Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Advantages  Economies of scale  Can use today’s rockets  Rapid generation of custom propulsion stages  Engine reuse possible  Does not require propellant transfer in microgravity

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Scalability of number

Engine design essentially independent of burn duration Use identical engines, just burn longer for greater impulse Optimize propellant storage and pressurization system for given engine Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Propellant N2O4/UDMH Pros • Space storable • Heritage • Hypergolic • Liquid temperature overlap • Cheaper than MMH • Mass ratio nearly 2

Cons • Low Isp (~316 s) • Toxic

The Long March 2F uses N2O4/UDMH for all stages

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Analytical Strategy Mass estimates for one building block Operating pressure Operating temperature Propellant mass

Propellant volume Check shroud diameter

Helium volume

Propellant tank mass

Helium mass

Helium tank mass

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Analytical Results

(N2O4)

Mass estimates for one building block

60 psi

Operating pressure Operating temperature

290 K

Propellant tank mass

Propellant volume

Propellant mass

Check shroud diameter

20,000 kg

86.5 kg

d = 2.98 m  OK Helium volume

Helium mass

169 liters

(aluminum alloy)

19.8 kg (including solid propellant grains)

Helium tank mass

23.4 kg (titanium liner, composite overwrap)

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Estimates for other components

(N2O4)

Mass estimates for one building block

Item

Mass (kg)

Helium regulators (2)

6

Exoskeleton truss

100

Rails

75

Powered liquid disconnects (2)

40

Fluid lines

10

MMOD shielding

30

Residual propellant (2%)

400 Total dry mass = 785.7 kg Total mass = 20,785.7 kg  OK

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Analytical Results

(UDMH)

Mass estimates for one building block

60 psi

Operating pressure Operating temperature

290 K

Propellant tank mass

Propellant volume

Propellant mass

Check shroud diameter

17,778 kg

140.8 kg

d = 3.51 m  OK

(mass ratio of 2.25)

Helium volume

Helium mass

275 liters

(aluminum alloy)

22.7 kg (including solid propellant grains)

Helium tank mass

35.0 kg (titanium liner, composite overwrap)

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Estimates for other components

(UDMH)

Mass estimates for one building block

Item

Mass (kg)

Helium regulators (2)

6

Exoskeleton truss

110

Rails

90

Powered liquid disconnects (1)

20

Fluid lines

10

MMOD shielding

40

Residual propellant (2%)

356 Total dry mass = 830.5 kg Total mass = 18,608.5 kg  OK

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Results Based on the estimates given, the following payloads could be delivered to Trans Lunar Injection (TLI): Earth departure stage

No. fuel tanks

No. oxidizer tanks

Space storable?

Payload to TLI*

Modular Building Blocks 1

2

yes

23,300 kg

Modular Building Blocks 2

4

yes

50,500 kg

Modular Building Blocks 3

6

yes

77,800 kg

Delta IV Heavy

no

11,000 kg

Falcon Heavy

no

16,000 kg

Saturn V

no

47,000 kg

*Payload estimates for Modular Building Blocks were found using the rocket equation and assume a basic Hohmann transfer along with the following values: specific impulse of 316 seconds; delta-v of 3.18 kilometers per second; engine module mass of 4,000 kilograms; usable propellant fraction of 92%)

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Extensions of the concept Integrated fuel/oxidizer (Increases scalability)

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Extensions of the concept Hydrogen propellant

Deliver departure stage to LEO with hydrogen only

Rendezvous with depot to pick up oxygen building blocks

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

References Bilstein, Roger E. Stages to Saturn. 1996. Gainesville: University Press of Florida, 2003. N. pag. Print. Brown, Charles D. Elements of Spacecraft Design. Ed. John S. Przemieniecki. Reston: American Institute of Aeronautics and Astronautics, 2002. N. pag. Print. AIAA Education Series. Chaffee, Norman. Personal interview. 7 Oct. 2011. Eichstadt, Frank. Personal communication. 15 Mar. 2011. Grush, Gene. Telephone interview. 17 Oct. 2011. Huzel, Dieter K., and David H. Huang. Modern Engineering for Design of Liquid-Propellant Rocket Engines. Ed. A. Richard Seebass. Washington DC: American Institute of Aeronautics and Astronautics, 1992. N. pag. Print. Ring, Elliot. Rocket Propellant and Pressurization Systems. Englewood Cliffs: Prentice-Hall, 1964. N. pag. Print. Sutton, George P., and Oscar Biblarz. Rocket Propulsion Elements. 8th ed. Hoboken: John Wiley & Sons, 2010. N. pag. Print.

Sasakawa International Center for Space Architecture, University of Houston College of Architecture

Thank you For Listening.

Feedback appreciated. James Doehring [email protected] Sasakawa International Center for Space Architecture, University of Houston College of Architecture

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