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The Use of SpaceX Hardware to Accomplish Near-Term Human Mars Mission

posted May 16, 2011, 6:50 AM by Michael Stoltz   [ updated May 20, 2011, 11:12 AM ]

Robert Zubrin, Pioneer Astronautics, 05.15.11

The recent announcement by the entrepreneurial Space Exploration Technologies Corp. (SpaceX) that it intends to field within two years a heavy lift rocket capable of delivering more than twice the payload of any booster now flying poses a thrilling question: Can we reach Mars in this decade?

I believe the answer is yes. In this paper, I will lay out a plan to make use of the soon-to-be-available SpaceX systems to accomplish near-term human Mars exploration with minimal technology development. First, I will layout a baseline mission architecture and plan. In the next section, I will discuss various technology alternatives available within the selected mission architecture. Then, in the following section, I will discuss alternative mission architectures. I will then conclude with some overall observations bearing on the question of sustained exploration and settlement of Mars.

It may be noted that the author is not an employee of the SpaceX company, and does not have detailed knowledge of the SpaceX systems. It will take the hard work and ingenuity of the SpaceX engineers to develop configurations and systems that can make these ideas a reality. Nevertheless, it is apparent that if an approach such as that recommended here is adopted, the requirements and capabilities numbers can be made to converge. We can reach Mars in our time.

            1.         Baseline Mission Plan
 
Here’s how it could be done. The SpaceX Falcon Heavy will have a launch capacity of 53 metric tons to low Earth orbit. This means that if a conventional hydrogen-oxygen chemical rocket upper stage were added, it could have the capability of sending about 17.5 tons on a trajectory to Mars, placing 14 tons in Mars orbit, or landing 11 tons on the Martian surface. The same company has also developed and is in the process of demonstrating a crew capsule, known as the Dragon, which has a mass of about 8 tons. While its current intended mission is to ferry up to 7 astronauts to the International Space Station, the Dragon’s heat shield system is overdesigned, and is capable of withstanding reentry not just from Earth orbit, but from interplanetary trajectories. It’s rather small for an interplanetary spaceship, but it is designed for multiyear life, and if we cut its crew from 7 to 2, it should be spacious enough for a pair of astronauts who have the right stuff.

Using these basic tools, a Mars mission could be done utilizing three Falcon Heavy launches. One would deliver to Mars orbit an unmanned Dragon capsule with a kerosene/oxygen chemical rocket stage of sufficient power to drive it back to Earth. This is the Earth return vehicle (ERV). 

A second launch will deliver to the Martian surface an 11 ton payload consisting of a Mars Ascent Vehicle (MAV) employing a single methane/oxygen rocket propulsion stage, a small automated chemical reactor system, 3 tons of surface exploration gear, and a 10 kilowatt power supply, which could be either nuclear or solar. The MAV would land with its propellant tanks filled with 2.6 tons of methane, but without the 9 tons of liquid oxygen required to burn it. This oxygen could be made over a 500 day period by using the chemical reactor to break down the carbon dioxide that composes 95 percent of the Martian atmosphere.  Since the reactor and the power system together  only weigh about 2 tons, using such technology to generate the required oxygen in-situ rather than transporting it saves a great deal of mass, and offers the further benefit of providing copious power and unlimited oxygen to the crew once they arrive.  Combined, the 11.6 tons of methane/oxygen propellant is sufficient to deliver a 2 ton crew cabin (equal in dry mass to the lunar ascent vehicle used during the Apollo missions) from the Martian surface to high Mars orbit where it can rendezvous with the ERV.

Once these elements are in place, the third launch would occur, which would send a Dragon capsule with a crew of two astronauts on a direct trajectory to Mars. The capsule would carry 2500 kilograms of consumables, sufficient, if water and oxygen recycling systems are employed, to support the two-person crew for up to three years. Given the available payload capacity, a light ground vehicle and several hundred kilograms of science instruments could be taken along as well.

The crew would take six months to reach Mars, after which they would land their Dragon capsule near the MAV. They would then spend the next year and a half exploring Mars. Using their ground vehicle for mobility and the Dragon as their home and laboratory, they could search the Martian surface for fossil evidence of past life that may have existed in the past when the Red Planet featured standing bodies of liquid water. Going further, they could set up drilling rigs to bring up samples of subsurface water within which native microbial life may yet persist to this day. If they find either, they will prove that life is not unique to the Earth, but is a general phenomenon in the universe, thereby answering a question that thinking men and women have wondered upon for millennia.

