09ICES-0072 Starship Life Support Copy

09ICES-0072

Starship Life Support

Harry Jones NASA Ames Research Center

Copyright © 2009 SAE International

ABSTRACT

The design and mass cost of a starship and its life support system are investigated. The mission plan for a multigenerational interstellar voyage to colonize a new planet is used to describe the starship design, including the crew habitat, accommodations, and life support. Only current technology is assumed. Highly reliable life support systems can be provided with reasonably small additional mass, suggesting that they can support long duration missions. Bioregenerative life support, growing crop plants that provide food, water, and oxygen, has been thought to need less mass than providing stored food for long duration missions. The large initial mass of hydroponics systems is paid for over time by saving the mass of stored food. However, the yearly logistics mass required to support a bioregenerative system exceeds the mass of food solids it produces, so that supplying stored dehydrated food always requires less mass than bioregenerative food production. A mixed system that grows about half the food and supplies the other half dehydrated has advantages that allow it to breakeven with stored dehydrated food in about 66 years. However, moderate increases in the hydroponics system mass to achieve high reliability, such as adding spares that double the system mass and replacing the initial system every 100 years, increase the mass cost of bioregenerative life support. In this case, the high reliability half food growing, half food supplying system does not breakeven for 389 years. An even higher reliability half and half system, with three times original system mass and replacing the system every 50 years, never breaks even. Growing food for starship life support requires more mass than providing dehydrated food, even for multigeneration voyages of hundreds of years. The benefits of growing some food may justify the added mass cost. Much more efficient recycling food production is wanted but may not be possible. A single multigenerational interstellar voyage to colonize a new planet would have cost similar to that of the Apollo program. Cost is reduced if a small crew travels slowly and lands with minimal equipment. We can go to the stars!

INTRODUCTION

A preliminary analysis of the life support system for a multigeneration interstellar starship is developed. It

includes mission plan analysis needed to define the life support requirements. For humans to go to stars while constrained by current physics, the travel speed must be significantly less than the speed of light. The travel time necessarily extends to many decades or centuries. Multiple generations of humans will make the voyage.

Two concepts indicate the value of investigating life support for a multigeneration starship. It is often assumed that growing the large quantity of food needed will be necessary on a multigeneration starship voyage. Bioregenerative life support, using crop plants in a closed hydroponics system to provide food, oxygen, and water, appears cost effective on decades-long missions. (Jones, 2006-01-2082) Recently it has been found that it is possible to develop ultrareliable physicochemical life support systems to recycle spacecraft oxygen and water with a relatively small cost in spacecraft mass. Ultrareliable bioregenerative support equipment can be similarly designed. We can build life support systems to take us to the stars. (Jones, 2008-01- 2160) Consideration of starship life support clarifies the effect of reliability mass provision on the attractiveness of bioregenerative life support.

This paper first provides an overview of interstellar human colonizing missions that includes new planets and the travel time, propulsion systems, and a possible mission scenario. This is followed by a discussion of the appropriate crew size. Then reliability and Equivalent System Mass (ESM) considerations are summarize. The mass cost of physicochemical and bioregenerative life support systems is investigated and their behavior with increasing mission duration described. To assess the context and reasonableness of the life support conclusions, the crew habitat and interstellar spacecraft mass requirements are investigated and an overall description of the starship given. Finally the results and heir implications are discussed.

POSSIBLE NEW PLANETS FOR HUMANS

Stars that are not part of a multiple star system are expected to have planets, which conserve the angular momentum of the originating gas and dust cloud. In recent years hundreds of extrasolar planets have been observed. Most are large gas giants close to their primaries, but Earth-like planets may soon be found.

NEW PLANETS’ TRAVEL DISTANCE, SPEED, AND time - The nearest star system that might have habitable planets is Alpha Centauri at 4.3 light years distance, although it is a triple star. There are only ten stars within 9 light years, and only three of these are single stars. However, there are 55 stars within 16 light years, including 31 single stars. (Kraus, p. 2-5, 6) The earliest human interstellar colonization voyages will probably traverse from 4.3 to 16 light years distance, with 10 light years a reasonable estimate.

Traveling at or near c, the speed of light, appears impossible according to current physics. If a multigenerational starship achieved 0.10 c, a 10 light year trip would take 100 years, and at 0.01 c, it would take 1,000 years. A voyage at 0.001 c would take 10,000 years and appears practically impossible. Travel speeds of 0.01 to 0.10 c seem needed for interstellar colonization.

ROCKET PROPULSION AND TRAVEL SPEED

Chemical rockets cannot take us to the stars, but nuclear rockets using currently understood technology could. A

rocket is thrust forward by expelling mass backward at

high speed. The rocket’s final velocity is determined by the amount of mass expelled and its exhaust speed.

The rocket equation for final velocity, V, is

V = E ln (Mt/Mv)

where E is the rocket exhaust speed, ln is the natural logarithm, Mt is the total initial mass of the rocket propulsion fuel plus Mv, the vehicle mass after all the fuel is burned. The natural logarithm, ln, is a slowly increasing function of its argument (Mt/Mv). Therefore, achieving a rocket final velocity, V, that is much larger than the rocket exhaust speed, E, requires a propulsion mass (Mt – Mv) much larger than the vehicle mass (Mv). See Figure 1.

Doubling V/E from 1.5 to 3.0 requires a little more than four times the mass, since Mt/Mv increases from 4.5 to 20.

The inverse function is

Mt/Mv = exp (V/E)

See Figure 2.

Figure 1. Rocket final velocity/rocket exhaust speed (V/E) versus total mass/vehicle mass (Mt/Mv).

V/E=ln(Mt/Mv)

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Figure 2. Rocket total mass/vehicle mass (Mt/Mv) versus rocket final velocity/rocket exhaust speed (V/E).

It is clear that increasing V/E beyond 2 or 3 incurs an increasingly large mass cost. To achieve the required rocket speed V of 0.01 to 0.1 c, E must be only a factor of 2 or 3 less.

CHEMICAL ROCKETS - Current chemical rockets cannot achieve the required final velocity. E is not more than 4,500 m/s. (HSF, p. 768) For Mt/Mv = 20, V = 3 E, or 13,500 m/s. The speed of light, c, is 3 * 108 m/s, so V = 4.5 * 10-5 c. The highest practical velocity for a chemical rocket is less than one ten-thousandth of the speed of light, so that crossing one light year takes more than ten thousand years. At V = 4.5 * 10-5c, the time to Alpha Centauri, 4.3 light years away, is 95,000 years.

There is essentially no chance of successfully completing an interstellar mission using chemical rockets. Much higher velocities are needed and they can be provided by nuclear propulsion.

NUCLEAR PROPULSION - Nuclear rocket propulsion has been studied but not developed, for example in Project Orion and the nuclear salt-water rocket. Project Orion was conducted around 1960 and developed the concept of nuclear pulse propulsion. Small nuclear bombs are exploded behind the spacecraft, accelerating inert mass placed between the bomb and the spacecraft into a spring mounted pusher plate on the rear of the spacecraft. Project Orion’s pulsed nuclear power can provide higher exhaust speed than chemical and most other nuclear propulsion systems and can be constructed entirely with current technology.

The plasma debris exhaust speed can range up to 3 * 107 m/s = 0.10 c. Orion nuclear pulse propulsion using atomic fission explosions can achieve a rocket speed of 0.03 to 0.05 c. A thermonuclear fusion Orion starship

might reach 0.08 to 0.1 c. (Wikipedia, 7/25/08, http://en.wikipedia.org/wiki/Project_Orion_%28nuclear_propulsion%29) (Czysz and Bruo, 2006) A more conservative estimate of Orion top speed is 0.01 c. (Dyson, in Hart and Zuckerman, p. 41) The maximum spacecraft velocity depends on the propellant mass, as defined by the rocket equation and Figures 1 and 2.

A nuclear salt-water rocket was proposed by Robert Zubrin. The exhaust speed can reach 4.7 * 106 m/s = 0.016 c, and the rocket speed can reach 0.036 c. A nuclear salt-water rocket is somewhat similar to the Orion propulsion system, except that it generates continuous rather than pulsed thrust. It may be much smaller than the smallest Orion designs. Project Orion nuclear pulse propulsion systems may mass hundreds of tons due to the mass of the shock-absorber system and the minimum size of efficient nuclear explosives. (Wikipedia, 7/25/08, http://en.wikipedia.org/wiki/Nuclear_salt-water_rocket) (Zubrin, 1991.)

Nuclear propulsion systems seem capable of exhaust speeds of 0.01 to 0.1 c, and final rocket velocities several times higher. They are achievable with current or near future technology. The time for a multigenerational starship to travel 10 light years is 100 years at 0.10c and 1,000 years at 0.01 c. This range is 4 to 40 generations, assuming a generation is 25 years.

MULTIGENERATION INTERSTELLAR MISSIONS

The mission scenario determines starship life support requirements. Suppose a target planet is identified circling a star at 10 light years distance. The spacecraft is assembled in Earth orbit and boosted to the outer solar system, where the foundational crew joins. Nuclear

Mt/Mv = exp (V/E)

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http://en.wikipedia.org/wiki/Project_Orion_(nuclear_propulsion)

http://en.wikipedia.org/wiki/Project_Orion_(nuclear_propulsion)

http://en.wikipedia.org/wiki/Nuclear_salt-water_rocket

propulsion is used with an exhaust speed of 0.01 c and a final velocity of 0.02 c. V = 2 E, and Mt/Mv = 7.4. The spacecraft is gradually accelerated to maximum velocity and reaches the vicinity of the new planet on 10 ly/0.02 ly/y = 500 years, or 20 generations.

If the same nuclear propulsion system was used to decelerate the starship from 0.02 c to orbit around the new planet, the propellant required would be Mt/Mv = 7.4 times the mass of the spacecraft. To send this propellant to the vicinity of the new planet would require 7.4 times as much propellant, making the original propellant requirement equal to 7.42 = 54.8 times the original spacecraft mass. This is obviously prohibitively costly in terms of mass in Earth orbit.

