Addendum A added below 7-20-19
Addendum B added below 8-19-19
Addendum C added below 8-19-19
Tuesday July 16 is the 50th anniversary of the launch of Apollo 11 to the moon. Saturday July 20 is the actual 50th anniversary of the landing. Those of us old enough, remember where we were and what we were doing, when the news of the landing broke.
I was attending the US Naval Academy at the time of that landing. The moonwalk took place in the wee hours of the morning, in that time zone. We otherwise-incommunicado midshipmen were awakened and taken to see the televised event of history being made.
This truly historic event was the culmination of a cold war space race with the Russians, that started with their orbiting of Sputnik in 1957. I remember watching that first satellite in the evening sky, and I have watched a great many since.
The US seemed to play catch-up for several years afterwards, as the Russians orbited the first man in space, took the first spacewalk, flew the first woman in space, and flew the first multi-person crews. Until 1958, the US did not have a NASA at all.
With some exceptions, our effort was a cost-is-no-object, take-some-serious-risks “crash program”. The Christmas 1968 Apollo 8 mission around the moon was an example of taking some very serious risks. It was really the first manned test flight of the Saturn-V rocket.
But it worked. Their giant rocket had failed too many times while ours did not, which is what led to Apollo 11 and the other 5 landings. The goal was to beat the Russians to the moon, and it was accomplished.
Consequently, what we did was less about science and exploration, and more about experimental test flight work. This was true from the very first Mercury shot (Alan Shephard in 1961) onward. All but one of our moon-walking astronauts was an engineering test pilot.
Only Harrison Schmidt on the very last landing was originally educated as a geologist. Our astronauts were trained to set up experiments, and to collect certain types of rocks, so that other scientists back home could do the actual science work later. That work is still going on.
Why We Never Went Back All These Decades
There was no longer any “race” to win. Because there was no “race” to win, there was no support for spending large amounts of government money. Typically, the funding for science and exploration projects here on Earth was far smaller, and most of it was historically non-governmental.
Government funding is very highly politicized. John F. Kennedy committed us to a moon landing before 1970, but he was not a fan of the space program, until he witnessed a Saturn-V engine test. It was Lyndon Johnson who was the real believer in a space program, and he convinced Kennedy that beating the Russians to the moon was really something possible and worthwhile.
The presidential leadership continuity that got this 8-year effort accomplished was mainly because Johnson succeeded Kennedy. Richard Nixon, who was not a big supporter of the space program, succeeded Johnson, and he killed the Apollo moon program at the 6th landing (Apollo 17).
We were supposed to have flown another 5 landing missions all the way through Apollo 22. Leftover unused rockets and capsules are why there are so many of these items on display around the country today.
The big-ticket manned space items since, have been the Space Shuttle and the International Space Station. These have been marvelous things, to be sure, but they are as much about “workfare” for the businesses that do this, and porkbarrel politics for the congressional districts where they are located, as anything else.
Why We Should Go Back (And Farther)
In the long run, it’s not about winning some race, and it’s not so very much about doing pure science just for the sake of knowledge. It’s about exploration of the unknown, something hard-wired into humans. In centuries past, this was exploration of the unknown parts of the Earth.
“Exploration” is a really an emotionally-loaded code word. What it truly means is you go there to find out what all is there (resources), and where exactly it is (how hard to obtain). Then you stay a while to figure out how to use what you found, to cope with living in the local environment.
Unless you do that correctly, there is no possibility of future settlements and the associated future economies! There is no way to accomplish anything else, except just the act of going there and returning (which is the bulk of what Apollo accomplished).
Is there anything worthwhile to accomplish out there? Yes, definitely!
In the longer term, there are those future off-world settlements and the associated future economies. I cannot tell you the details of how this might benefit us, because it has yet to be done. But it has always proven beneficial in prior centuries here on Earth.
In the shorter term, there are the possibilities of space resource businesses, and of planetary protection against rogue asteroid and comet impacts. There is simply no better reason for continuing both unmanned and manned space programs than finding ways to protect the folks back home!
