Thursday, July 19, 2012

Rough-Out Mars Mission with Artificial Gravity


The purpose of this particular study was to show how easy it is to supply artificial gravity to the crew of any reasonable candidate design for a manned Mars mission. No gigantic “Battlestar Galactica” or complicated Rube Goldberg cable-connected designs are necessary.

All you need is a radius exceeding 56 m from the center of gravity of a long, slender orbit transport to the habitat at one end. The gee level achievable this way is 1 full gee at 56 m for a “fuzzy” maximum rotation rate speed limit of 4 rpm. Longer radii allow lower, even more tolerable, rotation speeds.

Mission Architecture and Vehicles

The mission plan outlined here uses two vehicles sent by Hohmann transfer from low Earth orbit (LEO) to low Mars orbit (LMO). One is the manned vehicle, spun for artificial gravity, the other is an unmanned transfer of the return propellant supply for the manned ship. Both are propelled by solid-core nuclear thermal rockets, using liquid hydrogen propellant.

Two mass allowances for otherwise-undefined landers are included in the manned ship. The landers, empty tanks, and one supply storage module (loaded with wastes) are left in LMO. The manned ship is much smaller in mass for the return to LEO, but still about the same overall length. Modular vehicle design is what allows this mission architecture (a key point with this design study).

Why Do It This Way Instead of Direct?

There are two excellent features of plans like this: (1) staging from LMO allows more than one landing to be made during the one trip, and (2) returning much of the manned vehicle to LEO allows re-use of major components for other missions, components that do not need to be launched again, only re-supplied. Both of these dramatically increase the usable results obtainable from the one mission.

This particular mission design also has two serious downsides. One is that if rendezvous fails in LMO, the crew dies. Their return propellant supply must be rendezvoused-with in LMO, there is no other option. I would not seriously propose this particular plan for actual implementation, precisely because of that lack of “a way out”, even though that event is very likely low-probability.

However, this mission design study is a “representative plan” for illustrating the incorporation of the artificial gravity that we already know the crew will need.

The other downside is lack of a crew return vehicle to be used for emergency free return, should the final return-to-LEO burn fail. Being a multi-engine propulsion module, that event is also low probability, but there is no other option for reaching LEO or Earth. If that event should occur, again the crew is lost. So, for that reason as well, I would not propose this particular architecture as a serious plan. Again, it merely provides representative vehicles for examining the artificial gravity issue.

There is one other mild “downside” to this particular architecture: there must be propulsion on the landers, but also on the unmanned supply vehicle. There would be fewer” dead-head” tons to ship, if the lander propulsion could also be used to push part of the propellant supply to Mars. However, the design criteria for lander propulsion could well be incompatible with the propulsion needs for transit from LEO to LMO. With an undefined lander design, that option could not be utilized in this study.

Velocity Requirements

These were assumed to be the same as from my earlier studies (reference 1): 8 km/sec one-way, orbit-to-orbit, LEO to LMO, and the same again to return. This requirement would include escape and capture burns at average planetary orbital parameters, with crude allowances for plane changes, and for propellant boil-off losses. These are representative, and a bit conservative, and so are not “accurate” for any given scenario.

Crew Habitat Modules

My design requirements for habitat open space per person, and consumables per person, were just “good guesses”. I tend to use larger volumes of free space per person than other designers. This is based on my own living room’s volume at home, and the perception that I could not share that volume with more than one other person for around 2-3 years of forced confinement. Prisoners in most state prisons share smaller spaces, but psychologically, they would not make very good astronauts, even if they were not criminals. Some designers might find my habitats overly-spacious, but I don’t think they are. This is a “fuzzy choice” that must be made, that seriously impacts the sized of vehicle designs.

See Figure 1 below for the two-module habitat I came up with. By my criteria, this is big enough for a crew of 6, and would contain a large common room, exercise facilities, sleeping cubicles, the flight control station, “ from-orbit” science stations, and a galley area. The water and wastewater tanks should be arranged around the flight control station, so that it can also serve as the temporary solar flare radiation shelter for the entire crew. The dimensions and masses of the modules as I estimated them are compatible with single launch by Atlas-V-552, or both-at-once by Falcon-Heavy.

Engine Module

I assumed data from the final tests of the old “NERVA” nuclear engine from 1972-1973, just as Project Rover shut down. This was the engine that was essentially flight-qualified, but never actually got to fly. There are better designs, but none of those got the testing required to be flight qualified. Data are shown for a multi-engine module in Figure 2 below, to take advantage of multi-engine reliability. This could be 2, 3, 4, or even 5 engines. Module sizing requires at least 0.05 gees vehicle acceleration in order to be impulsive enough not to suffer the velocity losses of long-burn / low-thrust schemes.

This module sized-out at not quite 12 metric tons. That would be compatible with an Atlas-V launcher, or a Falcon-9 (by Spacex’s 2012 website data), if dimensions can be resolved for a practical partial shroud size. Otherwise, the individual engines and components could be launched separately and assembled in LEO. Two of these modules are necessary: one for the manned vehicle, the other for the unmanned vehicle.