However, by the time of arrival at the new planet, much of the starship mass will consist of used spares and consumables, waste and trash, that need not be decelerated. A small part of the starship may orbit the new planet and support colonization. To minimize mass, it is assumed that only small descent vehicles are detached from the interstellar spacecraft and landed. These landing craft could be similar to the Apollo/Orion capsules, with parachutes and ablative shields in addition to retro rockets and navigation. The interstellar spacecraft speeds through and past the new planetary system.

Major aspects of the mission remain to be determined; the spacecraft mass, the habitat design, and especially the life support system, which is expected to be the largest and most costly element of the mission. The size of the crew is considered next.

NUMBER OF CREWMEMBERS

The probability of colonizing success depends on the number of the starship crew that land on the new planet. A very small crew, even a few individuals, might succeed in colonizing a new planet if they had use of reproductive and other technology and if the environment was very favorable. Clearly a much larger crew would have a much better chance of success. But increasing the crew size gives diminishing returns, as measured by the additional probability of success per extra crewmember. Determining the number of crewmembers requires

balancing higher individual mission probability of success against more efficient allocation of crew and mass resources between missions. The number of crewmembers determines the mission mass, but the life support design trade-offs can be made independently, using mass per crewmember.

ESTIMATED PROBABILITY OF COLONIZING SUCCESS AND NUMBER OF CREW - Suppose that a crew size of N crewmembers is maintained during the multigenerational voyage and is landed on the new planet as the starship passes through the extrasolar planetary system. A family with three living generations would have at least six members. It seems plausible that 2 or 3 families with 12 to 18 members would have roughly a 0.50 probability of success.

The probability of success for a crew of N = 15 is assumed to be 0.5, 50 percent. Their probability of failure is also 0.5. If N = 30, it would be possible to land two crews of 15, each with the same expected 0.5 failure rate. The probability that both crews fail is the original failure rate for 15 crew squared, which is equal to 0.52 = 0.25. The probability of success for a crew of 30 is the 1.00 – 0.25 = 0.75. For N = 60, there could be four independent crews of 15, and the failure rate would be 0.54 = 0.937. If a crew of 15 has a 50 percent chance of successful colonization, very much larger crews are not necessary to achieve a good probability of success.

It can be assumed the starship landing crew has five families and 30 individuals. The crew number could easily vary by a factor of 2, down to 15 or up to 60. The crew number seems very unlikely to vary by a factor of 10, down to 3 or up to 300.

The failure rate for N crew has the form of function(N) = ConstantN, since doubling N squares the failure rate. The probability of colonizing success Pc = 1 - ConstantN. The probability of success for a crew of N = 30 is 0.75, as computed above. The Constant is determined by Pc = 1 - Constant30 = 0.75, Constant30 = 0.25, Constant = exp[(ln 0.25)/30] = 0.955. The probability of colonizing success is

Pc = 1 - 0.955N

This is shown in Figure 3.

Figure 3. Probability of colonizing success, Pc, versus number of crew, N.

Pc is 0.5 for N = 15, 0.75 for N = 30, 0.875 for N = 45, and 0.937 for N = 60. For N = 100, Pc = 0.99. Pc is 0 for N = 0 and is shown dotted for N = 0 to 5, considered not feasible.

EFFICIENT CREW USE INDICATES SMALLER CREW SIZE - Figure 3 and the equation for Pc, the probability

of colonizing success, indicate that there are diminishing returns with increasing number of crew. The increase in probability of colonizing success is less for each added crewmember. Conversely, the greatest probability of success per crewmember is achieved at the lowest number of crew. This is shown in Figure 4.

Figure 4. Probability of colonizing success divided by number of crew, Pc/N, versus N

Pc = 1 - 0.955^N

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If the crew size is large enough to give a reasonably good chance of success, it is not an efficient use of crew to greatly increase the crew size. Suppose as above that the probability of success for a crew of 30 crewmembers is 0.75, 75 percent. If the total crew size is doubled to 60, the probability of success will be larger. Suppose the crew of 60 is divided into two crews of 30, and that each crew of 30 mounts a completely independent colonization effort on the same planet. Each crew of 30 has a 1 – 0.75 = 0.25 probability of failure, and the probability that both fail is 0.25 squared, 0.0625. The probability of colonizing success with two independent crews of 30 is then 1 – 0.0625 = 0.9375. This is an underestimate of the potential benefit of the larger crew, since two independent efforts lack the synergism of the larger crew working together. But even if the larger crew lands and works together, and no matter how large the crew and how synergistic their interaction, their probability of success cannot exceed 1.0, a certainty. The options presented are to send a crew of 30 with a 75 percent chance of success or a crew of 60 and have nearly a 100 percent chance of success. But suppose instead we send one crew of 30 to one planet and the other crew of 30 to a completely different planet. Each crew still has a 75 percent chance of success, so the expected number of successful colonizations is 1.5. There is a 0.0625 chance that both fail, a 0.375 chance only one succeeds, and a 0.5625 chance both succeed. If the probability of success for a crew of 30 is 0.75, doubling the crew to 60 can gain at most another 0.25 chance of success, but sending a second crew of 30 to different planet gains another 0.75 percent chance of a second success. Either way, there is a 0.9375 chance of at least one successful colonization.

As long as the probability of colonizing success for a crew of N is greater than 0.5, it is more efficient to send out a second independent expedition of the same size N than to double the original crew to 2 N. No matter how large a single crew, and how valuable its group synergism, its expected number of colonized planets cannot exceed one. Two crews each with probability of success of 0.5 + x (where 0 < x < 0.5) have an expected number of successfully colonizations of 1 + 2 x planets.

HIGH PROBABILITY OF COLONIZING SUCCESS REQUIRES LARGER CREW SIZE - It is unfortunately clear that achieving a reasonably high probability of success for a particular mission will require a large crew size and reduced efficiency in use of the potentially available crew. Many small high risk missions give the highest number of expected successes but are unacceptably wasteful of human potential. The compromise between high probability of success for large missions and the greater expected number of successes with small missions requires a balance of important conflicting values.

If the available resources are lavish, many large missions can be sent and the problem disappears. In the likely case that resources are scarce, the first choice will be between one mission with higher probability of

success and two with lower probability. Suppose that a crew of N has probability of failure f and probability of success 1-f. A crew of 2 N will have a probability of success of 1 - f2. Similarly, a crew of N/2 will have a probability of success of 1 – f1/2. The probability of colonizing success for X crew, Pc(X), is shown in Table 1 for X = N/2, N, 2 N over a range of f = Pf(N), the probability of failure for a crew of N. The expected number of colonizing successes, #, is also shown for crews of N/2, N, and 2N.

Table 1. Probability of colonizing success, Pc(), and expected number of successes, #.

Pf(N) Pc(N/2)

# for N/2

Pc(N)

# for N

Pc(2 N)

# for 2 N

f= 1 – f

1/2= 2

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Pc(N)= 1 – f 2

= 0.5

Pc(2 N)

0.00 1.00 2.00 1.00 1.00 1.00 0.500.05 0.78 1.55 0.95 0.95 1.00 0.500.10 0.68 1.37 0.90 0.90 0.99 0.500.15 0.61 1.23 0.85 0.85 0.98 0.490.20 0.55 1.11 0.80 0.80 0.96 0.480.25 0.50 1.00 0.75 0.75 0.94 0.470.30 0.45 0.90 0.70 0.70 0.91 0.460.35 0.41 0.82 0.65 0.65 0.88 0.440.40 0.37 0.74 0.60 0.60 0.84 0.420.45 0.33 0.66 0.55 0.55 0.80 0.400.50 0.29 0.59 0.50 0.50 0.75 0.38

Table 1 suggests in bold an acceptable range for the probability of mission success. Consider the first row for f = 0.00, where the probability of success is 1.00 for crews of N/2, N, and 2 N members. Clearly N/2 is better, since the N crewmembers will conduct two missions and achieve two successes. For the second row, with f = 0.05, N/2 provides 1.55 expected successes, even though each of the two missions has only 0.78 probability of success. N provides 0.95 expected successes in one attempted mission. The entries that seem most attractive in each row are in bold. For the third row, with f = 0.10, N/2 provides 1.37 expected successes with two missions each having only 0.68 probability of success. This seems slightly better than one mission of N crew with 0.90 expected successes, but the choice is difficult. For the next five rows, f = 0.15 to 0.35, one mission of N crew seems best. Two missions with N/2 have too low individual mission probability of success, while a larger mission of 2 N has a significantly lower expected number of colonizing successes than N crew. For the last three rows, f = 0.40 to 0.50, a larger mission of 2 N seems necessary to gain a sufficient individual mission probability of success, and causes a relatively small reduction in the expected number of successes. This subjective survey suggests that the crew size should be large enough to achieve a probability of colonizing success in the range 0.65 to 0.85, say about 0.75. The expected number of colonizing successes normalized for N crewmembers, #, is always

largest for two missions of N/2 crew, since smaller crews are always more efficient.

Sending a very large crew is wasteful, since the resources could be split to mount two independent missions. Sending a very small crew has high probability of failure, even though it is efficient in resources. The crew should be large enough to have a roughly 0.75 probability of success, to flourish 3 times in 4 tries. It seems plausible that this can be achieved with a crew of 30.

LIFE SUPPORT SYSTEM RELIABILITY

The crew requires a life support system that operates reliably for generations. Reliability can be increased by system design and by providing redundancy and spares. High reliability can be achieved with a relatively small increase in system mass.

Suppose that the mission has a travel time of T years and a constant starship failure rate of l failures per year. The probability that the spacecraft does not fail before time T is the reliability after mission travel time T, R(T). For a constant failure rate, l, the reliability at time T, R(T), is given by the exponential function.

R(T) = e-λT = exp (- λT)

The failure probability, F(T), is the probability that the system does fail before time T.