Change in Approach is Required
This cannot be done properly as a “NASA-does-it-all” set of projects. It is likely not possible for the US to do it all alone, either. This cannot be done as a “crash program”, because that is far too expensive. But exploration is risky, you cannot be too risk-averse, either.
Ethically, you do have to address the known risks as best you can, because we have learned to our chagrin that there is nothing as expensive as a dead crew! If addressing those risks means you design vehicles and hardware different from anything ever seen before, then so be it. See addendum A below.
This definitely requires thinking way outside the boxes where we have been before, with Apollo and Space Shuttle and ISS. This is going to require a far stronger and more equal partnership between the government and the companies in the business. And it requires looking far beyond the traditional contractor base. Pork-barrel politics-as-usual and big-corporate “workfare” simply have to go away. See addendum B below.
My Suggestions for the Near Term
Establish a continuous human presence on the moon, the first item. Start small and expand it slowly over time. Do the “exploration” thing right, this time.
Send humans to Mars as the fulfillment of a dream centuries old, probably the second item. When we go, do the “exploration” thing right, from the very first landing.
But, any vehicle capable of taking crews to Mars can take a crew to near-Earth asteroids and comets. Visit those asteroids and comets and properly explore them, in order to learn how to defend against their impacting Earth.
That’s the third item, but it is just as easily done, and at least as important, as going to Mars. Maybe we do them at the same time. See addendum C below.
Update 7-20-19: An edited version of this article appeared as a guest column today in the Waco "Tribune Herald".
Addendum A 7-20-19: Ethically and Responsibly Addressing Known Risks For Spaceflight
In the main article, I said we are ethically bound to address known risks of spaceflight as best we can, but did not explore that topic. Here are the details of that topic.
There is a whole long list of safety risks associated with any sort of spaceflight. Three come to mind as the most truly credible risks: (1) reliability of, and escape from, spacecraft and booster rockets, (2) microgravity diseases, and (3) exposure to radiation.
The first one has cost us three American crews totaling 17 people (Apollo 1, shuttle Challenger, and shuttle Columbia). Each caused a year-or-more stand-down, and very expensive investigations, plus very expensive changes. The two shuttle losses were ultimately caused by bad management decisions. This is why I claim “there is nothing as expensive as a dead crew, especially one dead from a bad management decision”.
Making spaceflight more safe from a reliability standpoint is also something we already know how to address! This takes careful design allowing for failure modes, redundant systems, and copious verification testing. Mitigation efforts will never be perfect, but they can be quite good.
The other two have been long studied in low Earth orbit, where microgravity exposure is inherent in everything we have done there, and radiation exposure is somewhat more than on Earth’s surface, but less than outside the Van Allen radiation belts, and far less than inside the belts themselves.
Microgravity has proven to affect the human body in a variety of expected, and unexpected, ways. The longer one is exposed, the worse the various diseases become. Beyond the bone decalcification and muscle-weakening we expected, there are also degradations of the heart and circulatory system, degradation of vision from eye geometry changes due to the fluid pressure redistribution, and immune system degradations. No doubt more will be discovered, as that has been the trend.
The longer exposed, the longer it takes to recover upon returning home, with full recovery actually still in doubt for some of the effects. The practical time limit seems to be only a bit more than a year. Usual practices on ISS call for 6 months to a year’s exposure at most.
For operations anywhere in Earth-moon space, one-way flight times are in the “4 days to a week or so” class. Out of 6 months-to-a-year allowable, that leaves at least several months available “on site” anywhere in Earth-moon space to do whatever needs to be done, before too-serious ill effects occur.
We do not know if low gravity, such as the 16% on the moon, is therapeutic enough to reverse the effects, or at least extend the feasible exposure time. That is because we never built the spinning space stations in low Earth orbit, in which to experiment with different levels of artificial spin gravity. In hindsight, that lack was rather stupid of us, now wasn’t it?
We do know that something near one full Earth gravity (one “gee”) is therapeutic, because that is what we evolved in. So, until we know better, any artificial spin gravity schemes need to supply very near one gee, in order to obtain the full Earthly benefits that we know work.
Destinations outside of Earth-moon space are very much further away: one-way travel times range from near 6 months to multiple years. This is pretty much outside the limits of microgravity exposure that we have established on ISS (no more than a year, only 6 months preferred).