At the end of their 18-month surface stay, the crew would transfer to the MAV, take off, and rendezvous with the ERV. This craft would then take them on a six-month flight back to Earth, whereupon it would enter the atmosphere and splash down to an ocean landing.

            2.         Technical Alternatives within the Mission Architecture

a.                  MAV and associated systems  

In the plan described above, methane/oxygen is proposed as the propulsion system for the MAV, with all the methane brought from Earth, and all the oxygen made on Mars from the atmosphere. This method was selected over any involving hydrogen (either as feedstock for propellant manufacture or as propellant itself) as it eliminates the need to transport cryogenic hydrogen from Earth or store it on the Martian surface, or the need to mine Martian soil for water. If terrestrial hydrogen can be transported to make the methane, about 1.9 tons of landed mass could be saved. Transporting methane was chosen over a system using kerosene/oxygen for Mars ascent, with kerosene coming from Earth and oxygen from Mars because methane offers higher performance (Isp 375 s vs. Isp 350 s) than kerosene, and its selection makes the system more evolvable, as once Martian water does become available, methane can be readily manufactured on Mars, saving 2.6 tons of landed mass per mission compared to transporting methane, or about 3 tons per mission compared to transporting kerosene. That said, the choice of using kersosene/oxygen for Mars ascent instead of methane oxygen is feasible within the limits of the mass delivery capabilities of the systems under discussion. It thus represents a viable alternative option, reducing development costs, albeit with reduced payload capability and evolvability.

b.                  ERV and associated systems.

A kerosene/oxygen system is suggested for Trans-Earth injection. A methane/oxygen system would offer increased capability if it were available. The performance improvement is modest, however, as the required delta-V for TEI from a highly elliptical orbit around Mars is only 1.5 km/s. Hydrogen/oxygen is rejected for TEI in order to avoid the need for long duration storage of hydrogen. The 14 ton Mars orbital insertion mass estimate is based on the assumption of the use of an auxiliary aerobrake with a mass of 2 tons to accomplish the bulk of braking Delta-V. If the system can be configured so that that Dragon’s own aerobrake can play a role in this maneuver, this delivered mass could be increased. If it is decided that the ~1 km/s Delta-V required for minimal Mars orbit capture needs to be done via rocket propulsion, this mass could be reduced  to as little as 12 tons (assuming kerosene/oxygen propulsion). This would still be enough to enable the mission. The orbit employed by the ERV is a loosely bound 250 km by 1 sol orbit. This minimizes the Delta-V for orbital capture and departure, while maintaining the ERV in a synchronous relationship to the landing site. Habitable volume on the ERV can be greatly expanded by using an auxiliary inflatable cabin, as discussed in the Appendix.

c.         The hab craft.

The Dragon is chosen for the primary hab and ERV vehicle because it is available. It is not ideal. Habitation space of the Dragon alone after landing appears to be about 80 square feet, somewhat smaller than the 100 square feet of a small standard Tokyo apartment. Additional habitation space and substantial mission logistics backup could be provided by landing an additional Dragon at the landing site in advance, loaded with extra supplies and equipment. Solar flare protection can be provided on the way out by proper placement of provisions, or by the use of a personal water-filled solar flare protection “sleeping bag.” For concepts for using inflatables to greatly expand living space during flight and/or after landing, see note in Appendix.

          3.         Alternatives to the Selected Mission Architecture

a.         Direct Return.

In an ideal world, direct return from the Martian surface using in-situ produced propellants is the way to go. This, of course, is the basis of the Mars Direct plan, which other things being equal, would be my preference. However, under the assumption that this is a near-term mission using soon-to-be-available systems with minimal technology development, that is not feasible. For example, direct return of a Dragon capsule from the Martian surface in one stage using hydrogen/oxygen propellant produced from Martian water would require about 50 metric tons of propellant. This would require 50 kilowatts of 24-hour power to produce, which, assuming a nuclear reactor is not available, means a solar array of about 5000 square meters.  Such an array would likely weigh at least 10 tons, thereby blowing the mission mass budget, and be difficult to deploy by automated systems as well. In addition, assuming a water concentration of 4% by weight in the soil, obtaining 50 tons of Martian water would require mining 1200 tons of soil, which is a non-starter. Using Martian water in combination with atmospheric CO2 to produce methane/oxygen instead of hydrogen/oxygen would cut the power requirement by about 40% and the mining requirement by 60%, but the plan still remains unfeasible within the limits of the available systems. Thus the use of a lightweight LEM-type vehicle to perform Mars ascent and rendezvous with a Dragon placed in a highly elliptical Mars orbit is necessary if the mission mass requirements and delivery capabilities are to converge.