F(T) = 1 - R(T) = 1 - e-λT

The mean time between failures (MTBF) is 1/l. The expected number of failures during the mission duration T is λT. If the MTBF = 1/λ< T, then λT > 1, and several failures are expected to occur during the mission. On the other hand, if the MTBF = 1/λ>> T, then λT << 1, and a failure is unlikely during the mission. For λT << 1, exp (-λT) ~ 1 – λT, and F(T) ~ 1 – (1 - λT) = λT. For λT << 1, doubling the mission duration T approximately doubles the failure probability, F(T) ~ λT.

IMPROVING RELIABILITY WITH REDUNDANCY- The simplest way to improve system reliability is to provide a second fully redundant system in parallel with the first. If we do this, the system function does not fail unless both redundant systems fail. If the probability that one system fails is 10-2, the probability that both systems fail is 10-2 * 10-2 = 10-4. Dual redundancy doubles the required system mass and squares the failure probability.

Reliability can be further improved by adding more redundant systems. Each time that the total mass is redoubled by again doubling the number of redundant subsystems, the failure probability is squared. If the probability that the original system fails is 10-2, the probability that a three times redundant system fails is 10-2 * 10-2 * 10-2 = 10-6, and the probability that an N redundant system fails is (10-2) N = 10-2 N. To achieve any required probability of failure using redundancy, the

redundancy factor is log (required new lower probability of failure)/log (original single system probability of failure). To go from 10-x for a single system to 10-x N using N redundant systems, the mass increase is –x N/-x = N times. Since N times as many systems have N times as much mass, the mass increase ratio is equal to the redundancy factor, which equals the ratio of the logs of the failure probabilities. The mass required increases slowly for higher reliability.

USING REDUNDANCY WITH INCREASING MISSION DURATION - The above analysis of redundancy assumed that the system failure probability F(T) ~ λ T was unchanging as reliability was increased. Suppose we have a system with an acceptably low failure rate F(T1) ~ λ T2 at time T1 but the mission duration increases to T2. The same approach can sometimes be used. Suppose that a system designed for T1 = 10 years has F(10) = 10-4 ~ λT, so λ~ 10-5. Suppose that the mission duration T2 is increased to 1,000 years. Now F(T2) ~ λT2 = 10-5 * 103 = 10-2. Providing a second parallel system squares the failure probability to F(T2) * F(T2) = 10-4, as before.

Using redundancy does not easily achieve high reliability if the original system has too high a failure probability, so that the original system is unreliable at the extended mission duration. Suppose that a system designed for T1 = 10 years has F(10) = 10-2 ~ λT, so λ~ 10-3. Suppose that the mission duration T2 is increased to 1,000 years. Now F(T2) ~ λT2 = 10-3 * 103 = 1. The approximation F(T) ~ λ T does not apply unless F(T) is small, and providing parallel redundant systems will increase reliability at any level, but highly reliable systems are needed if redundancy is to achieve very high reliability.

Designing high reliability systems requires providing redundancy or spares at the component level.

IMPROVING RELIABILITY WITH SELECTED SPARES - Providing selected component spare parts is more cost and mass effective than implementing full system redundancy. Yakut and Barker determined spare part requirements as a function of reliability and mission duration. The required failure probabilities F(T) were 10-

1, 10-4, and a variable intermediate point near 10-3, for missions of 400 and 2,000 days They found that the mass including spares for a failure probability of 10-4 is usually very close to 2.0 times the mass including spares needed to achieve a failure probability of 10-1for the same mission duration. In seventeen cases, this ratio varied from 1.79 to 2.27 and averaged 2.01. (Yakut and Barker) If the reliability had been increased using full system redundancy rather than selected component spares, going from F(T) = 10-1 to 10-4 would require a mass increase of -4/-1 = 4 times, double the ratio achieved. Because the original systems without spares had failure probability of more than 10-1 at the mission duration, spares were added to achieve 10-1, so the total mass for a failure probability of 10-4 was 2.54 to 3.24 times the original system mass. More spares are needed to achieve the target failure probability for longer

durations. Low failure rates were achieved for multiyear missions. For F(T) = 10-4 and T = 2,000 days = 5.5 years, l = 5 * 10-8/day or 2 * 10-5/year, an MTBF of 50,000 years. (Yakut and Barker)

Using selected component spares is much more mass-effective than providing full system redundancy. Typical system design places the most failure-prone functions in small replaceable components. Providing spares for these components requires less mass than full system redundancy. A mathematical model can show that any required reliability can be achieved merely by doubling the system mass. The system is divided into many components of equal reliability and each component is provided one spare. The reliability increases without limit as the number of components increases. Any reliability is reachable with two times the original mass. Both the practical examples of Yakut and Barker and the mathematical models show that ultrareliable systems can be built using spares by roughly doubling the mass needed for typical 0.90 to 0.99 reliability systems. (Jones, 2008-01-2160)

High reliability can be maintained over very long durations without increasing system mass beyond 3 to 4 times the original system mass. Extended mission duration does not impose a significant system mass penalty through increasing failure rate. The failure rate can be reduced as necessary with small mass penalty. This removes one potential driver toward shorter mission duration, greater spacecraft velocity, and increased propulsion mass. The spacecraft seems assured of arriving safely and functioning nominally at its destination even after very long missions.

MASS ALLOWANCE FOR HIGH RELIABILITY - Increasing reliability requires additional mass. How much should be allowed? The discussion of achieving high reliability noted that reliability can be increased from a moderate level to ultra high reliability by doubling the mass. Initial designs that had low reliability at the longer mission durations required three or more times as much mass. Three times the mass for high reliability seems adequate, only two times optimistic, and no increase unreasonable.

Starship mission durations will extend being 10 or 50 years to 100’s or a thousand years. Even with high reliability, most technical systems have limited life. Automobiles last much more than 5 years, with 20 years not uncommon, but much longer times are impractical, achievable only by well maintained, protected, seldom used antiques. Household appliances and heating/cooling systems have similar lifetimes. Major simple structures, dams and bridges, can have a 100 year mean time to failure, but they may last 1,000 years with occasional repair. This suggests that starship systems should be completely replaced on some schedule, even if they have high reliability.

How much of a starship system should be routinely replaced and how frequently? When the initial mass of a

system is doubled or tripled to achieve high reliability, the approach is not simple full system redundancy. Instead, many duplicates of the least reliable components are provided as spares, with fewer replicas of the more reliable components, and none of very high reliability components, such as structural members. Only the original mass of the system need be routinely replaced, not the doubled or tripled mass of spares provided for high reliability. Replacing the original system mass every 50 years seems adequate, every 100 years optimistic, and never replacing the original system for 100’s or a thousand years unreasonable.

There are thus three possible mass allowances to achieve high reliability.

1. Providing 3 times the original system mass for high reliability and replacing the original system mass every 50 years (3/50).


3. Providing only the original system mass and never replacing the system (1/∞).

Work below will examine the different mass costs of candidate life support systems, including the effects of these three mass allowances for high reliability.

BREAKEVEN EFFECT OF MASS INCREASE FOR HIGH RELIABILITY - The additional mass for reliability will extend the breakeven durations of candidate systems. This is because initially higher mass systems are paid for by their lower continuing logistics mass requirements. Higher initial masses have higher differences that take longer to pay back, and replacing the initial mass costs more for higher initial mass systems.

The mass requirement of a system consists of its initial fixed mass plus its yearly logistics supplies, X kg + p kg/year * t years. Suppose a second system has higher initial mass and lower logistics, Y kg + q kg/year * t years, with Y > X and q < p. The higher initial mass of the second system, Y – X, is paid for by its lower logistics rate, p – q, in T years, where T = (Y-X)/(p-q) is the breakeven date.

Clearly, if the initial fixed masses, X and Y, are tripled, their difference, Y – X, is tripled, and the breakeven date T = (Y-X)/(p-q) is tripled. Allowing more mass for high reliability increases the breakeven date proportionally.

Suppose that the initial fixed masses X and Y, are replaced every 50 years. The mass requirement of the system becomes X kg + X kg/50 years * t years + p kg/year * t years = X kg + (X kg/50 years + p kg/year) * t years. The system replacement adds an additional time dependent logistics factor. Since Y > X, the revised logistics pay back rate, X/50 + p - Y/50 – q, will be lower than the original logistics payback rate, p – q.

Periodically replacing the initial mass increases the breakeven date.

MULTIGENERATIONAL STARSHIP LIFE SUPPORT

This section describes the life support system for a multigenerational starship. The key design decision is the method of providing food, by either provision of stored food or by growing food plants in hydroponic systems.

EQUIVALENT SYSTEM MASS (ESM) - In life support analysis, the Equivalent System Mass (ESM) is used to compare the launch cost of alternate systems that perform the same functions. ESM includes the mass of the system itself and the logistic supplies it requires and adds to these the masses needed to provide the pressurized volume, power, and thermal cooling that the system requires to operate.

Computation of ESM - Equivalent System Mass (ESM) is based on the system mass, m, volume, v, power, p, cooling, c, logistics mass per year, l, and d, the mission duration. ESM determines the system mass launch cost.

ESM (m, v, p, c, l, d) = m + v * me(v) + p * me(p) + c * me(c) + l * d

me(v) is the mass equivalent of volume, me(p) is the mass equivalent of power, and me(c) is the mass equivalent of cooling. (Jones, 2003-01-2635, Appendix C)

Volume mass equivalent - It is difficult to estimate spacecraft structural mass without doing a detailed design. On average, structures and mechanisms account to about 21 percent of spacecraft dry mass. (HSF, p. 669) Approximately this value was used in a lunar south pole mission and a Mars base mission. (HSF, pp. 444, 998) The percentage is typically greater for a crewed spacecraft, because of the need to contain atmospheric pressure. The Spacelab flown by the space shuttle had 27 percent structural mass. (HSF, p. 669)

The life support BVAD gives a volume mass equivalent of me(v) = 215.5 kg/m3, for a shielded inflatable module on a Mars transit mission. (BVAD, p. 11) The mass per unit volume for conventional structures can be estimated from earlier and suggested designs for long missions. Skylab had a mass of 77,088 kg and a habitable volume of 10,000 ft3 (283.2 m3). This is a mass per unit volume of 272 kg/m3. (http://en.wikipedia.org/wiki/Skylab) The International Space Station (ISS) has a mass of 2.78 * 104 kg, and a habitable volume of 15,000 ft3 (424.8 m3). This is a mass per unit volume of 655 kg/m3. (http://en.wikipedia.org/wiki/International_Space_Station) The Russian Mir has a mass of 1.24 * 104 kg, and a habitable volume of 350 m3. The mass per unit volume of 354 kg/m3. (http://en.wikipedia.org/wiki/Mir) A Mars base mission by 6 crew for 680 days would have a total landed mass of 45,284 kg and a volume of 85.5 m3.

