These are some spacecraft designs that are based on reality. So they appear quite outlandish and undramatic looking. In the next page will appear designs that are fictional, but much more breathtaking. Obviously the spacecraft on this page are all NASA style exploration vehicles, they are not very suited for interplanetary combat (well, most of them at least).

For slower-than-light star ships, go here.

Many of these spacecraft have a table of parameters. You can find the meaning of many of them here. Numbers in black are from the documents. Numbers in yellow have been calculated by me using the document numbers, these might be incorrect.

This is not actually "real", but the science is admirably hard.

PUTTING THE TRAMP IN TRADER

     “It’s a steel box.”
     “It’s a fully functional – well, mostly functional, but all primary systems are functional – Hariven-class free trader. Just what you want when you’re starting out in this business.”
     “It’s a steel box with a plasma torch welded on the back.”
     “And a generous cargo capacity for its displacement, regenerative life support, ah – adequate crew quarters and food vats, and docking room for a single surface-orbit shuttle.”
     “And it’s –”
     “– a steel box, yes. If you wanted to pay for stylish, would you be shopping for starships in a wreckyard?”
     “Show me the contract again.”

– overheard in Kathar orbit, Cilmínár system

HARIVEN-CLASS FREE TRADER

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• 

So, I got a request from a reader for a few specs on the Hariven-class free trader. Well, why not?

(Sadly, they were imagining something like Vaughan Ling’s Planetes-inspired debris collector with comparable dimensions, capacity, etc. Sorry to say it, but that ship? Had some style. The Hariven? Really doesn’t.)

HARIVEN-CLASS FREE TRADER

Operated by: Desperate free traders, just starting-out bands on tour, your sketchy brother, refugees, space hobos, and anyone else who can’t afford a better ship.
Type: Basic freighter.
Construction: Under open-source license; produced by multiple manufacturers, most of whom would prefer not to admit it, along with various backyard fab shops.

(And when I say “desperate free trader”, I don’t mean, say, the people who fly around in a Firefly-class in Firefly. Those people, in this verse, own something like a Kalantha-class. This is down from there at the true ass end of space travel.)

Length: 46m, of which 30m is the hold.
Beam: 8m (not including radiators)

Gravity-well capable: No.
Atmosphere-capable: No.

Personnel: 3, as follows:

Flight Commander
Flight Director
Flight Engineer

(This assumes you’re following the typical regulations which require – since the Hariven has no AI, and only dumb automation – that at least one qualified person be on watch at all times, hence a minimum of three. In practice, a Hariven can be flown by one and very often is, if they don’t mind violating the rules of navigation of every halfway sane polity in space.)

Drive (typical; may vary from build to build): Nucleodyne Thrust Applications “Putt-Putt” fusion pulse drive.
Propellant: Deuterium pellets. (dirty D-D fusion)
Cruising (sustainable) thrust: 0.6 standard gravities (0.56 g)
Peak (unsustainable) thrust: 1.2 standard gravities (1.12 g)
Delta-v reserve: (Not yet calculated, but limited; if you’re flying a Hariven, you ain’t going brachy unless you devote a lot of your hold space to extra tanks. Be prepared to spend much of your voyage time on the float.)
Maximum velocity: 0.02 c (based on particle shielding)

Drones:

Not supplied as standard, but buy some. You’re gonna need ’em.

Sensors:

Orbital Positioning System sensors
Inertial tracking platform
Passive EM array
Short-range collision-avoidance and docking radar

Weapons:

None.

Other systems:

Omnidirectional radio transceiver
Communications laser
Whipple shield (habitable area only)
Mechanical regenerative life support (atmosphere/water only)
Algi-prote vat
2 x information furnace data systems
Sodium droplet radiators

Small craft:

Not supplied as standard, but a common as-supplied variant adds a partition to convert part of the forward hold into a bay with docking clamps suitable for many surface-to-orbit vehicles.

DESCRIPTION

It’s a classic tail-lander layout of the crudest form: a 30m steel box welded on top of an 8m steel cylinder welded on top of a cheap fusion pulse drive, the latter two surrounded by pellet containers. It couldn’t look more brutalist/functional if it tried. At least most Hariven owners try to give it a bright paint job.

The hold is up front, a big steel box roughly the size of eight standard shipping containers. (Indeed, sometimes it’s made from eight standard shipping containers.) Putting it right for’ard has the advantage of simplifying construction greatly – all the machinery is at one end – and giving Hariven captains the assurance that if they ram their junker into anything accidentally, at least there’s 30m of other stuff between them and whatever they hit.

The hold opens up along its entire length on the port side to permit access. Responsible captains who convert their Hariven for passenger transport (the aforementioned touring bands, refugees, and space hobos, for example) by attaching deck partitions inside the hold and adding canned air have these welded shut. Less responsible captains simply pray for a lack of wiring faults.

The habitable section (the cylinder at the back) is wrapped in auxiliary engineering machinery and fuel storage, to the point that it’s only 4m in internal diameter. (If you need to fiddle with most of the engineering systems, you’re going to need a drone, or to take a walk outside.) It’s divided into four decks, from the bow down:

The bridge, which shares space with most of the avionics;

A small living area, which contains the food vat, a tiny galley, the inner door of the airlock, and any luxuries you see fit to squeeze in there. Like chairs;

The crew quarters, which means four vertically-mounted sleep pods, and maybe room for another luxury or two if they’re small;

And a tiny workshop, for any repairs that need doing.

That all sits right on top of the shadow shield and the business end of the drive. If you need to adjust anything below that – well, hope you brought a drone.

But enough of this. You buy this ship, treat her proper, she’ll be with you the rest of your life.

Ain’t sayin’ how long that’ll be, mind.

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Ru said:

     Not that long ago, I spent quite some time running the numbers on fusion-pulse torch drives, and working out various performance figures and limitations. It was all quite informative and interesting. You haven’t given nearly enough information about the ship (eg. dry mass) or the engine (exhaust velocity ranges, reaction mass) for me to hazard a guess at its parameters, so it’ll be interesting to see what else you have for us…
     There’s a lot of steel mention in its design. Sounds pretty heavy. Also, metal shells plus charge particle radiation equals bremsstrahlung delight (meaning crew will be constantly irradiated by deadly x-rays) (and it makes for a poor neutron shield, which this sort of drive badly needs). Carbon is probably easier to come by, and much easier to push around.
     You don’t mention reaction mass, but with those performance figures you won’t be using pure fusion for peak acceleration. Presumably the drive expends additional deuterium for that purpose (though lithium might be a better choice).
     You’re using pure deuterium fusion, but that’s a terrible choice for spacecraft fuel, really. For flights much less than the half-life of tritium, D-T offers easier ignition, lower neutron flux and more charged particles to thrust against. D-³He would be the fuel of choice, but if you want stable and conveniently mineable fuel p-⁶Li or p-¹¹B would be a much better choice than pure deuterium.
     Delta-V reserves for even a fairly conservative fusion spacecraft design are pretty generous. You might not be tooling around at cruising speed for long, but it should be able to sustain a centigee for weeks (or even months if the drive is good enough) at a time with a 3 or 4:1 wet to dry mass ratio.
     (on reflection, I am of course wrong that D-T offers a lower neutron flux or a higher proportion of charged reaction products than D-D, but it does offer a significantly lower x-ray output and a higher exhaust velocity)

Alistair Young said:

     Ah, but you’ve got to bear in mind the target market, and therefore the design paradigm. If this were higher up the scale of starships, it’d have all the fancy carbon-composite hulls, high-efficiency fuel blends, etc., etc., one could possibly desire.
     It’s steel, though, because it’s designed to be repaired – and in some cases, even built – by a monkey with a wrench, a backyard welding kit, and duct tape, not by professional yard dogs with all the nanowhatsits in the catalog.
     (Same reason the neutron protection is a slug of paraffin in the lower hull space rather than proper formed HICAP.)
     Likewise, it uses D pellets to power an old-style fusion pulse drive rather than D-He3 slush to power a new-style fusion torch because that drive needs much less maintenance, any backplanet schmuck can separate deuterium from water, and the calibration is rough enough that in a pinch, you can stuff just about anything that’ll fuse in there and it’ll mostly work for a while.
     (Basically, you want to picture the spacegoing equivalent of the beat-to-hell jalopy that’s been driven around the rainforest for forty years, being fixed with banana peels and duct tape and occasionally run on rough home-cooked rum when gas was short. It’s a sh*tbox, but it’s a sh*tbox that’s hard to kill by design.)

This is mostly from Aviation Projects Review volume 1 number 3 which has more details than given here. Additional material from Helios pulsed nuclear propulsion concept (1965) which discusses the Lawrence Radiation Laboratory (LRL) Helios.

Again, there were several spacecraft designs that all wanted to use the name "Helios", which is confusing. Almost as many as the designs who all want to use the name "Orion."

This Helios is closely related to the Project Orion designs, in as much as they both used tiny nuclear bombs as propulsion. Sadly the Helios concept had some fundamental design problems that it never overcame.

The basic idea was created by visionary Dandridge Cole who was then working at the Martin corporation. Mr. Cole was unaware of the nuclear-shaped-charge innovation, so he thought the Project Orion design was wasting 90% of the bomb energy. He figured he could do better than that. The more you surround the bomb, the less energy you will waste. Since most material objects fare poorly when hit by a nuclear blast, Mr. Cole used three strategies:

• The reaction chamber surrounding the bomb was given a huge radius. This spreads the ravening energy of the blast over more chamber wall area, so each square meter of wall has to deal with a smaller portion of the total blast. Keeping in mind that when he said "huge", he wasn't fooling. The first design had a reaction chamber diameter of a whopping 40 meters (130 feet).
• The bombs were much weaker than the Project Orion pulse units, so the total blast was less. Project Orion units were 1 kiloton, Helios units were 0.01 kiloton, or one hundred times weaker.
• 390 kilograms of water propellant was injected into the chamber prior to each bomb. The pious hope was that the water would soak up the blast and go shooting out the exhaust nozzle at high velocity, instead of the chamber walls. Hopefully the water would also cool off the chamber walls so they wouldn't melt.

The Cole model I had engine performance that can be charitably described as "disappointing". Specific impulse was 931 seconds, which is in the upper range of conventional solid core nuclear thermal rockets. At one pulse per second the engine had a thrust-to-weight ratio of only 0.25, enough to land on Luna but not enough to lift-off from Terra. By comparison small first generation Project Orion ships were expected to have a specific impulse of 2,500 seconds and a thrust-to-weight ratio of at least 4.0.

One little inconvenient detail that Mr. Cole glosses over is the problem with tiny nuclear bombs. You see, fission reactions have that tiresome "critical" mass requirement. Meaning that if you use less than the critical mass there will be no fission chain reaction. The problem is that a critical mass of uranium-235 or plutonium will ordinarily make a much bigger bang than 0.01 kiloton. Damping the bomb down to 0.01 kiloton means that most of the uranium or plutonium does not enter the reaction. Instead they are merely volatilized into glowing radioactive vapors of death and spread to the four winds at high velocity. This makes it difficult to get permission to launch this monster from Terra's surface.

Even ignoring the radioactive contamination the inefficient use of fissionables is unconscionable. Weapons-grade uranium and plutonium are monstrously expensive, and this design will use tons of the stuff.

A more ambitious (and utterly insane) version was Cole's nuclear pulse jet. This would be a titanic airbreathing version that utilizes Terra's atmosphere as propellant until the ship climbs into space. The radioactive fallout would be only slightly less horrific than that from Project Pluto. The difference was that Pluto was supposed to be a weapon.

Cole and the Martin corp stopped working on the concept in the early 1960s, because of the lack of interest on the part of the USAF, NASA, and Martin higher management. There were a few amusing "artist conceptions" of the concept created by the advertising department of other aerospace companies that wanted to appear new and trendy.

AMERICAN BOSCH ARMA CORP HELIOS

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ATOMIC PULSE ROCKET

American Bosch Arma Corporation

This is the Atomic Pulse Rocket, a pot-bellied space ship nearly the size of the Empire State Building, propelled by a series of atomic blasts.

The enormous rocket (weighing 75,000 tons fully loaded) is designed to leave Earth with a thrust of 100,000 tons. Altogether a thousand atomic blasts—each equal to 1,000 tons of TNT—are fired from a low velocity gun into a heavy steel rocket engine at a rate of one per second until the vehicle leaves Earth's atmosphere. Then steam and vaporized steel maintain the thrust. After transit speed is reached, and the propulsion system shut off, power is provided by solar batteries plating the wing and body surfaces.

Inside the rocket. living quarters are situated in the rim of a pressurized wheel-like cabin which revolves to provide artificial gravity. Radio and radar antennae revolve with it. Tubular hydroponic "gardens" on either side of the rim grow algae to produce oxygen and high protein food.

The Atomic Pulse Rocket could transport payload to the Moon at $6.74 per lb., less than one quarter the prevailing air freight charges over equivalent distance. A similar project is past the pilot-study stage in the Defense Department

(ed note: This is vaguely based on the Cole study, but is more public relations than a real engineering design study)

Helios Nuclear Pulse

In 1963 the Lawrence Radiation Laboratory started working on their own version under the name of Project Helios. This was for a crewed mission to Mars. Mass in low Earth orbit (IMLEO) was to be 680,000 kg, delta-V of 18,000 m/s, delivering a 45,000 kg Mars lander into Mars orbit (total payload 91,000 kg).

The reaction chamber would have a diameter of 9.2 meters (radius 4.6 m); into which would be introduced 100 kg or so of hydrogen propellant, a small nuclear explosive charge, and a sacrificial positioning framework to hold the nuke at the center. This will be added with each detonation, at intervals of 10 seconds or longer. Of the hydrogen propellant, nuke, and framework mass; the fraction that is hydrogen propellant is called χ.

The nuclear pulse units were one meter in diameter. The core is a 2 kg sphere of weapons-grade uranium. It is coated by 5 kg of high density chemical explosive, and the entire clanking mess is jacketed by 15.7 kg of low density chemical explosive. The nuclear explosive yield is a miniscule 0.0051 kilotons (5.1 tons).

Nozzle

The nozzle sticking out of the chamber is conical with a 20° half-angle. The mass of the nozzle is approximately:

M_N: mass of nozzle
k: a constant, report does not specify its value
ϵ: area expansion ratio of the nozzle
p₀: initial pressure within the chamber
r_t: radius of the nozzle throat
(ρ/σ)_N: weight/strength ratio for nozzle material

Pressure Vessel

The minimum mass of a spherical pressure vessel that can withstand a steady internal pressure p without exploding into a zillion pieces is:

A factor of 4 is then included because the engine is NOT subject to a steady pressure, the pressure pulsates. Then an additional safey factor of 2 is added. So the equation becomes:

M_s: mass of pressure vessel
V: cavity volume in the shell
ρ: density of shell material
σ_s: shell material yield stress
p: steady pressure
p₀: initial pressure within the chamber

Payload

The analysis used the payload mass M_F to "hide a multitude of sins." It includes the mass of the nozzle throat valve, shock absorbers, shadow shields, life support, observational equipment, Mars excursion vehicle, and Terra atmospheric reentry vehicle. They figure that the sum of the nozzle throat valve, shock absorbers, and shadow shields will come to a total of less than 9,100 kg.

Propellant

The analysis assumes that the liquid hydrogen propellant will require an additional 8% of propellant mass for tanks, insulation, and boil-off. The ratio of hydrogen tankage mass to useful hydrogen mass is α.

Nuclear Charges

Each nuclear charge and the sacrificial positioning framework is assumed to have a combined mass of 32 kilograms. There will be an additional 2.3 kg per unit for storage and handling in the pulse unit magazine. The ratio of the charge storage/handling mass to the total mass of the charges is β

Total Mass of Vehicle, Propellant, and Nuclear Charges

Remember that each pulse start with the pressure chamber containing hydrogen propellant, a nuclear pulse unit, and a sacrificial framework holding the nuke at the chamber center. The nuke and the framework will be volatilized in the explosion, and the volatilized gas plus the propellant will be heated and sent out the exhaust nozzle. Of the combined mass, the fraction that is hydrogen propellant is called χ.

δM_H: mass of hydrogen propellant
δM_c: mass of charge debris: volatilized nuclear charge and sacrificial framework
χ: propellant fraction

If it takes N pulses total to accelerate the vehicle to the mission delta-V, then the total amount of effluent mass is:

Remember that the ratio of hydrogen tankage mass to useful hydrogen mass is α and the ratio of the charge storage/handling mass to the total mass of the charges is β

Total Initial Vehicle Mass

Using the equation to determine mass ratio (μ) from delta-V and specific impulse (or exhaust velocity) we can make an equation that will spit out the total vehicle mass (M₀) given the mission delta-V (ΔV) and engine specific impulse (I_sp)

Combining the effluent mass equation and the total vehicle mass equation we can create three new equations:

Vehicle "Cost"

The cost of the vehicle is assumed to be $91 per kilogram (cost of delivering vehicle components into LEO) plus $50,000 per nuclear charge. Both in 1960 US dollars.

Above graph is number of pulse units (N) vs plenum chamber radius (r). Superimposed on top is a grid of chamber pressure (p) vs propellant-to-total-chamber-contents fraction (χ).

Plotted are the family of curves for vehicle cost COST (10⁹$) in units of billion of 1960 US dollars.

• Payload Mass (M_F) = 9,100 kg
• Nozzle expansion ratio (ϵ) = 200
• Chamber temperature (T) = 6000 K
• Delta-V (ΔV) = 18,000 m/s

The cost curves close on the left because the mass of the chamber increases rapidly with pressure, due to the thick-shell correction.

The cost curves close on the right because the enthalpy and specific impulse decrease with decreasing pressure for a fixed expansion ratio and initial temperature.

HELIOS

      As was mentioned, the HELlOS concept is by no means new. In its simplest form HELlOS is a large cavity, or pressure vessel, filled with high-temperature and pressure gas. This gas is obtained by introducing it cold to the chamber and then adding a large amount of energy. A plug in the throat of a nozzle attached to the vessel is removed and the gas flows out the nozzle, producing thrust.

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     What we feel is reasonable to include on the rocket for which HELlOS is the engine is shown in Figure 1, which may or may not bear any relation to reality. Forward of the engine is shown a warehouse for the energy sources. Next are propellant tanks, biological compartments, re-entry vehicles, mission vehicles, and what-not. Since most of the thrust is produced by the nozzle, this has to be a fairly substantial structure. Shown in Figure 1 are superstructures surrounding the nozzle which would be dead, wasted weight to be carried along.

• 

     Figure 2 is a rather exotic device by which we bring this dead weight back to life. We use the nozzle itself as part of the structure. The biological comparimetitS, mission vehicles, and propellant storage tanks are carried on the edge of the nozzle. The skirt on the end of the nozzle was placed there by the artist to serve as a warehouse for the nuclear devices. The nozzle expansion ratios we talk of are on the order of 100 in area.

     In all of the schemes that have been discussed so far in the meeting, there are two questions that must be answered. First, is it possible to build the machine, and the second, is it worth building? I believe that we can categorically state that it is possible to build this engine given enough mass. Naturally, our primary concern is seeing just how light we can build the engine—and still make it work.

     As was mentioned, the idea has been discarded several times previously. The reasons for this were two-fold. First, until recently, reliable steels had yield strengths below 100,000 psi, and second, nuclear device technology dictated that reproducible yields were in excess of 20 tons of H.E. equivalent energy, and that in order to make efficient use of the nuclear fuel the yields had to be in the kiloton range.

     How these two figures tend to fix the weight of the pressure vessel is explained as follows. If we assume that the nuclear energy is transformed into. the internal energy of a perfect, γ-law gas contained within an elastic, thin-shelled, spherical container, then it is quite simple to show that the total weight of the pressure vessel is a linear function only of the specific internal energy of the gas. A few years ago it was estimated that the constant of proportionality between shell weight and nuclear yield was 100 tons. of pressure shell per ton of yield. This, coupled with a minimum of 20 tons of yield (8.37×10¹⁰ J), indicated excessive engine weights. Recent advances in materials technology have led us to believe that a value of 20 tons/ton is realizable. This will be discussed in a bit more detail later. Additionally it is felt that it is not unreasonable to consider nuclear yields down to around 2 tons of H.E. equivalent (8.37×10⁹ J).

     Soon after the energy of the nuclear device is released the vessel will be subjected to nuclear radiation, recurrent shock waves, high pressures, thermal radiation, and energy transfer by conduction and convection from high temperature gas. In order to arrive at quantitative estimates of the internal environment of the shell we ran a series of problems on one of the existing digital computer codes at LRL. The code used was designed for the weapons program, and the equations-of-state are known to be incorrect in our region of interest, although not grossly so. The greatest inaccuracy is in the computation of radiation transport. However, we expect that the results are not grossly in disagreement with nature.

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     Figure 3 shows the computed pressure on the inside of the vessel wall as a function of time. The first spike is the initial shock wave that hits the wall and reflects back inwards to the center of the cavity and reverberates repeatedly between the wall and the center. There may well be structural problems created by the shock waves traveling through the material of the pressure vessel. We are just beginning an experimental program to determine the internal behavior of metals under repeated, low-intensity shocks.

• 

     Figure 4 indicates the response of the shell to this sort of a pressure history. The wave labeled "b" represents response of the shell to a step in pressure to the final steady-state pressure. The curve labeled "a" is the response of the shell to the same pressure step with a short, square-wave pulse near time t = 0. We have shown that the pressure profile of the previous slide yields about this same response with minor corrections for the later reverberations.

     We expect the pressure in the vessel to start dropping noticeably in a few milliseconds. From Figure 3 we see that the gas is not too far from hydrodynamic equilibrium at times of this order. This exhaust-time scale is varied by the area of the exhaust port. At present the nozzle throat is fixed such that about 90% of the gas will be exhausted in around a tenth of a second. From Figure 4 we see that there is an overshoot of the wall displacement from what would be expected from a step in the pressure. It is something like two and a half times the mean pressure expansion. The weight/yield constant of 20 tons/ton quoted earlier comes from increasing the weight of the pressure shell by a factor of four over that which would just contain the steady-state pressure without yielding. This slide indicates that we only need a factor of 2.5; however, this would leave no safety margin. The factor of four is used throughout all of our performance and mission analyses.

     Perhaps our most serious and least understood problem is the energy transfer from the gas to the vessel. Very crude calculations indicate that there will be something like 100 cal/cm² deposited on the inside of the wall of the pressure vessel. About 10% of this comes from the initial fire ball, and will be deposited within the first milliseconds. We need very much to have a physical understanding of the processes that go on here. We are working on this from a purely theoretical standpoint at the moment. The high-energy deposition rate from the fireball may necessitate the use of some kind of ablative coating on the inside of the wall.

     At any rate, we believe that there will be between 3 and 4% of the total nuclear energy, of the total energy released, deposited in the metal itself. Perhaps half a percent of this will be in the form of dynamic ringing of the vessel.

• 

     This 3 to 4% of the energy that gets into the wall will account for something like a 15 to 20°K rise in the temperature of the wall per pulse. Sooner or later, it will be necessary to remove the heat from the wall as demonstrated in Figure 5. Here is plotted the effective strength as a function of temperature and pressure. The q is that "safety" factor four; omega (ω) is the tons/ton of yield which, it will be remembered, was claimed to be completely constant, dependent only on the energy contained within the pressure vessel. This is true only for a perfect γ-law gas which hydrogen is not, at temperatures in the dissociation region. The formula for omega includes a factor which is the ratio of the PV energy of the gas to its internal energy. The fact that this ratio is a complicated function of both P and E generates the three different curves shown. From Figure 5 we see that the weight of the pressure shell needed to contain this energy is fairly flat as a function of temperature up to around 700°K, then it starts rising rapidly. In order to keep the weight to a minimum, we must keep the temperature as low as possible. Since energy cannot be radiated to space from a body at 700°K rapidly enough to be commensurate with the pulse rates we contemplate, it seems that the only reasonable place to deposit this energy is in the propellant itself. Even there, energy may not be removed from the wall unless there is a temperature difference maintained between the propellant and the wall. It will take several pulses to get up to a reasonable temperature difference, and we feel that we can maintain this difference by propellant flowing through the pressure shell wall. Naturally, if we were to expel the gas faster, we wouldn't have as much of a heating problem; but this would say that there would be more peak thrust from the engine which would demand that the nozzle and shock absorbers be increased in strength and weight. There would thus appear to be some optimum peak thrust dictated by the cooling requirements of the pressure shell and the mass requirements of the nozzle.