Mars is 6-to-9 months away one-way, and we do not know how therapeutic its lower gravity (38%) really is. Other destinations are further away still, and all those we can reach are even lower gravity than Mars. That situation says quite clearly that we need to provide artificial gravity (no matter how inconvenient that might otherwise be) at something near one gee (until we actually know better), during these one-way transits, in order to best preserve the health of the crews.
Ethically, you simply cannot argue with that conclusion, no matter how inconvenient for design purposes, or for total mission cost purposes.
Supplying Artificial Gravity
There is as yet no such thing as “Star Trek”-type artificial gravity. The only physics we have to serve that purpose is centrifugal force. You must spin the vehicle, to generate centrifugal force as an equivalent to gravity. If the spin rate is low, then Coriolis forces (something everyone experienced on a merry-go-round) become less important, and so fewer folks can tell the difference between this and real gravity.
The physics says that the acceleration you feel is proportional to the radius of spin and to the square of the spin rate. The actual physics equation says
a = R w2 where a is the acceleration, R the spin radius, and w the spin rate
Another form expressed in gees, and not absolute acceleration units is
gees = 1.00 * (R, m / 55.89 m) (N, rpm / 4 rpm)2
Experience with spin rates says that normal untrained and unacclimatized people can tolerate 3 to 4 rpm immediately, for long-term exposures, without getting motion sick. People extensively trained might tolerate higher spin rates in the 8-12 rpm class without getting motion sick from long exposures. Still-higher spin rates (16+ rpm) induce blood pressure gradients head-to-toe that are just unacceptable for long term exposures.
The upshot of that is that the required spin radius (half of a crucial dimension of the craft you must build) will be about 99-56 meters at 3-4 rpm, and about 14-6.2 meters at 8-12 rpm. These are inconveniently large dimensions, at least at first glance.
3-dimensional objects typically have 3 axes. About these axes these objects have a property called “moment of inertia” that relates to spin dynamics. Usually, higher moment of inertia correlates with larger dimension along a perpendicular axis to the spin axis.
There are two (and only two) stable spin modes for most objects: about the axis for highest moment of inertia (longest dimension), and about the axis for lowest moment of inertia (shortest dimension). The first case is exemplified by a baton twirler’s spinning baton, and the second case is exemplified by a spinning bullet or artillery shell. There are no other stable modes of spin. See Figure A-1.
Clearly, building a “spinning rifle bullet” 112 m in diameter at 4 rpm for one full gee at its outer girth is not so very feasible. But spinning a smaller-diameter “something” that is 112 m long, end-over-end at 4 rpm, for 1 gee at each end, would indeed be a feasible thing to attempt.
We already know a lot about the transient dynamics of spinning rigid objects, something important for spin-up and spin-down, as well as for applying thrust while spinning. This got started balancing steam locomotive wheels more than 200 years ago. It continues today balancing wheels and tires on cars and trucks. There would be no engineering development work to design a long, narrow spacecraft that spins end-over-end for artificial gravity. There would only be proving-out the specific design before we use it.
The most-often-proposed alternative is a cable-connected structure, because it is conceptually easy to reel-out long cables between two small objects. Cables only support tension loads, not compression, nor side-directed loads. Transient dynamics for spin-up and spin-down, and especially for applying thrust while spinning, are incredibly complex and still not very well-known. This is not something we have much experience with, at all. So there is a huge engineering development effort needed, beyond just proving-out the actual design to be applied.
What this really says is that the preferred near-term spacecraft design is a long and rigid, more-or-less cylindrical shape, to be spun end-over-end, baton-style. This will generate varying gee from a maximum near the ends, to zero at the spin center. That is very likely the lowest-weight rigid-body design for any given spin condition, and it is consistent with our long history rigid-body spin experience, thus eliminating the huge engineering development effort required of cable-connected systems.
We know that microgravity vs gravity has no impact while prone sleeping, or else Earthly bed rest studies would not be a decent surrogate for some in-space microgravity effects. That means you can put the sleeping quarters in the low gravity section of the spacecraft near the spin center, and just put the daily workstations in the full-gravity sections of the spacecraft near the ends. See Figure A-2.