(HSF, pp. 998, 997) The mass per unit pressurized volume is 530 kg/m3.

The average mass per unit volume of the above four conventional metal designs is 450 kg/m3. Assuming that the structures and mechanisms are 27 percent of the spacecraft dry mass, they would require 122 kg/m3.

Power mass equivalent - The BVAD gives a power mass equivalent of me(p) = 237 kg/kW, for a solar power on a Mars transit mission. (BVAD, p. 11) Hanford earlier computed a power mass equivalent of me(p) = 90.7 kg/kW, for nuclear power. The power is provided by an SP100 nuclear reactor using Brayton conversion with shielding and cooling. The SP100 has a mass of 9,070 kg, produces 100 kW, and has a life of 15 years. About half the mass, 4,250 kg, is radiation shielding which is here assumed reusable. (Hanford, p. 11) Other sources give the reactor life as 7 years. (HSF, p. 655) Taking the life as 10 years, the mass cost of power is me(p) = 42.5 kg/kW for shielding plus a recurring power logistics cost of 4.82 kg/kW-year to replace the reactor, conversion, and cooling.

Cooling mass equivalent - The life support BVAD gives a cooling mass equivalent of me(c) = 60 kg/kW, for an internal flow loop and external flow-through radiators on a Mars transit mission. (BVAD, p. 11) Cooling can be easily accomplished during interplanetary or interstellar transit by radiation into cold space.

Logistics mass - The logistics mass per year can include consumables such as food or filters and their packaging and any materials used to operate or maintain equipment. For a multigeneration starship, incremental resupply is not possible. All logistics consumables are provided initially.

PHYSICAL/CHEMICAL (P/C) LIFE SUPPORT WITH SUPPLIED FOOD

For physical/chemical life support with supplied food, we consider both an upper bound on ESM derived from the International Space Station (ISS) approach and a closer estimate. In physical/chemical life support, water and oxygen are recycled from wastewater and carbon dioxide.

UPPER BOUND ESM ON INTERNATIONAL SPACE STATION (ISS) ENVIRONMENTAL CONTROL AND LIFE SUPPORT SYSTEM (ECLSS) WITH HYDRATED FOOD - To develop an upper bound on the ESM of a physical/chemical life support system with supplied food, we consider the International Space Station (ISS) Environmental Control and Life Support System (ECLSS) with fully hydrated packaged food. The entire assembly complete ECLSS will have 10 racks of about 500 kg and 1.5 m3 each. Total power is 5.4 kW. Annual resupply is 1,868.5 kg, exclusive of oxygen and water since these are to be provided by recycling. (Carrasquillo et al.) We conservatively assume the system supports only four crewmembers. The mass, volume, power, and

http://en.wikipedia.org/wiki/Mir

http://en.wikipedia.org/wiki/International_Space_Station

http://en.wikipedia.org/wiki/Skylab

cooling, ESM, and logistics are shown in Table 2, where crewmember is CM.

Table 2. Upper bound ESM of the ISS P/C ECLSS with hydrated food.

quantity massvolum

e power power logistics thermaltotal ESM logistics

unitskg/CM m3/CM

kW/CM kW/CM kg/CM

kg/CM-y

ISS P/C ECLSS 1,250 3.8 1.4 1.4 1,846 467power logistics 7packaged hydrated food 1,314mass equiv 1 122 42.5 4.82 60 units kg/kg kg/m3 kg/kW kg/kW-year kg/kW totals 1,846 1,788

The yearly mass cost for hydrated food is included under logistics. This cost is computed in Table 3.

Table 3. Logistics mass cost for hydrated food. kg/CM-

dayPercent

Food solids 0.62 17Water in food 1.15 32Packaging, containers

0.94 26

Mass subtotal 2.71 Volume mass penalty 0.89 25Total food ESM 3.60 100

The standard crewmember daily requirement for hydrated food is 1.77 kg/CM-day, consisting of 0.62 kg of food solids and 1.15 kg of water in the food. (Wieland, p. 6) (Reed and Coulter, p. 122) In addition, there is a need for about 0.53 kg of packaging and containers for each kg of hydrated food. (BVAD, p. 56) The total mass for packaged hydrated food is 2.71 kg/CM-day. There is also a need for about 0.0027 m3 of volume per kg of total mass of hydrated, packaged, containerized food. (BVAD, p. 56) Using the volume mass equivalent of 122 kg/m3, the volume mass penalty is 0.89 kg/CM-day, and the total mass for packaged hydrated food including its storage volume is 3.60 kg/CM-day. The logistics ESM for hydrated food is 1,314 kg/CM-year. Water accounts for about two-thirds of the hydrated food mass cost.

Referring to Table 2, with the volume, power, and cooling mass equivalents, the ISS ECLSS has an ESM of 1,846 kg/CM. The ECLSS logistics, packaged hydrated stored food, and power logistics total 1,788 kg/CM-y. The major mass costs are the ECLSS mass, volume, and logistics and the food supply logistics. For multiyear missions, the supplied food dominates the mass cost.

LOWER BOTTOM-UP ESTIMATE OF ISS P/C ECLSS ESM WITH DEHYDRATED FOOD - The above estimate of the ESM of the ISS P/C recycling system is about five times larger than other estimates for P/C systems using identical air and water technologies. Estimates of about 550 kg/CM can be obtained using the BVAD (Table 7.3) or Doll and Eckart (HSF, Tables 17-9 and 17-10, summarized in HSF, p. 998). Since these sources do not define the logistic requirements of the P/C systems, which are most of the mass cost for long duration missions, a bottom-up estimate of ESM is developed for the ISS ECLSS that includes logistics.

The ECLSS is again assumed to support four crewmembers. The life support functions and the implementing P/C technologies are given in Table 4.

Table 4. Lower estimate ESM of ISS P/C ECLSS with dehydrated food.

quantity massvolume power power logistics thermal

total ESM logistics

unitskg/CM m3/CM

kW/CM kW/CM kg/CM

kg/CM-y

function tech-nology Oxygen generation

SPWE

28.3 0.04 0.37 0.37 71 3Carbon dioxide removal

4BMS

50.3 0.10 0.22 0.22 85 0Carbon dioxide reduction

Sabatier

4.5 0.19 0.01 0.07 32 0Trace contaminant conrol

charcoal

19.6 0.07 0.04 0.04 32 41Waste water processing

multi-filtration

119.0 0.56 0.08 0.08 196 120Urine processing

VCD

32.0 0.09 0.02 0.02 45 44Oxygen storage

tank

272.0 0.00 0.00 0.00 272 136Water storage

tank26.5 0.13 0.00 0.00 42 0

25% package mass

138.1 138 10% package volume

0.12 14 power logistics 4 packaged dehydrated

food 279 totals 690.3 1.3 0.74 0.80 928 627 mass equiv 1.0 122 42.5 4.82 60

units kg/kg kg/m3 kg/kW kg/kW-year kg/kW

The data are from (Carrasquillo et al.), except for the Sabatier which is currently only scarred for on ISS. The Sabatier data are from (Eckart.) The total mass is increased by 25 percent and the total volume by 10 percent to allow for system packaging. The mass equivalents are those used previously. The yearly mass cost for dehydrated food is included, as computed below in Table 5.

Table 5. Logistics mass cost for dehydrated food. kg/CM-

dayPercent

Food solids 0.62 8110 % for packaging 0.06 8Volume mass penalty 0.08 11Total ESM 0.76 100

The standard crewmember daily requirement for food solids is 0.62 kg/CM-day. (Wieland, p. 6) (Reed and Coulter, p. 122) Food solids are assumed to require only

10 percent additional mass for containers and storage, a total of 0.68 kg/CM-day. Assuming that the packaged

hydrated food has the density of water, 103 kg/ m3, 0.68 kg occupies 0.68 * 10-3 m3. Using the volume mass of 122 kg/m3, the volume mass penalty is 0.08 kg/CM-day, and the total system mass for packaged dehydrated food is 0.76 kg/CM-day. The dehydrated food logistics ESM is 279 kg/CM-year. Food solids make up 81 percent of this mass cost, where they were only 17 percent of the mass cost of hydrated packaged food.

Referring to Table 4, with the volume, power, and cooling mass equivalents, the lower bound ISS ECLSS has an ESM of 928 kg/CM, 50 percent of the upper bound in Table 2. The ECLSS logistics, packaged dehydrated food, and power logistics total 627 kg/CM-y, 35 percent of the upper bound in Table 2. As before, the major mass costs are the ECLSS mass, volume, and logistics and the food supply logistics, and logistics dominate the long term cost. ECLSS logistics are 56 percent and food logistics 44 percent.

BIOREGENERATIVE LIFE SUPPORT (BLS)

In bioregenerative life support, crop plants are grown to provide the crewmembers’ food. The crop plants also

provide the needed oxygen and water recycling. We first consider a system providing nearly all the food, and then a more mass efficient system that provides only about one-half the food, with the remainder supplied.

BIOREGENERATIVE LIFE SUPPORT PROVIDING NEARLY ALL THE FOOD - The ESM of a space plant growth chamber, in kg per m2 of growth area, is shown in Table 6. (BVAD, p. 33) The mass equivalents are as above.