• 

     As the gas expands out the pressure vessel, it follows the curve given in Figure 6, if an infinite nozzle, complete recombination, and 100% isentropic efficiency are assumed. We have performed some hand calculations to determine the effect of a fixed expansion ratio nozzle. We found that this actually tends to sharpen up the shoulder of the curve. Naturally, it drops the whole I_sp curve; but the I_sp does hold up flatter a little longer before it starts dropping off. We must also include the effects of recombination and the effects of a fixed amount of contaminant from the energy source, perhaps 70 lb, in the propellant. This also has to be included in the hydrodynamic calculations. We must know what effect the contaminants have on the recombination rate. We suspect that it will increase the rate by catalysis. This is something that we have not done much work on and will have to look into.

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     The basic question in our minds is the value of building the engine—as it seems possible to do so. The answer to this entails a certain amount of mission analysis. Very preliminary, but indicative results are given in Figure 7. It has been stated that mission analyses are to be avoided for this conference; however, we feel that they are necessary for an understanding of the problems of HELlOS.

     We have assumed for this a payload of 50 tons. The equilibrium temperature of the gas would be 6000°K and the pressure 100 atmospheres. We have assumed a rather optimistic Martian mission ΔV of 60,000 feet/sec (18,000 m/sec) for our analysis.

     Included in the analysis are analytic, dissociating-gas, thermodynamic properties for hydrogen, the rocket equation, material properties of steel for the pressure shell, and fairly optimistic estimates of tankage and nozzle weight. The thin-shell approximation was used to obtain the vessel radius and mass. Exclusive of specific impulse, the quantities obtained from this analysis are, in part, listed in Figure 7. It turns out that there are three independent variables which must be specified to completely determine a given system; and these we chose to be temperature, pressure, and specific. impulse. One might wonder how specific impulse could be an independent variable here. This obtains because we fix the amount of energy-source debris we have in the propellant at 70 lb. The rest of the gas is hydrogen. The debris is treated as a perfect, γ = 5/3 gas of molecular weight 12. Figure 7 is a typical page of computer output and shows, for example, that if we want 1800 sec specific impulse from the given conditions, the mass fraction of hydrogen would then have to be 96%. Now with 70 lb of debris, the mass of gas in the cavity is quite large. The vessel radius would be something like 30 ft and the yield around 40 tons, which means that the pressure vessel would have to be quite heavy. The number of pulses for this particular case is around 2000. This is a fairly low nuniber, because of the high specific impulse and the large amount of propellant exhausted per pulse. The total mass of the rocket is quite large because of the large engine weight.

     Now at the other extreme in I_sp we see that we need only a small yield and a low weight percentage of hydrogen in the bottle. However, the low mass expelled per pulse and low I_sp demands many such pulses for a given mission ΔV. In this case a large total rocket mass is produced from the rocket equation with low I_sp rather than a high engine weight, as in the high I_sp case. As shown, between these extremes there is a minimum in total weight.

     We have devised a figure-of-merit for a given set of parameters which we guardedly call the mission cost. It is expected that the amount of fissionable fuel is around 1 kg per pulse. We take $50,000 to be the cost per pulse. NASA indicates that $200 per pound is not an unreasonable freight charge to place the total mass in orbit. Our figure-of-merit is the sum of these two. We find that the minimum in mission cost occurs not too far from the minimum in total mass. Thus, for these particular parameters the least expensive mission would start with around 700 tons in orbit, have average I_sp a bit above 1500 sec, and use between 4000 and 5000 pulses.

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Question: And the I_sp is at equilibrium?

Answer: It is the equilibrium value multiplied by some number less than one to account for the effects of a finite nozzle, finite recombination, and pulsing. We have used 0.8 for the results given in Figure 7.

Answer (Cooper): The point is that if the energy of the bomb were added to pure hydrogen, the temperature would be over 10,000°K. However, the energy is divided between the hydrogen and the bomb material which has a molecular weight of 10 or 12. This extra material thus lowers the temperature and raises the average molecular weight resulting in a lower equilibrium specific impulse

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     One good feature of this machine is that it can be tested underground. In this connection one can conceive of many good physics experiments that can be carried out with a pressure vessel which could contain a nuclear shot.

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Question: What is the time between shots?

Answer: There is no a priori fixed time between pulses. This depends upon the cooling rate of the vessel and the complexity of introducing the next device. We want to make the pulse rate as fast as possible. Typically, in the absence of shock absorbers, the acceleration is about a 2 g pulse lasting for a tenth of a second. Now, if you wait 10 sec between pulses the average thrust is then 0.02 g. We would like as high an average thrust as possible.

Question: Would you assume that your charges would need no shielding for radiation?

Answer: That's right. We have so far.

Question: Do you think this is realistic?

Answer: Perhaps not.

Question: You mentioned 1 kg of nuclear fuel. Is that a critical mass in this case?

Answer: Well, it depends upon how you build these devices. I can't go into that deeply, but I think 1 kg is about as low as you're going to go.

Answer (Cooper): I think that the point is, that it's kilogramish and that it's not 2 to 500 grams. A factor of two one way or the other is reasonable if the fuel is plutonium and somewhat larger, naturally, if you use uranium as the fuel.

Question: Does some of the propellant go through the wall? You can see that the dynamics of the shell would be changed when you have some flow of the coolant through the wall.

Answer: We have not yet considered the dynamic response of any structure other than a thin spherical shell. We would like, if possible, to spray the inside or have flow holes in the pressure vessel, which is going to increase the weight. Practically any modification of this sort is going to increase the weight of the pressure vessel. We should also be worried about the thermal stresses on the inside. wall. This is something that we have not looked at closely and may well be one of the more serious problems.

Question: What do you think will be the total mission cost?

Answer: The last slide will give that to us.

Answer (Cooper): These are based on something like 50 tons round trip through 60,000 ft/sec, and the cost generally turns out to be divided about equally between the fuel and the weight that you have to put into orbit and the minimum cost is about the minimum weight. So, it's about $148,000,000 to put the thing into orbit, and it's something like a $140 to $200 million for the charges.

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     The total cost is thus around $500 million (in 1965 dollars. About $4 billion in 2020 dollars).

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Question: How much hydrogen do you use per pulse?

Answer: Each case is different. Shown in this slide is the percent of hydrogen necessary to produce a given I_sp· As you will remember, we fix the weight of non-hydrogen additive at 70 lb. From this you can easily get the total mass of gas expelled per pulse. The numbers presented here are merely indicative of our general analysis.

Question: What's the average thrust to weight ratio for the engine alone?

Answer (Cooper): Well, he really answered the question before. It depends on the rate at which you fire; and in particular, it doesn't matter. I think they want to fire at such a rate that they have better than 0.1 g, on the average. We made some computations on ground take off, and then it appeared to be quite a problem. It was necessary to fire one per second or faster.

Answer: Well, if you wanted to take off, there are two problems. The first is a political one, and the second is how fast the engine can be pulsed.

Question: Was the mass number for the pressure shell based on the properties of maraging steel?

Answer: That's right. It's just about the best we can do. We don't know if the yield strength we assumed is an over or under estimate. We are just getting into the experimental study of the problem.

Answer (Cooper): Right. Oh, incidentally, let me mention that there's an unclassified report that Bob Fox did on this, in I guess about 1957, which has almost all of this material.

R. H. Fox, "A Study of the Nuclear Gaseous Reactor Rocket,". UCRL-4996, Oct. 31, 1957 (Uncl.)

FINAL CALCULATION

1. Improvements in Computations

     The performance code described in AFN 32 and 45 has undergone two additional, important improvements. For the most part, the code employs tables of real gas properties to determine the thermodynamics of the propellant gas. However, in three crucial steps of the performance calculation it employs an approximation which assumes the gas to behave like an ideal, γ-law gas.

     Our approach to approximating real gas behavior for each of these steps was to keep the functional form of the equations used in each step unchanged. We then used three different γ's which were functions of T_o , p_o, χ, and ϵ and forced a fit to the real behavior of the gas.

     The second major improvement was to include the temperature dependence of the yield strength of the pressure vessel material.

2. Results

     Figure 2 is a plot of the number of pulses vs the radius of the cavity with grid of p_o vs χ (the mass fraction of hydrogen in the gas) overlaid. We note a marked difference in the optimum value of pressure and chamber radius from the values previously published. This is due almost entirely to the second term in equation (l) which tends to favor small radii in order to keep the chamber wall strength high. The inclusion of a temperature dependent material strength also allows us to optimize with respect to temperature as seen in Figure 3.

• 

• 

     The vehicle also has a minimum in cost and total mass with respect to the discharge time constant. As we decrease τ the nozzle throat, and thus the nozzle mass becomes large; however, as we increase τ the pressure vessel heating and thus mass becomes large. This minimum is graphically shown In Figure 4a.

     For the rest of the plots of Figure 4 we have held temperature, pressure, χ and ϵ fixed at their near-optimum values of 6750°K, 1000 Atm, 0.7, and 300, respectively. δM_c is the mass of the nuclear charge, assumed to be pure carbon for calculations purposes. β is the hydrogen tankage fraction and η is the "tankage" fraction for the nuclear charges. Figure 4e depicts the sensitivity of the cost and total mass on the yield strength of the pressure vessel. We have assumed that the temperature dependence is unaltered and have merely multiplied the strength values given in Figure 1 by the factors shown as the abscissae in Figure 4e. Figure 4f clearly indicates the sensitivity of the overall vehicle performance to the amount of energy absorbed per pulse in the pressure vessel vail. The abcissa point labeled "Base + 1%” indicates that the constant multiplying Y in the first term of equation (l) has been altered to 0.0271, etc.

• 

• 

     Figure 5 shows the dependence of total vehicle mass on the mission demanded of the vehicle. Each point represents the conditions obtaining when the pressure and expansion ratio is held fixed but temperature and χ are chosen such that $ is minimum with respect to both of these. "Mission" is defined here as the specification of both the payload (fixed mass) and the ideal velocity Increment, ΔV, through which this payload must be accelerated.

• 

     In Table 1 we present the various vehicle parameters obtained uhen the Helios vehicle is optimised with respect to pressure, temperature, χ, and expansion ratio for a mission of payload = 100 tons and ΔV = 60,000 ft/sec. Of interest is the indication that the diameter of the pressure vessel is about 10 ft, the nozzle throat radius is 1 ft and the nuclear yield required is 3.6 tons of H. E. equivalent energy.

Table 1
Properties of an
Optimized Hellos Vehicle

Extrinsic (assumed) Variables
MF	100 tons
ΔV	60,000 ft/sec
τ	0.2 sec*
δMc	70 lb.
β	0.08
η	5/70
Optimised Variables
po	1,000 Atm
To	6750°K
χ	0.70
ϵ	500
Intrinsic (calculated) Variables at Optimum
Mo	1,005 tons
N	6,052 pulses
$	0.7045 x ($109)
MN	25.7 tons
Ms	118.2 tons
Îsp	1,538 sec
Mp	494.2 tons
r	1.933 meter
rt	52.4 Cm
Y	3.6 tons

* We see from Figure 4a that this should also be an "optimised variable" having an optimum value of about 0.27 sec.

The Martian movie is based on the novel of the same name. Both have the Atomic Rockets Seal of Approval. Enough said.

Warning: this section contains spoilers for the novel and the movie.

Hermes in the Novel

Author Andy Weir based the original mission on Robert Zubrin's Mars Direct proposal. Weir updated Zubrin's chemical rocket to a nuclear-reactor-powered ion drive using argon propellant. You see, a puny chemical rocket has to use Hohmann transfer orbits which have launch windows tied to the synodic period of Mars. That mission would have had a required stay time on Mars of a little over a year. For dramatic reasons, Weir needed the mission capable of being aborted at any time with a return to Terra. The ion drive allowed this.

In Andy Weir's original conception, the Hermes is cone-shaped so it can aerocapture at Mars and at Terra, saving precious propellant mass.

The Hermes has an acceleration of 0.002 m/s² (2 millimeters per second, per second). Andy Weir said that the delta V budget for the return trip was about 5,000 m/s.

The spacecraft is split down the middle parallel to the long axis. This allows the two halves to separate, attached with cables, so they can spin like a bola to provide artificial gravity. The halves are called "Semicone-A" and "Semicone-B".

The main airlock/docking port is located in the customary place, on the nose.

Andy Weir mentioned that the movie version of the Hermes has quite a different design. But he also noted it was "way cooler-looking than the version I imagined."

THE MARTIAN: A TECHNICAL COMMENTARY

An Ares mission begins with 14 uncrewed launches (probably with an Atlas or similar sized booster) dropping airbag-cushioned payloads on Mars. These would each weigh about 1000kg on launch, with up to 600kg of payload to the surface. This includes parts of the Hab and supplies.

The crewed part of the mission is mediated by the Hermes, a large vehicle for deep space with a nuclear powered ion drive designed to fly between Earth and Mars and back. The Hermes is used by every mission and was assembled in Earth orbit at (no doubt) astronomical expense.

The six astronauts of Ares 3, together with their supplies are launched from Earth to Hermes. The Mars Descent Vehicle is launched separately towards Mars at about this time.

Hermes and the MDV travel to Mars, parking in Mars orbit after 124 days in deep space. Hermes remains in orbit, uncrewed, while the MDV flies the astronauts to the surface.

On the surface, the astronauts build their Hab from airbag cushioned cargo drops and perform their mission. After 30 days on the surface, they climb into the Mars Ascent Vehicle and fly back up to orbit, where they meet the Hermes and fly back to Earth, taking 208 days to return.

Hermes parks in Earth orbit and the crew return to Earth in some re-entry vehicle like Orion or Dragon.

The MAV was launched years before, made its return fuel on Mars using electricity and ambient atmosphere, then was used for about 6 hours to get back to the Hermes.

This mission architecture is very credible, given a nuclear powered ion drive, which is technically possible but politically problematic. IMO, the architecture is inefficient given most of the hardware is used only once, Hermes is not self sufficient, and the astronauts spend only 30 days on the surface.

Mars Ascent Vehicle (MAV) is quite large. It had to be soft landed, but even empty weighs much more than all 14 presupply missions combined. Given that NASA has (in the story) developed soft-landing capability for tens of tonnes, it's not clear why stuff as mission critical as the Hab is landed relatively inaccurately in lots of parts. It could be that this simply reduces mission cost and complexity, or that there was no practical way to land something as bulky as the hab (even disassembled) in one piece.

The MAV employs In-Situ Resource Utilization (or ISRU) to make fuel and oxidizer for the return flight. Two (Earth) years of power from a 100W Radioisotope Thermoelectric Generator (RTG) is enough to make 13kg of fuel (methane) and oxidizer (oxygen) from every 1kg of hydrogen (H2) precursor brought from Earth, for a total of nearly 20T of fuel.

At various points of the novel, Weir describes the MAV as weighing 32 metric tons when fully fueled, and standing 27m tall. This implies that it is very long and skinny, which is unnecessary in the thin Martian atmosphere. Not only that, this means a lot of rocket mass relative to the amount of fuel it can carry (spherical rockets are vastly more efficient, absent significant atmosphere). Needless to say that's a bad thing. By comparison, the Falcon 9, a long and skinny rocket by usual standards, is about 70m tall and weighs about 600T on launch. The MAV could easily be a conical shape perhaps 5m wide and 10m tall.

Weir states that it has two stages, though one stage is perfectly adequate for the relatively low delta-V required to reach Low Mars Orbit (4.1km/s). Nevertheless, with a 325s Isp methane-oxygen engine, a two stage system would have a 16T first stage, a 8T second stage, and a 8T orbital module, with an implied mass fraction of 81% fuel vs 19% metal in each stage.

Towards the end of the novel, engineers at JPL describe the MAV as having an unrealistically low launch weight of 12,600kg (12.6T) — similar to a fully-loaded Dragon capsule. So we'll assume this is the dry mass. Let's assume, then, that the orbital module is 8T, the first stage is 3T, and the second stage is 1.6T, empty. The 19.397T of fuel is distributed accordingly, implying an engine Isp of 405s in order to reach 4.1km/s of Low Mars Orbit. This is low for H2/O2 engines, but extremely high for a methane-oxygen engine. Even SpaceX's planned monster Raptor engine has a notional vacuum Isp of 380s.

In order to get to 5.8km/s and intercept the Hermes, the mass of the orbital module needs to be reduced from 8000kg to 4280kg, a reduction of 3720kg. This takes into account adding 780kg of fuel, removing 500kg from the first stage (pulling off an engine), and so on. The accuracy of the numbers indicate that Andy Weir did the math, but it's not clear on what metrics he designed the MAV and its launch system.

More generally, given that the total delta-V needed to get from Mars to Earth is *only* 7.8km/s, a MAV that flies all the way back to Earth is completely possible, though it would probably need to be bigger than the MAV presented here to have adequate life support. But given that the fuel/delta-V is most easily obtained on the surface of Mars, rather than brought from Earth, a direct ascent architecture actually makes a lot of sense.

On Sol 68, Watney points out that NASA never used large RTGs on crewed missions before Ares, but during the Apollo program RTGs were deployed by astronauts to power lunar seismometers. On Sol 69, Watney states that Lewis had buried the RTG for safety reasons. A RTG stashed somewhere on the surface, however, is much less likely to overheat.

Mars Descent Vehicle (MDV). In the Sol 7 log entry, Watney mentions that the Mars Descent Vehicle (MDV) is useless to him for escaping, since its thrusters cannot even lift its own weight. This, of course, refers to its weight when fully fueled. Before landing, much of its fuel has burned off and it can achieve neutral thrust for a hovered landing. Nevertheless, it lacks (by far) the fuel capacity, thrust, and delta-V necessary to fly anything back to orbit!

In Chapter 8, Bruce and Teddy discuss potential MDV modifications. It is strongly implied, though not stated, that the design would not admit the addition of more engine clusters, and they don't have the time to invent a bigger engine. It is likely that this is a narrative device.

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Orbital Mechanics

On the first page, Watney states that he was six days into the best two months of his life. Evidently he was confused, as later it's made clear that the surface operations of the mission were only 30 days long.

30 days on the surface is the extent of surface stay permitted under an "opposition" class mission, wherein the astronauts fly by Venus on either the outbound or inbound leg. While shortening the overall mission, opposition class missions significantly lengthen the time spent in space, and also bring the spacecraft much closer to the Sun, increasing the crew's exposure to radiation.

The alternative mission design is the "conjunction" class mission, wherein the crew takes a relatively short 4-6 month Hohmann transfer flight either way, with a ~560 day stay on the Mars surface in between launch windows. Obviously, if Watney had been stranded on a conjunction mission, he would have had no shortage of snacks!

One additional detail is that Weir's spaceship, the Hermes, employs ion thrusters throughout the mission, enabling a wider class of missions and trajectories than the traditional point-and-shoot orbital mechanics described in the previous paragraphs.

In Chapter 16, the Purnell maneuver is discussed, by which the crew can return to Mars fewer days than the 404 it would take Iris to get there. It's probably worth noting that there is a very similar delta-V requirement for Iris to get to Mars vs a resupply probe to reach the Hermes. The advantage of the Hermes approach is that Iris had to be able to do entry, descent, and landing. If this is the case, Iris could also get there faster by borrowing a basic ion thruster package from, say, the Asteroid Redirect Mission (ARM) spacecraft. It's also not clear why all the crew need to return to Mars (aside from narrative reasons) - most or all could return to Earth in the entry vehicle while Hermes takes the Purnell maneuver to Mars to pick up Watney before he starves. Although the remaining crew would then depend on a new entry vehicle being sent up to meet them on their eventual return to Earth.

In Chapter 20, Annie somewhat incongruously asks why Hermes can't wait at Mars for Watney to get there, when it seems he'll be slowed by the dust storm. Venkat points out that Hermes is on a fly by and can't slow down enough to be captured into orbit, but this is not entirely true. On Sol 505, Bruce says to Venkat that Hermes is flying by Mars at 5.8km/s. Mars escape velocity is only 5.5km/s (the Earth's is 11.2km/s by comparison) meaning that a delta-V of only 300m/s is needed to capture into orbit. Given that Hermes can accelerate at 2mm/s/s, a two day burn would be sufficient to capture into a big elliptical orbit, drastically increasing their margin of error. Perhaps, if Hermes slows enough for an orbital capture, its launch window to return to Earth will close too quickly to be useful.

Of course, the MAV was designed to reach Low Mars Orbit, with a delta-V of 4.1km/s. Getting to 5.8km/s is highly non-trivial, as discussed in the previous section describing the MAV. Of course the unmodified MAV has life support, so Watney could wait while Hermes spirals down to 4.1km/s to pick him up (~30 days, because Hermes can't exploit the Oberth effect), while in the modified MAV he gets close to 5.8km/s, making it much easier for Hermes to rendezvous. A MAV that got to, say, 5.2km/s would split the difference nicely. Either way, the most likely explanation is that maneuvering Hermes to do this would make them miss the Earth launch window.

On Sol 543, Beck mentions that the modified MAV will hit 12gs during launch. While they have lightened the primary payload by about 16%, Watney also removed a spare engine, suggesting that the unmodified MAV would hit at least 10gs during launch, which is unlikely for a rocket designed to fly humans! Later, Johanssen reads out a velocity of 741m/s at an altitude of 1350m, which is staggeringly fast, implying an acceleration of 20.7gs. Perhaps she dropped a zero?

When Johanssen and Vogel talk about getting Watney to orbit, what they mean is solar orbit, since Watney will have to escape Mars entirely in order to be intercepted by Hermes.

During the intercept procedure, Ares 3 crew have to think fast to find additional sources of delta-V to move the Hermes close enough to catch Watney as he flies by. The distances and velocities mentioned during this passage in Chapter 26 are correctly calculated and almost entirely realistic.

Watney suggests making a small breach in his suit and using the stored gas as a rocket to close the velocity mismatch of 42m/s. Assuming he has 5kg of gas on board (including reserve tanks) and an exhaust velocity of 400m/s (unlike rocket exhaust, it's not hot) this confers about 17m/s of delta-V, which is just not enough. This idea is transferred to the Hermes, which will spit out its atmosphere to slow down. Assuming Hermes weighs 100T, it would have to lose about 5T of air to make up the required 29m/s of delta-V. At sea-level atmospheric conditions, this implies that Hermes has a volume of 4000 cubic meters, or a floor area of 1300 square meters, or 13,000 square feet, which is the same as a very large house. Perhaps Hermes has large pressurized volumes that aren't used much for habitation? Martinez estimates that the air will take 4 seconds to leave, which implies a relatively small hole, since the shockwave would take about 0.1s to cross a Hermes-sized volume of air. A realistic concept is that Hermes is composed of two large Bigelow inflatable modules each with a diameter of about 12m, such as the BA 2100 habitats. Also worth mentioning that the process of blowing the "Vehicular Airlock" (VAL) will send lots of airlock fragments into space, hopefully missing Watney.

francisdrakex's version of Hermes

francisdrakex is a talented space artist who took a stab at designing the Hermes. He did an outstanding job if I say so myself, and not just because he was assisted by some data from this website.

The entire spacecraft was designed to fit inside a 5 m payload fairing for easy boosting into LEO.

His design used the "tumbling pigeon" method of artificial gravity, which is always a good choice to reduce the spin rate below nausea levels.

The ion engine array is mounted at the spin center, a classic technique from Stuhlinger's Ion Rocket.

As per standard best practices, the dangerously radioactive nuclear reactor is mounted as far as possible from the habitat module and the crew. The reactor has a set of heat radiators to reject waste heat. The radiators are in a triangular pattern, to keep them inside the shadow cast by the anti-radiation shadow shield.

The habitat module is a standard TransHab inflatable module.

Movie version of Hermes

The movie Hermes has the engines mounted aft the reactor, and a centrifuge to provide artificial gravity.

Ship's power is apparently from a set of 12 solar cell arrays, straight off of the International Space Station (the brown elongated rectangles). Which seems a bit redundant if you already have a nuclear reactor.

Rhett Allain did some calculations about the gravity centrifuge on the movie version of Hermes, and not surprisingly discovered that there was a bit of artistic license involved.

