Space Settlements - A Design Study 1977

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A second version of a passive catcher is a circular disk, 10,000 $m^2$ in area, of crushable material such as rigid foam or bonded, glass wool boards. A payload would penetrate this material dissipating its energy and lodging in the material from which it could be retrieved at a later time. Theoretical analysis shows that a typical payload would penetrate about 1.3 m into FR type polystyrene foam (density of 28.4 $kg/m^3$). The foam catcher could be foamed in place. After collecting for a period of time it could be melted down with a solar furnace; the desired material extracted; and the catcher refoamed in space. It has the advantage of being very simple in conception, but its 500 t of mass is a disadvantage, as is the fact that, at least initially, the plastics for making the catcher would have to come from Earth. Eventually it would be possible to use mostly lunar materials such as bonded glass wool. Like the other passive catcher, the foam catcher requires very high precision in the launchings.

Figure 4-28 — Passive bag-catcher concept.

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APPENDIX M: 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 $\Delta v$ of 4084 m/s for a one-way transfer. The NERVA nuclear rocket gives a specific impulse ($I_{sp}$) of 825 s, or, with operational cooldown losses, 800 s. The SSME has $I_{sp}$ of 460 s. But with a 6:1 mixture ratio and extraterrestrial oxygen resupply available at L5, the effective $I_{sp}$ 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 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 $1/\sqrt{\text{molecular weight}}$. 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.

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

  1. Transfer $\Delta v$, following recommendations from NASA
  2. Mass-ratio for $H_2/O_2$, $I_{sp} = 460$ s
  3. Mass-ratio for $H_2/O_2$ with $O_2$ resupply available either at L5 or at the lunar surface; mixture ratio, 6:1
  4. Mass-ratio for ion propulsion, $I_{sp} = 3000$ s

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

$\Delta v = g_0 I_{sp} \ln \mu$

where $g_0$ = acceleration of gravity = 9.81 $m/s^2$, $\mu$ = mass-ratio.

The following mass factors express the ratio, (initial mass in LEO)/(payload delivered to destination). Rocket engine uses $LH_2/LO_2$ at 6:1 mixture ratio; structural mass fraction is 0.1; $I_{sp} = 460$ s.

TABLE 4-17 — ESTIMATING FACTORS FOR SPACE TRANSPORT

| Mission | $\Delta v$, m/s (1) | $\mu$ (2) | $\mu$ (3) | $\mu$ (4) | | :--- | :--- | :--- | :--- | :--- | | LEO to L5 | 4084 | 2.47 | 1.78 | 1.18 | | L5 to LEO | 4084 | 2.47 | 1.78 | 1.18 | | L5 to LPO | 2370 | 1.69 | 1.40 | 1.10 | | LPO to L5 | 2370 | 1.69 | 1.40 | 1.10 | | LPO to LS | 2100 | 1.59 | 1.35 | — | | LS to LPO | 2100 | 1.59 | 1.35 | — |

NOTES: (1) Gives $\Delta v$'s for impulsive (Hohmann) transfers and may be less than $\Delta v$'s for ion propulsion. A 30-percent increase in $\Delta v$ has been assumed to account for this effect. (2) Gives mass-ratio, $\mu$, for $H_2/O_2$ assuming $I_{sp} = 460$ s. (3) Gives mass-ratio, $\mu$, assuming an extraterrestrial supply of oxygen. (4) Gives mass-ratio, $\mu$, for ion propulsion.

I. Round trip, LEO-L5, vehicle returned to LEO

  • a. No resupply, all propellants carried to LEO: 4.0
  • b. Resupply at L5 for down trip only: 2.83
  • c. Resupply at L5 for both legs of trip: 1.97

II. Delivery to the Moon

  • a. One-way flight with single-SSME modular vehicle: 5.06
  • b. Chemical tug LEO-L5-LPO-L5-LEO, with NASA-recommended lunar landing vehicle based in parking orbit; only $LH_2$ from Earth, all $O_2$ at L5: 3.34 (LPO refers to lunar parking orbit).
  • c. 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: $4.16 \times 10^6$ kg Propellant ($LH_2/LO_2$, 1:6 mixture ratio): $3.03 \times 10^6$ kg Structural mass: $0.31 \times 10^6$ kg Engine and avionics: $9.1 \times 10^3$ kg Payload: $0.82 \times 10^6$ kg Mass delivered to Moon: $0.87 \times 10^6$ kg Thrust/weight at landing: 1.16 Estimating factor: 5.06

These characteristics assume a two-stage operational mode, with empty tankage staged off (jettisoned) upon reaching lunar parking orbit (LPO), prior to final descent. The propellant and tankage masses are as follows: LEO-LPO: Propellant, $2.49 \times 10^6$; structure, $0.25 \times 10^6$ kg LPO-LS: Propellant, $0.56 \times 10^6$; structure, $0.055 \times 10^6$ 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.

Figure 4-29 — Altitude distribution of HCl deposition rate from a Space Shuttle launch vehicle (A-JPL, 1975, B-Oliver, 1973).

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TABLE 4-18 — ANTICIPATED LAUNCH RATES IN LAUNCHES PER YEARᵃ

| Year | Launches | | :--- | :--- | | 1 | 100 | | 2 | 200 | | 3 | 400 | | 4 | 600 | | 5 | 800 | | 6 | 1000 |

ᵃ No provision for launching propellant is included in these numbers; its inclusion will more than double the annual launch requirements.

APPENDIX N: IMPACT OF EARTH LAUNCH VEHICLES ON THE OZONE LAYER

Hydrogen chloride gas (HCl) produced from the exhaust of the Space Shuttle booster motor (see fig. 4-29) dissociates to produce free chlorine which, in turn, reacts to remove ozone from the stratosphere by the following catalytic reactions:

$Cl + O_3 \rightarrow ClO + O_2$ $ClO + O \rightarrow Cl + O_2$

One-dimensional models of HCl deposition, vertical transport, and chemical production and removal of participating trace stratospheric constituents (OH, O, O(¹D), $O_3$, $CH_4$, $H_2$, and NO) have been used by NASA to simulate ozone depletion (refs. 45, 46). Steady-state solutions were obtained simulating 60 shuttle launches per year given that the emissions were spread uniformly in the horizontal over a hemisphere and over a 1000-km wide zone. The levels of ozone reduction computed were about 0.3 percent and 1.0 percent, respectively. More recently, Whitten has revised the ozone depletion calculation for the hemisphere downward to less than 0.1 percent (personal communication, July 1975). Launch rates that might be anticipated for the initial colonization program are shown in table 4-18.

The reduction in ozone concentration in the upper levels of the atmosphere allows the molecular oxygen dissociating radiation to penetrate lower before producing ozone; hence, a primary effect is a downward shift in the ozone distribution. A reduction in the total ozone concentration, which would appear to be very minor, results only as a secondary effect.

Advanced launch vehicles using liquid oxygen-liquid hydrogen ($LOX-LH_2$) propellants above 30 km would eliminate the emission of hydrogen chloride into the stratosphere; however, there are also potential problems with hydrogen fuel which produces water. Water is dissociated as:

$H_2O + O(^1D) \rightarrow 2OH$

in which the $O(^1D)$ results from ozone photolysis at wavelengths shorter than 310 nm. The OH reacts with odd oxygen in a catalytic cycle:

$OH + O \rightarrow H + O_2$ $H + O_3 \rightarrow OH + O_2$

However, compared to the 2 ppm of water in the stratosphere, increases due to hydrogen combustion may be negligible. Further study of the problem is required (R. Whitten, NASA-Ames Research Center, personal communication, August 1975).

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