Table 6. ESM of a plant growth chamber.

component massvolume power power logistics thermal

total ESM logistics

units kg/ m2 m3/ m2 kW/ m2 kW/ m2 kg/ m2 kg/ m2-y

chamber 101.5 1.03 2.6 2.6 494 3.8

power logistics 12.5

mass equivalent 1 122 42.5 4.82 60

units kg/kg kg/m3 kg/kW kg/kW-year kg/kW totals 494 16.3

Studies of growing plants to provide a balanced diet show that at nominal productivities a mixture of peanut, potato, and wheat can provide 86 percent of the calories required using 41.2 m2 per crewmember. The missing calories can be supplied by about 50 grams of oil per crewmember per day (0.050 kg/CM-d). (Jones, 2000-01-2397)

The required plant chamber ESM is thus 41.2 m2/CM* 494 kg/m2 = 20,353 kg/CM. The total logistics to provide the grown food are 41.2 m2/CM * 16.3 kg/m2-y = 672 kg/CM-y.

The food supplement oil of 0.050 kg/CM-d can be treated as dehydrated food per Table 5, so its ESM is 0.050 *(0.76/0.62) = 0.061 kg/CM-d, which times 365 d/y gives the logistics cost of food supplement oil as only 22 kg/CM-y. 14 percent of the calories are provided by 3.2 percent of the total logistics mass.

The transpiration of the growing plants serves as the primary means of water purification. Any additional water processing equipment will be relatively small and is neglected. The plants absorb carbon dioxide and use it

to produce carbohydrates and oxygen. The plants produce about twice the oxygen required by the crew. This is because the harvest index of plants grown for food, the ratio of edible to total biomass, is roughly 50 percent. The 50 percent of inedible biomass must be

oxidized to provide carbon dioxide for future plant growth. (Jones, 2003-01-2357)

For a fully bioregenerative system, a composter is used to slowly oxidize the waste biomass. Table 7 shows the ESM of a composter which requires no power or logistics and little cooling. (Maxwell and Drysdale) The composter ESM is 495 kg/CM.

Table 7. ESM of a composter.

massvolume power

power logistics

thermal

total ESM

logistics

unitskg/CM

m3/CM

kW/CM

kW/CM

kg/CM

kg/CM-y

totals 174.4 0.7 0 0.1 266 0mass equivalent 1 122 42.5 4.82 60 units kg/kg kg/ m3 kg/kW kg/kW

The ESM of a bioregenerative life support system providing nearly all the food is shown in Table 8. Table 8 includes the plant chamber, food supplement oil, and composter.

Table 8. ESM and logistics of bioregenerative life support providing nearly all the food.

quantity ESM logistics

units kg/CM kg/CM-ychamber 20,339 673food oil 0 22composter 266 0totals 20,605 695

The plant chamber mass, volume, power, and cooling and the power logistics are significant.

BIOREGENERATIVE LIFE SUPPORT PROVIDING ABOUT HALF THE FOOD - Parametric studies that vary the percentage of food grown from 0 to 100 have shown that a shallow minimum of ESM occurs when about 50 percent of the food is grown. (Hanford) Growing all the food requires using more of the calorie inefficient, higher protein and fat producing crops such as soybeans and peanuts. Growing about 50 percent of the food provides all the air and water recycling needed by the crew, so no physical/chemical air and water processing systems are needed. (Jones, 2003-01-2357) Also, if less than 50 percent of the food is grown, there is no need for (or oxygen available for) waste biomass recycling by composting. Since about 50 percent of the crop biomass is inedible, growing about 50 percent of the food eliminates the need for air, water, or inedible biomass recycling equipment.

A mixture of wheat, potato, and sweet potato can provide about 43 percent of the calories required using 20.6 m2 per crewmember. The required plant chamber ESM is thus 20.6 m2/CM* 494 kg/m2 = 10,176 kg/CM. The total logistics to provide the grown food are 41.2 m2/CM * 16.3 kg/m2-y = 336 kg/CM-y. An additional 7 percent of the calories can be supplied by about 25 grams of oil per crewmember per day, 0.025 kg/CM-d, with a mass including packaging of 0.030 kg/CM-d, or 11.2 kg/CM-y. The remaining 50 percent of the calories are supplied by 0.31 kg/CM-d of food solids, with a total mass cost of 0.38 kg/CM-d with packaging, per Table 5, or 140 kg/CM-y.

The required ESM with bioregenerative life support providing half the food is equal to the plant chamber ESM of 10,176 kg/CM. The total logistics are 336 kg/CM-y to support the plant chamber, 11.2 kg/CM-y for food oil, plus 140 kg/CM-y for the other half of the food, a total of 487 kg/CM-y. No additional air, water, or waste processing is needed.

MASS COSTS OF MORE RELIABLE LIFE SUPPORT SYSTEMS

The mass requirements of the above four life support systems will be compared over time and the lowest mass systems identified. The initial hardware and initial logistics ESM per crewmember are shown in Table 9, for the four life support approaches investigated above.

Table 9. Initial and logistics ESM/CM of different life support systems.

Full BLS

50% BLS -

dehydrated

ISS upper bound (P/C) - hydrate

d

ISS lower bound (P/C) -

dehydrated

Initial ESM (kg/CM)

20,605 10,176 1,846 928

Initial logistics ESM (kg/CM-y) 695 487 1,788 6272 times initial ESM (kg/CM)

41,210 20,352 3,692 1,856

Logistics ESM, replace mass at 100 years (kg/CM-y) 901 589 1,806 6363 times initial ESM (kg/CM)

61,815 30,528 5,538 2,784

Logistics ESM, replace mass at 50 years (kg/CM-y) 1,107 691 1,825 646

The four approaches are full bioregenerative life support (BLS), 50 percent BLS with dehydrated food, ISS upper bound physicochemical (P/C) life support with hydrated food, and ISS lower bound physicochemical (P/C) life support with dehydrated food.

The initial hardware ESM and the logistics ESM are taken from the summaries above. The initial hardware mass was doubled or tripled to achieve high reliability, as described the review of reliability. The initial mass was also replaced either every 100 years or every 50 years. It was assumed that the initial logistics material is inert and has reasonably high reliability, and that it can be stored in a large number of independent packages. In this case, a small percentage of spare packages can provide high reliability with small additional mass. (Jones, 2008-08-2160)

The ESM of the four life support systems is compared for three cases.



3. Providing only the original system mass and never replacing the system (1/∞).

COMPARISON OF LIFE SUPPORT MASS REQUIREMENTS AND SYSTEM SELECTION

Figure 5 shows the ESM/CM versus duration, for four life support systems, with only the original system mass and not replacing the system (1/inf), for mission duration up to 200 years.

The life support system with lowest ESM/CM is the ISS lower bound physicochemical (P/C) life support with dehydrated food, out to the breakeven mission duration of 66 years. Beyond that time, 50 percent BLS with dehydrated food has lower ESM/CM.

At shorter durations up to about 20 years, the full bioregenerative system growing most of the food has the highest total ESM because of its high initial hardware ESM. The total ESM of the ISS upper bound P/C increases very rapidly because of the high mass of the hydrated food. The bioregenerative system growing 50 percent of the food always has lower ESM than the full bioregenerative system growing most of the food, since it has both lower hardware and lower logistics ESM.

Figure 6 shows the ESM/CM versus duration, for four life support systems, with 2 times original system mass and replacing the system every 100 years (2/100), for mission duration up to 800 years.

Figure 5. ESM/CM with only the original system mass and not replacing the system, for full BLS, 50% BLS with dehydrated food, ISS upper bound P/C with hydrated food, and ISS lower bound P/C with dehydrated food.

ESM/CM versus duration with no added mass for reliability (1/inf)

0

50,000

100,000

150,000

200,000

0 50 100 150 200

Duration, years

ESM/CM

Full BLS

50% BLS -dehydrated

ISS upper (P/C) -hydrated

ISS lower (P/C) -dehydrated

Figure 6. ESM/CM with 2 times original system mass and replacing the system every 100 years, for full BLS, 50% BLS with dehydrated food, ISS upper bound P/C with hydrated food, and ISS lower bound P/C with dehydrated food.

The life support system with lowest ESM/CM is the ISS lower bound physicochemical (P/C) life support with dehydrated food, out to the breakeven mission duration of 389 years. Beyond that time, 50 percent BLS with dehydrated food has lower ESM/CM.

In the case of the four life support systems, with 3 times original system mass and replacing the system every 50 years (3/50), the life support system with lowest ESM/CM is always the ISS lower bound physicochemical (P/C) life support with dehydrated food. This can be seen from Table 9, where this system has the lowest 3 times original system mass and the lowest logistics with replacing the system every 50 years.

Achieving high reliability over long durations seems to require 3 times original system mass and replacing the system every 50 years (3/50). If this is the case, then physicochemical life support with dehydrated food has significantly lower ESM at all mission durations. If only 2 times the original system mass and replacing the system every 100 years (2/100) is sufficient, then a 50 percent BLS with dehydrated food is competitive at most mission durations, as shown in Figure 6.

Anticipating an initial mission duration of less than 389 years, and observing that the choice does not significantly affect the required ESM over the range of expected durations, the ISS lower bound physicochemical life support with dehydrated food system will be assumed for the starship life support.

LOGISTICS AND HARDWARE REPLACEMENT MASS DOMINATES LONG DURATION MASS COST - Over long durations, the logistics and hardware replacement mass dominates the total ESM, so that the life support system with the lowest logistics per crewmember-year has the lowest total ESM. Because of their similar logistics and hardware replacement ESM’s, for the case with only the original system mass and not replacing the system (1/inf), shown in Figure 5, the full bioregenerative system and the lower estimate ISS with dehydrated food have approximately equal ESM slopes and differ only by their initial hardware ESM’s at long mission durations. Similarly, for the case with 2 times original system mass and replacing the system every 100 years (2/100) shown in Figure 6, the ISS lower bound physicochemical (P/C) life support with dehydrated food and the 50 percent BLS with dehydrated food has lower ESM/CM have similar slopes and ESM due to roughly equal logistics and hardware replacement mass.