The novel states that the artificial gravity is 0.4 g. Mr. Allain did some measurements from the movie and figured the centrifuge is spinning at about 1.08 rotations per minute (0.109 radians per second). Unfortunately to produce 0.4 g the radius of the centrifuge would have to be an outrageous 329 meters! According to one of the graphics in the flight center, the Hermes is only 80 meters long.

Mr. Allain made some further measurements from the movie and concluded the centrifuge was about 9.0 to 14.5 meters in radius. To produce 0.4 g it would have to spin at an angular speed of 4.97 to 6.30 rotations per minute (0.52 to 0.66 rad/second). Which is right at the nausea limit.

Anyway 1.08 rpm is six times slower than 6.30 rpm, which is where the artistic license comes in.

Using the movie figures of 14.5 meters in radius and a spin rate of 1.08 rpm, the artifical gravity would be a pathetic 0.02 g, not 0.4.

Another difficulty is that the spacecraft is supposed to slow down at Mars and Terra by using aerobraking. This will require something like the ballute from 2010 The Year We Make Contact. Two of them, one for each braking.

Human Outer Planet Exploration (HOPE) is from the NASA report TM-2003-212349 by Melissa McGuire, Stanley Borowski, Lee Mason, and James Gilland (2003). . Revolutionary Concepts for Human Outer Planet Exploration (HOPE) { slide show }.

This was a given as a design problem for rocket scientists.

The problem was to design a manned mission to the Jovian moon Callisto, transporting a given payload, and returning the crew and scientific samples back to Terra. The payload included an In-situ resource utilization (ISRU) plant capable of cracking Callistonian water ice into hydrogen and oxygen rocket fuel. They assumed that space probe precursor missions had mapped Callisto's surface so that landing sites could be selected in advance, with due respect toward safety, operations, and scientific gain.

Calllisto was chosen as a destination because it is outside of Jupiter's radiation belts, and it has water ice on the surface for propellant production. The purpose of the mission was to establish an outpost and propellant production facility near the Asgard impact site on Callisto.

Several design teams entered the challenge, each basing their spacecraft around a different propulsion system for comparison purposes. The idea was to promote apples-to-apples comparison, as opposed to the sad proliferation of apple-to-oranges comparisons.

Transportation of the specified payload is left up to the mission designers.

Some designs use several unmanned spacecraft to deliver all the payload except the crew and TransHab module. Those arrive on a separate manned spacecraft, which is only dispatched upon successful arrival of the unmanned spacecraft.

Other designs using more potent propulsion systems have a single spacecraft carrying all the payload.

PAYLOAD

The standard HOPE payload is a TransHab crew quarter for six (including consumables and the crew), a Lander to ferry three crew and supplies to and from the surface of Callisto, a surface habitat module to house the three surface explorer crew members, and an In-situ resource utilization (ISRU) plant. The ISRU plant package includes an ISRU factory to crack Callistian ice into fuel for the lander, a reactor to power the ISRU plant and surface hab, and two rovers.

Some of the designs that use weaker propulsion systems and thus have longer mission lengths use two TransHab modules to reduce risk and increase available storage for the increased consumables required.

Payload: TransHab Module

This is pretty much a bog-standard TransHab habitat module, right off the shelf.

Crew quarters for six crew. Pressurized volume is about 333 cubic meters. Typically includes 15 metric tons of consumables, but varies according to mission length of the particular design.

Payload: Lander

The common base section carrying a three person crew pod. Can transport three crew to Callisto surface and back. It can carry down 40 tons to the surface.

Payload: Surface Habitat Module

The common base section carrying the inflatable surface habitat module. Can house three crew members on the surface of Callisto. It provides shelter and serves as a laboratory.

Payload: In-Situ Resource Utilization Plant

The common base section carrying the ISRU kit. This lands on Callisto the nuclear reactor, two rovers, and the ISRU plant to crack Callistian ice into LH₂/O₂ fuel for the lander.

This is a 1 MW-thermal reactor using a Brayton power converter to produce 250 kilowatts of electricity. It supplies power to ISRU plant and surface habitat module. Reactor is sited one kilometer away from habitat due to radiation. Alternatively tractors can be used to create hills out of local material to act as radiation shielding and reduce the mass required for reactor shielding and long cables.

Material on Callisto's surface is about 55% water ice and 45% rock. The ISRU plant will consume 215 kW of electrical power while processing 21 kilograms of water per hour into liquid hydrogen and liquid oxygen fuel for the lander. This will produce enough lander fuel for one lander sortie mission between the base and the orbiting ship every 30 days. The created fuel is stored in the fuel tanks of the common base sections of the ISRU plant and surface habitat. The engines of those common base sections are used as spares in case the lander's engines need repairs.

The rovers are equipped with bulldozer shovels in order to scoop and transport ice to the ISRU plant.

This section has been moved here

This is from Future space transportation systems systems analysis study, phase 1 technical report, section 3.3, page 139 (160 in PDF) Boeing, 1975.

When they say "independent", they mean the missions are totally self-contained. They do no rely upon logistics flights, support missions, orbital propellant depots, or anything like that. A few of the mission modes require a space shuttle flight to return the crew to Terra from orbit, but that's it.

There are several configurations, each with advantages and disadvantages. The baseline mission for all configurations is:

1. Departure from Terra orbit
2. Transfer to Luna
3. Entry into Lunar orbit
4. Lands on Luna four crew and 4,500 kg of payload
5. Supports a 14 day exploration stay on Luna
6. Lift-off into Lunar orbit
7. Transfer to Terra
8. Entry into Terra orbit
9. Return crew to Terra surface

The spacecraft is basically an Orbital Transfer Vehicle (OTV) transporting a Lunar Lander. Pretty much exactly like the Apollo Command/Service Module transporting the Lunar Module. Except much bigger.

There are four flight modes, distinguished by the lunar orbit used by the OTV while the lander is on the surface. The timelines for each are shown in the graph above.

• CREW and EQUIPMENT MODULE (CEM): habitat module and the crew
• CARGO MODULE(S) (CM): the surface payload
• LUNAR TRANSFER VEHICLE (LTV): transports the CEM and CM to the lunar surface, returns the CEM to lunar orbit at end of surface mission
• ORBITAL TRANSFER VEHICLE (OTV): transports the above between Terra orbit and Lunar orbit. And back.

CREW AND EQUIPMENT MODULE

This is basically the habitat module. The crew lives in this for the entire duration of the mission, except at the end of an Apollo-Mode mission (see below) where they briefly live in a reentry vehicle during aerobraking and splash-down.

The upper part is the crew section with a capacity of four crew members. The lower part is the equipment section containing electronics, life support, electrical power generators, and consumables. For the return-to-LEO option the typical life support endurance is 30 days (120 person-days), with a maximum capacity of 41 days (164 person-days). The direct-return option has a nominal life support endurance is 34 days (136 person-days).

One of the many options is the type of crew return method. The CEM can just Return to Earth Orbit (REO Mode) and have the crew retrieved and landed on Terra by a Space Shuttle mission. The alternative direct-return option (Apollo Mode) is to carry a Earth Entry Module (EEM) much like the Apollo Command Module and land the crew with an aerobraked reentry.

The return-to-Earth-Orbit option has a larger payload capacity, more life-support endurance, and more options for the Orbital Transfer Vehicle. The drawback is it relies upon an expensive Shuttle mission to retrieve the crew.

The Direct-Return-To-Earth option does not require an expensive Shuttle mission. The drawback is reduced payload capacity, reduced life-support endurance, and only one option for the Orbital Transfer Vehicle.

The Earth Entry Module has a mass of 6,870 kg with no crew, and a mass of 8,030 with crew.

CARGO MODULE

The two cargo modules contain the Surface Exploration Payload. For the 14 day surface mission, 4,535 kilograms of exploration equipment is provided. Any combination of equipment can be be accomodated. Each of the two payload pallets is approimately 4.4 × 4.4 × 6.0 meters (about 116 cubic meters).

This is from the novel Delta-V by Daniel Suarez. Yes, it is fiction but the author did his homework. The list of NASA scientists and rocketry experts in the acknowledgements is quite extensive.

The spacecraft is named after Konstantin Tsiolkovsky. Who you better have heard of or RocketCat is going to give you such an atomic wedgie.

The spacecraft is designed to travel to NEO asteroid 162173 Ryugu and spend four years mining the heck out of it. It will cannibalize three of its five rocket engines in order to construct three unmanned tugs. The tugs will carry large masses of ore back to cis-Lunar space, where they will accelerate the industrialization of space. The ores are of common elements and metals, but the point is they will already be out of Terra's gravity well. This increases their value a thousand-fold.

In the science-fiction novel, the five methalox rocket engines were purchased from Burkett's "Starion Aerospace" (i.e., five BE-4 methalox engines purchased from Jeff Bezos' Blue Origin). Each engine has 400,000 pounds of thrust (1,993,000 N, total of 9,965,000 N). They are fed from a cluster of oxygen and methane bladder tanks. The tanks are shielded from solar heat by a mylar mirror shield, to avoid fuel boil-off losses. The tanks are clustered around the ship's refinery, used to separate the asteroid material into various elements.

The inflatable habitats were built by Ray Halser's company (i.e., TransHabs built by Robert Bigelow of Bigelow Aerospace)

The spacecraft is just over 250 meters from prow to stern.

The prow of the spacecraft has a large solar photovoltaic array and a communication antenna.

Next is a hub containing four docking port, each with a Mule class utility vehicle. The hub also has four suit ports with suits, and four teleoperated drones.

The mid-section has an artificial gravity centrifuge, with three arms each with an inflatable habitat on the end. Two habs are for the crew to live in, the Fab hab is a factory with 3D printers and other manufacturing equipment. The spin radius is 106 meters, the artificial gravity at the bottom of each habitat is 1 g. I calculate the rotation rate to be 2.9 rotations per minute, which is within the spin-nausea limit. The centrifuge has three arms because two arms are not spin-stable. The three arms are at asymmetric angles because the Fab hab is much more massive than the two crew habs. Each arm has a rail-mounted ballast weight used to balance the spin as crew and material moves from hab to hab. The arms must spun down, folded "upward", and locked to the ship's spine before each burn of the main engines. Otherwise the arms might shear off. This is not a problem since there are only two engine burns during the Terra-Ryugu trip.

They need spin gravity because the mission plan calls for them to stay at Rhyugu for four freaking years. Zero gee for that long might not be survivable.

Connecting each habitat to the hub are 2 meter wide tunnels, with electrical/data/plumbing conduits along three sides and composite ladder rungs on the fourth. To save on mass and maintenance an elevator is not used. Instead there is a steel cable with carabiners on a planetary gear. The crew hook their suit harness to the carabiners. The cable can transport a crewperson the length of the tunnel in five minutes.

Next comes a bay carrying four Honey Bee mining robots.

After that comes an ore refinery, surrounded by bladder tanks of methalox fuel, protected from solar heat by a mylar sun shield.

At the bottom is the engine cluster of five methalox engines.

The spacecraft has a dry mass of 346 metric tons and a wet mass of 563 metric tons (mass ratio of 1.63). It was secretly assembled in Lunar DRO for plot related reasons, at a cost of nine billion dollars. The fully amortized cost including construction and R and D was more like twenty-four billion

Every century or so there is an orbital window for an ideal Terra-Ryugu trajectory. According to the novel the next is December 13, 2033. Starting at Lunar DRO during the window a Trans Ryugu Injection will cost only 1.7 km/sec delta-V, have a trip time of 48 days. and a Ryugu Orbital Insertion of 1.9 km/sec. Total of about 3,600 m/s delta-V. The ship was designed to carry the bare minimum propellant for the return trip, but that would not be needed if they can demonstrate an ability to extract oxygen and methane from Ryugu in-situ resource utilization within ten days of arrival. If successful, they can stay and fill their tanks to the brim.

While mining Rhyugu, the Konstantin will hide behind the asteroid, to help shield the crew from deadly solar flares. Only the solar photovoltaic array and the communication antenna will peek over the asteroid. Rhyugu does not have enough gravity to hold the Konstantin in a proper orbit so the ship is forced to use reaction control thrusters in "bang-bang" mode.

HABITAT MODULE

Habitat module are pretty close to standard TransHab modules. On the Konstantin, there are two hab modules of 38 metric tons each, both of which house four crew. Modules are 11 meters in diameter, 8.5 meter tall, and are divided into two levels by static-free decking (plus an upper "attic" and a lower "bilge"). The hab walls were fashioned from a 50-centimeter-thick laminate of Kevlar, Nextel, Nomex, Viton, polyethylene foam, and insulation, designed to withstand the impact of micrometeorites. The upper level contained the ship’s galley, medical bay, and living area, while the lower level held four crew workstations, a shower, a bathroom, and life support systems. The aluminum-walled water-lined core contained the crew quarters and doubles as the anti-radiation storm cellar. The storm cellar walls contain 4,000 liters of water as radiation shielding.

The workstations are for among other things the ship's controls. They are basically chairs, the controls are created inside virtual-reality goggles worn by the crew. This saves mass, avoiding the necessity to boost tons of display screens, dials, buttons, wiring, and switches into orbit—not to mention replacement parts.

MINING ROBOTS

The mining robots are APIS™ and Honey Bee™ designs created by the real-world TransAstra Corporation, and used in the novel by permission. APIS stands for Asteroid Provided In-situ Supplies. It is a pun on "apis" which is the genus that includes the species honeybee (because like bees Apis efficiently gathers and returns useful resources and then utilizes those resources to perform useful work). A PDF report with more detail about the Honey Bees is available here.

The problem is that asteroids such as Ryugu are more like a pile of gravel in free fall than it is a solid lump of rock. So if you used jack-hammers or explosives you'd just scatter the gravel all over the solar system. Then comes the problem of refining the blasted stuff.

Honey Bees use optical mining, bagging rocks in large sacks and using beams of concentrated sunlight to spall the rock into tiny pieces. Ryugu's orbit has a perihelion of 0.9633 AU and an aphelion of 1.4159 AU. So the solar flux varies from 1.08 to 0.5 that of Terra, or from 1.48 to 0.683 kilowatts per square meter. Each Honey Bee has a pair of 7.5 m radius parabolic mirrors. Therefore each can gather up to 382 to 242 kilowatts of sunlight to spall the ore.

The ship carries drone robots that look like three-legged spiders, with the legs being 18 meters long. These are based on technology created for NASA’s canceled Asteroid Redirect Mission. They walk around on Ryugu's surface looking for likely bolders up to 10 meters in diameter possessing surface spectra of useful elements. When one is found, the drone stradles the bolder with its legs, drills into it with small mandible-like arms, then pushes off Ryugu. Once it has the bolder off Ryugu's surface, it leaves it there and returns to Ryugu to prospect some more.

A Honey Bee then rendezvous with the bolder. It opens up a large bag and wraps it around the bolder. It then focuses its twin parabolic mirrors on the Sun, and directs the beam on the bagged bolder. Frozen volatile ices are vaporized. The vapors are prevented from escaping by the bag. A cold trap condenses the volatiles in a side pouch. The process of vaporizing embedded ice spalls the surface of the bolder into tiny gravel. Eventually the entire bolder has been reduced to a form suitable for the ship's refinery to work on. All the volatile ices have been extracted, but the volatiles that are chemicaly bound in minerals (the "hydrates") are still there.

The ore is taken out of the Honey Bee's bag and put into the spacecraft refinery's reaction chamber. First the refinery heats the gravel to 500°C, causing the hydrates to release their volatiles. This is mostly water vapor, but also includes ammonia, carbon monoxide, nitrogen, methane, hydrogen cyanide, and hydrogen sulfide. These are all valuable chemicals, especially the water. The reaction chamber is spun like a washing machine to condense the volatiles into a liquid and to get them out of the chamber. This liquid is piped into storage for later purification. There hydrogen peroxide and postassium permanganate is used to oxidize out the non-water chemicals, leaving pure water. The water can be used as is, or cracked into hydrogen and oxygen. Hydrogen peroxide and postassium permanganate reagents can be synthesised from asteroid elements.

By this stage all the volatiles in ices and volatiles in hydrates have been extracted. There remains some more stubborn volatiles.

The remaining rock in the reaction chamber now undergoes "benefication", which means concentrating the valuable stuff by throwing away the worthless stuff. This process also extracts the more stubborn valueable volatiles. The first step is to pressurize the reaction chamber with pure hydrogen and heated to 800°C. Stubborn volatiles such as hydrogen, carbon dioxide, sulfur, nitrogen, hydrocarbons, chlorine, sulfuric acid, hydrochloric acid, and assorted amino acids are extracted. These are also spun out from the chamber and stored for later filtration.

At this point the ore in the reaction chamber has had its mass reduced by about half because all the volatiles have been extracted. What remains is the involatile residue. About a third of the residue is valuable iron-nickel-cobalt alloy. This is now purified by an acid leach, which removes everything except the alloy and the silicates. Acid is injected into the chamber, causing a furious chemical reaction. The chamber is spun again to remove the acid and the impurities: phosphorus, sodium, potassium calcium, and magnesium. They are called "impurities" but all are useful elements, they are captured and stored. The acids used in the leach can be synthesised from elements obtained earlier in the process so this is sustainable.

The remainder is iron-nickel-cobalt alloy plus silicates. On Terra this would be purified by smelting in a furnace, but that's a very bad idea in a spacecraft. So an alternate method is used. The idea is to separate the iron-nickel-cobalt alloy into its component elements by successive gasification. The chamber is heated to 100°C, carbon monoxide is injected, and the pressure raised to two atmospheres. Nickel reacts first combining with four carbon monoxide molecules to form nickel tetracarbonyl gas (very toxic). This is removed by spinning the chamber and stored; leaving the iron, cobalt, and silicates. Next the pressure is increased. Now the iron reacts with five carbon monoxide molecules to become iron pentacarbonyl gas (also toxic). It too is spun out of the chamber and stored; leaving the cobalt and the silicates. Finally the chamber temperature is increased to 200°C and the pressure increased to 10 atmospheres. Now the cobalt combines with eight carbon monoxide molecules to form dicobalt octacarbonyl gas (yes, this is toxic too). It is also spun out and stored. At room temperature and pressure the stored nickel and iron carbonyls condense into liquids, which makes them easier to handle (easier than handling red-hot gas at any rate). The stored cobalt carbonyl becomes a powder, but can be turned back into a gas by heating it to a modest 52°C.

The bolder has now been rendered down into valuable water, iron, nickel, and colbalt; plus other assorted chemicals.

What's left is silicate residue, basically sand. The worthless stuff. It currently isn't useful for anything other than bulk radiation shielding. But in the future it may be used to synthesize silicon or glass. And later still there may be a method invented that can economically extract the tiny amount of platinum group metals.

A much more high-tech method of rendering down an asteroidal bolder is vaporizing the thing in a fusion torch and running the resulting plasma through a mass spectrometer (a disassembler). That will sort out all the component atoms nice and neat into separate bins. A pity it requires use of nuclear fusion, ultra-high electromagnetic fields, and other technologies that will take centuries to master. And even longer to make the equipment cheap enough so that this will be an economically sensible method to extract those elements.

EVA

Above the central hub is a cluster of four docking ports for the four Mule utility craft. The port cluster is located far enough above the central hub so that the 40° thrust cone from a Mule will not spray the centrufuge arms. Otherwise the exhaust could possibly damage the arms. Immediately above the docking ports are the four suit-ports for the space suits. In between the suits are stored the Valkyrie telepresence drones.

The space suits are entered through suit ports. This means the spacesuits never enter the habitable part of the ship, they are always outside of the spacecraft. The crew gets into the suit by opening the back pack like a fat airlock door.

Why? Because the asteroid regolith is an extreme biohazard. Asteroid regolith is five times finer than talcum powder, and under the microscope looks like tiny razor blades. In the lungs they can enter the blood stream, which will kill you. The regolith which stays in the lungs can cause silicosis, which will kill you. The regolith can abrade gaskets, jam valves, and short-out circuit boards, which can also kill you. Don't let that deadly crap get inside the ship. Since you cannot work in a spacesuit near an asteroid without being freaking covered with death-dust, you can't let your spacesuit inside either. That's why you use suit ports.

RocketCat sez

This thing looks really stupid, but it could be the key to opening up the entire freaking solar system. Orbital propellant depots will make space travel affordable, and these water Mosquitos are just the thing to keep the depots topped off.

Kuck Mosquitoes were invented by David Kuck. They are robot mining/tanker vehicles designed to mine valuable water from icy dormant comets or D-type asteroids and deliver it to an orbital propellant depot.

They arrive at the target body and use thermal lances to anchor themselves. They drill through the rocky outer layer, inject steam to melt the ice, and suck out the water. The drill can cope with rocky layers of 20 meters or less of thickness.

When the 1,000 cubic meter collection bag is full, some of the water is electrolyzed into hydrogen and oxygen fuel for the rocket engine (in an ideal world the bag would only have to be 350 cubic meters, but the water is going to have lots of mud, cuttings, and other non-water debris).

The 5,600 m/s delta-V is enough to travel between the surface of Deimos and LEO in 270 days, either way. 250 metric tons of H₂-O₂ fuel, 100 metric tons of water payload, about 0.3 metric tons of drills and pumping equipment, and an unknown amount of mass for the chemical motor and power source (probably solar cells or an RTG).

100 metric tons of water in LEO is like money in the bank. Water is one of the most useful substance in space. And even though it is coming 227,000,000 kilometers from Deimo instead of 160 kilometers from Terra, it is a heck of a lot cheaper.

Naturally pressuring the interior of an asteroid with live steam runs the risk of catastrophic fracture or explosion, but that's why this is being done by a robot instead of by human beings.

In the first image, ignore the "40 tonne water bag" label. That image is from a wargame where 40 metric tons was the arbitrary modular tank size.

There are more details here.

The name of the spaceship is LOX-augmented Nuclear Thermal Rocket Lunar Transfer Vehicle, mercifully abbreviated to LANTR LTV.

This is from Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies: A Preliminary Assessment (1995). Stanley Borowski of NASA Lewis Research Center was looking for a way to economically send manned expeditions to Luna, that is, with something cheaper than a non-reusable Saturn V rocket.

First off, he outlined the parameters for a true reusuable single-stage-to-orbit booster, instead of that Rube Goldberg Space Shuttle contraption.

Secondly, a more powerful engine that puny chemical rockets was indicated for the spacecraft. A NERVA-like solid core nuclear thermal rocket would be nice. Unfortunately while their specific impulse was a vast improvement over chemical engines, scaling up the blasted things from the putt-putt NERVA prototype so they had halfway decent thrust levels was a problem. It was a pity, since it only needed high thrust at certain parts of the mission. For the rest of the mission it could get by with already achieved levels of thrust. It's too bad there wasn't any way to make the engine shift gears... waitaminute!

There is a way to shift gears, the old LANTR trick! Just inject some supersonic oxygen into the exhaust nozzle like an afterburner and you could increase the thrust by up to 440%. That is good enough, and sure is easier than designing a monstrously huge reactor. Of course this degrades the specific impulse by a drastic 45% but you can't get something for nothing. You only need the afterburner for small parts of the mission, the rest of the time you can have normal specific impulse.

As it turns out, while Borowski didn't actually invent the LANTR concept, he helped develop it and has promoted it for lunar applications.

Thirdly there was that perennial problem of The Tyranny of the Rocket Equation. You can't have a reasonably sized payload as long as you are lugging along all your propellant. This looks like a job for In-situ Resource Utilization. That always gives the Tyranny a swift kick in the gonads.

Luna has a scarcity of hydrogen, but it has oxygen coming out of its ears. Which is just what a LANTR needs. As it turns out, there was a 1993 study looking into this, called LUNOX.

There had been earlier grandiose plans for huge lunar bases with titanic mining and refining installations to exploit Lunar resources. But LUNOX was trying to do this on the cheap, on a more modest scale. Read: on a scale that would NOT give NASA's funders in Congress an acute case of sticker-shock.

An initial lander delivers an oxygen production plant, storage tanks, and a nuclear reactor. A second lander delivers six remote-controlled tractors. Two of them are "loaders", which operators on Terra use to scoop up ilmenite-rich lunar soil and deliver it to the oxygen production plant. They figure that one plant could produce about 24 metric tons of LOX per year, which is enough for three manned missions.