Figure A-2 – Why Selecting Baton Spin Mode Is Wisest Choice
There are basically three radiation hazards to worry about: galactic cosmic rays (GCR), solar flare events (SFE), and the Van Allen radiation belts about the Earth. All three are atomic or subatomic particles, just at different speeds and quantities. The threats they pose are location-dependent.
GCR is a very slow drizzle of really high-speed particles moving at a large fraction of the speed of light. Particles that energetic are very difficult to shield against, because they penetrate deeply into shielding material, and quite often create “secondary showers” of other harmful radiation when they strike the atoms in the shield material. If the shielding atoms are low atomic weight, the secondary shower effect is greatly reduced.
GCR comes from outside the solar system. Its quantity is affected by the solar wind, in turn affected by the sun’s sunspot cycle, which is about 11 years long. The solar wind is stronger when sunspots are active, making GCR lower in the vicinity of the Earth-moon system at that time.
From NASA’s radiation effects website, I obtained these values that apply in the vicinity of the Earth-moon system. GCR maximizes at about 60 REM per year when the sun is quiet, and minimizes at about 24 REM per year, when sunspots are most active. To “calibrate” what may be unfamiliar units, the natural Earthly background radiation is about 0.3 REM per year, and a lethal dose would be 300 to 500 REM accumulated in a “short time” of hours to weeks.
The NASA astronaut exposure standards are set at about twice the levels allowed for Earthly nuclear workers. Those NASA standards are no more than 50 REM per year, no more than 25 REM in any one month, and a career limit that varies with age and gender, but maxes-out at no more than 400 REM accumulated over a lifetime. These are predicated upon a single-handful percentage increase in the likelihood of late-in-life cancer.
Clearly, with a very modest shielding effect, GCR is not the “killer” it is often portrayed to be.
SFE (solar flare bursts) are different. They are much lower speed particles, much easier to shield, but there is an incredibly-huge flood of them. They come in very-directional bursts from the sun at erratic intervals. There are usually more of them during times of active sunspots, but they can indeed happen when the sun is quiet. They come at irregular intervals measured in several months.
The intensity of a burst can vary from tens of REM over a few hours, to tens of thousands of REM over a few hours. Obviously, for unshielded persons, the great bulk of events like this would be fatal doses, and it is an ugly death. There was a fatal-level event in 1972 between the last two Apollo missions to the moon, and a low-intensity (non-fatal) event during one Apollo mission to the moon.
We had chosen to ignore this SFE threat during Apollo because the short duration of the missions (at most 2 weeks) was small compared to the typical interval between events. Had a large one hit an Apollo mission, the crew would have died in space. As it turns out, this was not a good assumption.
For an extended or permanent return to the moon, shielding is obviously imperative. On Earth, we are protected from these SFE’s (and the GCR) by both the Earth’s magnetic field and its atmosphere.
These are a very real threat anywhere outside the Earth’s magnetic field. In low Earth orbit, we are protected only by the magnetic field, and the background exposure is higher than on Earth, but much less than beyond the magnetic field.
The Van Allen belts are regions of these radiation particles trapped in the Earth’s magnetic field. The intensity is lethal on a scale of days-to-weeks, but tolerable on a scale of hours-to-a-day-or-so. The inner boundary is not sharp, but generally considered to become a problem at about 900 miles altitude, and extending many thousands of miles out from the Earth.
The exception is the “South Atlantic Anomaly”, where the inner side of the Van Allen belt dips down locally to low Earth orbit altitude (100-300 miles). Satellites and spacecraft in high-inclination orbits inherently pass through the South Atlantic Anomaly every several orbits. The ISS does encounter this threat, it being short “flashes” of exposure that accumulated over time still fall within the astronaut exposure standards.
Spacecraft traveling to the moon or elsewhere must transit the Van Allen belts. Because of the potential for lethal exposure if you linger within them, such transits must be made quickly. Apollo did this, transiting within only several hours. Given the state of today’s electric propulsion technology, this rules out using electric propulsion for people to leave Earth orbit for the moon or elsewhere, because the spiral-out time is measured in months. That would be lethal exposure, even with some shielding.