THE 50 PERCENT FOOD SYSTEM DISCUSSION - For very long durations, the bioregenerative system growing about 50 percent of the food and supplemented with

ESM/CM versus duration with 2 times initial mass and replacement at 100 years (2/100)

0

200,000

400,000

600,000

800,000

0 200 400 600 800

Duration, years

ESM/CM

Full BLS

50% BLS -dehydrated

ISS upper (P/C) -hydrated

ISS lower (P/C) -dehydrated

dehydrated food is the best alternative to the selected ISS lower bound physicochemical life support with dehydrated food. It is not exactly half a full bioregenerative system and half of the ISS lower bound physicochemical life support with dehydrated food. Suppose it is decided to combine half a full bioregenerative system with half a lower estimate ISS with dehydrated food system. Each half system could support half the crew. The logistics of this half and half combination is the average of the two systems, (695 + 627)/2 = 661 kg/CM-y.

The bioregenerative system growing about 50 percent of the food has logistic requirements of only 487 kg/CM-y. It does not use the P/C air, water, and other systems of the lower estimate ISS with dehydrated food system, but only the dehydrated food with logistics of one-half of 279 kg/CM-y. And it uses one-half of the bioregenerative system with food oil supplement, so its logistics are one-half of 695 kg/CM-y. The bioregenerative system growing about 50 percent of the food thus has logistics needs of (279 + 695)/2 = 487 kg/CM-y, as previously found. This is lower than the logistics ESM of the full bioregenerative system, the lower estimate ISS with dehydrated food system, or their half and half combination, at 695, 627, and 661 kg/CM-y.

The bioregenerative system growing about 50 percent of the food has a hydroponic plant chamber with an area of 20.6 m2 per crewmember. This area weighs about 2,090 kg without spares for reliability and fills a volume of 21.2 m3. The plant growth areas are shelves, at plant not human height. Lighting and root zone water are provided. The average power required is 53.6 kW, and is provided by a nuclear generator. Wheat, potato, and peanut are grown and provide about 43 percent of the calories required. Oil provides another 7 percent and dehydrated food the remaining 50 percent of required calories.

The approximately half food growing system is reasonably attractive. It provides two distinctively different food sources, plant growth and stored food. It provides variety and bulk in the diet, with the opportunity to grow small amounts of unusual vegetables and spices. And it gives the crew vitally important, pleasant, natural, and challenging work to occupy them during the long voyage.

However, the approximately half food growing system is only the second least expensive system. Growing food will be an expensive comfort.

THE ISS LOWER BOUND P/C LIFE SUPPORT WITH DEHYDRATED FOOD DISCUSSION - The ISS lower bound physicochemical (P/C) life support with dehydrated food has lowest ESM. This minimized ISS system is best, and much better than ISS upper bound P/C with hydrated food, because of the much lower logistics cost. Hydrated packaged food requires 3.60 kg/CM-d with only 17 percent food solids, while dehydrated minimally packaged food requires only 0.76

kg/CM-d. It is the cheapest life support option, but least attractive in diet quality.

WASN’T BIOREGENERATIVE THE EXPECTED RESULT? - Full bioregenerative was expected to have lowest ESM for long missions. Previous work by this author found that physical/chemical systems had the lowest ESM only for missions less than 29 years. 50 percent food growing systems had the lowest ESM for mission durations from 29 to 80 years, and full bioregenerative systems had the lowest ESM beyond 80 years. Earlier detailed studies found that full bioregenerative was better than 50% food growing much sooner at 10 to 30 years. (Jones, 2006-01-2082)

Three things have changed from the previous assumptions, the increased initial mass for high reliability, inclusion of system replacement logistics, and the use of extremely dehydrated food. All of these changes favor the selected system, the ISS lower bound physicochemical (P/C) life support with dehydrated food.

The power and cooling mass equivalents used here are four and three times lower than those used previously from the BVAD, the change favoring bioregenerative life support over P/C. (Jones, 2006-01-2082) However the BVAD does not include a power logistics mass cost representing replacement of the power supply. It is apparently reasonably assumed that the power supply will last the mission duration. The 15 year operating life of the SP-100 nuclear reactor was mentioned by Hanford, but its mass was treated as initial hardware because the mission duration was equal to the 15 year operating life. The SP-100 was replaced in an analysis that showed bioregenerative breaking even with 50% food growing at 30 years. (Hanford)

The hydrated food supply is not considered part of life support design, but rather an external given, even a gift of water, determined by crew needs and preferences. Short missions have negligible food costs. On longer missions, the water in hydrated food can be recovered by recycling P/C life support and provides margin for inefficiencies and losses. Saving significant food cost by using severely dehydrated food is possible and necessary only on long missions. Using severely dehydrated and minimally packaged food solids strongly favors P/C compared to bioregenerative life support.

The three changes favoring physicochemical systems, increased initial mass for reliability, system replacement logistics, and dehydrated food, all reflect solutions to problems of very long duration, multigeneration missions that were not important in shorter mission planning.

FUTURE LOWER MASS LIFE SUPPORT SYSTEMS? - This investigation is limited to currently feasible technology. But, given the prohibitively high mass cost of even the least expensive current approaches to providing food, it is reasonable to consider fundamental limits and possible future improvements.

The best current system for the longest missions is a bioregenerative system growing about half the food with the other half provided by stored dehydrated food. The bioregenerative system providing 50 percent of the food actually grows 43 percent of the food, provides another 7 percent with 0.025 kg/CM-d of oil, and the remaining 50 percent with 0.31 kg/CM-d of food solids. The food mass to provide 57 percent of the calories is 0.335 kg/CM-day. With packaging allowances the oil requires 0.30 kg/CM-d and the food solids require 0.38 kg/CM-d, a total of 0.68 kg/CM-d. A potential saving of up to 0.34 kg/CM-d, 50 percent, could be achieved by eliminating all packaging. Future systems providing stored food can not save more than half the current logistics mass.

This part bioregenerative system grows 43 percent of the food using growing space of 20.6 m2 per crewmember. The corresponding logistics to provide this food are 20.6 m2/CM * 16.3 kg/m2-y = 336 kg/CM-y, or 0.92 kg/CM-d. Compared to stored food, the bioregenerative part requires more mass, 0.92 versus 0.68 kg/CM-d, for less food energy, 43 versus 57 percent of the calories.

The bioregenerative power required for this 50 percent system is 20.6 m2/CM * 2.6 kW/ m2 = 54 kW/CM, so each day 54 * 24 = 1.3 * 103 kW-hours/CM-day of energy is input. This can be compared to the food energy produced, 43 percent of the 2,677 calories/CM-day corresponding to food solids of 0.62 kg/CM-day, or 1,150 calories/CM-day. (Wieland, p. 6) (Reed and Coulter, p. 122)

1 joule = 1 watt-second, so 1 kW hour = 3.6 * 106 Joules. Since 1 calorie = 4.19 Joules, 1 kW hour = 8.59 * 105 calories. The 1.3 * 103 kW-hours/CM-day of energy provided for plant growth is equal to 1.12 * 109 calories. The plants provide 1,150 calories/CM-day. The input energy is 106 times, six orders of magnitude, higher than the food energy output. That this energy input is provided at a mass cost similar to that of the food output reflects the vast difference between nuclear and chemical energy density.

This energy ratio shows that bioregenerative systems have a vast potential for improvement. Some parts of the system, nuclear power output, lighting efficiency, or plant energy capture, might have their efficiency significantly improved. Even a factor of two improvement would make bioregenerative food production more mass efficient than dehydrated food could ever be. Possibly microbial, nanotechnology, or chemical means will be found to provide food. However, any fundamental gains in power, lighting, or food production would be widely useful and could transform the terrestrial as well as space economy. Any fundamental improvements would probably originate outside space life support research. Space programs are usually based on currently feasible technology.

SPACECRAFT DESIGN AND MASS ESTIMATE

The spacecraft description will continue beyond the crew and life support, to include the crew accommodations,

crew habitat, structures and mechanisms, power and cooling, and planetary descent vehicles and supplies. The mass, volume, and power per crewmember and the logistics mass per crewmember per year will be estimated.

LIFE SUPPORT - The life support mass, volume, power, and logistics are given in Table 10, for the ISS lower bound physicochemical (P/C) life support with dehydrated food. The initial mass, volume, and power, and the logistics mass and volume are derived from Tables 4 and 5 because the power, cooling, and structure (volume) masses are estimated later. The power supply replacement logistics are not included here, but are considered later. The cooling load will be assumed equal to power input The mass equivalent of volume is subtracted from the dehydrated food logistic ESM of Table 5 and the logistics volume included instead.

Table 10. ISS lower bound physicochemical (P/C) life support with dehydrated food.

MassVolume

Power

Logistics

mass

Logistics

volume kg/

CMm3/CM

kW/CM

kg/CM-y

m3/CM-y

ISS lower bound P/C- dehydrated

690 1.3 0.74 592 0.17

3 times mass and volume

2,070

3.9

Replacing the original system mass every 50 years

41 0.08

Final totals

2,070

3.9 0.74 633 0.25

The initial mass and volume are multiplied by three to support high reliability. The initial mass and volume are replaced every 50 years also to support high reliability. The cooling load is not listed and will be assumed equal to the power input.

CREW ACCOMMODATIONS - Initial estimated crew accommodation masses for a multigeneration interstellar colonizing starship, are listed in Table 11. They are adapted from the needs for six crewmembers in a Mars surface habitat for 500 days or on a 180-day return trip from Mars to Earth. (HSF, pp. 582, 597, 603)

Table 11. Crew accommodations mass.

Element SubelementEquipment,

kg/CMLogistics, kg/CM-

y

Galley

Freezers 67.0

Ovens 20.0

Sink, dishwasher 9.0

Cook, eat, clean supplies 5.0 15.2

Waste Management

Toilets 15.0

Cleaning, supplies 18.3

Vacuum 2.2

Trash compactor 25.0

Trash bags 84.0

Hygiene

Shower 12.5

Sink 1.3

Toiletry kit 0.3

Hygiene supplies 27.4

Clothing

Clothes 99.0

Washer, dryer 26.7

Maintenance

Hand tools 50.0

Test equipment 83.3

Machine tools 166.7

Recreation

Equipment 50.0

Other 20.0

Health Care

Exercise 24.2

Medical equipment 166.7

Medical consumables 83.3

Totals 828.0 243.8

The crew accommodations equipment and logistics mass of Table 11 are included with further information in Table 12.

