SPACE TRANSPORTATION SYSTEMS

Earth Surface to Low Orbit

Figure 4-14 gives specifics of lift vehicles proposed for transport to 500-km orbit. They are: a) the standard Space Shuttle; b) its two solid rocket booster (SRB) heavy-lift-vehicle (HLLV) derivative, obtained by replacing the Shuttle Orbiter with a payload fairing and by packaging the three Space Shuttle main engines (SSME) in a recoverable ballistic-entry body, and c) the four SRB Shuttle derivative, with four SSMEs.

Transport Beyond Low Earth Orbit

The nominal mission is the round-trip from low Earth Orbit (LEO) to L5 with of 4084 m/s for a one-way transfer. The NERVA nuclear rocket gives a specific impulse (Isp) of 825 s, or, with operational cool-down losses, 800 s. The SSME has Isp of 460 s. But with a 6:1 mixture ratio and extraterrestrial oxygen resupply available at L5, the effective Isp for the nominal mission is raised to 721 s. Consequently the SSME was selected. Its characteristics are as follows:

- Thrust - 2.09 MN
- Emergency power - 109 percent
- Chamber pressure - 20.4 MPa
- Area ratio - 77.5 (1975)
- Specific impulse - 460 s
- Mixture Ratio - 6.0:1
- Length - 4.24 m
- Diameter - 2.67 m X 2.41 m, powerhead, 2.39 m nozzle exit
- Life - 7.5 hr; 100 starts
- Weight - 2869 kg

Electric-propulsion technology rests upon the use of the 30-cm Kaufman thruster and its derivatives. Nominal characteristics when used with mercury propellant are as follows:

- Thrust - 0.14 N
- Specific impulse - 3000 s
- Input power - 2668 W
- Power efficiency - 79.3 percent
- Propellant utilization efficiency - 92.2 percent
- Beam current - 2 A
- Beam potential - 1058 V

Because mercury cannot be obtained from the
Moon, it would be advantageous to use another propellant. For use with
propellants other than mercury, there is the relation, 1/2 (thrust) X
(exhaust vel.) = (power), and exhaust velocity scales as:

Ionization potentials as high as 15 eV are admissible, when
operating with propellants other than mercury. Gaseous propellants are of
interest because they obviate the need for heating the thrust chamber to
prevent condensation of propellant. The 30-cm thruster has been run at
high efficiency with xenon, krypton, and argon. Use of oxygen is of
interest because of its ready availability and moderate ionization
potential (13.6 eV) and molecular weight (32). Its use would require the
cathode and neutralizer element to be of platinum to resist oxidation. It
would be preferable to use large numbers of such thrusters rather than to
develop very large single thrusters; a 10,000-t vehicle accelerated at
10^-5 g would require 6000 mercury thrusters or 20,000 oxygen thrusters. See Table 4-17
for estimating factors.

TABLE 4-17 (gif format)

To | |||||
---|---|---|---|---|---|

From | Lunar parking orbit | Low Earth orbit | L4/L5 | Geosynch.orbit | Lunar surface |

Lunar (1) parking (2) orbit (3) (4) | - - - - |
4084 2.47 1.25 1.20 | 686 1.16 1.03 1.03 |
1737 1.47 1.08 1.08 | 2195 1.63 1.10 - |

Low (1) Earth (2) orbit (3) (4) | 4084 2.47 - 1.20 | - - - - | 4084 2.47 - 1.20 | 3839 2.34 - 1.19 | - - - - |

L4/L5 (1) (2) (3) (4) | 686 1.16 1.03 1.03 | 4084 2.47 1.25 1.20 |
- - - - - | 1737 1.47 1.08 1.08 | - - - - - |

Geo-synch. (1) orbit (2) (3) (4) | 1737 1.47 1.08 1.08 |
3839 2.34 1.23 1.19 | 1737 1.47 1.08 1.08 | - - - - |
- - - - |

Lunar (1) surface (2) (3) (4) | 1895 1.51 1.09 - |
- - - - |
- - - - | - - - - | - - - - |

NOTES:

- Gives v's for impulsive (Hohmann) transfers and may be less than v's for ion propulsion. A 30-percent increase in v has been assumed to account for this effect.
- Gives mass-ratio,, for H2/O2 assuming = 460 s.
- Gives mass-ratio, , assuming an extraterrestrial supply of oxygen.
- Gives mass-ratio, , for ion propulsion.

Table 4-17 gives the following estimating factors for use in space transport where appropriate:

- Transfer v, following recommendations from NASA
- Mass-ratio for H2/O2, Isp = 460 s
- Mass-ratio for H2/O2 with O2 resupply available either at L5 or at the lunar surface; mixture ratio, 6:1
- Mass-ratio for ion propulsion, Isp = 3000 s

For a multi-leg mission, the total is the sum of the individual 's and the total mass ratio is the product of the individual mass ratios. Mass ratio is found from:

The following mass factors express the ratio, (initial mass in LEO)/(payload delivered to destination). Rocket engine uses LH2 /LO2 at 6:1 mixture ratio; structural mass fraction is 0.1; Isp = 460 s.

- Round trip, LEO-L5, vehicle returned to LEO
- No resupply, all propellants carried to LEO: 4.0

- Resupply at L5 for down trip only: 2.83

- Resupply at L5 for both legs of trip: 1.97

- Delivery to the Moon
- One-way flight with single-SSME modular vehicle: 5.06

- Chemical tug LEO-L5-LPO-L5-LEO, with NASA-recommended lunar landing
vehicle based in parking orbit; only LH2 from
Earth, all O2 at L5: 3.34
(LPO refers to lunar parking orbit).

- One-way trip; L5 oxygen used to maximum extent: 3.08

Specifications for a single-SSME modular vehicle assembled in LEO are:

- Initial mass in LEO: 4.16 X 10^6 kg
- Propellant (LH2 /LO 2, 1:6 mixture ratio): 3.03X10^6 kg
- Structural mass: 0.31 X 10^6 kg
- Engine and avionics: 9.1 X 10^3 kg
- Payload: 0.82 X 10^6 kg
- Mass delivered to Moon: 0.87 X 10^6 kg
- Thrust/weight at landing: 1.16
- Estimating factor: 5.06

LEO-LPO: Propellant, 2.49 X 10^6; structure,0.25 X 10^6 kg

LPO-LS:
Propellant, 0.56 X 10^6; structure, 0.055 X 10^6 kg (LS = lunar surface)

The vehicle requires a multi-burn injection mode to reach Earth escape velocity. This offers the possibility of further mass savings since expended tankage can be staged off after each burn. This factor is not considered here.

There is much discussion of reusable lunar transporters. But where propellant must be brought from Earth, the tankage of such a transporter cannot be refilled. The reason is that cryopropellants must be brought to orbit in their own tankage, and zero-g propellant transfer offers no advantages.

The transporter carries 6 standard payload modules, 136,900 kg each, in a hexagonal group surrounding a central core module. This carries engine and propellant for a lunar landing. Hence, the vehicle which lands on the Moon is of dimensions, 25.2 m diam X 30 m long. Additional propellant for Earth escape is carried in modules forward of the payload; four 8.41 m diam X 30 m long modules are required. Total vehicle length in LEO is 60 m.