The same NASA radiation site has data regarding the shielding effects of typically-considered materials. Those are hydrogen, water, and aluminum. Mass of shielding above a unit exposed area turns out to be the correlating variable, and 15-20 g/cm2 seems to be enough to generally address the worst SFE.
Hydrogen has the lowest density, requiring the thickest layering, but also has the least secondary shower potential against GCR.
15-20 cm of water is 15-20 gm/cm2, same shielding effect as a really thick layer of hydrogen. Water molecules are still light enough not to have much secondary shower risk.
Aluminum would be the thinnest layer, but with the greater secondary shower effect. However, of the practical metals, its atoms are the lightest, and this secondary shower effect is deemed tolerable with it. 6-8 cm thick aluminum plate would be required. That is quite out-of-line with current spacecraft hull design practices: something near a millimeter.
Other materials based on polymers, and even most rocket propellants, are light enough molecules to be effective shielding with a low secondary shower risk, yet with densities in the same ballpark as water, for a thinner layer thickness. Any of these could be practical shielding materials.
What you have to do is not simply add shielding weight to your design, but instead rearrange the distribution of masses you already otherwise need, so that they can also serve as radiation shielding. You will need meteoroid shielding and thermal insulation, and any manned craft will have water and wastewater on board, as part of the life support system. All spacecraft will need propellant for the next (and subsequent) burns. You use a combination of these, acting together.
The real suggestion here is to use water, wastewater, and next-burn propellant tankage as shadow shields, in addition to the meteoroid protection and thermal insulation materials that the manned modules require anyway. It doesn’t take much of this to cut the worst 60 REM/year GCR to under 50 REM/year. It takes only a little more to cut worst case SFE to safe short-term exposure levels.
If you cannot protect the whole manned interior, then the flight control station becomes first priority, so that maneuvers can be flown, regardless of the solar weather. Second priority would be the sleeping quarters, to reduce round-the-clock GCR exposure further.
See Figure A-3 for one possible way to do this, in an orbit-to-orbit transport design concept. This would also be a baton-spin vehicle for artificial gravity during the long transit. Plus, it requires a lot of interior space for the mental health of the crew. Somewhere between 100 and 200 cubic meters per person is needed, and it must be reconfigurable as desired by the crew. That is a topic out of scope here.
Spin-up is likely by flywheels in the center module. The vehicle is spun-up after departure, and de-spun before arrival. If a mid-course correction is needed, the vehicle could be de-spun for that, and spun back up for remainder of the transit.
Note how the arrival propellant and the water and wastewater tankage has been arranged around the manned core to provide extra shadow shielding for really effective radiation protection. The manned core modules are presumed insulated by polymeric layers that also serve as meteor shielding (while adding to the radiation protection without being driven by that issue). The pressure shell on the inside of this insulation should be unobstructed by mounted equipment, so that easy and rapid access for patching is possible.
At departure, the vehicle can be propelled by a different propellant and engine choice, since departure is a short event. The arrival propellant is likely a storable to prevent evaporation losses. Return propellant tankage sets can be sent ahead unmanned, for docking in orbit at the destination.
There is an emergency return capsule (or capsules) mounted at the center module, enough for the entire crew. “Bailout” at destination presumes a rescue capability there. Emergency bailout, upon a failed burn for returning to Earth orbit, is the main function of this capsule. Routinely, it could return a crew from the spaceship, once parked safely in Earth orbit.
This kind of orbit-to-orbit transport design could serve to take men to Mars or to the near-Earth asteroids and comets. For Mars, the lander craft could be sent ahead unmanned to Mars orbit. None are needed to visit asteroids. The design of Mars landing craft is out-of-scope here.
By refueling and re-supplying in Earth orbit, such a manned core design could easily be used for multiple missions, once built. Care must be taken in its design and material selection to support many thousands of cycles of use. Thus the craft could safely serve for a century or more, updated with better propellants and engines as the years go by.
There I went and wrote a basic “how-to” document for practical interplanetary spaceship design!
