1. 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

2. Radial Flow Hybrid Engine
H2O2 Tank
Helium Tank
Radial Flow Hybrid Engine
3 ft.
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3. 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

4. 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

5. 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

6. 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

7. Fuel grain dimension definitions
Lunar Lander Propulsion – Isp analysis approach
Fuel grain dimension definitions
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8. 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