV02 RV01 SV04 SV03SV05 100g and Large payload cases10kg payload case

5. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion Figure X: Propellant mass vs. I sp trade 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 4

6. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion 5 As area ratio, ε, increases M nozzle increases, but I sp increases also As I sp increases M prop decreases for a given thrust and burn time Wrote Matlab script that used Matlab CEA interface to compute multiple I sp ’s for different area ratio’s and the corresponding M prop and M nozzle 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 Lunar Lander Propulsion - Nozzle area ratio and mass optimization  Used CEA to compute I sp for given nozzle area ratio All other inputs constant  Empirical nozzle mass equation

7. AAE450 Spring 2009 Fuel grain dimension definitions Lunar Lander Propulsion – I sp analysis approach Thaddaeus Halsmer, Propulsion 6

8. AAE450 Spring 2009 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 I sp 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 a. Chamber dimensions approximated from fuel grain size and additional room for insulating materials Thaddaeus Halsmer, Propulsion 7