Figure A-3 – Using Otherwise-Required Materials To Also Serve As Radiation Shielding
Addendum B 8-19-19: “Corporate Workfare” and Political Pork Versus Worthwhile Projects
The pork barrel aspect got started with the early days of NASA itself. NASA labs and centers got situated in the districts of powerful senators and representatives, in order to get the votes for the funding to carry out the Kennedy mandate to go to the moon. Among many examples is the manned spaceflight center in Houston, Texas. This is essentially political corruption, just never referred to by that particular word.
Once this happened, it became the norm, and is seemingly not changeable. A huge base of contractors and factory locations got developed to build the Saturns and the Apollo vehicles. Once that program was done, then in the minds of those congressional figures, there needed to be something else for these contractors and factories to do, because they were located in the districts of those powerful people. Essentially the same contractor and factory base did the space shuttle, and the ISS, so the funding essentially continued to flow to those same districts, for all these decades since.
The current space launch system (SLS) giant rocket, and the Apollo-on-steroids Orion spacecraft, derive from that same contractor and factory base, located in those same districts of those same powerful congressional seats. Only the specific seat holders have changed. The projects which NASA can take on have to use those assets in those districts, lest funding dry up. That is how congress dictates the details of what rockets and spacecraft get developed and what the flagship missions are. NASA does not get to make a truly logical choice, only the political choice.
SLS got started as “Constellation” fpr a return to the moon, but is essentially the same rocket and capsule system, just under multiple names and programs. It is really just Saturn/Apollo redone with space shuttle technology and hardware. This is now very old and inherently-expensive technology, compared to what the new entrants (initially Spacex) in the business have been doing.
In NASA’s early days, the Mercury and Gemini capsules were supplied by McDonnell-Douglas, and the Apollo capsule by North American Aviation. All have since been gobbled-up into Boeing. Rocket stages were supplied by Boeing, General Dynamics, McDonnell-Douglas, and many others.
Now there are just Boing and Lockheed-Martin (sometimes together as ULA), and Orbital ATK, plus the new entrant Spacex and perhaps soon Blue Origin. All of the big-motor solid rocket plants (Thiokol, Hercules, CSD, and others) are now part of Orbital ATK as an effective solid propellant monopoly.
The maker of the Apollo lunar lander module (Grumman) is now part of Northrup-Grumman, the third remaining main airframe provider, and usually relegated to team member status, in teams headed by either Boeing or Lockheed-Martin. There is no longer any effective competition.
In effect, the many competing contractors in NASA’s early days have consolidated to a rather noncompetitive very few, an oligopoly. Without effective competition, there is little incentive to actually go and be successful flying anything. This shows in the track record of late: it was 8 years from Kennedy’s mandate before any Saturns had flown at all, to the first lunar landing using the final big Saturn 5. Compare that to the totality that is SLS / Orion: started under G. W. Bush, continuing today, a decade later, billions over budget, and still yet to fly at all. This is about adapting existing engines; the original Saturns were about developing new engines “from scratch”, a much tougher job to do.
Between the powerful pork-barrel aspect, and the non-competitive oligopoly aspect, it should be no surprise at all that what was once a powerful conglomeration of American know-how has devolved into little more than a welfare system for corporate giants sucking at the public tit, getting their contract payments, without really having to succeed at anything.
Addendum C 8-19-19: Overall Mission Architecture and Vehicle Concepts For Mars
Fully covering this topic is way too large for an addendum to this posting. It is essentially a new posting defining a planned mission to Mars, and likely another separate posting describing how to use the same hardware to visit a near-Earth asteroid. This 2019 Mars mission will be an update to an earlier posting titled “Mars Mission Outline 2016” and dated 28 May 2016. Watch for these new postings. They are in work as of this writing.
The new 2019 version of the Mars mission uses a larger orbit-to-orbit transport than the 2016 version, and it also recovers the solar-electric tugs that preposition unmanned assets at Mars for the manned mission. It uses similar landers as the 2016 version, and it still jettisons the Earth departure stage without recovery.
That non-recovered Earth departure stage could be addressed in future versions by fitting a larger departure stage with a second propulsion system, possibly electric, and putting it into a 2-year-period solar orbit after stage-off. Then it could be captured into Earth orbit for reuse.