Table 12. Crew accommodations summary.

MassVol-ume Power

Logis-tics

mass

Logis-tics vol-ume

kg/CM

m3/CM

kW/CM

kg/CM-y

m3/CM-y

Initial system 828 8.0 2.47 244 1.763 times mass and volume 2,484 24.0 Replacing the original system mass every 50 years 17 0.16Final totals 2,484 24 2.47 260 1.92

Additional information is included. The volume for the

crew accommodation equipment is about 8 m3/CM.

(HSF, pp. 582, 603, 997) The power for crew accommodation is about 2.47 kW/CM, largely for the galley freezers, oven, dishwasher and for the clothes dryer. (HSF, pp. 583, 997) The volume for the crew

accommodation logistics is about 1.76 m3/CM-y. (HSF,

pp. 582, 603) (Airlocks, spacesuits, and electrical equipment for communications, navigation, computation, etc., have negligible mass. HSF, pp. 732, 998)

The initial mass and volume are multiplied by three for high reliability. The total crew accommodation equipment mass is 2,484, after multiplying by 3 for high reliability. The crew accommodation mass is slightly larger than the life support mass. The initial mass and volume are replaced every 50 years also to support high reliability.

CREW HABITAT VOLUME - The space shuttle provides 10 m3 of pressurized volume per crewmember (CM). (HSF, p. 404) This is the lower volume “performance limit” for good performance on six month or longer missions, but more volume, 18 or 20 m3/CM, would be “optimal.” (HSF, p. 149) (NASA-STD-3000, p. 8-12) This optimal volume corresponds roughly to a small office or cubicle. The six month to year long missions of Skylab, Mir, and International Space Station (ISS) provide about 100 m3/CM (HSF, p. 404), and this value will be used for the multigenerational spacecraft. 100 m3 is about the volume of a small apartment.

100 m3/CM of crew habitat volume will be provided in addition to the life support and crew accommodations volume. Other spacecraft elements, such as power and cooling, will have no dedicated volume allowances, assuming that they can be accommodated in the habitat volume. The mass needed to provide the total volume is

considered under spacecraft structures and mechanisms.

STRUCTURES AND MECHANISMS MASS PER UNIT VOLUME - The computation of the ESM volume mass equivalent estimated the mass per unit volume. The average mass per unit volume of four conventional metal designs was 450 kg/m3. Assuming that the structures and mechanisms are 27 percent of the spacecraft dry mass, they would require 122 kg/m3.

POWER AND COOLING SYSTEMS - The mass required for power and cooling is estimated using the mass equivalents for power and cooling that were developed for the life support ESM computations. These are shown in Table 13.

Table 13. Power and cooling mass equivalents.Mass Logistics

kg/kWkg/kW-

yearPower mass equivalent 42.5 4.82Cooling mass equivalent 60.0Power plus cooling mass equivalent 102.5 4.82

The combined power and cooling hardware mass is estimated using the power plus cooling mass equivalent. Because of the recurring power logistics, a yearly mass cost also applies.

PLANETARY DESCENT VEHICLES AND SUPPLIES - Small spacecraft will separate from the starship and land in craft similar to the Apollo/Orion vehicles. A Command Module takes the crew to the surface after a Service Module decelerates the Command Module and separates.

The Apollo Command Module weighed 5,809 kg and the Service Module weighed 24,523 kg, a total of 30,332 kg. The Command Module accommodated 3 crewmembers, and reentered the atmosphere using parachutes and a heat shield. The Service Module housed the propulsion system and propellant, as well as power, communications, and air and water storage. (http://en.wikipedia.org/wiki/Apollo_Command/Service_Module#Command_Module_.28CM.29)

Early planning for the new Crew Exploration Vehicle (CEV) had a similar architecture, with the Command Module weighing 9,695 kg and the Service Module

49,750 kg, for a total of 59,445 kg accommodating 6 crewmembers. (ESAS, p. 98)

Both Apollo and the CEV have a mass total mass of about 10,000kg/CM. However, the starship planetary descent vehicle will have to decelerate from much greater relative speed and will probably require nuclear propulsion similar to that used to originally accelerate the starship. The propulsion system mass will depend on the starship speed.

The Apollo Command Module weighed 1,936 kg/CM and the CEV Command Module could weigh 1,615 kg/CM. To provide colonization supplies, the descent capsule weight is set at 3,000 kg/CM.

The total weight including the propulsion system and propellant is Mt = Mv-pd exp (V/E), where Mv-pd is the planetary decent mass of 3,000kg/CM, E is the rocket exhaust speed, and V is the starship approach velocity. Planetary descent vehicle and supply mass, including propulsion is 3,000 exp (V/E) kg/CM. The planetary descent vehicles will be external to the main spacecraft.

MASS ESTIMATE - A partly independent and bottom-up mass estimate for a multigeneration interstellar colonizing starship, scaled in kilograms per crewmember, is detailed by elements in Table 14.

Table 14. Multigeneration spacecraft mass estimate.

MassVolum

e PowerLogistics

massLogistics volume

kg/CM m3/CMkW/CM kg/CM-y m3/CM-y

Life support - ISS lower bound P/C - dehydrated 2,070 3.9 0.74 633 0.25Crew accommodations 2,484 24.0 2.47 260 1.92Crew habitat 100.0 Totals 4,554 127.9 3.21 893 2.17

Structures and mechanisms mass, 122 kg/m3 15,604 265 Power & cooling mass, 102.5 kg/kW + 4.82 kg/kW-year 329 15 Total spacecraft mass 20,487 1,173 Planetary descent vehicles and supplies mass 3,000 exp

(V/E)

The spacecraft vehicle mass to be sent to the new planet is Mv-sc = 20,487 kg/CM + 1,173 kg/CM-y * T+ 3,000 exp (V/E), where T is the mission travel time, E is the rocket exhaust speed, and V is the starship velocity.

It would be interesting to set reasonable values for the parameters, perhaps T = 200 y, V/E = 2, and see the results. But the largest contribution to mission mass is propulsion. Shorter travel time reduces the spacecraft logistics mass but increases the spacecraft propulsion and planetary descent propulsion mass. The trade off between logistics and propulsion mass is examined to minimize total mission mass.

MINIMUM TOTAL MISSION MASS - The amount of propellant required for the spacecraft depends on the distance and travel speed to the new planet. The total mass including the propellant is Mt = Mv-sc exp (V/E), where Mv is the spacecraft and planetary descent mass,

E is the rocket exhaust speed, and V is the starship approach velocity. Substituting the expression for the spacecraft vehicle mass,

Mt = exp (V/E) *[20,487 kg/CM + 1,173 kg/CM-y * T+ 3,000 exp (V/E)]

The ratio V/E, starship velocity divided by rocket exhaust speed, is a selectable design parameter that strongly determines the mission mass. Mass increases rapidly with V/E. Limiting propulsion mass requires relatively small V/E, 2 or 3. See Figures 1 and 2.

The mission travel time T is equal to the new planet distance, D, divided by the spacecraft velocity, V, so T = D/V. Since V is usually a small multiple of E, 2 or 3, T is typically a large fraction of D/E, ½ or 1/3. A reasonable estimate for D is 10 light years. The nuclear propulsion system exhaust speed could be E = 0.05 c. D/E is 10/0.05 = 200 y, and T would be ½ or 1/3 of that, 100 or 67 years.

The ratio D/E is a fundamental constraint on interstellar mission design. D/E depends on the chosen destination and available rocket technology. The mission travel time T is usually a large fraction of D/E. For early interstellar colonization missions, D could be 5 to 15 light years and E 0.01 to 0.10 c, giving a range of D/E of 50 to 1,500 years.

The substitution T = D/V = (D/E)/(V/E) allows Mt to be expressed in two variables, normalized distance D/E and normalized velocity V/E, both with constrained range.

Mt = exp (V/E) *[20,487 kg/CM + 1,173 kg/CM-y * (D/E)/(V/E)+ 3,000 exp (V/E)]

Mt is graphed versus D/E and V/E, in Figure 7.

Figure 7. Mission mass, Mt, versus normalized distance D/E and normalized velocity V/E.

Figure 7 shows that the mission mass, Mt, strictly increases with increasing normalized distance, D/E. Mt has a shallow minimum for normalized velocity, V/E, of

0.50

0.75

1.00

1.25

1.50

1.75600 550 500 450 400 350 300 250 200 150 100 50

- 500,000 1,000,000 1,500,000

2,000,000 2,500,000

3,000,000

Mt

V/E

D/E

Mt versus D/E and V/E

approximately 1.0 at higher D/E and about 0.75 for the very smallest D/E. For the efficient use of total mission mass, the spacecraft velocity, V, is equal to or slightly less than the rocket exhaust velocity, E.

For D = 10 ly, E = 0.05 c, and D/E = 200, the minimum mission mass, Mt = 7.10 * 105 kg at V/E = 0.90 and T = 222 y. For D = 10 ly, E = 0.02 c, and D/E = 500, the minimum mission mass, Mt = 1.67 * 106 kg at V/E = 0.95 and T = 526 y.

TOTAL MISSION MASS BREAKDOWN - The total minimum mission mass is shown in Table 15, for D = 10 ly, E = 0.05 c, and D/E = 200. The minimum mission

mass, Mt = 7.10 * 105 kg at V/E = 0.90 and T = 222 y.

Table 15. Total minimum mission mass for D/E = 200, at V/E = 0.90 and T = 222 y.