Before a LUNOX site had been established, the project would have make do with ordinary NTR using no LANTR. Due to the state of the art, these would be use-once-and-throw-away spacecraft. But once LUNOX was up and running they could switch to reusable LANTRs and enjoy a much more economical trip to Luna. Using LUNOX would cut the mass to be boosted into LEO in half! And a resuable LANTR LTV can perform up to 20 missions before its nuclear fuel rods become spent.

Expendable LANTR LTV would transport as payload a Lunar landing / Earth return vehicle (LERV) with a crew module. After the LANTR LTV was disposed of, the LERV would land the astronauts on Luna, and after the mission was over transport the crew back into Lunar orbit and send them on their way back home to Terra.

Resuable LANTR LTV would assume there were a supply of LERVs at the lunar site. These would have their name changed to Lunar Landing Vehicles (LLV) since they would not be used as Earth Return Vehicles. Instead, the LANTR LTV would just transport the naked crew modules and surface payload modules. The LLVs would ferry the crew and surface payload from orbit down to the surface, and later ferry crew modules and a tank of refueling LUNOX back up to the orbiting LANTR LTV. One of the surface payload module types would be liquid hydrogen from Terra. Remember that Luna has vast supplies of oxygen, but precious little hydrogen. The LLVs need the stuff.

The expendable LANTR LTV would operate their engines at an oxygen-to-fuel (O/H) mixture ratio (MR) of 4.0, while the resuables would use an MR of 6.0

Design Bravo and Design Delta have almost identical designs (Delta entries that say "BRAVO" are identical to Bravo entries). Design Charlie is one design with two columns, for two different payload mixes (2nd column Charlie entries that say "CHARLIE" are identical to 1st column Charlie entries).

ΔV LUNOX means delta-V without using LANTR afterburner, and not taking into account refueling at Luna with LUNOX lunar oxygen. MR ΔV LUNOX means delta-V with LANTR, but still not taking into account LUNOX.

DESIGN ALFA (expendable)

This is a pathetic Lunar Transfer Vehicle using only a putt-putt NERVA engine with no LANTR afterburner. It is presented for comparison purposes, so the LANTR designs can point at it and laugh. It was designed to have its components boosted into orbit by a conventional Space Shuttle or a Titan IV rocket.

Alfa has an inadequate delta V of 3,6000 m/s. This means it does not have enough ΔV to do a Lunar Orbit Insertion burn. Instead, unlike the other designs, the poor LERV has to separate and do the burn itself. The extra propellant required really cuts into the LERV payload mass. Alfa has a higher listed payload mass than the other designs, but more of it is LERV fuel and less of it is LERV hardware and payload.

DESIGN BRAVO (expendable)

This was designed to be boosted into orbit by a hypothetical new single-stage-to-orbit (SSTO) rocket with a 9.2 m (30 foot) cargo bay. So the design is split into three 9.2 m long parts (9 m propulsion module, 9 m propellant module, and the LERV).

Unfortunately the only way to make everything fit was to put the propellant module liquid oxygen tank inside the liquid hydrogen tank. This is a bad idea. In September of of 2016 a SpaceX Falcon 9 rocket blew up on the launch pad because a liquid helium tank immersed in a liquid oxygen tank froze the oxygen into solid oxygen. Design Bravo has extra insulation around the oxygen tank (incidentally cutting into the payload mass) but it is still a matter of concern.

Also in a desperate attempt to make everything fit into the booster, they had to use a methane (CH₄) fueled LERV instead of the more efficient hydrogen (LH₂) fueled LERV (methane tanks are smaller because methane is more dense).

DESIGN DELTA (expendable)

This is basically Design Charlie with the assumption that NASA can get Congress to approve funding for a more spacious SSTO booster with a 13.7 m (45 foot) cargo bay. This allows relocating the liquid oxygen tank outside of the liquid hydrogen tank, so it is no longer a ticking time bomb. It also allows using the more efficient LH₂ LERV.

DESIGN CHARLIE (reusable)

There are two columns for Design Charlie, but they are the same spacecraft with two different payloads. It assumes that there is a supply of reusable LLVs on site at the LUNOX base to ferry crew and cargo back and forth (delivered by prior expendable missions). So Charlie just carries the naked payloads, it does not lug along the LLVs as well.

The "crew delivery" payload has 6.8 metric tons of crew and crew module, plus 2 metric tons of surface supplies for the LUNOX base. The "cargo delivery" payload is just 12 metric tons of surface supplies. Among the surface supplies are liquid hydrogen fuel for the LLVs and other equipment at LUNOX base. Remember the Lunar soil is jam-packed with oxygen but hydrogen is very hard to come by.

The design shown is based on Design Bravo, with the oxygen tank ticking time bomb.

Since the LLVs at LUNOX base can also refuel Charlie with oxygen, Charlie is reusable. It can perform 20 missions before the nuclear fuel runs out. The other designs are disposed of, criminally wasting 95% of their costly nuclear fuel rods. Actually even Charlie is wasting 85% of its fuel rods, but NASA figures attempting to remove the rods for reprocessing is just begging for a nuclear disaster in space.

Charlie just runs its engines at 6.0 MR for the entire mission, to stay within the LH₂ and LOX propellant limits.

This is from Technology Requirements For Future Earth-To-Geosynchronous Orbit Transport Systems (1979). The unmanned Large Cargo Orbital Transfer Vehicle (LCOTV) transports cargo from Low Earth Orbit to Geosynchronous Orbit. The report also described a chemically powered LEO to GEO transport for priority cargo.

The report was trying to put a price tag on an transport system capable of handling the construction of a large solar power station.

According to a later report, for various reasons, the report concluded the project would require a fleet of 13 LCOTVs. Not all of the fleet would be on line, some would be undergoing maintenance. The ships in the fleet would be allocated such that the SPS project would receive a total of 56 flights, transport a total payload of 29,860 metric tons, make from 1 to 13 flights a year, and transport from 33 to 2010 metric tons per year.

Two designs were considered. The "Normal Growth" design used conservative extrapolations of the state-of-the-art, the "Accelerated Technology" assumed additional money was invested to increase the state of the art.

This is from An Integrated Hybrid Transportation Architecture for Human Mars Expeditions (2015)

Everybody in rocketry is waiting for somebody to invent a torchship.

For a typical Mars mission, designers want an engine with high thrust so the Trans-Mars insertion burn takes only a few minutes. Instead of taking a year, with most of that year spent inside the Van Allen radiation belt irradiating the hapless crew with atomic death-rays.

But at the same time designers want an engine with a high specific impulse, otherwise the design will get eaten alive by the The Tyranny of the Rocket Equation. The result will be a spacecraft looking like a skyscraper composed of fuel with a wretchedly tiny Mars expedition perched on the top.

Chemical engines have fantastic thrust and woefully minuscule specific impulse ("muscle" class). Electric propulsion like ion drives are the exact opposite ("fuel-economy" class). A torch drive would be high in both, but currently there ain't no such animal.

Most Mars missions using chemical propulsion are forced to rely upon staging to cope with the huge fuel requirements. This turns the spacecraft into a disintegrating totem pole, throwing away fuel tanks and engines. Oh, and did I mention that designers wanted the spacecraft to be totally reusable? Staging is almost criminal if you are using nuclear engines, but there are plenty of Mars missions that do just that. Nothing like jettisoning strongly radioactive fission reactors still full of nuclear fuel into eccentric solar orbits where they will be a radiation hazard for generations to come. Tends to earn harsh words from the future generations, it does.

So lacking a torch drive, how does a designer get out of this mess? Presently the best idea is to make a hybrid spacecraft, with two types of engine. The unstoppable Stanley Borowski has a design using solid core nuclear thermal rockets along with an ion drive, using the clever Bi-Modal NTR trick. The Langley design manages to make a chemical-ion hybrid.

The Langley design dramatically lowers the fuel requirements (which is the same thing as raising the payload size), allows the spacecraft to be reusable, and manages to get away with using a Solar Electric Propulsion (SEP) engine with a remarkably low power requirement of only 300 kilowatts. The latter is important because 300 kW can be easily suppled by a low-mass solar array, a megawatt class electric drive would need a nuclear reactor and those things weigh tons.

The chemical engines take care of just the high-thrust burns, and the fuel-economy ion drive performs the burns where thrust is not quite so urgent. If you arrange the trajectory with sufficient cleverness, the fuel mass savings will more than take care of the mass penalty of having two separate rocket engines. And also have no significant increase in the flight times, even if you are using a low-thrust ion drive.

The spacecraft does require a cis-lunar Basecamp in order to reduce the orbital energy, but you can't have everything.

Bottom line: larger mission payload and a totally reusable spacecraft.

To size the spacecraft the designers used software called Glenn Research Center Collaborative Modeling for Parametric Assessment of Space Systems (GRC COMPASS).

One of the design parameters was the ability to do Mars missions with [1] 300 day (10 month) or longer Mars orbit dwell time and [2] less than 1,100 day (3 year) round trip heliocentric mission duration.

The cargo and crew are launched to the basecamp. The crew uses a fast "taxi" transfer which takes eight to ten days, 300 to 400 m/s ΔV, and a lunar gravity assist (LGA). The cargo uses a slow ballistic lunar transfer which takes 120 to 200 days and 0 to 40 m/s ΔV (to maximize cargo delivery capability).

The heart of the system is the Hybrid Propulsion Stage (HPS).

The ion drive uses components designs that were created for the Asteroid Redirect Mission (ARM). The ARM engine buss uses 150 kW of electrical power for 12 Hall Effect ion engines utilizing xenon propellant. The engines can operate in high specific impulse (3000 sec) or low specific impulse (2000 sec) modes, shifting gears.

For the Hybrid, two busses are to be used, with a power requirement of about 318 kW_e, 24 engines, and double the thrust.

The ion engines are 12.5 kW HERMeS Hall Thruster, as described in "Development Approach and Status of the 12.5 kW HERMeS Hall Thruster for the Solar Electric Propulsion Technology Demonstration Mission" IEPC-2015-186. HERMeS is an acronym for Hall Effect Rocket with Magnetic Shielding.

Option 1 uses the ARM engines and the ARM structure. Option 2 uses the ARM engines but a new structure is designed so that the HPS can fit into the SLS narrower 8 meter diameter shroud instead of the larger 10 m shroud. Out of an abundance of caution Option 1 was chosen.

Both options use the same solar arrays which provides 435 kW_e at Terra's distance from Sol, but only 184 kW_e at Mars. The ion engines are not used near Mars so the fact that the solar array does not provide enough power for the ion engines is not a problem. Remember the engines use 318 kW_e

Both options use chemical engines: array of x10 Aerojet R-42 engines using storable MMH/NTO fuel with oxidizer-to-fuel ratio of 1.65, nominal specific impulse of 303 seconds and total thrust of 890 Newtons.

This is from Mission performance potential of regeneratively cooled gas core nuclear rockets (1971)

The report notes that while solid-core nuclear thermal rockets have twice the specific impulse of chemical rockets, this isn't enough of an increase for high-energy trajectories with very high payloads. On the other end of the spectrum, ion drives have superb specfic impulse, but their pathetic thrust lead to undesirably long mission times.

The report decided to stop playing around and look at a rocket engine that is more in the middle between solid NTR and ion drives in both specific impulse and thrust: open-cycle gas-core nuclear thermal rocket. Gas-core has about four to twelve times as much specific impulse as solid-core, about sixty times as much thrust, and a hideously deadly radioactive exhaust plume. About as radioactive as if a solid-core rocket had a total nuclear meltdown at a rate of one solid-core rocket per second.

The report ignores the radiation, figuring it is not their department to worry about it. They wanted to analyze the performance of the gas-core NTR to find the circumstances where the performance was so superior that it warranted a more detailed study (including how to deal with the lethal exhaust).

They didn't bother analyzing the safer closed-cycle NTR because it was a typical compromise that wound up with the disadvantages of both and the advantages of neither. For one thing the specific impulse was half that of an open-cycle engine.

Spoiler Alert: the open-cycle gas-core nuclear rocket is so superior to the solid-core that it isn't even a contest. The real eye-popping gains show up at 3,000 seconds of specific impulse, but even 1,500 seconds is impressive. It is worth it to develop the GCNR further.

Their baseline was a pretty standard regeneratively-cooled open-cycle gas-core engine. They assume a specific impulse in the range of 1,000 to 3,000 seconds (exhaust velocity of about 10,000 to 30,000 m/s) which is about seven times that of chemical rockets and about four times that of solid-core NTRs.

But unlike prior reports, they looked into optimizing the thrust levels for a given mission.

Engine Mass

The mass of the gas-core NTR engine is obviously the sum of its parts:

M_e = M_mod + M_ps + M_tp + M_n

M_e: mass of engine (kg)
M_mod: mass of nuclear moderator (kg)
M_ps: mass of pressure shell (kg)
M_tp: mass of turbopump (kg)
M_n: mass of nozzle (kg)

They assumed the following relationships:

M_H2: hydrogen propellant flow rate (kg/sec) Note the "M" should have a dot over it (because m-dot means mass flow). Unfortunately many computers do not have unicode fonts capable of rendering the character "Ṁ".
F: thrust (N)
I_sp: specific impulse (sec)
g: standard value of gravity acceleration (m/sec²) = 9.80665 m/sec²
I_sp × g: exhaust velocity (m/s)

M_U = M_H2 × ( 1 / H₂/U)

M_U: uranium fuel flow rate (kg/sec)
M_H2: hydrogen propellant flow rate (kg/sec)
H₂/U: hydrogen-to-uranium flow-rate ratio = 100
1 / H₂/U: reciprocal hydrogen-to-uranium flow-rate ratio = 0.01

V_U: volume of uranium in core (m³)
V_c: volume of core (m³)
M_H2: hydrogen propellant flow rate (kg/sec)
M_U: uranium fuel flow rate (kg/sec)

Moderator Mass

Moderator is assumed to be 0.762 meters thick.

M_mod: mass of nuclear moderator (kg)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
ρ_mod: density of moderator material (kg/m³) = 1,150 kg/m³

Pressure Shell Mass

P: pressure (atm)
M_cr: critical mass in reactor (kg) = 48 kg
F: thrust (N)
I_sp: specific impulse (sec)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
M_U: uranium fuel flow rate (kg/sec)
M_H2: hydrogen propellant flow rate (kg/sec)

t: thickness of pressure shell (m)
P: pressure (atm)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
σ: allowable stress in pressure shell (atm) = 13,600 atm

M_ps: mass of pressure shell (kg)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
t: thickness of pressure shell (m)
ρ_ps: density of pressure-shell material (kg/m³) = 8,000 kg/m³

Turbopump Mass

M_tp: mass of turbopump (kg)
M_H2: hydrogen propellant flow rate (kg/sec)
P: pressure (atm)
ρ_H2: density of hydrogen (kg/m³) = 72 kg/m³

Nozzle Mass

M_n: mass of nozzle (kg)
ε: area ratio of nozzle = 300
F: thrust (N)
P: pressure (atm)

Hydrogen Propellant Temperature

This does not help calculate the mass of the engine, but it is needed with other parts of the design.

T: propellant temperature (°C? K?)
P: pressure (atm)
F: thrust (N)
I_sp: specific impulse (sec)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
M_U: uranium fuel flow rate (kg/sec)
M_H2: hydrogen propellant flow rate (kg/sec)

Constants

Note that the thrust level has been optimized for each mission.

D: outer diameter of the core/moderator = for this study 3.66 m
M_cr: critical mass in reactor = 48 kg
H₂/U: hydrogen-to-uranium flow-rate ratio = 100
σ: allowable stress in pressure shell (atm) = 13,600 atm
ρ_ps: density of pressure-shell material (kg/m³) = 8,000 kg/m³
ρ_H2: density of hydrogen (kg/m³) = 72 kg/m³
ε: area ratio of nozzle = 300
ρ_mod: density of moderator material (kg/m³) = 1,150 kg/m³

Mission delta-Vs

The report took the simplistic ideal mission delta-Vs and made a function to account for gravity losses, since in the real world rocket impulse burns are not instantaneous (indeed, with ion drive a burn can take weeks). The function figured in acceleration levels, parking orbit eccentricity, and final hyperbolic excess velocity.

This has to be done iteratively, since once you calculate the real delta-V, you optimize the thrust level to the real delta-V, which means you have to recalculate the real delta-V, which means you have to reoptimize the thrust level… You keep iterating until the function converges on a value.

(M_p)_i: propellant mass of i^th maneuver (kg)
(M₀)_i: mass at beginning of i^th maneuver (kg)
e^x: antilog base e or inverse of natural logarithm of x
ΔV: delta-V (km/sec)
I_sp: specific impulse (sec)
g: standard value of gravity acceleration (m/sec²) = 9.80665 m/sec²
I_spg: exhaust velocity (m/s)

(M₀)_i+1: mass at beginning of next maneuver after i^th maneuver (kg)
(M₀)_i: mass at beginning of i^th maneuver (kg)
(M_p)_i: propellant mass of i^th maneuver (kg)
(M_pstr)_i: propellant-structure mass of i^th maneuver (kg)
M_jettison: mass jettisoned, such as Mars Lander (kg)
(M_is)_i: interstage structure mass of i^th maneuver (kg)

(M₀)_imax: mass at beginning of last maneuver (kg)
(M_p)_imax: propellant mass of last maneuver (kg)
M_pstr: propellant-structure mass (kg)
M_ts: thrust-structure mass (kg)
M_pay: payload mass (kg)
M_e: engine mass (kg)

where:

M_ts: thrust-structure mass (kg)
F: thrust (N)

where:

(M_pstr)_i: propellant-structure mass of i^th maneuver (kg)
(M_p)_i: propellant mass of i^th maneuver (kg)

where:

(M_is)_i: interstage structure mass of i^th maneuver (kg)
(M₀)_i: mass at beginning of i^th maneuver (kg)

where:

M_e: engine mass OF A SOLID-CORE NUCLEAR ROCKET as an approximation. For GCNR use this equation. (kg)
F: thrust (N)

For Mars and Jupiter missions M_jettison = 136,100 kg

Uncrewed Lunar Ferry Mission

In the uncrewed lunar ferry mission the GCNR starts parked in LEO, does a Hohmann transfer to deliver various amounts of payload (M_jettison) to be placed into a lunar orbit. It still has 50,000 kg of Terra return payload. It then does a Hohmann back into LEO parking orbit. After refueling it is ready to deliver another payload.

The graph in figure 5 compares the performance of the gas core rocket with a 825 sec Isp solid-core Nerva-II, and a theoretical maximum 825 Isp Nerva-II (meaning it has zero engine mass). A specific impulse of 2,000 seconds was chosen for the GCNR. The GCNR leaves both NERVAs in the dust. It needs much less initial mass in LEO for a given payload. Actualy for the 50,000 kg payload to Luna and 50K kg back (blue line in graph) the GCNR can carry close to twice the payload (red line in graph).

The graph in figure 6 shows fuel/propellant consumption, where the GCNR shames the NERVAs even more. With a 50,000 kg payload the GCNR uses about half the the propellant needed by the NERVAs.

Figure 7's graph shows, among other things, that raising the specific impulse from 2,000 seconds to 3,000 seconds only reduces the initial mass in LEO (IMLEO) requirement by 11%. Simply because the propellant requirement goes down with higher specific impulse.

Figure 8a shows how the moderator mass varies with the core diameter. You probably should use the "best guess" curve for required critical mass.

Figure 8b shows that changing the hydrogen-propellant to uranium-fuel ratio (H₂/U) doesn't change the IMLEO very much. The report uses H₂/U = 100. The IMLEO goes up as H₂/U rises because the pressure shell weight increases (see equation), of course as H₂/U rises the amount of expensive uranium fuel needed goes down.

Uncrewed Slingshot Mission

In the uncrewed slingshot mission the GCNR starts parked in LEO, the spacecraft accelerates into some amount of hyperbolic excess velocity (V_∞), releases a payload of mass M_jettison, waits until it comes back to Terra, and then burns like heck to circularize into LEO. It still has 50,000 kg of Terra return payload. It can be reused after being refueled.

Figure 9 shows how various factors affect the IMLEO. Mass increases with increase in amount of V_∞, mass decreases with increase of specific impulse. At V_∞=0 if you double the GCNR I_sp from 1,500 to 3,000 seconds you will save 20% on IMLEO. At V_∞=5 km/s doubling I_sp will save 30% on IMLEO.

The GCNR can point its finger and laugh at the optimistic solid-core NTR. SCRN cannot reach V_∞=5.5 km/s at all, not without staging at any rate. Which coincidentally is the delta-V requirement for a 300 day Mars round trip. A single-stage 1,500 I_sp GCNR can manage it. A 3,000 I_sp GCNR can do that without even working up a sweat.

Crewed Interplanetary Missions

For crewed interplanetary missions, it assumed that 136,100 kg (M_jettison) of payload is delivered into a 0.9-eccentricity planetary parking ellipse with a periapsis at 1.1 planet radii. An additional 90,700 kg is the command module / habitat module plus the Terra reentry vehicle. The GCNR does not brake into orbit to be reused, it goes streaking by Terra into a far Solar orbit. The crew bails out in the reentry vehicle that aerobrakes to the surface. The analysis did not set a limit on entry velocity, though in reality if it is much over 11 km/sec it is difficult to make a reentry vehicle that won't melt into a fiery blob of molten impure aluminum.

Figure 10 below shows the results for the Mars round-trip mission.

The analysis simplified things by assuming circular coplanar orbits, so this is more of the average performance regardless of synodic period.

In figure 10a, if you fix the IMLEO at 10⁶ kg for all cases, 3,000 sec I_sp GCNR can do the mission in 225 days, 1,500 sec GCNR needs 360 days, and the weakling 825 sec SCNR needs a whopping 430 days.

If you fix the trip time at 200 days, the 3,000 sec GCNR needs only one-fifth the IMLEO of the 1,500 sec GCNR. The IMLEOs are all higher than 10⁶ kg, but it might be worth it if you are trying to spare the crew from excessive radiation doses.

Figure 10b shows among other things that the optimum thrust is in the range of 400,00 to 5 million Newtons (4×10⁵ to 50×10⁵ N).

Figure 11 below shows the results for the Jupiter mission.

The possible missions are the circles, ignore the lines connecting them (those are just to label circles with their specific impulse). A mission can be performed in 580 days, or 1,020 days or 1,420 days but not at any other intermediate time values. The circle values are when Jupiter and Terra are in opposition or conjuction at mid-say. At other times the delta-V becomes impossibly high. If you really want the scary mathematical details they can be found in Approximate trajectory data for missions to the major planets.

Because of the circle limits, for the Jupiter mission raising the specific impulse can only reduce the IMLEO. For the Martian mission raising the specific impulse can both reduce the IMLEO and the trip time. But the GCNR can perform the mission faster than the SCNR because the latter is too weak to have access to the columns of circles with lower trip times. It can only do the 1,420 day trip.

In the 1,420 day column, the 3,000 sec I_sp GCNR requires only 15% the IMLEO of the SCNR, and 60% of the 1,500 sec GCNR.

Figure 11b shows among other things that the optimum thrust is in the range of 10⁶ to 2×10⁷ Newtons, about twice that of the Mars mission.

This is from a 1965 study by the Lewis Research Center entitled Space Flight Beyond The Moon.

This is a standard ion rocket powered by a nuclear reactor. The reactor is at the nose, behind a shadow shield and separated from the crew compartment by a 120 meter long boom in order to use distance as extra radiation shielding. The heat radiators are trimmed in order to not extend outside of the radiation shadow, their narrow aspect indicates the shadow shield is minimal in order to save on mass.

Higher ion drive exhaust velocities come at the expense of higher power requirements, which means higher power plant mass, which means lower acceleration. The report mentions that one can calculate the optimal exhaust velocity/power plant mass, but does not go into details. As a general rule a rocket's acceleration should not be below 0.05 m/s² (5 milligee) or it will take years to change orbits.

Some of the designs depict the crew cabin as two modules, which probably means they spin around the ship's spine for artificial gravity.

MOECKEL DESIGN

(ed note: Notice that this is for a 1,100 day mission, not 500 days like the Lewis design. That is why this has 39,600 kg for crew provisions, instead of only 18,000 kg.)