b.         Double rendezvous

An alternative to the plan described here might be to fly the crew to Mars in the same Dragon used for the ERV (i.e., a “mothership”), and fly another Dragon to the Martian surface to provide a surface hab. The crew would then rendezvous with the MAV, and take it down to land near the surface hab, which they would live in for 1.5 years, after which they would ride the MAV back up to the ERV.  This architecture is feasible in principle, but inferior to the one selected because it requires two orbital rendezvous per mission instead of one, does not allow the ascent propellant to be made in advance of the launch of the crew, and lands the crew separate from substantial living quarters or extended life support capability, without any countervailing advantages.

          4.         General Observations

The proper goal of a human Mars mission program should be sustained exploration followed by settlement. This can only be done if costs are kept low. This plan creates sufficiently low cost mission architecture to enable sustained exploration. Falcon Heavy launches are priced at about $100 million each, and Dragons are presumably even cheaper. Adopting such an approach, we could send expeditions to Mars at half the mission cost currently required to launch a Space Shuttle flight. In addition, both Dragons employed in the mission are re-used: one remaining on site to contribute to the growing Mars base, and the other returned to Earth. It will be observed that no orbital infrastructure, advanced orbital operations, advanced propulsion, or even surface nuclear power systems (although the 10 kilowatt Topaz demonstrated by the Soviet program would fit the bill) are required to enable the mission.  This, plus the fact that the mission can be done using a booster soon to be available minimizes development cost and time, and moves the potential timeframe of the mission from the indefinite future to the near-present.

For settlement, cheap one-way transportation to Mars is required. In addition, cargos larger in scale both in mass and in dimension need to be delivered. This will require development of a true heavy lift vehicle, with at least an 8 meter and preferably a 10 meter fairing, and launch capabilities of over 100 tonnes to orbit. Furthermore, if costs are to be lowered, reusability is desired. However reusability needs to be placed in perspective. The most important part of a space transportation system to make reusable is the lowest stage, since this is the most massive (therefore offering the greatest reusability savings), and adding mass to it (to make it reusable) does not cause any increase in the mass of the stages above it. On the other hand, making upper stages or interplanetary transfer systems reusable only saves a small amount of hardware, but causes the mass of the stages below them to increase. Thus reusability needs to be implemented in steps from the bottom-up, rather than from the top-down (as was unfortunately done in the Shuttle.)

Using the mission architecture described here, and the soon to be available Falcon Heavy and Dragon, the first human missions could be done and an initial outpost could be established on Mars during the present decade. With the advent of a heavy lift vehicle capable of delivering ~9 m diameter hab modules in the 30 ton class one-way to Mars, the subsidized settlement of Mars could begin, with such return flights as remain necessary continuing to be conducted by the FH/Dragon-derived systems. If the heavy-lift vehicle can evolve to reusability, starting with its lowest stages, costs of one-way transport to Mars could be lowered further, eventually reaching the point where individuals of fairly ordinary means would be able to pay their own way, freely venturing forth to start new lives on a new world. 

Appendix: Notes Concerning Various Mission Issues

1.         Zero Gravity Health Effects.

There is no need for zero gravity exposure. Artificial gravity can be provided to the crew by tethering the Dragon off the TMI stage, in the same way as is recommended in the baseline Mars Direct plan.

2.         Radiation.

Cosmic ray radiation exposure for the crew is precisely THE SAME as that which would be received by those on any other credible Mars mission, all of which would use the six month Conjunction class trajectory to Mars, both because that is the point of diminishing returns (the "knee of the curve") where Delta-V trades off against trip time, and because it is uniquely the trajectory that provides a 2-year free return orbit after launch from Earth. Assuming the baseline mission, the total cosmic ray dose would be no greater than that already received by a half-dozen cosmonauts and astronauts who participated in long duration missions on Mir or ISS, with no radiation induced health effects having been reported. (Cosmic ray dose rates on ISS are 50% those of interplanetary space. The Earth's magnetic field does not shield effectively against cosmic rays. In fact, with a crew of six, the current planned ISS program will inflict the equivalent of 30 man-years of interplanetary travel GCR doses on its crews over the next decade. This is an order of magnitude more than that which will be received by the crew of the mission proposed here.) There are enough consumables on board to provide shielding against solar flares.