MassVolum

e PowerLogistics mass at T

= 222Logistics volume at

T = 222

kg/CM m3/CMkW/CM kg/CM m3/CM

Life support - ISS lower bound P/C - dehydrated 2,070 3.9 0.74 140,526 56 Crew accommodations 2,484 24.0 2.47 57,720 426 Crew habitat 100.0 Totals 4,554 127.9 3.21 198,246 482 Structures and mechanisms mass, 122 kg/m3 15,604 58,772 Power & cooling mass, 102.5 kg/kW + 4.82 kg/kW-year 329 3,435 Totals 20,487 260,453

Spacecraft mass 280,940 Planetary descent vehicle mass, 3,000 exp (V/E) at V/E = 0.90 7,379 Spacecraft and planetary descent mass 288,319 Total mass at V/E = 0.90 709,150

Table 16 shows the mass contributions related to the spacecraft mass and the total mission mass.

Table 16. Mission mass contributions.Mass, kg/CM

% of spacecraft

mass

% of mission mass

Life support logistics 140,526 50 20 Crew accommodations logistics 57,720 21 8 Structures 74,376 26 10 Spacecraft 280,940 100 40 Planetary descent 7,379 3 1 Propulsion 420,831 150 59 Mission 709,150 252 100

Table 16 shows that the mission mass is 59 percent propulsion, 1 percent planetary descent vehicles, and 40 percent the spacecraft. The spacecraft mass is 26 percent structures (close to the assumed 27 percent), 21 percent crew accommodations, and 50 percent life support logistics.

Table 15 shows that most of the structures mass is due to logistics volume, and most of that is due to crew accommodations logistics. The 26 percent of mass due to structures includes 26 *(58,772 /74,376)*(426/482) = 18 percent due to crew accommodations logistics volume. Crew accommodations logistics should be charged with 39 rather than 21 percent of the spacecraft mass. Referring to Table 11, clothes and trash bags are the major components of crew accommodations logistics volume, 41percent and 34 percent.

Referring to Table 4, the largest component of the life support logistics is packaged dehydrated food, 44 percent. Using current approaches to life support and crew accommodation on a multigeneration starship

requires vast quantities of food, clothes, and trash bags. As always, the major mass cost is propulsion.

MASS COST AND DOLLAR COST ASSESSMENT - The example minimum mission mass, Mt = 7.10 * 105

kg/CM at V/E = 0.90 and T = 222 y. The International Space Station (ISS) has a mass of 2.78 * 104 kg, so the mass per crewmember is about 26 space stations. A mission with a crew of 30 would require a mass of 2.10 * 106 kg, 766 space station masses.

Launching this mission mass to low Earth orbit would be very expensive. The Space Shuttle cost to launch to low Earth orbit is typically quoted as $25 k/kg. (Wertz and Larson 1996, p. 125) A yearly Space Shuttle budget of 4 billion dollars for 10 launches of 16,000 kg to LEO corresponds to a cost of $25 k/kg. The cost to launch the mass for the example multigeneration mission with a crew of 30 would be $25 * 103/kg * 2.10 * 106 kg = $5.3 * 1010, 53 billion dollars. The overall mission cost for design, development, and launch might be two or three times the launch cost, approximately equal to the cost of the Apollo project of $135 billion in current dollars, or to the current NASA budget for a decade. (http://en.wikipedia.org/wiki/NASA_Budget) We can go to the stars!

MULTIGENERATIONAL INTERSTELLAR MISSION OBSERVATIONS

A multigenerational starship will be massive and costly. The mission cost is approximately the cost of the Apollo project or ten years of the current NASA.

Most of the cost is propellant. Decelerating at the target is especially costly, as the velocity reducing propellant must be launched with the starship. The crew can decelerate and land on planet in several small lander capsules, but the mass budget could allow decelerating a part of the starship without waste mass from the trip.

The probability of a mission’s success increases for a larger number of colonizers, but with diminishing returns. Some minimum number of crew is needed to achieve a reasonable probability of success. Sending many smaller, riskier missions to different targets increase the expected number of planets successfully colonized.

Since life support systems can achieve high reliability with mass slowly linearly increasing with time, but decreasing the transit time requires exponentially increasing mass, the lowest total mass is achieved with relatively slow speeds.

CONCLUSION

Life support and crew accommodations are a small cost for brief missions but totally dominate the cost of multiyear and multigeneration missions. The time dependent logistics mass is the major cost on long duration missions.

Life support traditionally includes atmosphere and water processing while crew accommodations includes food, waste management, clothing and laundry. The similarities and interactions between these areas strongly suggest that they be analyzed and optimized together.

Physicochemical life support with supplied dehydrated food is requires less mass than bioregenerative food growing, using current technology, for missions lasting hundreds of years.

Bioregenerative food growing is a million to one inefficient in energy use and would be practical only if some major breakthrough occurs in power, lighting, or plant productivity.

The implication for current and near term life support work is that physicochemical life support with supplied dehydrated food is least expensive for the Moon, Mars, and far beyond. Closed system food production is not competitive without some fundamental breakthrough, which are not anticipated and most probably would result from unrelated basic research. Targeted space life support research is needed now on high reliability physicochemical water and oxygen recycling systems and on low mass dehydrated food supply, but would be unproductive in bioregenerative lie support.

REFERENCES

BVAD, Advanced Life Support Baseline Values and Assumptions Document, NASA/CR—2004–208941, August 2004.

Carrasquillo, R. L., J. L., Reuter, and C. L. Philistine, “Summary of Resources for the International Space Station Environmental Control and Life Support System,” Proceedings of the 27th International Conference on Environmental Systems, ICES 972332, 1997.

Czysz, Paul A., and Claudio Bruno, Future Spacecraft Propulsion Systems: Enabling Technologies for Space Exploration (Springer Praxis Books/Astronautical Engineering), 2006.

HSF, W. K. Larson, and L. K. Pranke, eds., Human Spaceflight: Mission Analysis and Design, McGraw-Hill, New York, undated (2000).

Eckart, P., Spaceflight Life Support and Biospherics, Kluwer Academic, Dordrecht, 1996.

Hanford, A. J., “Advanced Regenerative Life Support System Study,” JSC 38672, National Aeronautics and Space Administration, Lyndon B. Johnson Space Center, Houston, Texas, 1997.

Jones, H. W., "Matching Crew Diet and Crop Food Production in BIO-Plex,” SAE Technical Paper No. 2000-01-2397, Proceedings of the 30th International Conference on Environmental Systems, 2000.

http://en.wikipedia.org/wiki/NASA_Budget

Jones, H. W., “Design Rules for Space Life Support Systems,” SAE Technical Paper No. 2003-01-2357, Proceedings of the 33rd International Conference on Environmental Systems, 2003.

Jones, H. W., “Comparison of Bioregenerative and Physical/Chemical Life Support Systems,” SAE Technical Paper No. 2006-01-2082, Proceedings of the 36th International Conference on Environmental Systems, 2006.

Jones, H. W., “Equivalent Mass versus Life Cycle Cost for Life Support Technology Selection,” SAE Technical Paper No. 2003-01-2635, Proceedings of the 33rd International Conference on Environmental Systems, 2003.

Jones, Harry, “Ultra Reliable Space Life Support Systems,” SAE Technical Paper No. 2008-01-2160, Proceedings of the 38th International Conference on Environmental Systems, 2008.

Kraus, J. D., Radio Astronomy, Cygnus-Quasar, Powell, Ohio, 1986.

Maxwell, S., and Drysdale, A., “Assessment of Waste Processing Technologies for 3 Missions,” SAE Paper No. 2001-01-2365, 31st International Conference on Environmental Systems, 2001.

NASA-STD-3000, Man-Systems Integration Standards, NASA, 1995.

Reed, R. D., and G. R. Coulter, “Physiology of Spaceflight,” in W. K. Larson, and L. K. Pranke, eds., Human Spaceflight: Mission Analysis and Design, McGraw-Hill, New York, undated (2000).

Weiland, P. O., Designing for Human Presence in Space: An Introduction to Environmental Control and Life Support Systems, NASA Reference Publication RP-1324, 1994.

Wertz, J. R., and W. J. Larson, eds., Space Mission Analysis and Design, third edition, Space Technology Series, Kluwer, Dordrecht, 1999.

Wikipedia, http://en.wikipedia.org/wiki/

Yakut, M. M, and Barker, R. S., Parametric study of manned life support systems, Volume 2 - Parametric relations and scaling law, Final report, McDonnel Douglas CR 3283, Jul. 1967- Aug. 1968.

Zubrin, R., "Nuclear Salt Water Rockets: High Thrust at 10,000 sec ISP", Journal of the British Interplanetary Society 44, 371-376, 1991.

CONTACT

Harry Jones

Mail Stop 239-8NASA Ames Research CenterMoffett Field, CA 94035-1000Phone: 650-604-5518e-mail: [email protected]

ACRONYMS AND SYMBOLS

Here is the Definitions section. This is an optional section.

1/∞: only the original mass and never replace the system2/100: 2 times the original mass and replace the system every 100 years 3/50: 3 times the original mass and replace the system every 50 years 4BMS: Four Bed Molecular Sieve∞: infinityBLS: bioregenerative life supportBVAD: Baseline Values and Assumptions Document (reference)c: cooling, speed of lightCEV: Crew Exploration VehicleCM: CrewmemberD/E: normalized distanced: day, mission durationD: new planet distanceE: rocket exhaust speedECLSS: Environmental Control and Life Support SystemESM: Equivalent System Massexp: exponential functionF(T): failure probability at time Tf: probability of failureHSF: Human Spaceflight (reference)inf: infinityISS: International Space Station kg: kilogramkW: kiloWattl: logistics mass per yearLEO: Low Earth Orbit ln: natural logarithmly: light year m: meterm: system massme(c): mass equivalent of coolingme(p): mass equivalent of powerme(v): mass equivalent of volumeMt: total mission mass including propellantMTBF: mean time between failuresMv: vehicle massMv-pd: planetary decent vehicle massN: Number of crewmembersP/C: Physical/Chemicalp: powerPc: probability of colonizing successPf: probability of failureR(T): reliability at time TSPWE: Solid Polymer Water ElectrolysisT: travel time, breakeven dateV/E: normalized velocityV: rocket velocity

mailto:[email protected]

http://en.wikipedia.org/wiki/

v: volume,VCD: Vapor Compression Distillationy: yearl: failure rate

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