ELECTRICAL POWER GENERATION

     Weight estimates made by members of the NASA Lewis Library staff for several electric power generation systems are shown in Fig. 4. Along the abscissa are shown the electirc power levels appropriate for various space propulsion missions. At the higher power levels required for manned interplanetary expeditions, only three of the many possible generating methods were found to be competitive on a specific weight basis—the nuclear-fission turboelectric, solar turboelectric, and the thermonuclear or fusion electric systems. The nuclear turboelectric system weights were based on use of a sodium-vapor cycle, with turbine inlet temperature of 2500° R and radiator temperature of 1800° R. The cycle efficiency for these temperatures was 20%. The design was based on minimization of radiator size without exceeding reasonable temperatures in other parts of the system. Even with this minimization, the radiator was the heaviest part of the generating system except at low power levels, where shielding weight became dominant.

     The solar turboelectric system weights were based on the use of half- silvered polyethylene balloons as collector of solar radiation (as suggested by Krafit Ehricke). These balloons replace the reactor and shielding of the nuclear turboelectric system.

     Solar batteries were the lightest of the low-power systems, but were not competitive at the power levels of interest for high-payload space missions.

     The solar turboelectric system was found to be very close in weight to the unshielded fission turboelectric system at all power levels. The straight line which represents both systems at high power levels corres- ponds to a specific powerplant weight of about 5.5 lb per electric kilowatt. If the electric power could be converted to jet power with 100%, efficiency and no accelerator weight, the specific powerplant weight at for electric propulsion systems would then be 5.5 lb per jet kilowatt. Allowing for some inefiiciency in conversion, and some accelerator weight, it appears that a specific powerplant weight between 5 and 10 lb per jet kilowatt should be attainable. These values, together with the curves of Fig. 2, indicate that electric propulsion systems should operate with initial acceleration between 10^-4 and 2×10^-4 g at a specific impulse near 10,000 sec to accomplish a round-trip satellite-to-satellite Mars mission with a payload ratio near 0.5.

INITIAL WEIGHT COMPARISION FOR MARS JOURNEY

     To illustrate the magnitude of weight savings that are possible with some of the nonchemical systems described in the preceding section, the total initial weights that must be placed in orbit near the Earth to undertake a rather elaborate Mars expedition are compared in Fig. 8. This comparison is for an eight-man expedition with landing and exploration equipment. The basic payload, which includes crew, cabin, navigation and communication equipment, environment control, etc., was assumed to be 50,000 lb. (22,700 kg) Exploration equipment, including glide landing vehicle and take-off rocket for six men and supplies, was estimated to be about 60,000 lb (27,200 kg). Food, water, and oxygen allowance of 10 lb (4.5 kg) per man-day is included. The total-initial payload, which is basic payload, exploration equipment, and survival supplies, is, therefore, about 200,000 lb. (91,000 kg). With a four-stage chemical rocket having specific impulse of 300 sec and a structure and motor weight of 0.05 per stage, the total-initial weight required in orbit is about 6.5 million pounds. For the advanced chemical rocket, with specific impulse of 420 sec, the initial weight is reduced to about 2,000,000 lb (907,000 kg). A nuclear heat-transfer rocket with specific impulse of 800 sec, a thrust-weight ratio of 1.0, a tankage weight equal to 8% of the propellant weight, and a powerplant weight of about 50,000 lb (23,000 kg) can accomplish the mission with an initial weight of about 800,000 lb (360,000 kg). If the powerplant weight is 20,000 lb (9,000 kg) the initial weight required is about 600,000 lb (270,000 kg).

     Some improvement in initial weight is possible, for the same powerplant weights, if a specific impulse of 1400 sec is achieved with an initial acceleration of 10^-1, but the initial weight increases if the initial acceleration with I_sp=1400 goes to 10^-2 g.

     An electric propulsion system with initial acceleration of 10^-4 g and a specific powerplant weight of 10 lb/kW (4.5 kg/kW) can do the job with about 400,000 lb (180,000 kg) initial weight.

     The nonchemical propulsion systems therefore offer the possibility of weight reductions by a factor of 10—15 over conventional chemical rockets and by a factor of 3—5 over high-energy chemical rockets. When it is remembered that each pound in orbit requires from 10 to 100 lb (4.5 to 45 kg) in launching weight, it seems clear that the development of nonchemical propulsion systems is essential for high-payload space missions.

HYPOTHETICAL SPACE VEHICLE

• 

• 

     An example of the type of space vehicle that might be feasible using a nuclear turboelectric propulsion system is shown in Fig. 9. The mission for which this vehicle was designed is similar to that used in the comparison shown in Fig. 8. The components of the nuclear-electric system are scaled in accordance with the size and weight estimate made for the curves of Fig. 4. The design is based on shadow shielding, with rigid tubular separation of the major components. The entire vehicle is to be spun about its axis to provide artificial gravity both for the crew and for the separation of liquid and vapor phase in the radiator.

     Needless to say, there are some unattractive features in this design, particularly in the magnitude of the radiator required. This size, and the need for light construction make it vulnerable to meteoroid damage. The frequency of penetration can not yet be accurately predicted, but rough estimates based on available meteoroid distribution data indicate that penetrations of the order of several per week may be expected with the thickness of tubing assumed. The weight penalty to insure against these penetrations would be prohibitive. Most penetrations, of course, will be microscopic, but they will nevertheless have to be repaired to avoid significant mass loss of the heat—transfer fluid.

     Assuming that the engineering problems associated with such a space vehicle can be solved, the propulsion system is seen to be a very good one indeed on the basis of payload capacity. The weight that must be launched into orbit is less than twice the initial payload weight. Similar weight ratios seem feasible with nuclear heat-transfer rockets. If launching rockets are developed which are capable of placing 45,000 kgs of payload into orbit, only four launchings would be required to start the expedition on its way. In terms of current technology, this is a staggering undertaking, but certainly few will deny that such projects are well within the limits of future capability.

This is from Manned Mars landing mission by means of high-thrust rockets (1966).

Yet another nuclear engine design that tries to squeeze out enough delta V so it can actually land astronauts on Mars, instead of attempting to do months of science in five minutes as they go whizzing by. Since they were designing with underpowered NERVA engines with only 850 seconds of specific impulse, they were forced into the irresponsibly cuckoo "nuclear staging" solution just like the Boeing IMIS. Except the Lewis design does not jettison five dangerously radioactive nuclear reactors like used beer cans. It jettisons seven.

And like the IMIS, it cannot afford a NERVA for braking into Terra orbit. They come in hot at full velocity, trusting to the miracle of aerobraking. The crew of seven enters a winged reentry craft and gets to dice with death, gambling that they will run out of delta V before they run out of ablative heat shield. If they lose, perhaps some kid will make a wish on the shooting star which is all that is left of their incinerated bodies.

The mission had a crew of seven. Depending upon the types of optimization, either four or all seven made excursions to the Martian surface. A minimum of three crew was needed for emergency operation of the spacecraft. Normal operation required six crew, operating in pairs. 1/3rd of the day was for spacecraft operational duty, 1/3 was for sleep, and 1/3 was off-duty (recreation, personal chores, scientific duties, and study).

The propulsion is by a series of NERVA style solid-core nuclear thermal rockets using liquid hydrogen as propellant. They have a specific impulse of 850 seconds and a thrust-to-weight ratio of 10. Well, actually T/W of 8 because the liquid hydrogen tanks need thermal protection from the engines or they will boil-off dry in no time. And T/W of 3 if you must have biological anti-radiation shields. The crew is protected from engine radiation by a combination of propellant tanks, separation distance, and command module storm cellar but no biological shields (but offhand it looks like that will not be enough shielding). Somewhat arbitrarily the study authors assumed each engine would have a minimum mass of 3,630 kg (7,260 kg with thermal and biological shields.) The initial acceleration of each engine with a full propellant tank is 0.2 gs.

There was a variant with chemical engines, but it was pretty pathetic.

The report has zillions of trajectories, each optimized for one factor or another. For mission with durations of 600 days and below, the total delta V requirement vary from 26,070 to 41,360 m/s.

COMMAND MODULE

As with most designs this is a combination of the spacecraft control room and the anti-radiation storm cellar. Because nobody wants to die horribly by manning an unshielded control room during a solar proton storm, and leaving the control room unmanned is a Really Bad Idea. During normal operations a crew of two or three occupies the command module. All seven can be contained during radiation events for short periods (1 day), four of the crew have to stand.

The command module has a mass of 4080 kg (9,000 pounds), not counting the radiation shielding. It has a volume of 12.7 cubic meters, of which 1.4 m³ are radiation sensitive operating equipment.

The radiation shielding is the chemical rocket fuel and oxidizer used by the Earth reentry vehicle. The fuel is pumped into the vehicle at the end of the mission.

LIFE SUPPORT SYSTEM

The life support system is wrapped around the command module. The food and oxygen cycles are open, but the water cycle is closed (totally recycled). It assumes each crew consumes per day 1 kg of food, 1.8 kg of water, and 0.95 kg of oxygen (which is a little skimpy on the water). Plus each crew has 9 kg of recycled water per day for washing and other utility purposes.

Cabin air leakage is assumed to be about 0.68 kg per day, and complete air changes are made at three month intervals.

The total life support requirements for 7 crewmembers per day is 1608 kg plus 260 watts of power.

LIVING MODULE

The living module has a mass of 4080 kg (9,000 pounds) not including radiation/meteor shielding and a volume of 156 cubic meters. The breathing mix pressure is 48 kPa (7 psi). To improve reliability the module is divided into two pressure independant units.

The walls hold about 29 kilograms per square meter of reentry vehicle chemical fuel as meteor and mild radiation protection. For full-blown radiation storms the crew retires to the storm cellar. The chemical fuel is held in multiple independant loops. Since the sun only heats the sun-side of the module, temperature is equalized by rotating the meteor shield or circulating the fuel. The desired level of temperature is maintained by adjusting the angle of the module with respect to Sol and by surface coatings.

The living module is connected to the Earth deceleration system by a long boom. The boom rotates fast enough to provide the living module with 0.3 g of artificial gravity (close to the Martian surface gravity of 0.376g). A smaller counter-rotating centrifuge balances the angular momentum. It provides up to 10 g's for the crew to exercise because it would be real nice for the crew to be able to walk when they return to Terra.

The scientific crew might spend their entire duty cycle (all day) in the living module. The control crew have to spend part of their duty cycle in the command module.

SOLAR POWER GENERATOR

The spacecraft requires 7.5 kilowatts of power. A solar Rankine system was selected because nuclear power reactors have too much mass and are too radioactive to be repairable. The Rankine was assumed to have an alpha of 91 kilograms per kilowatt for a mass of 680 kg. 1,360 kg since they carry along a spare.

During the mission the generator dish does its best to shield the cryogentic hydrogen propellant tanks from the burning rays of the sun. Otherwise all the hydrogen will boil off.

Solar photovolatic panels were considered, but in 1966 they were not exactly "mature" technologies.

EARTH DECELERATION SYSTEM

As previously mentioned the design cannot afford the additional NERVA engines and propellant to brake the spacecraft into LEO. So the designers took the cheap way out and used aerobraking.

The system has a mass of 24,490 kilograms, including the 16,780 kg of chemical fuel which is stored during the mission around the command and living modules as radiation shielding. It carries a payload of 1,090 kg (the "real mission payload") consisting of the crew, the Mars surface samples, and the data.

The system is a winged reentry vehicle with a thick ablative heat shield on its belly (much like NASA's Space Shuttle), a retro-rocket with the chemical fuel, and a meteor shield to protect everything during the long mission.

The winged reentry vehicle has a mass of 4,990 to 7,480 kg for atmospheric entry velocities of 7,930 to 19,810 m/s. It has 1.7 meters of unobstructed interior depth to accomodate the seated crew. Internal volume of 40 cubic meters. Leading edges swept 60°, small radius nose, and maximum attack angle of 23°. Maximum G load of 10 g's, entry corridor depth is 48 kilometers. The planned entry maneuver calls for 8 g deceleration at supercircular speeds and 4 g at subcircular speeds.

MARS LANDING VEHICLE

Each landing vehicle can carry two crew.

Since the landers use aerobraking most of their mass is the propellant needed to lift off from Mars and return to the orbiting spacecraft.

While the spacecraft is in Martian orbit, the landers separate and shed their meteor/thermal shields. The retrorocket fires to put the lander on entry trajectory, then is jettisoned so it does not obstruct the aerobraking heat shield. The shield burns its little heart out because the Martian atmosphere is like making love in an airlock exceedingly close to vacuum (1/100th of Terra's atmospheric density does not help much with braking). The landing/hover rocket extends from the top and burns at 2 gs of deceleration. It has enough fuel to hover and "translate" for only two minutes so the pilot has to pick the landing spot quickly. Hopefully the shock absorbers are up to the task of absorbing the landing impact.

The takeoff tanks are kept cool enough to avoid boil-off by foils in a vacuum jacket. Ordinarily you couldn't do this on a planet with an atmosphere, but as previously mentioned the Martian atmosphere is not that far from being a vacuum already. The fuel has a specfic impulse of 430 seconds, they suggest Diborane fuel with Oxygen difluoride oxidizer (B₂H₆ + OF₂). Well, at least they didn't do a jackass maneuver like try to use FLOX for oxidizer.

The crew does as much science as they can possibly cram into 40 days. Life support and electrical power have 40 days worth of consumable. If you need a longer mission stay time, you'll have to replace the power and life support with a system as massive as the spacecraft's.

Just before takeoff the lander is stripped of every possible gram of excess mass, because it is not carrying much in the way of extra fuel. This includes ripping off the tank insulation, detaching the landing/hover rocket, ditto the remains of the heat shield and landing system. The lander takes off on a bare minimum delta V trajectory into orbit. Two more tiny boosts allow rendezvous with the spacecraft.

The takeoff section has a mass of 1,540 kg, including two crew plus 227 kg of Mars samples and data. If one lander fails, the other can carry all four crew but no Mars samples.

EQUIPMENT LANDERS

These are two unmanned landers on a one-way trip to land exploration equipment. Each lander has a mass of 1,360 kg and carries 1,810 kg of payload. The total mass for both loaded landers is 6,340 kg. The payload includes scientific equipment, land roving vehicles and their fuel.

Reference Source

This is from Nuclear thermal rocket workshop reference system Rover/NERVA (1991). The author is Dr. Stanley K. Borowski, who helped design several realistic spacecraft in this section of the website.

The point of this paper was to present a "reference design" to use to measure other design proposals presented at the conference. There was a Mars mission reference and a Lunar mission reference, both using solid-core nuclear thermal rocket propulsion based on NERVA technology.

The standard NERVA is basically liquid hydrogen heated in a nuclear reactor then emitted through a converging/diverging exhaust nozzle to create thrust. The hydrogen is compressed to high pressures by a turbopump. It is then "preheated" by cooling the nozzle, reflector, control rods, peripheral shield, and core support structure. Finally it is injected into the reactor.

One of the problems is where to get the energy to run the turbopump. There are two solutions: the "hot-bleed" cycle and the "full flow topping" or "expander" cycle.

In Hot-bleed, about 3% of the hot hydrogen exhaust emitted from the reactor is diverted (left green arrow) to run the turbopump. It is then either used for roll control or reintroduced into the exhaust nozzle. So it is called "hot-bleed" because it is bleeding off some of the hot stuff and using it to make the turbopumps spin.

In Full Flow Topping, all the preheated hydrogen is diverted (right green arrow) to run the turbopump. Then it is injected into the reactor to create thrust. Full Flow Topping has superior specific impulse compared to Hot-bleed, I presume it is much more difficult to engineer.

The Lewis Mars Reference mission was developed by Borowski in 1991, aimed at the Mars launch opportunity in the far-future year of 2016 (heh). It examined both old-school 1972 and more modern 1991 NERVA engines.

In 1969, Werner von Braun described a Mars mission where the spacecraft had triple NERVA engines, using a 640-day opposition class mission with an 80-day stay at Mars and inbound Venus swingby. The Lewis Mars Reference mission uses a spacecraft with a single NERVA, using a of 434-day opposition class mission with a 30-day stay at Mars and an inbound Venus swingby. The Lewis mission is much easier on spacecraft stress and astronaut exposure to galactic cosmic rays.

The Lewis mission came in two options. The "all propulsive" profile uses extra propellant so at the end of the mission the spacecraft can be braked into Terra orbit for future re-use. The other option is to forgoe the extra propellant, carry seven metric tons of Earth Crew Capture Vehicle (ECCV), and at end of mission have the crew use the ECCV to do a aerobraking landing while the abandoned spacecraft sails off into an eccentric heliocentric orbit, never to be used again. Though presumably in future decades the authorities would want to capture and properly dispose of derelict spacecraft with still radioactive engines littering the solar system.

The base assumptions and ground rules for designing the mission are as follows:

2016 MARS MISSION ASSUMPTIONS/GROUND RULES

PAYLOAD OUTBOUND:

• 73.12t MARS EXCURSION MODULE (MEV)
• 34.94t MARS TRANSFER VEHICLE (MTV)
• 7.00t EARTH CREW CAPTURE VEHICLE (ECCV)

PAYLOAD RETURN:

• 34.94t MTV
• 7.00t ECCV (USED ONLY WITH "EXPENDABLE MODE")
• 0.50t MARS RETURN SAMPLES

PLANETARY PARKING ORBITS:

• 407 km CIRCULAR (EARTH DEPARTURE)
• 250 km x 1 SOL¹ (MARS ARRIVAL/DEPARTURE)
• 500 km x 24 hr² (EARTH ARRIVAL)

g-LOSSES MODELED FOR EARTH DEPARTURE ONLY

EARTH DEPARTURE PLANE CHANGE ΔV PENALTIES:

• 340 m/s (dia > 28.5°)
• 100 m/s (dia < 28.5°)

MARS APSIDAL ALIGNMENT ΔV PENALTIES:

• 560 m/s

PLANETARY TRAJECTORIES OPTIMIZED FOR "ALL PROPULSIVE" MISSION SCENARIO. FOR 2016 OPPORTUNITY, TRIP TIMES RANGE FROM 120 TO 434 DAYS

SINGLE BURN AND "3-BURN" PERIGEE DEPARTURES FROM EARTH EXAMINED

¹ 250 km x 33,852 km = 1 SOL ORBIT = 24.66 HOURS
² 500 km x 77,604 km = 24 HOUR ORBIT

PROPULSION SYSTEM/PROPELLANT/TANKAGE ASSUMPTIONS

RESERVE / COOLDOWN PROPELLANT / BOILOFF RATES: 2% / 3% / 0.65 kg/m²/mth

PROPELLANT TANKS JETTISONED AFTER TMI AND MOC BURNS

³ CHAMBER PRESSURE = 1000 psia, NOZZLE EXPANSION RATIO (ε) = 500:1
⁴ ASSUMED VALUE - DETAILED CALCULATIONS REQUIRED TO VERIFY ADEQUACY/INADEQUACY
⁵ INCLUDES MASS FOR RCS ATTITUDE CONTROL WHILE ON STATION, MAIN PROPELLANT FEEDLINE FROM TANK LINES TO ENGINE, RUN TANK, TRUSS, AND INTERSTAGE[I'HRUST STRUCTURE)

The reference mission's optimization is focused on reducing the Initial Mass in Low Earth Orbit (IMLEO) {which can be thought of as the wet mass}. This is the reason the spacecraft and mission is built around using a single NERVA engine, since those things are heavy. The engine has a thrust of 334 kilo-newtons (75 klb_f).

One engine instead of three means the thrust-to-weight (T/W) ratio goes way down. The spacecraft has a lower acceleration, which means it takes longer to escape Terra's gravity, which means the gravity losses become larger, which means more delta V is needed, which means more propellant is needed.

The way to avoid this vicious cycle is to use the magic of the Oberth Effect. By doing the Terra departure burn at perigee (at the point in the orbit when closest to Terra, more generally periapsis) you actually get some delta V for free (actually the extra delta V comes from the potential energy from the mass of the propellant expended).

In this case, you want to do three burns at periapsis to minimize gravity loss. Refer to the graph above. If the spacecraft has a thrust-to-weight ratio of 0.05, escaping Terra with a single burn at periapsis will cost you 1,500 m/s of gravity loss. But if you do the escape with three separate burns at periapsis, the gravity loss is only 350 m/s.

Naturally if you increase the engine thrust in such a way that the T/W ratio goes up, this will also lower the gravity loss penalty. This is tricky since higher thrust engines generally also have a higher mass. But in this case it is almost impossible since the optimization is focused on lowering IMLEO. You'd somehow have to increase the thrust while keeping the mass the same. Maybe by shifting gears.

The report points out if you swap the 334 kilo-newton NERVA engine for a honking monsterous 1,112 kilo-newton Phoebus engine the spacecraft could do a single burn escape with gravity loss of only 400 m/s (the T/W ratio rises to 0.15). The price is the IMLEO rises from 615 metric tons to 750 metric tons.

⁶ AI/Li VERSUS SiC/AI METAL MATRIX TANKS ON BOEING REF., G-LOSS AS FUNCTION OF VEHICLE THRUST-TO-WEIGHT (FROM LOOK-UP TABLE) VERSUS ASSUMED CONSTANT VALUE (200 m/s), ETC.

The table above compares the IMLEO spacecraft masses of two older Mars reference missions, and the Lewis all-propulsive (resuable) reference mission with a NERVA operating at a specific impulse of 925 seconds. The impressive part is how the optimized trajectory of the Lewis mission saves about 150 metric tons of IMLEO.

On the left is the Lewis all-propulsive optimized reference ship. On the right is the NASA reference ship. The differences are in the sizes of the various propellant tanks, and the IMLEO. The Trans-Mars Injection Drop Tanks are limited by the payload shroud dimesions of anticipated heavy launch vehicles. The report assumes the limit is 10 meters in diameter by 30 meters in length.

For the four NERVA engine types, the initial mass in low Earth orbit (IMLEO) and the total engine burn time was calculated for the mission. The carbide core has the lowest mass, but the 1,112 kN composite core has the shortest burn time.

The 2016 propulsion-optimized 434-day mission was assumed, along with engine thrust of 334 kilo-newtons (75 klb_f) or 1,112 kN (250 klb_f), 1000 psia chamber pressure, 500-to-1 nozzle expansion ratio, 3 perigee burn Terra departure (1 burn for 1,112 kN composite), and in reuse mode (i.e., spending extra delta V to avoid discarding the ship).

TMI = Trans Mars Injection, MOC = Mars Orbital Capture, TEI = Trans Earth Injection, EOC = Earth Orbital Capture

The report looked at other missions across the synodic period. The chart below assumes the spacecraft uses the 334 kN (75 klb_f) composite engine with an I_sp of 925 seconds. The chart shows how increasing the initial mass in low Earth orbit (IMLEO) shortens the trip time.

"All Prop" is the all-propulsive mission where extra delta-V is spent to capture the spacecraft into Terra orbit for reuse.

"ECCV" is the mission where the extra delta-V is NOT spent, the crew abandons the spacecraft in the Earth Crew Capture Vehicle (ECCV) and aerobrakes to Terra landing, but the spacecraft goes sailing off into the wild black yonder.

"Split-sprint" is where the cargo is sent in an unmanned spacecraft on a Hohmann conjunction-class trajectory, while the crew goes in a manned spacecraft on a faster high-energy opposition-class trajectory.

Not shown is the dangerous "Hohmann tanker/dual vehicle" mission. This is where the unmanned cargo ship also carries the manned spacecraft's return propellant. Which means if the manned ship arrives only to discover that all the return propellant has leaked out, the crew is doomed.

As you can see, the 2018 All Prop mission has an IMLEO of about 700 metric tons at a 434 day mission. If you wanted to decrease the mission time to 365 days (1 year) you'll have to almost double the IMLEO to about 1350 metric tons.

This is from History of rocketry and astronautics : proceedings of the twenty-fourth Symposium of the International Academy of Astronautics, Dresden, Germany, 1990, starting at page 33.

It is about Lockheed Missiles and Space Company's study of the Early Manned Planetary-Interplanetary Roundtrip Expeditions (EMPIRE) Mars mission.

Lockheed Missiles and Space Company

     Before becoming contractually involved in the EMPIRE studies with the NASA-Marshall Space Flight Center, the Lockheed Aircraft Corporation (later the Lockheed Corporation) of Sunnyvale, California, had performed several in-house research studies germane to manned Venus and Mars fly-by missions. As early as I960, the Lockheed Missiles and Space Division produced a report on space mechanics by C. M. Petty. Another pertinent study, published a month later by Leighton F. Koehler, concerned the orbital parameters of a manned satellite orbiting Mars.