The spreadsheet inputs for velocity requirements, engine module, and habitat modules, are imaged in Figure 3 below. User inputs are highlighted yellow. User interventions-to-converge are blue.

Storage Supply Modules

The basic masses and dimensions of the two storage supply modules are given in Figure 4 below. These two can support a 27-month round trip with consumables masses that seem realistic, even by some of the online mission resource calculators that now exist (see reference 2). The packaging density is far higher than the habitat modules, as there is almost no free volume inside. This affects their shape.

The excess consumables mass that I used (relative to the online calculator results) should “cover” the use of bulky, heavy frozen food supplies. These are required for a 27 month round trip, since the conventional non-frozen dehydrated or sealed astronaut foods cannot last the mission time without decomposing. Those are typically “gone” after about a year to 18 months.

Propellant Modules

Storage of liquid hydrogen over 2-3 year timelines is an engineering issue that is still unresolved, although it could be resolved very quickly. I am assuming these propellant tank models include a dewar-as-inner-tank, external insulation that includes foam/foil layered meteoroid armor, a deployable sunshade, and deployable solar power wings to power cryo-cooler equipment. Accordingly, the volumetric load-out factor is lower, and the inert weight factor higher, than most folks in the spacecraft business would assume. These are “good guesses” only, to cover the extra equipment.

The mass and dimensional data are given in Figure 5 below. It is presumed that one simply stacks up enough of these modules to achieve the velocity requirement. Uniquely, that stack of modules may be multiple linear arrangements tied together in parallel, in order to achieve the necessary form factor without being excessively long or short. The modules are deliberately sized to launch one at a time on an Atlas-V-552 configuration, or two-at-once on a Falcon-Heavy.

The inputs to the spreadsheet regarding the storage supply modules and the propellant tank models are imaged in Figure 6 below. User inputs are highlighted yellow. User interventions-to-converge are blue.

Manned Transport, Outbound Configuration (to Mars)

Sizing this vehicle to an acceleration requirement set the mass of the engine module and the total engine thrust. This was an iterative process. The as-sized vehicle is in the 840 metric ton class, at ignition departing from LEO. It departs at 0.05 gees initially, which is impulsive enough to avoid low-thrust losses. It is depicted in Figure 7 below.

This vehicle includes two undefined Mars landers, in order to make two separate landings at different sites during the one mission trip. These are included as mass allowances only, depicted schematically in Figure 8 below. Each lander is 60 metric tons, assembled from two items docked in LEO that are in the 30-ton class. These components probably require Falcon-Heavy launchers.

Figure 9 below is a spreadsheet image that gives the as-sized details of masses and module count for the manned transport in this outbound configuration. There are a total of 23 propellant tank modules, arranged as a triple stack 7 units long plus a double stack one unit long. The habitat and storage modules are a linear string at one end, with the engine module at the other.

Vehicle center of mass is near the middle of the overall length. The radius from there to the middle of the habitat is long enough to provide one full gee artificial gravity at about 3 rpm, spun “head-over-heels”. This is a rigid slender baton, not too slender to be structurally sound. This is stable dynamically, and easy to de-spin for maneuvers, and re-spin afterwards.

This demonstrates just how easy it is to incorporate one-full-gee artificial gravity into a design not otherwise overtly driven by artificial gravity concerns. It just can naturally happen. No giant structures, no space trusses, no spinning crosses, no cable-connected modules. Just a rigid slender baton-like structure spinning end-over-end. Spin-up and de-spin are just thruster firings, anytime needed.

Manned Vehicle Return Configuration (to LEO)

For the return, half the supplies are already consumed, so one storage module can be left at Mars in LMO, containing wastes not otherwise disposed-of. The landers are left in LMO, as is the entire outbound propellant tank cluster, now empty. The two habitat modules, one storage module, the return propellant tank stack of 9 modules, and a NERVA module comprise the return vehicle configuration, depicted in Figure 10 below. (The other NERVA module and a cluster of empty tanks from the unmanned vehicle also remains in LMO.) The spreadsheet data corresponding to this configuration are imaged in Figure 11 below. Note that the NERVA engine module, operated at the same design thrust as outbound, now provides 0.13 gee acceleration at ignition in LMO.

This configuration is a simple linear stack, no parallel clustering. It is far less massive (330 ton class), but about the same overall length as the manned outbound vehicle, and also just about the same radius from center-of-mass to middle of the habitat. Again, one full gee of artificial gravity is available at about a 3 rpm spin rate, “head-over-heels”. This demonstrates again just how easy it is to incorporate one gee artificial gravity into a representative very-practical design, without driving that design to extremes by the artificial gravity requirement. The modular approach provides the form-factor tailorability that is absolutely necessary to achieve this result.

Unmanned One-Way Propellant Supply Vehicle

The return from Mars depends upon 9 propellant modules sent there separately. Those modules, plus a small guidance and control unit, plus another of the same NERVA engine modules plus a propellant tank cluster for the trip, comprise the one-way propellant supply vehicle. Being unmanned, this vehicle does not spin for artificial gravity. It is depicted in Figure 12 below. The as-sized mass and module numbers are in the spreadsheet image of Figure 13 below.