3.         Aerocapture.

The preferred method of Mars capture is aerocapture, rather than direct entry. This means that the Dragon aeroshield, which has some lifting capability, may well be adequate. This is a complex problem, but a back of the envelope calculation indicates that the Dragon’s shield size is in the ballpark. Thus, consider a loaded Dragon system with an entry mass of 17000 kg, an effective shield diameter of 4 meters, a drag coefficient of 1, coming in with an entry velocity of 6 km/s at an altitude of 25 km, where the Mars atmospheric density is 1.6 gm/m3.  Setting drag equal to mass times deceleration, it can be seen that that the system would decelerate at a speed of 42 m/s2, or a little over 4 gs. It could thus perform a 1 km/s deceleration in about 25 seconds, during which time it would travel about 140 km. This deceleration is sufficient to capture the spacecraft from an interplanetary trajectory into a loosely bound highly elliptical orbit around Mars. If the perigee is not raised, the craft will reenter again, and again, progressively lowering the apogee of its orbit, until either a desired apogee for orbital operations is achieved or the craft is committed to entry for purposes of landing. That said, if a larger aerobrake were desired, this could be created by adding either a flex-fabric or inflatable skirt to the Dragon core shield.

4.         EDL.

Using just its aeroshield for deceleration, the Dragon would have a terminal velocity of around 340 m/s on Mars at low altitude (air density 16 gm/m3). So we could either give it a rocket Delta-V capability of 600 m/s (a 20% mass hit assuming storable or RP/O2 propulsion, Isp~330 s) to land all propulsive, or we could use a drogue to slow it down (a 20 m diameter chute would slow it to ~70 m/s) and then employ a much smaller rocket Delta-V for landing.

5.         Living Volume.

The habitable volume of the Dragon capsule is admittedly lower than optimal. However it should be noted that with 5 cubic meters per crew member, it is 2.5 times higher than the 2 cubic meters per crew member possessed by Apollo crews. Alternative comparisons include 9 cubic meters per crew member on the Space Shuttle, or 8 cubic meters per crew member on a German U-Boat (Type VII, the fleet workhorse) during WWII. This would be uncomfortable, but ultimately, workable by a truly dedicated crew.

However these limits can be transcended. The Dragon has a 14 cubic meter cargo area hold below the aeroshield. Into this we could pack an inflatable hab module, in deflated form, but which if inflated, could be as much as 6 m in diameter and perhaps 8 m long, thereby providing an additional three decks, with added useful volume of 226 cubic meters and a total floor space of 85 square meters, 85% as much as that in the Mars Society's MDRS or FMARS stations, which have proved adequate in size for crews of six. After Trans-Mars injection, the Dragon would pull away from the cargo section and turn around, then return to mate its docking hatch with one in the inflatable. It would then pull the inflatable out of the cargo hold, much as the Apollo command module pulled out the LEM. The inflatable could then be inflated. The other end of the inflatable would be attached to the tether, which is connected to the TMI stage, for use in creating artificial gravity.

Upon reaching Mars, the inflatable could either be expended, along with the tether system and TMI stage, prior to aerocapture. Alternatively, and optimally, the tether and TMI stage alone would be expended, but the inflatable deflated and retained for redeployment as a ground hab after landing.

Extra space could be also be provided on the ground by using a 4th launch to pre-land another Dragon loaded with supplies, including one or more inflatable modules which could be set up by the crew after they land.

6.         Overall Risk.

The mission architecture is much safer than any based on complex mega systems requiring orbital assembly, since the quality control of orbital assembly does not compare with that which can be accomplished on the ground. It would be better to have a crew of four, but if we are to do it with Falcon Heavys, a crew of two is all we can do.  While such a crew size lacks a degree of redundancy otherwise desirable, it also offers the counter benefit of putting the fewest number of people at risk on the first mission. It's quite true that not flying anywhere at all would be safer, but if you want to get to Mars, you have to go to Mars.  

[Image: SpaceX]

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