     Detailed studies on interplanetary transportation systems supporting EMPIRE were conducted during 1962 and 1963 under contract NAS8-2469 by Lockheed Missiles and Space Company’s (LMSC) Flight Mechanic Group, Aerospace Sciences Laboratory. Supervision was exercised by Marshall’s Future Projects Office and Aero-Astrodynamics Laboratory. The final report, which was submitted on 30 April 1964, covered LMSC’s study of analytic derivations and numerical data applicable to interplanetary and planetocentric flight mechanics, navigation, and guidance. In the words of the report, “the general information provides the background and means for extended mission studies; the specific [information] illustrates its use and significant applications.”

     This study examined five areas of importance to EMPIRE planning: non-stop interplanetary round-trips, stopover interplanetary round-trips, missions launched normal to the ecliptic, nonstop trips passing Mars and Venus, and precise calculations and investigations of requisite guidance sensitivities.

     EMPIRE studies conducted under NASA contract NAS8-5024 were directed to Marshall’s Future Projects Office during 1962 and 1963. The first phase of this effort was concerned with spacecraft and launch vehicle sensitivities for manned flyby missions carried out during the Venus 1974 conjunction and the 1975 Mars opposition. The contract, which was subsequently extended from 1 April 1963 to 1 January 1964, emphasized missions most likely to be realized using launch vehicles proposed at the time.

     Study objectives established for LMSC by Marshall’s Future Projects Office included:

1. A detailed definition of suitable mission profiles not requiring the development of major new chemical or nuclear propulsion systems.
2. Sufficient investigation of subsystems to delineate requirements and possible pacing items.
3. Preliminary design of spacecraft capable of undertaking early Venus and Mars round-trips based on the capabilities of the Saturn launch vehicle family.
4. Identification of launch vehicle requirements and comparison with current programs.
5. Development plan and funding schedule.

     The method of approach was described in these terms:

By an iterative process, mission velocities, and their effects on booster and re-entry requirements, were compared with mission times and their effects on life support and environmental control weights for ‘nearly open’ and ‘nearly closed’ systems for different crew sizes. When it became clear that Earth departure velocity was the dominating factor for the missions under considerations, the analysis to select missions of minimum mass on Earth orbit was simplified to the selection of missions with minimum Barth departure velocities.

The many systems were investigated sufficiently to provide (within limits set by study fund limitations) an indication of applicability of the manned interplanetary program; development status; weight, size, and power characteristics; and cost. Vehicle concepts were developed and masses determined for different missions considering mission time and its effects on life support requirements, and also (with the influence of solar proximity) on shielding requirements. The spacecraft masses were compared with the velocity/payload performance capability of different Earth orbit boosters and the resulting combined weight effect for Earth surface launch boosters.

     Principal assumptions developed for the EMPIRE study took account of a number of constraints. Among them:

1. The Earth surface launch vehicle was to be the Saturn V with the S-II serving as its second stage.
2. Systems carried to Earth orbit by the Saturn V were to rendezvous (where needed) to form the escape vehicle.
3. If possible, the mass required in Earth orbit was to be held to that achievable by two Saturn V launches (one rendezvous) so that a third launch pad could accommodate a backup launch vehicle and two payloads ready in the event either Saturn V launch were to fail.
4. The Earth orbit escape was to use a chemical propulsion system or a nuclear propulsion system developed from NERVA technology (either a single stage or a two stage configuration).
5. The spacecraft was to carry probes to gather direct data on planetary atmospheres and surface conditions.
6. Earth reentry was to be accomplished using an Apollo capsule modified to meet new mission requirements and increased entry speeds.
7. Unless and until it were established that an artificial gravity field was not required, vehicle systems capable of providing simulated gravity were to remain under investigation.

     Study limitations were an important aspect of the LSMC effort. Considering the capability of the Saturn V, nuclear propulsion likely to be developed by then available technology, and the use of the Apollo spacecraft, the only interplanetary missions that appeared to be achievable were flybys. Limited funding precluded detailed investigations of the many subsystems such missions demanded.

     LMSC made its first EMPIRE progress report on 6 August 1963 at the Marshall Center. A second report was delivered on 2 October 1963 at Lockheed’s Palo Alto, California, facility. It was concluded that launch vehicle and spacecraft performance and design assumptions were valid and that major subsystems had been more clearly identified and delineated than had earlier been the case.

     At the conclusion of the study, the LMSC team under the direction of Benjamin P. Martin delivered its final report to the Marshall Space Flight Center in several volumes. The first of these was an unclassified summary of the complete study. The second volume consisted of unabridged findings in two parts, Part A containing all unclassified material and Part B incorporating classified data on the nuclear propulsion systems and their associated launch vehicles. The third volume was a condensed, unclassified summary of the entire project.

     The follow-on phase of the EMPIRE studies emphasized manned flyby missions of Venus and Mars during the 1974 conjunction of the former planet and the 1975 opposition of the latter. For the Venus mission, a departure on 11 November 1973 was proposed, a cruise duration of 370 days, and an approach to within 500 nautical miles of the planet. For the Mars mission, launch would occur on 24 September 1975; and, following a 670-day cruise, the spacecraft would also pass within 500 nautical miles of its target.

     Martin and his associates stressed that such flyby missions would be very useful as precursors for later manned landings on the planets. Not only would landings provide the opportunity to gather valuable planetary data by on-board sensors and crew observations but would further define man-machine relationships and roles and crew requirements for long-term space voyages.

     A number of conclusions were drawn from studies dealing with launch from parking orbit. Among them were:

1. The problem of launch window (the time available for launch from a particular parking orbit including plane-change requirements) must be a major consideration in selecting orbit launch systems and determining related operations. Launch windows would depend on the vehicle’s ability to absorb a plane change and could range from several days to several hours.
2. A dual burn technique might be useful in reducing the plane change penalty.
3. The penalty of maintaining the departure asymptote with the orbital plane was found to be greater than that of absorbing the plane change during launch.
4. Selection of an orbit inclination slightly higher than the declination of the departure asymptote would keep the plane change penalty low over the longest period.

     Most of the work involving crew utilization and requirements was accomplished in the first phase of the EMPIRE effort and was so reported in earlier documentation. However, general recommendations for life support and environmental control systems were made for a physico-chemical system of the “nearly closed” type shown in Figures 17 and 18. Assumptions supporting the selection are shown in Table 9. The crew compartment as a subsystem of the complete spacecraft and environment control system is shown in Figure 19. Food and water requirements for the subsystem are presented in Table 10, and the electric power requirements for various size crews are listed in Table 11. To protect the crew from ionizing radiation, sufficient shielding (consisting of aluminum and polyethylene) was felt to be necessary to limit the probability to 0.0001 of exceeding 200 rad to the blood-forming organs, with doses from cosmic and nuclear reactor sources both assumed to be 20 rad/hour.

     In carrying out the guidance and control portion of the study, Martin and his associates divided the interplanetary mission into the following phases: (a) injection into the heliocentric trajectory; (b) midcourse; (c) planetary approach; and, where stopover missions were considered, (d) the stopover orbit. Depending on the mission, some phases would be repeated on the return leg as shown in Figure 20.

     During the initial study period, navigation techniques, accuracy requirements, guidance procedures, and the conceptual development of an integrated spacecraft guidance system were considered for the injection, midcourse, and planetary approach phases. The follow-on study was directed toward a refinement of techniques, extending spacecraft guidance to include stopover capability as well as the development of an instrumented landing probe to aid in defining stopover orbit guidance requirements and surface reconnaissance techniques. A diagram of a proposed system is shown in Figure 2l.

     Estimated weight and power requirements are summarized in Table 12. It is noted that environmental control equipment and attitude torquers are not included. Computer memory and clock are continuous while other items operate intermittently.

     Interplanetary reconnaissance as envisioned by Martini and his LMSC EMPIRE study team was to require a total, integrated system of data sensing, processing, correlation, and evaluation geared to supply planetary data for future designs related to manned landings on and exploration of Mars. It was anticipated that these data would supplement a logical sequence of earlier Earth-based astronomical and unmanned reconnaissance missions. However, the intimate knowledge of the landing site and its environment necessary to commit a manned lander appeared attainable only through the coordinating efforts of an on-board space crew. The crew was to perform this role only if it were possible to design an adequate, on-board, data processing and display system.

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     Preliminary review on the reconnaissance goals, capabilities, and problems of three interplanetary missions—flyby, orbiting, and landing—led to the following conclusions:

1. The flyby would be a limited, one-shot, narrow-swath reconnaissance tool, subject to continuous scale and angular-rate variation detrimental to optimum sensor design and to data interpretation. Its best use might be for operational testing of future-mission equipment and possibly for accurate insertion of a reliable unmanned reconnaissance orbiter and/or lander.
2. The manned orbiting mission would provide a stable reconnaissance platform, capable of mapping the complete planet, sensing vast quantities of detailed surface data at near-uniform scale, and—if properly equipped—collating atmospheric and surface data. Major problems were felt to involve mass-data handling, automatic processing and displays required to aid the crew in tracking, correlation, sensor control and evaluation.
3. The landing mission would require a higher order of data-handling and correlation sophistication than the orbiter. It was noted that in the course of repetitive orbits, preferred exploration areas would have to be located and landing site details (including the immediate environment) closely inspected by means of on-board and probe-borne sensors. The safety of manned landing operations would be judged from orbit. Furthermore, landing and return-rendezvous sensor systems would have to be provided along with surface exploration aids. The total equipment weight (including probes) was found to vary from 1,600 pounds for the flyby mission to 20,000 pounds (plus storage and transmission equipment) for the landing mission.

     Studies of electric power requirements for three-man Venus/Mars flyby missions took into account nuclear and solar-activated primary systems as well as auxiliary or emergency systems and trade-offs. Auxiliary power plant investigations considered Gemini- and Apollo-type hydrogen-oxygen fuel cells and reciprocating engines. A power output of 5 to 8 kilowatts was assumed for each such system. Among the nuclear options studied, the preference was for an 8-kilowatt nuclear/dynamic system using a SNAP 2/8 reactor and a Brayton gas-cycle energy conversion unit.

     At the time the electric power supply and system portion of the LMSC EMPIRE study was completed, a number of observations could be made:

1. The leading candidate for the spacecraft was a nuclear power system.
2. No firm requirement for auxiliary power could be established in conjunction with the nuclear power plants.
3. There were weight penalties inherent in the nuclear power plant.
4. At the same time, there were several operational constraints in the solar/dynamic power plant.
5. It was felt that multiple solar/dynamic systems might be required.
6. An auxiliary power plant would be needed in conjunction with the use of a solar/dynamic primary power plant for Earth escape.

     The LMSC investigators proposed a slowly rotating spacecraft that would produce an artificial gravity of 0.4 g. They reported that “two detailed spacecraft configurations were designed for the interplanetary flyby missions with differences brought about largely because of the electrical power system concepts. One design employs a solar dynamic power system, and the other a nuclear dynamic power system. The first configuration, which is a modification of the spacecraft described in the initial EMPIRE report, incorporates a larger, more spacious solar shelter in the hub and provides for a heavier Earth-entry system. The second configuration is a more detailed study of a spacecraft mentioned briefly in the initial report.” In both cases, the spacecraft would comprise a command module, a mission module, a power system, and a mid-course propulsion unit.

     The command module for the spacecraft fitted with a solar power system would be a modified Apollo with an internal volume of about 8.5 cubic meters. It would serve as the crew’s launch vehicle, emergency escape vehicle during launch, Earth entry vehicle during an emergency in Earth orbit, command/control center during interplanetary cruise, and Earth entry vehicle at the completion of the mission. For one configuration using a nuclear power system, the command module would be a special purpose design integrating the command/control center with the solar shelter. A second configuration would em-ploy a command module similar to the solar power system configuration’s command module.

     The mission module would house the crew’s living quarters, the dining and recreational area, and the environmental control equipment, food, water, and spares for the spacecraft. Internal volume of the module would be 113 cubic meters.

     The spacecraft was designed with a protective shelter to shield the crew from radiation during periods of intense solar activity.

     This configuration, shown in Figure 22, consisted of command and mission modules connected by 25-meter rigid spokes to a hub. Crew access to the modules would be through these spokes. The command module would be a modified Apollo spacecraft with an attached retro-propulsion pack to decrease Earth entry velocity to design values. The mission module would be 3.66 meters in diameter and 12.1 meters long. The solar shelter, shown in Figure 23, would have an internal volume of approximately 5.6 cubic meters and would be located at the rotational hub with the electric power system and the mid-course propulsion unit.

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     For launch, the spacecraft would be stowed as shown in Figure 24. The sequence for deploying the craft once in Earth orbit was described by the LMSC EMPIRE team thus:

The spacecraft is launched as a compact package on the booster and then erected in Earth orbit. The crew is launched into orbit in the command module to which the emergency escape rocket is attached. This rocket is jettisoned at the end of the first stage burnout. In orbit, the sequence is as follows: (1) the fairings and interstage adapters are jettisoned, (2) the command module with the attached retrorocket is rotated 180 degrees and secured to its spoke, (3) this assembly is rotated 90 degrees and secured to the hub, (4) the opposite spoke with the mission module attached is rotated 90 degrees and secured to the hub, and then (5) the mission module is rotated and secured to the spoke. A cable system can be used to move the spokes and modules into proper position, one section of the system would ‘pull’ the modules into position and the other section would ‘restrain’ these modules from closing too rapidly.

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     All systems within the spacecraft were to be checked out in Earth orbit before proceeding with the interplanetary cruise. Power during the checkout phase was to be supplied by an auxiliary chemical system, either a thermal piston or turbine system, which would operate on storable propellants from the mid-course propulsion system, or by fuel cells. The solar concentrator was not to be erected in Earth orbit because the large diameter reflector could not withstand escape loads during the injection phase.

     The nuclear-powered configuration is shown in Figure 25. Basically, the arrangement of the modules is the same as that of the solar-powered version, except that a third spoke is required to support the nuclear-power system and a radiation shield is added. The solar shelter located in the hub, as seen in Figure 26, is of a somewhat different design. A variation shown in Figure 27 would shift the Apollo Earth-entry module to the hub. Its solar shield is illustrated in Figure 28.

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     At launch the spacecraft would be folded as shown in Figure 29. The crew would accommodate itself in the entry body and the erection sequence on orbit would proceed as follows: (a) the entry body would rotate 180 degrees and two crewmen would transfer into the command module/solar shelter, (b) the entry body would next be maneuvered to the side of the hub and docked, (c) the command module/solar shelter would be rotated, attached to the spoke, and this assembly would then rotate 90 degrees and be secured to the hub, (d) the mission module and spoke would rotate 90 degrees, the spoke secured to the hub and the module to the spoke, and (e) the nuclear power system and spoke would rotate 90 degrees and the spoke would be secured to the hub. The nuclear power system would then be activated to supply power to the spacecraft.

     Injection payload weights for the various spacecraft configurations studied are summarized in Table 13. Detailed weight breakdowns for the Venus flyby are given in Table 14, and the same data for the Mars mission appear in Table 15.

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     Since the 1970-1980 timeframe was specified for the study, all evaluations of launch vehicles had to be predicated on the Saturn V and its various stages, actual and proposed. as shown in Figure 30.

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     Flyby missions were analyzed using both an all-chemical Saturn V launch vehicle and a variation incorporating a nuclear-powered upper stage. One alternative studied for the former configuration consisted of two S-IVB stages while another would be capable of orbiting an S-II stage. Both approaches would require modifications to the stages, orbital rendezvous and on-orbit fueling operations.

     Figure 31 shows NERVA nuclear stage configurations considered by the Lockheed Missile and Space Company team. S-NA is the designation given to any nuclear stage powered by a Mod 1 engine. Two sizes were studied: (a) S-NA1 for stages with a propellant capacity of less than 64,000 kilograms, and (b) S-NA2 for stages with propellant capacities in excess of that amount. The name S-NB applies to any stage using Mod 2 engines. Engine characteristics assumed are given in Table 16. Venus and Mars flyby configurations for nuclear stages are shown in Figure 32 while Figure 33 depicts the tank structure for a typical nuclear stage.

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     Conclusions reached by the LMSC team during the course of the EMPIRE study program include:

1. A manned Venus flyby based on a single Saturn V/N (nuclear upper stage) launch could not be completely ruled out.
2. The Venus flyby definitely appeared to be possible based on two Saturn V launches (single rendezvous). The mission could be initiated from orbit by: (a) staging of either of two S-IVB’s, two nuclear stages, or an S-IVB/nuclear combination; or (b) one large nuclear stage.
3. The liquid oxygen/liquid hydrogen propellant required for the Venus flyby mission was beyond the amount that could be orbited in an S-IVB stage by a Saturn V. Therefore, a flyby initiated with one S-IVB would not be possible unless propellant were added through refueling, or providing additional tanks, or developing another suitable propulsion stage.
4. The ability to perform a Mars low-energy flyby with two Saturn V launches to orbit could not be resolved at the time.
5. It was determined that three two-stage booster configurations would yield the highest performance: S-NA1/S-NA1, S-NA1/S-NA2, and S-IVB/S-NA1. Substituting a Mod 2 or superior engine would provide an even greater capability. However, the S-NA1/S-NA1 would require close to a 50-S0 split in the payload launched to orbit, greatly complicating rendezvous operations and systems. The S-NA1/S-NA2 would require two different nuclear stage sizes. The S-IVB/S-NA1 would achieve high performance because the thrust-to-weight ratio for each stage would be sufficient to reduce gravity loss to a minimum. However, it presented configuration problems arising from a first stage of smaller diameter than the second stage.
6. The Saturn V appeared incapable of launching a single nuclear stage into orbit with sufficient propellant to accomplish a Mars low-energy flyby. Thus, either the payload mass would have to be reduced or some other configuration or operational change would need to be devised.

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     There were, of course, ways to overcome the discrepancy between payload requirement and booster capability. Three Saturn V launches, for example, with two orbiting S-NB nuclear stages and a third orbiting the intact payload, would virtually guarantee Mars flyby capability. This was seen as one of the simplest and least costly methods offering the greatest chance of success.

     Development planning undertaken by the Lockheed team was restricted by both time and budget; therefore, only the Mars mission was considered.

This is from Manned Mars missions: A working group report (1986). This was mainly to update earlier Mars mission data, examine the impact of new technologies, and identify new issues that needed new research. In other words a lot of this report is a re-hash of old stuff.

The report starts with the old, tired, golly-gee-whiz stuff about how the Space Shuttle was only the beginning. You ain't seen nuttin' yet, folks! This was followed by the obligatory word-salad of miscellaneous excuses why Mars was the logical destination for our valiant astronauts. Things like "going to Mars is an endeavor that can captivate and motivate a generation of young people from throughout the world, just as the American frontier motivated generations past".

PAYLOAD

The payload is basically:

• Mission Module: the habitat module the astronauts live in
• Mars Excursion Module: vehicle that transports the explorers to the Martian surface, then returns them to the spacecraft in orbit at the end of the surface stay.
• Scientific Equipment: equipment used for scientific observations on the journey to and from Mars, plus on the Martian surface, in an attempt to cram as much science into the expensive Mars mission as they possibly can.

PROPULSION SYSTEM

The report figured that anything with an acceleration below 0.05 g's (dotted line in figure 4.22) would increase the trip time over acceptable duration, so should be eliminated from consideration. So the only acceptable propulsion systems are Orion nuclear pulse, Solid-core nuclear thermal rockets (NTR), or Chemical rockets (cryogenic liquid oxygen/liquid hydrogen)

Orion drive with its fuel tanks full of nuclear bombs is a non-starter. Solid-core NTR would be a vast improvement over chemical rockets, but the technology is not quite ready for prime time. Chemical rockets have a lackluster specific impulse, but at the time the report was written they had been around for half a century. It certainly was a mature technology. Most of the report focuses on chemical rocket propulsion (with occasional sunday-newspaper science-supplement comments about pie-in-the-sky stuff like antimatter rockets. Hey kids, ask your grandparents what a newspaper was).

MISSION PROFILES

Just like every other Mars mission design in the universe, there are three basic trajectories: Fly-by, Opposition Class, and Conjunction Class.

Fly-by mission are a waste for a manned mission, since there really isn't anything a manned fly-by can do which couldn't be performed vastly less expensively by a robot probe.

Conjunction class missions have a longer mission duration than Oppostion class (three years as opposed to two). They require a longer stay on Mars than Opposition class (one year as opposed to two months). The advantage is that Conjunction class requires less propellant and less initial mass in LEO (IMLEO).

The report suggests that Opposition class is suited for unmanned cargo delivery missions, and for initial short duration manned missions (when the bugs are still being worked out). Later manned missions might be better suited to use Conjunction class.

This is from NASA's Nuclear Thermal Propulsion (NTP) Project (2018) and STMD: Nuclear Thermal Propulsion Update (2019)

NASA is no stranger to atomic rockets. The Nuclear Engine for Rocket Vehicle Application (NERVA) ran for almost two decades before it got the axe in 1972. Shifting priorities, political winds and space budget cutbacks lead to NERVA's downfall.

The problem is trying to make a Mars mission spacecraft using conventional chemical rockets is like attempting to make a rubber-band powered passenger airplane.

A Mars expedition spacecraft with only chemical rockets can only abort back to Terra within five days of the start of the mission. After that it is committed to the mission. By contrast, a spacecraft using NTP can abort any time from the start up to three months into the mission.

As a ray of hope, in 2018 NASA started a new atomic rocket program, the Nuclear Thermal Propulsion Project, lead by Doyce “Sonny” Mitchell. NASA is partnering with BWXT Nuclear Energy, Inc. on engine development.

One of the obstacles is that the original NERVA project used highly-enriched uranium. Otherwise know as "weapons-grade." The powers-that-be are hysterically afraid of Highly Enriched Uranium (aka "Weapons-Grade") falling into the Wrong Hands. HEU drastically increases the development costs and security regulations.

BWXT Nuclear is well aware of this, and has figured out how to make a solid core nuclear thermal rocket engine using Low Enriched Uranium (LEU). As it turns out there is so much energy in uranium that using HEU is a bit of over-kill, if you pardon the expression. LEU has more than enough power to send a spacecraft to Mars while simultaneously soothing the fears of the powers-that-be.

Chemical engines are hard pressed to make the Terra-Mars section of the mission is less than 258 days. A trio of 111 kN BWXT engines can do it in 160 days flat. That's a bit more than three months less, which is three less months of freefall muscle atrophy and cosmic radiation exposure. Not to mention three months less of life support consumables mass that can be reassigned to mission scientific payload. Actually six months less if you count the return trip.

Abbreviations

• NTP: Nuclear Thermal Propulsion. Another name for Nuclear Thermal Rocket
  
  
• 4% FPR: Flight Performance Reserve, the extra propellant carried in case things go wrong. In this case it is an additional four percent of the propellant mass.
  
  
• OMS: Orbital Maneuvering System, engines used for orbit-to-orbit manuevers
  
  
• RCS: Reaction Control System, the spacecraft's attitude jets
  
  
• TMI: Trans Martian Insertion, engine burn to depart Terra for Mars
  
  
• MOI: Mars Orbit Insertion, engine burn to brake into Mars Orbit
  
  
• TEI: Trans Earth Insertion, engine burn to depart Mars for Terra
  
  
• EOI: Earth Orbit Insertion, engine burn to brake into Terra orbit
  
  
• NRHO: Near Rectilinear Halo Orbit, a stable spot to assemble spacecraft components or for a ship to await the arrival of the crew. The object sort of orbits around EML1 or EML2
  
  
• LD-HEO: Lunar Distant High Elliptical Orbit. An orbit with a perigee near LEO and an apogee near Luna. Starting in this orbit helpfully reduces the delta-V needed for TMI.

The report has a vague reference to an Orbital Maneuvering System (OMS) engine that has a specific impulse of 500 seconds. It is used for manuevers within Terra and Martian spheres of influence. I'm not sure if this is a separate engine or a different mode of the nuclear engine.

Inspired by a post by Retro Rockets I took a look at the fictional spaceship Luna from the movie Destination Moon (1950). With Robert Heinlien as technical consultant, this movie was the most scientifically acurate one since Frau im Mond (1929). It held the throne for 18 years, until it was supplanted by the movie 2001: A Space Odyssey (1968).