There are 9 propellant modules that are the payload. It takes another 19 such modules to send this payload from LEO to LMO, again with the same NERVA module at the same design thrust. That vehicle is arranged as a 10-module center stick, surrounded by three 6-module sticks spaced equally circumferentially. The result is a vehicle about 160 m long, in the 700 ton class at ignition in LEO. Initial acceleration is 0.06 gees. Other stack configurations are possible, of course.

Launcher Count and ROM Program Costs

Figure 14 below provides an image from the spreadsheet of the module count and two options for the corresponding launch rocket requirements. Launches for transferring the crew to and from the vehicle in LEO are included (either as something 6-man-caopable on an Atlas-V, or as a Falcon-9/Dragon).

The two options maximize either United Launch Alliance (ULA) launcher usage or Spacex launcher usage. See reference 3 for data regarding launch prices. The “truth” lies in a mix of launchers in between these extremes, so that the ultimate rough-order-of-magnitude (ROM) program costs reported here are the average of the two numbers computed for the two options. Launch costs were arbitrarily assumed to be 20% of total program outlays. That would be realistic for a program managed by a lean focused agency, and conducted by lean, focused contractors, with little or no new technology development, only implementation-design efforts.

Not only are the ROM estimates of total program cost in the figure informative, but also the costs per landing site visited, and per person sent. These are good “bang-for-the-buck” measures. Any way one looks at these numbers, they are startlingly lower than the preconceptions or preferred beliefs of most folks. They are radically lower than most of the Mars mission proposals coming from NASA in the last couple of decades. This is a remarkable result, considering that this is not a “minimalist” mission design. We are looking at two vehicles in the 800-900 metric ton class, as assembled in LEO. We are looking at sending a crew of 6 to Mars here.


The mission architecture and vehicle designs of this study are not the architecture and designs that should actually fly. But, they are similar enough to what should really fly, that the issue of artificial gravity can be realistically investigated.

Very important: It is the modular spacecraft design approach that allowed tailoring vehicle form-factor at any given mass to obtain a convenient radius length, for easily designing-in artificial gravity at 1 full gee, by simple end-over-end spin of a “slender baton”.

Using a mix of ULA and Spacex launchers, direct launch costs are estimated somewhere near $6.4 B, based on current retail launch prices.

Overall program outlays were ROM-estimated from the assumption that launch costs are around 20% of total program outlay. Thus the entire mission cost is near $32 B, which is far lower than the $400-500 B numbers that Congress saw from NASA recently.

“Bang-for-the-buck”: program ROM costs are near $16 B per site visited, and near $5.8B per man sent to Mars. That last is about the same as those claimed for the minimalist mission designs, but the return here is far larger (in terms of sites visited and persons sent). More sites and bigger crews always mean that more and better information can be obtained and brought home.

It is the LMO basing that allows multiple landings, which in turn produces more return for the cost. Therefore, this should be the preferred architecture for exploration-type missions. (Establishing bases or experimental stations on the surface is a different type of mission entirely, although an exploration mission can also do some of those things.)

There is a mission architecture and vehicle design opportunity that I could not take advantage of in this study, with its undefined lander. That would be to use the lander propulsion as the transfer propulsion for sending some assets to Mars. That would be more like the designs outlined in reference 1.

This Study Achieved Its Overall Purpose vs the Artificial Gravity Issue

Achieving 1 full gee of artificial gravity need not drive vehicle designs into gigantic, complicated or unrealistic forms, nor need it drive launch and program costs up. The key is form-factor adjustment for a given mass, as enabled by modular vehicle design. The preferred form-factor is the “slender baton”, for end-over-end spin at low, tolerable rates (under about 4 rpm maximum).


1. Updated version of paper presented at 14th Mars Society convention in Dallas, August 2011. See, article dated 9-6-11, titled “Mars Mission Second Thoughts, Illustrated”.

2. On-line supplies calculator application: see

3. Launch cost data plots posted in:, article dated 5-26-12, titled “Revised, Expanded Launch Cost Data”.

Figure 1 – Crew Habitat Module, One of Two Total for a Crew of 6

Figure 2 – Multi-NERVA Engine Module, and the Guidance and Control Package

Figure 3 – Spreadsheet Inputs for Velocity Requirements, Engine Module, and Habitat Module

Figure 4 – Storage Supply Module, One of Two Required

Figure 5 – Data for Common Propellant Module

Figure 6 – Spreadsheet Inputs Image for Storage and Propellant Tank Modules

Figure 7 – Manned Transport in Outbound Configuration (to Mars)

Figure 8 – “Dummy” Mass Allowances for Undefined Mars Lander Vehicle, 2 Required

Figure 9 – Spreadsheet Image for As-Sized Manned Transport, Outbound Configuration

Figure 10 – Manned Transport in Return-to-LEO Configuration

Figure 11 – Spreadsheet Results for Manned Transport, Return-to-LEO Configuration

Figure 12 – Unmanned One-Way Return-Propellant Delivery Vehicle

Figure 13 – Image of Spreadsheet Data for Unmanned Return-Propellant Vehicle

Figure 14 – Launcher Counts and Costs, and ROM Program Costs

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