Main point here: if one does spin gravity in a baton-spin mode, the resulting orbital transport vehicle is ill-adapted for a direct entry at Mars, or a direct entry at Earth upon return. Such a design is far better-adapted as an orbit-to-orbit transport, with any Mars lander function relegated to a separate vehicle, sent ahead separately with its propellant supplies. Long-life reusability also points toward an orbit-to-orbit transport design, free of entry heat shield requirements. Such a concept was sketched in Figure A-3, located in Addendum A above.
This has inherently-higher velocity requirements, there is no way around that! But the direct-entry scenarios simply cannot provide this degree of safety for the crew in terms of radiation exposure, microgravity diseases, and having a “way out” at every step of the mission. That higher velocity requirement is just the price you have to pay to do this job “right”, in terms of ethical requirements!
The resulting mission architecture requires that both landers and the return propellant get sent ahead unmanned to parking orbit about Mars, with the manned orbit-to-orbit transport arriving afterward, and docking in Mars orbit with those items. This powerful concept is not unlike the Lunar Orbit Rendezvous architecture that made it possible to mount each Apollo landing mission with only one Saturn 5 booster. The concept is illustrated in Figure C-1, and its orbital velocity requirements in Figure C-2. The mass ratio-effective velocity requirements are given in Figure C-3.
For Mars arrival only, there needs to be an additional propellant allowance to cover rendezvous requirements with the assets sent ahead. As a wild guess, add 0.2 km/s to the Mars arrival delta-vee requirement in Figure C-3. There is no such corresponding allowance requirement for Mars departure, and none is needed for the Earth arrival.
The landers themselves are envisioned as one-stage reusable articles that make multiple flights, based out of low Mars orbit. Sending 3 landers allows one vehicle to make a landing, with another in reserve as a rescue craft. Thus, there is a “way out” even during the landings, unlike with Apollo. The presence of a third lander allows one vehicle to become unserviceable, while still maintaining the reserve rescue lander capability, without which landings so far from Earth become too ethically risky to attempt.
There is an insignificant velocity requirement for the deorbit burn, with aero-deceleration to about Mach (1 km/s) at a rather low altitude. From there the vehicle speed must be quickly killed with retro-propulsion (no time to deploy a chute, much less wait for it to do any effective speed reduction). That last requires a fairly large “jigger factor” to cover maneuvering, hovering, and diversion-away-from-hazards. Even so, the mass-ratio-effective velocity requirement for descent is not large at all. This allows larger descent payload fractions.
The ascent velocity requirement must cover full orbital speed, plus aero and gravity losses, and a final rendezvous allowance. This is a far larger velocity requirement, but the payload is smaller because most of the supplies are exhausted, some of the equipment will get left behind, and the weight of the collected samples is not much in comparison. The ascent payload fraction is quite a bit smaller. These requirements are illustrated as a part of Figure C-4.
Figure C-4 – Surface Landing Forays Based Out Of Low Mars Orbit
The unmanned transfers can be done more efficiently with solar electric propulsion (SEP) because of its far-higher specific impulse, and because there is no need to transit the Van Allen belts quickly. There is also no need to worry about reducing crew confinement times, because these transfers are unmanned.
This prepositioning of assets at Mars using SEP was also a part of my 2016 Mars mission posting. The difference here is that I recover the SEP “tugs” into Earth orbit, for reuse on future missions.
The propellant sent ahead to Mars is all storable (the same NTO-MMH), and comprises both the Mars departure and Earth arrival propellant, plus a supply for the landers. All this gets sent with the landers themselves, as three unmanned “cluster” vehicles. The Earth arrival propellant is arranged about the periphery of the manned spaces of the orbital transport, so that it can also serve as part of the radiation shielding during the transit home. The Mars departure tankage is on one end, and is staged-off after the departure burn, before the vehicle is spun up. See again Figure C-1.
For the assets sent ahead with SEP, mass ratio-effective design velocity requirements are much more problematic. There are no drag losses in vacuum, but the gravity losses are huge, since the “burns” are months long. For a rough rule-of-thumb estimate, just use twice the values in Figure C-3.
There is a lot of detailed work yet to do and to document for this version of the mission. That will be covered in the new posting. Watch for it.