For the specifications I used data from Spaceship Handbook and the Retro Rockets article. I then massaged the figures until they were internally consistent.

Spaceship Handbook calculated that a round trip mission to the surface of Luna would take about 16,480 m/s of delta V. So that's our performance limit for the mission. In addition, it will have to have a thrust-to-weight ratio greater than 1.0, since it has to lift off from Terra's surface. The movie specifies 5 gs, which translates to 11,000,000 newtons.

The movie specified that the reaction mass was water, not liquid hydrogen. While this does simplify the tankage, it does cut the exhaust velocity/specific impulse in half. Heinlein apparently wanted storable propellant, not liquid hydrogen with all its cryogenic coolers and other complications.

A solid-core nuclear thermal rocket engine is not going to be able to crank out enough delta V, not with water at the specifed mass-ratio it ain't. But the liquid-core Liquid Annular Reactor System (LARS) will do nicely. It can jet out liquid-hydrogen propellant at 20,000 m/s or better, so it can probably manage to hurl water at 10,300 m/s. That will give the Luna a delta-V of 16,600 m/s, just a tad larger than the required 16,480 m/s for the Lunar mission. More than enough, assuming you don't waste a lot of delta V during the landing.

The movie says the structural mass is 27 metric tons, which makes it 60% of dry mass. Nowadays NASA vessels typically have a structural mass of 21.7% of structual mass. 60% is a bit extravagant but believable with 1950's technology. If you made the structure NASA-light, you could add about 17 metric tons to the payload. The payload is the crew, equipment, life support, acceleration couches, and controls.

DESTINATION MOON

      “Loading is coming all right,” answered Bowles, “provided the trucks with the oxygen aren’t late.”
     “You should have flown it in.”
     “Quit jittering. The trucks are probably in Cajon Pass this minute.”
     “Okay, okay. Power plant, Doc?”
     “I haven’t broken Ned Holmes’ (Atomic Energy Commission's Tamper-Evident) seal on the atomic pile yet. The water tanks are filling, but they’ve just started.” (spacecraft uses water as propellant)

---

• 

     The great ship was ringed with floodlights spaced inside the bull pen. It was still framed by the skeleton arch of the Gantry crane, but the temporary anti-radiation shield which had surrounded its lower part down to the jets was gone; instead there were posted the trefoil signs used to warn of radioactivity—although the level of radiation had not yet become dangerously high.

     But the power pile was unsealed and the ship was ready to go. Thirteen-fifteenths of its mass was water, ready to be flashed into incandescent steam by the atomic pile, to be thrown away at thirty thousand feet per second.

(ed note: 13/15^th of wet mass is propellant mass. Means mass ratio = 7.69

exhaust velocity 30,000 fps = exhaust velocity 9,144 m/s = specific impulse of 932 sec

ΔV (delta-V) = 18,640 m/s. Spaceship handbook calculated Luna would need 16,480 m/s for the moon round trip so we have enough. )

     High up in the ship was the control room and adjacent air lock. Below the air lock the permanent anti-radiation shield ran across the ship, separating the pressurized crew space from the tanks, the pumps, the pile itself, and auxiliary machinery. Above the control room, the nose of the craft was unpressurized cargo space.
     At its base triangular airfoils spread out like oversize fins—fins they would be as the ship blasted away; glider wings they would become when the ship returned to Earth with her tanks empty.

---

     Time checks had been completed with Muroc, with White Sands and with their blockhouse. The control room was quiet save for the sighing of air-replenishing equipment, the low hum of radio circuits, and stray sounds of auxiliary machinery. The clocks at each station read 3:29—twenty-four minutes to H-hour.

• 

     The four were at their stations; two upper bunks were occupied by pilot and co-pilot; the lowers by power engineer and electronics engineer. Across the lap of each man arched a control console; his arms were supported so that his fingers were free to handle his switches without lifting any part of his body against the terrible weight to come. His head was supported so that he might see his instruments.
     Barnes could hear sirens, rising and falling, out on the field. Above him in the TV screen, the Luna stood straight and proud, her head in darkness.
     “Eleven!
     “And ten!
     “And nine!
     “And eight!”—Barnes licked his lips and swallowed.
     “Five—four——three———two—
     “Fire!”
     The word was lost in sound, a roar that made the test blast seem as nothing. The Luna shrugged—and climbed for the sky.
     Barnes felt himself shoved back into the cushions. He gagged and fought to keep from swallowing his tongue. He felt paralyzed by body weight of more than half a ton; he strained to lift his chest. Worse than weight was noise, a mind-killing “white” sound from unbearable ultrasonics down to bass too low to be heard.
     The sound Dopplered down the scale, rumbled off and left them. At five effective gravities they outraced their own din in six seconds, leaving an aching quiet broken only by the noise of water coursing through pumps.

(ed note: initial acceleration = 5g = 49 m/s² )

---

     Corley joined Bowles. Barnes hesitated; he wanted very badly to see, but he was ashamed to leave Traub working. “Wait,” he called out. “I’ll turn ship and we can all see.”
     Mounted at the centerline of the ship was a flywheel. Barnes studied his orientation readings, then clutched the ship to the flywheel. Slowly the ship turned, without affecting its motion along its course. “How’s that?”
     “Wrong way!”
     “Sorry.” Barnes tried again; the stars marched past in the opposite direction; Earth swung into view. He caught sight of it and almost forgot to check the swing.
     Power had cut off a trifle more than eight hundred miles up. The Luna had gone free at seven miles per second (11,265 m/s which is Terra escape velocity); in the last few minutes they had been steadily coasting upwards and were now three thousand miles above Southern California. Below—opposite them, from their viewpoint—was darkness. The seaboard cities stretched across the port like Christmas lights. East of them, sunrise cut across the Grand Canyon and shone on Lake Mead.

---

     Traub was saying in a steadier voice, “Spaceship Luna, calling White Sands. Come in White Sands.”
     At last an answer came back, “White Sands to Spaceship—go ahead.”
     “Give us a series of radar checks, time, distance, and bearing.”
     A new voice cut in, “White Sands to Spaceship—we have been tracking you, but the figures are not reasonable. What is your destination?”
     Traub glanced at Barnes, then answered, “Luna, to White Sands—destination: Moon.”
     “Repeat? Repeat?”
     “Our destination is the Moon!”
     There was a silence. The same voice replied, “ ‘Destination: Moon’—Good luck, Spaceship Luna!”

---

     Later the radio claimed his attention. “White Sands, calling Spaceship—ready with radar report.”
     The first reports, plus a further series continued as long as White Sands and Muroc were able to track them, confirmed Barnes’ suspicion. They were tracking “high,” ahead of their predicted positions and at speeds greater than those called for by Hastings’ finicky calculations. The difference was small; on the autopilot displays it was hardly the thickness of a line between the calculated path and the true path.
     But the difference would increase.
     “Escape speeds” for rockets are very critical. Hastings had calculated the classical hundred-hour orbit and the Luna had been aimed to reach the place where the Moon would be four days later. But initial speed is critical. A difference of less than one percent in ship speed at cutoff can halve—or double—the transit time from Earth to Moon. The Luna was running very slightly ahead of schedule—but when it reached the orbit of the Moon, the Moon would not be there.
     Doctor Corley tugged at his thinning hair. “Sure, the cutoff was mushy (when you turn off the engine the thrust does not instantly stop, it sort of gradually tapers off. This was not according to the design), but I was expecting it and I noted the mass readings. It’s not enough to account for the boost. Here—take a look.”
     Corley was hunched at the log desk, a little shelf built into the space between the acceleration bunks. He was strapped to a stool fixed to the deck in front of it. Barnes floated at his shoulder; he took the calculation and scanned it. “I don’t follow you,” Barnes said presently; “your expended mass is considerably higher than Hastings calculated.”
     “You’re looking at the wrong figure,” Corley pointed out. “You forgot the mass of water you used up in that test. Subtract that from the total mass expended to get the effective figure for blast off—this figure here. Then you apply that—” Corley hesitated, his expression changed from annoyance to dismay. “Oh, my God!”
     “Huh? What is it, Doc? Found the mistake?”
     “Oh, how could I be so stupid!” Corley started frenzied figuring.
     “What have you found?” Corley did not answer; Barnes grabbed his arm. “What’s up?”
     “Huh? Don’t bother me.”
     “I’ll bother you with a baseball bat. What have you found?”
     “Eh? Look, Jim, what’s the final speed of a rocket (the delta-V), ideal case?”
     “What is this? A quiz show? Jet speed times the logarithm of the mass ratio. Pay me.”
     “And you changed the mass ratio! No wonder we’re running high.”
     “Me?”
     “We both did—my fault as much as yours. Listen; you spilled a mass of water in scaring off that truckload of thugs—but Hastings’ figures were based on us lifting that particular mass all the way to the Moon. The ship should have grossed almost exactly two hundred fifty tons at takeoff; she was shy what you had used—so we’re going too fast.”

(ed note: wet mass = 250 tons = 226,800 kg. Shortly before lift-off they did a half-g burn of superheated steam to discourage a truckload of Soviet agents from sabotaging the launch. The expended propellant altered the mass ratio)

     “Huh? I wasted reaction mass, so we’re going too fast? That doesn’t make sense.” Barnes hooked a foot into the legs of the stool to anchor himself, and did a rough run-through of the problem with slide rule and logarithm table. “Well, boil me in a bucket!” He added humbly, “Doc, I shouldn’t have asked to be skipper. I don’t know enough.”
     Corley’s worried features softened. “Don’t feel that way, Jim. Nobody knows enough—yet. God knows I’ve put in enough time on theory, but I went ahead and urged you to make the blunder.”
     “Doc, how important is this? The error is less than one percent. I’d guess that we would reach the Moon about an hour early.”
     “And roughly you’d be wrong. Initial speed is critical, Jim; you know that!”
     “How critical? When do we reach the Moon?” Corley looked glumly at the pitiful tools he had with him—a twenty-inch log-log slide rule, seven place tables, a Nautical Almanac, and an office-type calculator which bore the relation to a “giant brain” that a firecracker does to an A-bomb. “I don’t know. I’ll have to put it up to Hastings.” He threw his pencil at the desk top; it bounced off and floated away. “The question is: do we get there at all?”
     “Oh, it can’t be that bad!”
     “It is that bad.”

     From across the compartment Bowles called out, “Come and get it—or I throw it to the pigs!”
     But food had to wait while Corley composed a message to Hastings (professor at the university who calculated the original course using room-sized computers). It was starkly simple: OFF TRAJECTORY. USE DATA WHITE SANDS MUROC AND COMPUTE CORRECTION VECTOR. PLEASE USE UTMOST HASTE—CORLEY.
     The Luna plunged on; Earth dropped away; radio signals grew weaker—and still no word from Hastings. Corley spent the time trying endlessly and tediously to anticipate the answer he expected from Hastings, using the tools he had.
     “Doctor Hastings calling.”
     “Oh, fine,” Corley acknowledged. “Slide out of there and let me take it.”
     Hastings’ report was short but not sweet. They would reach the Moon’s orbit where planned, but more than fifty hours too soon—and would miss the Moon by more than 90,000 miles!
     Barnes whistled. “Hot pilot Barnes, they call me.”
     Corley said, “It’s no joke.”
     “I wasn’t laughing, Doc,” Barnes answered, “but there is no use crying. It will be tragic soon enough.”
     Traub broke in. “Hey—what do you mean?”
     “He means,” Bowles said bluntly, “that we are headed out and aren’t coming back.”
     “On out? And out—out into outer space? Where the stars are?”
     “That’s about it.”
     “Not that,” Corley interrupted, “I’d estimate that we would reach our farthest point somewhere around the orbit of Mars.”
     Traub sighed. “So it’s Mars, now? That’s not so bad, is it? I mean, they say people live on Mars, don’t they? All those canals and things? We can get another load of water and come back.”
     Corley raised his voice. “Please! Everybody! I didn’t say we weren’t going to get back. I said—”
     “But you—”
     “Shut up, Red! I said this orbit is no good. We’ve got to vector west, toward the Moon. And we’ve got to do it at—” He glanced at a clock. “Good grief! Seven minutes from now.”
     Barnes jerked his head around. “Acceleration stations, everybody! Stand by to maneuver!”

---

     For forty hours they fell toward the Moon. The maneuver had worked; one could see, even with naked eye, that they were closing with the Moon. Some two hundred thousand miles out the Luna slid past the null point between Earth and Moon, and began to shape her final orbit. It became evident that the correction vector had somewhat overcompensated and that they were swinging toward the Moon’s western limb—“western” as seen from Earth: the Luna’s orbit would intersect her namesake somewhere on the never-yet-seen far side—or it was possible that the ship would skim the far side at high speed, come around sharply and head back toward Earth.
     Two principal styles of landing were possible—Type A, in which a ship heads in vertically, braking on her jets to a landing in one maneuver, and Type B, in which a ship is first slowed to a circular orbit, then stopped dead, then backed to a landing when she drops from the point of rest.
     “Type A, Jim—it’s simplest.”
     Barnes shook his head. “No, Doc. Simple on paper only. Too risky.” If they corrected course to head straight in (Type A), their speed at instant of braking would be a mile and a half a second and an error of one second would land them 8000 feet above—or below!—the surface.
     Barnes went on, “How about a modified ‘A’?”
     Modified Type A called for intentionally blasting too soon, then cutting the jets when the radar track showed that the ship hovered, allowing it to fall from rest, then blasting again as necessary, perhaps two or three times.
     “Confound it, Jim, a modified ‘A’ is so damned wasteful.”
     “I’d like to get us down without wrecking us.”
     “And I would like us to get home, too. This ship was figured for a total change of twelve and a half miles per second. Our margin is paper thin.”
     “Just the same, I’d like to set the autopilot to kick her a couple of seconds early.”
     “We can’t afford it and that’s that.”
     “Land her yourself, then. I’m not Superman.”
     “Now, Jim—”
     “Sorry.” Barnes looked at the calculations. “But why Type A? Why not Type B?”
     “But Jim, Type B is probably ruled out. It calls for decelerating at point of closest approach and, as things stand now, ‘closest approach’ may be contact.”
     “Crash, you mean. But don’t be so damned conventional; you can vector into a circular orbit from any position.”
     “But that wastes reaction mass, too.”
     “Crashing from a sloppy Type A wastes more than reaction mass,” Barnes retorted. “Get to work on a ‘B’; I won’t risk an ‘A.’ ”
     Corley looked stubborn. Barnes went on, “There’s a bonus with Type B, Doc—two bonuses.”
     “Don’t be silly. Done perfectly, it takes as much reaction mass as Type A; done sloppily, it takes more.”
     “I won’t be sloppy. Here’s your bonus: Type A lands us on this face, but Type B lets us swing around the Moon and photograph the back side before we land. How does that appeal to your scientific soul?”
     Corley looked tempted. “I thought about that, but we’ve got too little margin. It takes a mile and a half of motion to get down to the Moon, the same to get up—three miles. For the trip back I have to save enough mass to slow from seven miles a second to five before we dip into the atmosphere. We used up seven to blast off—it all adds up to twelve. Look at the figures; what’s left?”
     Barnes did so and shrugged. “Looks like a slightly fat zero.”
     “A few seconds of margin at most. You could waste it on the transitions in a Type B landing.”
     “Now the second bonus, Doc,” Barnes said slowly. “The Type B gives you a chance to change your mind after you get into a circular orbit; the straight-in job commits you beyond any help.”
     Corley looked shocked. “Jim, you mean go back to Earth without landing?”
     Barnes lowered his voice. “Wait, Doc. I’d land on the Moon if I had enough in tanks to get down—and not worry about getting up again. I’m a bachelor. But there’s Mannie Traub. No getting around it; we stampeded him. Now it turns out he has a slew of kids, waiting for poppa to come home. It makes a difference.”

---

     Traub’s features worked in agonized indecision. “But it’s not my business to decide!” Bowles spoke up suddenly. “You’re right; it’s not!” He went on, “Gentlemen, I didn’t intend to speak, because it never crossed my mind that we might not land. But now the situation demands it. As you know, I received a coded message.
     “The gist was this: our trip has caused grave international repercussions. The Security Council has been in constant session, with the U.S.S.R. demanding that the Moon be declared joint property of the United Nations—"
     “As it should be," Corley interrupted.
     “You don’t see the point, Doctor. Their only purpose is to forestall us claiming the Moon—we, who actually are making the trip. To forestall us, you understand, so that the United States will not be able to found a base on the Moon without permission—permission that is certain to be vetoed.”
     “But,” pointed out Corley, “it works both ways. We would veto Russia establishing a base on the Moon. Admiral, I've worked with you because it was a way to get on with my life's ambition, but, to be frank, using the Moon as a rocket launching base—by anybody—sticks in my craw.”
     Bowles turned red. “Doctor, this is not an attempt to insure the neutrality of the Moon; this is the same double-talk they used to stop world control of atomics. The commissars simply want to tie us up in legalisms until they have time to get to the Moon. We'll wake up one morning to find Russia with a base on the Moon and us with none—and World War Three will be over before it starts.”
     “But—Admiral, you can’t know that."
     Bowles turned to Barnes. “Tell him, Jim."
     Barnes gestured impatiently. “Come out of your ivory tower, Doc. Space travel is here now—we did it. There is bound to be a rocket base on the Moon. Sure, it ought to be a United Nations base, keeping the peace of the world. But the United Nations has been helpless from scratch. The first base is going to belong to us—or to Russia. Which one do you trust not to misuse the power? Us—or the Politburo?"
     Corley covered his eyes, then looked at Bowles. “All right,” he said dully. “It has to be—but I don’t like it."
     Traub broke the ensuing silence with “Uh, I don’t see how this ties in with whether we land or not?"
     Bowles turned to him. “Because of this: the rest of that message restored me to active duty and directed me to claim the Moon in the name of the United States—as quickly as possible. We would have what the diplomats call a fait accompli. But to claim the Moon I have to land!”

---

     Barnes locked eyes with him. “You can take your authority and—do whatever you think proper with it. I'm skipper and will stay so as long as I'm alive." He looked around. “All hands—get ready for approach. Doc, go ahead with trial calculations, Type B. Mannie, warm up the pilot radar. Bowles!"
     Finally Bowles answered, “Yes, sir."
     “Rig the autocamera in the starboard port. We’ll take a continuous strip as we pass around the far side."
     “Aye aye, sir.”
     Traub leaned from his couch and peered out the starboard port. “lt’s just like the other side.”
     Barnes answered, “What did you expect? Skyscrapers? Co-pilot, how do you track?”
     “Speed over ground—one point three seven. Altitude, fifty-one point two, closing slowly."
     “Check. I project closest approach at not less than twenty-one—no contact. What do you get?”
     “Closer to twenty, but no contact."
     “Check. Take over orientation. I'll blast when altitude changes from steady to opening."
     “Aye aye, sir!"
     The Luna was swinging around the unknown far face of the Moon, but her crew was too busy to see much of the craggy, devil-torn landscape. She was nearing her closest approach, traveling almost horizontally. She was pointed tail first, ready to blast back from a top speed of a mile and a half a second to a circular orbit speed of a mile a second. At Barnes‘ order Bowles gave his attention to placing her axis precisely horizontal.
     The television screen read “View Aft”; in its center was a cross mark lying over a picture of the mountainous horizon they were approaching. He jockeyed the ship against the re- action of the flywheel, then steadied her by gyros when one cross line held steady on the horizon.
     Barnes set his controls on semiautomatic, ready both to fire and cut off with one punch of the firing button. Into his autopilot he fed the speed change he wished to achieve. Altitude dropped to forty miles, to thirty, to less than twenty-five. “Power plant,” Barnes called out, “stand by for blasting!"
     “Ready, Jim,“ Corley reported quietly.
     “Electronics?”
     “Everything sweet, Skipper."
     Barnes watched ground speed with one eye, the radar altimeter with the other … twenty-three, it said … twenty-two … twenty-one and a half…
     Twenty-one point five … twenty-one point four—point four again—and again. Point five! and crawling up. His finger stabbed at the firing button.
     The blast was fourteen seconds only, then it cut off, but in the same mushy fashion which it had before. Barnes shook his head to clear it and looked at his board. Altitude twenty-one point five; ground speed, one plus a frog‘s whisker—they were in orbit as planned. He sighed happily. “That’s all for now, troops. Leave everything hot but you can get out of your hammocks."
     Bowles said, “Hadn’t I better stay and watch the board?"
     “Suit yourself—but they won’t repeal the law of gravitation. Doc, let’s see how much juice we have left." He glanced at a clock. “We've got an hour to make a decision. It will be almost half an hour before Earth is in sight again."
     “I don’t like the way she cuts off," Corley complained.
     “Quit fretting. I used to have a car that sounded its horn every time I made a left turn." Bowles got a container of coffee, then joined Traub at the starboard port. They peered around the automatic camera and watched the moonscape slide past. “Rugged terrain," Bowles remarked.
     Traub agreed. “There’s better stuff going to waste in California." They continued to stare out. Presently Bowles turned in the air and slithered back to his acceleration couch.
     “Traub!”
     Mannie came to the desk. “Mannie,” Barnes said, pointing at a lunar map, “we figure to land spang in the middle of the Earthside face—that dark spot, Sinus Medii. It's a plain.”
     He was interrupted by Bowles’ voice. “Captain! We are closing, slowly."
     “Are you sure?”
     “Quite sure. Altitude nineteen point three—correction: point two … closing.”
     “Acceleration stations!”
     Barnes was diving toward his couch as he shouted. Traub and Corley followed him. As he strapped down Barnes called out, “Co-pilot—get a contact prediction. All hands, stand by for maneuvers." He studied his own board. He could not doubt it; they were in something less than a perfect circle.
     He was trying to make a prediction from his display when Bowles reported, “I make it contact in nine minutes, Captain, plus or minus a minute.”
     Barnes concentrated. The radar track was jiggling as much as five or ten percent, because of mountains below them; the prediction line was a broad band. As near as he could tell, Bowles was right.
     “What now, Captain?” Bowles went on. “Shall I swing her to blast forward?" A slight nudge would speed up the ship, in effect, lift her, permit her to fall around the Moon rather than curve down.
     It would also waste reaction mass.
     Nine minutes . . . nine hundred miles, about. He tried to figure how many minutes it would be until they raised Earth over the horizon, ahead.
     Seven minutes, possibly—and Earth would be in sight. A landing at Sinus Medii was impossible but they still might land in sight of Earth without using more precious water to correct their orbit. Co-pilot, steady ship for deceleration. Sing out when you see Earth."
     “Aye aye, sir!"
     “There’s Earth!“
     Barnes glanced up, saw Terra pictured in the TV screen, rising behind a wall of mountains. Bowles went on, “Better land, Jim. You'll never get over those mountains."
     Barnes did not argue; their altitude was barely three miles now. He shouted, “Stand by. Red, start swinging as soon as I cut off."
     “Right!”
     “Fire!” He stabbed the button. This maneuver was manual, intended only to stop their forward motion. He watched his ground-speed radar while the ship shivered—nine-tenths. . . seven . . . five. . . four. . . three . . . two . . . one. . . six-hundredths.
     He jerked his finger off just before it dropped to zero and prayed that a mushy cutoff would equal his anticipation.
     He started to shout to Bowles, but the ship was already swinging.
     Earth and the horizon swung up in the TV screen and out of sight.
     For a crawling ten seconds, while they fell straight down, the Luna crept into position for a tail-first landing. They were less than three miles up now. Barnes shifted scale from miles to feet and started his prediction.
     Bowles beat him to an answer. “Contact in seventy-two seconds, Skipper."
     Barnes relaxed. “See the advantage of a Type ‘B’ landing, Doc," he remarked cheerfully. “No hurry—just like an elevator."
     “Quit gabbing and get us down,” Corley answered tautly.
     “Right,” Barnes agreed. “Co-pilot, predict the blast altitude."
     His own hands were busy to the same end.
     Bowles answered, “Jim, you going manual or automatic?”
     “Don’t know yet." Automatic firing was quicker, possibly more certain—but that mushy cutoff could bounce them like a pingpong ball. He steadied cross-hairs on his autopilot display and read the answer: Blast at five two oh feet. What do you get, Red?"
     “Check." Bowles added, “That’s less than three seconds blast, Jim. Better make it automatic."
     “Tend to your knitting."
     “My mistake."
     Nearly forty seconds passed and they had fallen to eleven thousand feet before he decided. “Power plant, set for manual landing. Co-pilot, cover me at five hundred feet.”
     “Jim, that’s too late,” Bowles protested.
     “You will be covering me all of a tenth of a second after I should fire."
     Bowles subsided. Barnes grabbed a glance at the TV screen; the ground under them seemed level and there was no perceptible drift. He looked back at his board. “Correction—cover at five ten.”
     “Five ten—right."
     The seconds clicked past; he had his finger poised over the button when Bowles shouted, “Jim—look at the screen!"
     He looked up—the Luna, still carrying a trifle of drift, was now over a long crack, or rill —and they were about to land in it. Barnes jabbed the button.
     He let up at once; the Luna coughed to silence. The rill, canyon, or crevasse was still in sight but no longer centered. “Co-pilot—new prediction!"
     “What happened?" Corley demanded.
     “Quiet!”
     “Prediction,” Bowles chanted, “blast at…at three nine oh."
     Barnes was adjusting verniers for his own prediction as Bowles reported. “Check,” he answered. “Cover at three seven oh." He threw one glance at the TV screen. The crevasse was toward the edge of the screen; the ground below looked fairly smooth. Unquestionably the ship had a slight drift. All he could do was hope that the gyros would keep them from toppling. “Brace for crash!”
     480—450—400— He jabbed the button.
     The terrible pressure shoved his head back; he lost sight of the altimeter. He caught it again—190—150—125—
     At “fifty" he snatched his finger away and prayed.
     The jet cut off sloppily as always. A grinding jar slammed him more deeply into the cushions. The ship lurched like an unsteady top—and stayed upright.
     Barnes found that he had been holding his breath a long time.

This is from Report of the 90-Day Study on Human Exploration of the Moon and Mars (1989) and from Astronautix.com.

This transport system has two components: the Lunar Transfer Vehicle (LTV) and the Lunar Excursion Vehicle (LEV). The LTV transports crew and cargo between Low Earth Orbit (LEO) and Low Lunar Orbit (LLO), part of the cargo could be a LEV. The LEV transports crew and cargo between LLO and the lunar surface.

Lunar Transfer Vehicle

The LTV is a "one and one-half" stage design, with a reusable core surrounded by expendable propellant tanks. This reduces the propellant load by about 10% compared to a single-stage reusuable vehicle. The core contains the propulsion/avionics module, the main propellant tanks, the aerobraking shield, the crew module (if any), and other assorted subsystems.

The LTV and LEV are boosted into orbit in a single heavy-lift launch vehicle. The LTV will be boosted with the core fully fueled, but the LEV will only be partially fueled due to the payload limit of the launch vehicle. The four fully loaded drop tanks will be boosted into orbit by two subsequent heavy-lift vehicles. The crew and any cargo modules would be boosted by the space shuttle.

Some in orbit assembly will be required: adding drop tanks to LTV core, the eight peripheral aerobrake segments attached to the LTV aerobrake shield core, and the cargo modules added to the LEV.

The LTV does a trans-lunar injection burn, and jettisons two empty drop tanks. It brakes into LLO and drops the two remaining empty drop tanks. It then acts as a staging base in LLO for the LEV.

If there is already an empty LEV in LLO parking orbit waiting to be reused, the LTV loads it with propellant, consumables, and attaches new cargo modules.

When the LEV has performed its mission, the LTV does a trans-Earth burn using the core propellant tanks.

It circularizes itself into LEO using aerobraking instead of propellant, at a considerable savings in initial mass required in LEO at mission start. After each mission the aerobrake shield is refurbished and verified at the International Space Station. The aerobrake shield can be reused for five missions.

The optional LTV crew module provides habitable support for the crew for the 4 day translunar trip and up to 7 days for the return to the space station. Naturally the crew can override the automatic rendezvous and docking system. Crew module obtains electricity from the LTV, has a two-gas open-loop environmental control and life support system, has a galley, zero-gravity toilet, and a personnel hygiene station.

The crew module has docking ports fore and aft, passing through one is an intravehicular activity (no space suit required). There is no airlock, so extravehicular activity requires all the crew to don space suits and depressure the entire module. There is enough repressurization gas carried for 2 EVAs.

The crew module carries a storm cellar with walls filled with water radiation shielding. The water is vented before aerobraking to save wear and tear on the aerobrake shield.

Lunar Excursion Vehicle

In "reusable" mode, the LEV can transport 15 metric tons of payload to the lunar surface (along with a crew and crew module), and return to LLO. It can be reused up to five missions.

In "expendable" mode the LEV can transport 33 metric tons of payload to the lunar surface (with neither crew nor crew module) and stays on the surface forever after.

If the cargo load is small enough, an unmanned LEV with no crew module has enough of an automatic pilot to be able to land, discharge cargo, return to orbit, and rendezvous with the orbiting LTV.

The LEV and LTV shares a lot of systems designs to reduce development and testing time (such as engines, cryogenic RCS, avionics, software, communiciation equipment, fuel cells, etc.).

When the LEV is parked in lunar orbit and abandoned, it is powered by solar arrays. On the lunar surface, the propellant system is designed for 30 days. For longer stays it will require surface support (from in-situ resource utilization).

The LEV's crew module is related to the LTV crew module, but with some differences. It has no storm cellar. It transports four crew members between the LTV and the lunar surface. During landing operations two crew members have landing control panels and windows, the other two are in shock webing and just have to be patient and stare at the walls.

The LEV's crew module's systems are in a quiescent state, except for 4 days during descent/ascent missions (2 days during descent and initial surface operations, 2 days for preparation and ascent to orbit). While quiescent the crew module has no interal power, thermal control, or propellant conditioning. Bottom line is either the descent/ascent mission only lasts 4 days, or there has to be support systems available on the lunar surface (a lunar base in other words).

Just like the LTV crew module, the LEV crew module has no airlock and only enough repressurization gas for 2 EVAs.

This is from Lunar Expedition Plan from the headquarters Space Systems Devision of US Air Force Systems Command (1961).

Please do not confuse this with LUNOX which has a similar name, that is about using lunar mined oxygen with LANTR nuclear thermal rockets.

This was a US Air Force project started in 1958 designed to seize the "high ground" of space, about three years before the start of NASA's Apollo program. The idea was to ensure that the first lunar landing was made by the US (showing those pesky Soviets who is top dog in space), establish a 21-crew lunar USAF base, and use the base like a spy satellite to keep an eye on US enemies. And maybe even site a few nuclear missiles at the lunar base.

The project was cancelled in 1961, mostly due to the USAF's increasing preoccupation with how the Viet Nam War was devouring their budget. Contributing factors were that it would cost x1.5 the Apollo program, as the study progressed it became obvious the scheduled time-line was incredibly overoptimistic, President Kennedy wanted the moon race to be a civilian effort not a military one, and only an idiot would put a spy satellite 400,000 kilometers away from Terra's surface on Luna when the satellite would have a much better view from a 200 kilometer Low Earth Orbit. The only thing more stupid is siteing nuclear missiles at a lunar base, since it provides zero advantage over siteing them on Terra but does inflict lots of severe handicaps.

So that the project wasn't a total waste, the USAF released a partially declassified report in order to disconcert the Soviets. Look what we were planning, and you didn't even know. Makes you wonder what else we are up to, eh Khrushchev? You better watch your step.

If the project had actually created a real rocket, the advantage was that the US would have the technology for a great launch vehicle very similar to the Space Shuttle. A booster with a LOX-LH₂ core, huge strap-on solid rocket boosters, and a flying re-entry vehicle. This would be a great set of technologies for future space programs. Instead we went with the Apollo program, which gave us the Saturn V rocket which will never be used again, the Lunar Module which isn't good for much except allowing three crew and a minuscule payload to visit the lunar surface, and the slightly less useless Apollo command/service module.

Having said that, the LUNEX design had some severe problems.

Mainly because it was using a direct ascent design. This is when you boost the lunar spaceship into orbit as one piece, land the entire clanking mess on Luna, then blast-off the entire clanking mess (less the landing rockets) directly into a return to Terra.

The point is that every gram counts. Particularly during landing and lift-off. Each extra gram is going to require tons more fuel. This is why the Apollo mission didn't send down to the surface the command/service module along with the lunar module, because it wasn't needed on the moon. Only an idiot would design a mission like that, or a designer terrified of having the astronauts being forced to perform an orbital rendezvous with the return spacecraft.

The fuel required for LUNEX's direct ascent was so outrageous that the spacecraft could not carry enough. The program would have to use an unmanned cargo rocket to land 20 tonnes of fuel first. Once they were in place, LUNEX could land next to the cargo rocket and refuel. Or accidentally land too far from the cargo rocket and inflict a public relations nightmare on the USAF as the stranded astronauts waited for their oxygen to run out. The cargo rocket also carries a landing beacon to provide terminal guidance for the crewed spacecraft. The cargo rocket is basically just the LUNEX's landing stage with the fuel tank cargo perched on top, it is missing the launch stage and the reentry glider.

Assuming everything went according to plan, the spacecraft would return to Terra. Shortly before arriving all parts of the spacecraft would be jettisoned except for the hypersonic lifting body. This would have to enter Terra's atmosphere at a blistering 11.3 kilometers per second, aerobrake while hoping they run out of excess velocity before they run out of heat shield, and land on a aircraft runway at Edwards Air Force Base. This would be a huge engineering task since basic data on such hypersonic reentry had not been explored yet. Did I mention already how incredibly overoptimistic the development time-line was?

At the start of the mission, the booster does not send the spacecraft in Terra orbit. No, it is boosted directly into trans-Lunar trajectory (the same way as the Apollo missions). The trip to Luna will take about and one-half days. The spacecraft will arrive pretty much still with a mass of 61,000 kilograms. All of which will have to be delta-Veed about 2,737 meters per second to land the entire contraption next to the uncrewed cargo ship, since the landing legs can only handle about 6 m/s of jolt.

The crew will spend five days exploring, then lift-off directly into their Terra return trajectory.

When aerobraking, strict accuracy is required. The re-entry angle must be within ± 2° of optimal. Too steep and the crew will be flattened by the gee forces while being incinerated by the heat. Too shallow and they will ricochet off the atmosphere into an eccentric orbit. By the time they can make another attempt, the crew will run out of air and/or suffer a lethal radiation dose while cruising the Van Allen belt.

The 2.5 day trip to Luna and 2.5 day trip back home was selected as optimum. Longer flights would have more problems with life-support and guidance. Shorter flights would need excessive amounts of fuel.

These two designs are from The Resources of the Solar System by Dr. R. C. Parkinson (Spaceflight, 17, p.124 (1975)). The Lighter ferries tanks of liquid hydrogen from an electrolyzing station on Callisto into orbit where waits the Tanker. Once the Tanker has a full load of tanks it transports them to LEO. All the ships are drones or robot controlled, there are no humans aboard. The paper makes a good case that shipping hydrogen from Callisto to LEO would eventually be more economically effective than shipping from the surface of Terra to LEO, with the break-even point occurring at 7.8 years. Please note that this study was done in 1975, before the Lunar polar ice was discovered, and probably before the ice of Deimos was suspected.

Warning: most of the figures in the table are my extrapolations from the scanty data in the report. Figures in yellow are sort of in the report. Use at your own risk.

The tanker uses a freaking open-cycle gas-core nuclear thermal rocket. This is an incredibly powerful true atomic rocket, but it is only fractionally more environmentally safe that an Orion nuclear bomb rocket. The report says it should be possible to design it so the amount of deadly fissioning uranium escaping out the exhaust is kept down to as low as one part per 350 of the propellant flow (about 300 grams per second), but I'll believe it when I see it. Since it is used only in deep space we can allow it, this time. The report gives it an exhaust velocity of 35,000 m/s, which is about midway to the theoretical maximum.

The lighter can get by with a more conventional hydrogen-oxygen chemical rocket. It will need an acceleration greater than Callisto's surface gravity of 1.235 m/s², for safety make it 1.5x the surface gravity, or about 1.9 m/s (0.6g).

The four major Galilean moons are within Jupiter's lethal radiation belt, except for Callisto. The black monolith from 2010 The Year We Make Contact only told us puny humans to stay away from Europa, so Callisto is allowed. If you want ice that isn't radioactive, you've come to the right place. It is almost 50% ice, and remember this is a moon the size of planet Mercury. That's enough ice to supply propellant to the rest of the solar system for the next million years or so. Europa has more, but it is so deep in the radiation belt it glows blue. Callisto is also conveniently positioned for a gravitational sling shot maneuver around Jupiter to reduce the delta-V required for the return trip to Terra.

The report says that the requirements for an economically exploitable resource are:

1. It is not available in the Terra-Luna system
2. It must provide more of it than the mass originally required to be assembled in Terra orbit at the outset of the expedition
3. It must be done within a reasonably short time (the break-even time)

Hydrogen fits [1], or at least it did until the Lunar ice was discovered. [2] and [3] depend upon the performance of the vehicle.

There are three parts. First is the Tanker, which is an orbit-to-orbit spacecraft to transport the hydrogen back to LEO and brings the expedition to Callisto in the first place. Next is an electrolysis plant capable of mining ice, melting it into water, cracking it into oxygen and hydrogen, and liquefying the hydrogen. Last is a Lighter which is an airless lander that ferries liquid hydrogen from the plant on Callisto to the orbiting Tanker.

The report decided to use modular cryogenic hydrogen tanks that would fit in the Space Shuttle's cargo bay. They would have to be about 18.3 meters x 4.57 meters, about 300 cubic meters capacity. The report has a filled tank massing at 26,000 kg, with 22,000 kg being liquid hydrogen and 4,000 kg being tank structural mass. Examining the drawing of the tanker, the front cluster is composed of four tanks while the rear has nine, for a total of thirteen. The tanker will have a length of two tanks plus the length of the rocket engine, 37 meters plus rocket. The rear has tanks arranged in a triangular array about four tanks high. So a diameter roughly 18 meters or so.

The lighter carries a single tank, so it is roughly one tank in diameter, and one tank long plus the fuel tanks+engine length. It will need a large enough liquid hydrogen/liquid oxygen chemical fuel capacity to lift off from Callisto to the tanker and land back on Callisto.

The report figures that the electrolysis plant can produce hydrogen for about 39 kW-h/kg, that is, each kilogram of hydrogen in the plant requires 39 kilowatt-hours. Figure it needs more electricity to liquefy the hydrogen, and more to produce the liquid oxygen needed by the lighter, for a total cost of 50 kW-hr/kg for liquid hydrogen delivered to the orbiting tanker. So a 2 megawatt nuclear reactor could produce 350 metric tons of hydrogen per year. Launch windows back to Terra occur every 398.9 days.

Once the lighter has made enough trip to fully load the tanker, the tanker departs for LEO. It will use some of the hydrogen for propellant, some will be the payload off-loaded at LEO, and enough will be left to return the tanker to Callisto. The amount of payload is specified to have a mass equal to 37% of the fully loaded mass of the tanker. It also specifies that the inert mass fraction of the tanker is 25% of the tankers fully loaded mass.

The report had an esoteric equation that calculated the mass of the lighter and electrolysis plant as a percentage of the tanker mass in order to be economically viable. It turns out to be 13% of the fully loaded mass of the tanker. When the expedition is launched the tanker will carry the lighter, the electrolysis plant, and enough propellant so that the total mass is 52.9% of the fully loaded mass (i.e., it departs half empty). The lighter will have its tanks full.

Five years later, upon arrival at Callisto, the lighter lands the electrolysis plant on a prime patch of ice. It then starts the cycle: patiently waiting for the plant to fill the payload tank and the fuel tanks, boost the payload to the tanker, then land back at the plant to start again.

In context: this was one of a series or articles I wrote for Spaceflight at the encouragement of the then editor, Ken Gatland, triggered off in the dark days following abandonment of the Apollo programme by a discussion at the BIS as to what would be needed to make spaceflight self-supporting. The first article was published in Spaceflight 1974 p.322 under the title "Take-Off Point for a Lunar Colony." There was then a second on "The Colonization of Space" (S/F 1975 p.88) and a couple of subsequent ones on Lunar Colonies (S/F 1977 p.42/103). Later, when they invented the first spreadsheets, I did some speculation on how the economics of everything might fit together economically in a big input-output model which got published as "The Space Economy of 2050 AD" in JBIS v.44, p.111 (1991) which also appeared in my book Citizens of the Sky (1989) later. It is unlikely that I was consistent through all of this — my opinions develop with time — and by the 1991 period I was heavily in to the economics or reusable launchers and what would happen if the models were pushed to very high flight rates.

Going back to "The Resources of the Solar System", I'm not sure how much detail I managed at the time. I remember that there were a couple of things influencing me at the time. One was the concept of a gas core nuclear engine (GCR) which might have a specific impulse of about 3500 sec (35 km/s). To really move around the Solar System you need a high thrust-to-mass engine with this sort of specific impulse, and GCR had the interesting property of using hydrogen as propellant. (Ion motors can meet the specific impulse, but to do a similar job would require a power-mass ratio several orders better than anything we could consider then or even today — VASIMIR suffers the same problem). Nowadays I might put my money more on a pulsed-fusion system (see “Using Daedalus for Local Transport,” JBIS, 62, p. 422-426 (2009)) — note NOT using helium-3, which would change the model significantly.

The second thing at the time that influenced me — at a time when the Space Shuttle was still a paper vehicle — was that the Space Shuttle payload bay was just about the right size (15 ft × 60 ft) to carry a full liquid hydrogen tank (there are reasons now why it wouldn't which led to the abandonment of design work using Centaur as an upper stage) — so my modular design was based around using that as a standard tank. For use in long duration space missions the tank would have to have some sort of active cooling system to keep the hydrogen from boiling away, but given that you could then ship LH₂ around the Solar System on slow, economical trajectories like modern oil tankers on Earth. Once you have rerfuelling stations at either end interplanetary flight becomes a lot easier and you can think of using higher speed trajectories for special cargo like human beings.

All the other figures in the table are ones I've extrapolated from the few figures given in the report.

A plausible figure for nuclear power generation is 0.12 Megawatts per ton of generator. This would make the electrolysis 2 MW power reactor have a mass of 16,000. This is close to the 25,000 kg mass of a payload tank. So to simplify, assume the electrolysis rig with liquefaction gear and all masses a total of 25,000. This also ensures that the lighter is capable of landing it.

The tanker's inert mass fraction is 25%, and hydrogen payload is 37%. This means the dry mass is 62%, which means the mass ratio is 1.61. With an exhaust velocity of 35,000 m/s, this yields a total delta-V of 16,730 m/s. I am unsure if this is enough for a Callisto orbit-LEO mission followed by a LEO-Callisto orbit mission. Not without a heck of a gravitational sling-shot it isn't. Or I could have made a mistake in math.

Note both the payload and the propellant is hydrogen, stored in the same array of tanks. If the inert mass fraction is 25%, then the payload+propellant mass fraction is 75%. If there are 13 tanks each of 25,000 kg, then the total is 325,000 kg. If this is 75% of the wet mass, the actual wet mass is 433,000 kg. If the payload is 37% of the wet mass, it is 160,000 kg. If a hydrogen tank is 87% hydrogen and 13% tankage, the amount of hydrogen payload is 139,000 kg.

On the initial trip, the tanker carries the electrolysis plant and the lighter (with no payload, but with full fuel tanks). This is 13% of the wet mass or 56,300 kg. If the electrolysis plant is 25,000 kg, the lighter (with no payload) must be 31,290 kg. The lighter payload is one payload tank at 25,000 kg. So the lighter wet mass is 56,290 kg.

The lighter needs a delta-V of 3,414 m/s (Callisto-surface-to-orbit + orbit-to-Callisto-surface). Chemical fuel has exhaust velocity of 4,410 m/s. This means the mass ratio has to be 2.17. This implies the dry mass is 25,898 kg. Subtract the 25,000 kg payload, and there is 898 kg for the structure and the engine. Seems a little flimsy to me, perhaps 25,000 kg is a bit to generous for the payload tank.

Tank is scaled to fit in Space Shuttle cargo bay. At least the the proposed size of the bay in 1975 when the report was written, it was later reduced in size.

• 4.55 m wide × 18.2 m long
• Mass 26 metric tons
• LH₂ Mass 22 metric tons

Notes, from left to right:

1. Docking Ring
2. Limited amount of pressurization equipment round head end, also radar transponder
3. Strong ring with attachment points
4. Recirculating and pressurization pipes
5. Strong ring with attachment points
6. Docking Ring
7. Fill valve connect. Associated propellant management equipment
8. Radiator

This is from a student project for Master of Architecture candidates at the University of Houston: Manned Mars Explorer project: Guidelines for a manned mission to the vicinity of Mars using Phobos as a staging outpost.

After weighting all the options, chemical propulsion was chosen. Nuclear electric had too many drawbacks.

The tyranny of the rocket equation led them to go with reliability over redundancy. Equipping the spacecraft with back-up units for all critical systems cuts too much into payload mass. Instead they went with single units that were super-duper fault tolerant.

Medical issues dictated supplying the crew with a full one-Terran-gravity. An elaborate bola system was designed. The system resists twisting via a unique spreader system and four tether configuration. Spin grav is used for the trans-Mars coast and the trans-Terra coast. The tether is reeled in before each propulsive manuever to prevent the spacecraft from destroying itself by the crack-the-whip effect.

THE MISSION

An opposition class Venus inbound swingby was used for the trajectory. About 300 days are spent travelling to Mars. It spends 60 days in Mars orbit. It uses the Venus inbound swingby leg to travel to Terra LEO which takes 210 days. This trajectory was chosen due to relatively short overall mission and Mars stay time. It does however require more delta V than conjunction class trajectories.

The sixty day exploration period is mostly focused on Phobos and Deimos, but there is a segment where a crew of three is sent to the surface of Mars for seven days. Staging bases will be set up on the moons, and they will be assesed for deposits of water ice and other valuable in-situ resource utilization goodies.

MPV is the Manned Planetary Vehicle, the spacecraft. CCV is the Crew Command Vehicle, a small auxiliary spacecraft carried as payload.

Mission Phases:

1. Low Earth Orbit construction
   1. Vehicle assembly
   2. Crew training
2. Trans-Mars injection
   1. Propulsive maneuver
   2. Communication satellite deployment
   3. Spin-up
   4. Power system deployment
   5. Tether system deployment
   6. Trans-Mars coast
   7. De-spin
   8. Power system retrieval
   9. Tether system retrieval
   10. Communication satellite retrieval
3. Mars circularization
   1. Propulsive maneuver
   2. CCV surface operations
   3. CCV return to MPV
4. Trans-Earth injection
   1. Propulsive maneuver
   2. Communication satellite deployment
   3. Spin-up
   4. Power system deployment
   5. Tether system deployment
   6. Trans-Earth coast
   7. De-spin
   8. Power system retrieval
   9. Tether system retrieval
   10. Communication satellite retrieval
5. Earth orbit capture
   1. Propulsion stage, CCV, and MPV separation
   2. Propulsion stage remains in hyperbolic orbit
   3. CCV propulsively circularizes at LEO with crew
   4. MPV aerobrakes into Space Operations Center orbit

MANNED PLANETARY VEHICLE (MPV)

The mission spacecraft is composed of the following components:

• Pressurized Environment System: the habitat module
• Power System: provides electricity
• Structural System: the ship's spine
• Folding Aerobrake System
• Four-Tether System: provides 1 gee of spin gravity
• Staged Propulsion System

It carries the following payload:

• Crew Command Vehicle (CCV): transports explorers to Phobos, Deimos, and Martian Surface
• Communication Satellite

Pressurized Environment System

This is the MPV's habitat module. It is composed of a habitation module, a laboratory module, a safe haven (storm cellar) and connecting tunnels. The large modules will have three airlock section, each with two means of egress: one to another pressurized airlock section, and the other to either the exterior or another airlock section. Space suits will be stored next to each exterior egress for use in a planned EVA or emergency escape to the CCV. The storm cellar will accomodate the 6 member crew for 12 days. The anti-radiation walls are 10 cm thick aluminum.

The mass budget for the system is 98,000 kg.

Power System

The power system mass is budgeted at 20,700 kg. It is specified to provide 150 kW constant power. It uses a set of solar dynamic power systems rated at 128 kg/kW. This system cannot be used during propulsive maneuvers or planetary eclipses. During these periods power is supplied by fuel cells.

Structural System